In-flight comparisons of boundary-layer and wake measurement probes for incompressible flow

NASA Technical Reports Server (NTRS)

Mertaugh, L. J., Jr.

1972-01-01

The results are presented of in-flight comparisons of a number of boundary-layer and wake measurement probes suitable for low-speed flight-test investigations. The tested boundary-layer probes included a traversing total-pressure probe and a hot-film probe mounted on an internally-mounted drive mechanism, a curved and a straight boundary-layer rake, and a traversing hot-film probe with an externally-mounted drive mechanism. The wake measuring devices included a traversing, self-aligning probe, a wake rake, and an integrating wake rate. The boundary-layer data are compared with a common reference velocity profile and comments given regarding the accuracy of the static-pressure and total-pressure measurements. Discussions on the various calibration presentations used with hot-wire and hot-film sensors and various aspects of improving the accuracy of hot-film sensor results are given in the appendix of this report.

Wall-pressure fluctuations beneath a spatially evolving turbulent boundary layer

NASA Astrophysics Data System (ADS)

Mahesh, Krishnan; Kumar, Praveen

2016-11-01

Wall-pressure fluctuations beneath a turbulent boundary layer are important in applications dealing with structural deformation and acoustics. Simulations are performed for flat plate and axisymmetric, spatially evolving zero-pressure-gradient turbulent boundary layers at inflow Reynolds number of 1400 and 2200 based on momentum thickness. The simulations generate their own inflow using the recycle-rescale method. The results for mean velocity and second-order statistics show excellent agreement with the data available in literature. The spectral characteristics of wall-pressure fluctuations and their relation to flow structure will be discussed. This work is supported by ONR.

An experimental study of the wall-pressure fluctuations beneath low Reynolds number turbulent boundary layers.

PubMed

Van Blitterswyk, Jared; Rocha, Joana

2017-02-01

A more complete understanding of the physical relationships, between wall-pressure and turbulence, is required for modeling flow-induced noise and developing noise reduction strategies. In this study, the wall-pressure fluctuations, induced by low Reynolds number turbulent boundary layers, are experimentally studied using a high-resolution microphone array. Statistical characteristics obtained using traditional cross-correlation and cross-spectra analyses are complimented with wall-pressure-velocity cross-spectra and wavelet cross-correlations. Wall-pressure-velocity correlations revealed that turbulent activity in the buffer layer contributes at least 40% of the energy to the wall-pressure spectrum at all measured frequencies. As Reynolds number increases, the low-frequency energy shifts from the buffer layer to the logarithmic layer, as expected for regions of uniform streamwise momentum formed by hairpin packets. Conditional cross-spectra suggests that the majority of broadband wall-pressure energy is concentrated within the packets, with the pressure signatures of individual hairpin vortices estimated to decay on average within traveling ten displacement thicknesses, and the packet signature is retained for up to seven boundary layer thicknesses on average.

F-16XL ship #1 - CAWAP boundary layer rakes and hot film on left wing

NASA Technical Reports Server (NTRS)

1996-01-01

This photo shows the boundary layer hot film and the boundary layer rakes on the left wing of NASA's single-seat F-16XL (ship #1) used for the Cranked-Arrow Wing Aerodynamic Project (CAWAP) at Dryden Flight Research Center, Edwards, California. The modified airplane features a delta 'cranked-arrow' wing with strips of tubing along the leading edge to the trailing edge to sense static on the wing and obtain pressure distribution data. The right wing receives data on pressure distribution and the left wing has three types of instrumentation - preston tubes to measure local skin friction, boundary layer rakes to measure boundary layer profiles (the layer where the air interacts with the surfaces of a moving aircraft), and hot films to determine boundary layer transition locations. The first flight of CAWAP occurred on November 21, 1995, and the test program ended in April 1996.

Flat Plate Boundary Layer Stimulation Using Trip Wires and Hama Strips

NASA Astrophysics Data System (ADS)

Peguero, Charles; Henoch, Charles; Hrubes, James; Fredette, Albert; Roberts, Raymond; Huyer, Stephen

2017-11-01

Water tunnel experiments on a flat plate at zero angle of attack were performed to investigate the effect of single roughness elements, i.e., trip wires and Hama strips, on the transition to turbulence. Boundary layer trips are traditionally used in scale model testing to force a boundary layer to transition from laminar to turbulent flow at a single location to aid in scaling of flow characteristics. Several investigations of trip wire effects exist in the literature, but there is a dearth of information regarding the influence of Hama strips on the flat plate boundary layer. The intent of this investigation is to better understand the effects of boundary layer trips, particularly Hama strips, and to investigate the pressure-induced drag of both styles of boundary layer trips. Untripped and tripped boundary layers along a flat plate at a range of flow speeds were characterized with multiple diagnostic measurements in the NUWC/Newport 12-inch water tunnel. A wide range of Hama strip and wire trip thicknesses were used. Measurements included dye flow visualization, direct skin friction and parasitic drag force, boundary layer profiles using LDV, wall shear stress fluctuations using hot film anemometry, and streamwise pressure gradients. Test results will be compared to the CFD and boundary layer model results as well as the existing body of work. Conclusions, resulting in guidance for application of Hama strips in model scale experiments and non-dimensional predictions of pressure drag will be presented.

A study of juncture flow in the NASA Langley 0.3-meter transonic cryogenic tunnel

NASA Technical Reports Server (NTRS)

Chokani, Ndaona

1992-01-01

A numerical investigation of the interaction between a wind tunnel sidewall boundary layer and a thin low-aspect-ratio wing has been performed for transonic speeds and flight Reynolds numbers. A three-dimensional Navier-Stokes code was applied to calculate the flow field. The first portion of the investigation examined the capability of the code to calculate the flow around the wing, with no sidewall boundary layer present. The second part of the research examined the effect of modeling the sidewall boundary layer. The results indicated that the sidewall boundary layer had a strong influence on the flow field around the wing. The viscous sidewall computations accurately predicted the leading edge suction peaks, and the strong adverse pressure gradients immediately downstream of the leading edge. This was in contrast to the consistent underpredictions of the free-air computations. The low momentum of the sidewall boundary layer resulted in higher pressures in the juncture region, which decreased the favorable spanwise pressure gradient. This significantly decreased the spanwise migration of the wing boundary layer. The computations indicated that the sidewall boundary layer remained attached for all cases examined. Weak vortices were predicted in both the upper and lower surface juncture regions. These vortices are believed to have been generated by lateral skewing of the streamlines in the approaching boundary layer.

On the Unsteadiness of a Transitional Shock Wave-Boundary Layer Interaction Using Fast-Response Pressure-Sensitive Paint

NASA Astrophysics Data System (ADS)

Lash, E. Lara; Schmisseur, John

2017-11-01

Pressure-sensitive paint has been used to evaluate the unsteady dynamics of transitional and turbulent shock wave-boundary layer interactions generated by a vertical cylinder on a flat plate in a Mach 2 freestream. The resulting shock structure consists of an inviscid bow shock that bifurcates into a separation shock and trailing shock. The primary features of interest are the separation shock and an upstream influence shock that is intermittently present in transitional boundary layer interactions, but not observed in turbulent interactions. The power spectral densities, frequency peaks, and normalized wall pressures are analyzed as the incoming boundary layer state changes from transitional to fully turbulent, comparing both centerline and outboard regions of the interaction. The present study compares the scales and frequencies of the dynamics of the separation shock structure in different boundary layer regimes. Synchronized high-speed Schlieren imaging provides quantitative statistical analyses as well as qualitative comparisons to the fast-response pressure sensitive paint measurements. Materials based on research supported by the U.S. Office of Naval Research under Award Number N00014-15-1-2269.

Interaction between a normal shock wave and a turbulent boundary layer at high transonic speeds. I - Pressure distribution

NASA Technical Reports Server (NTRS)

Messiter, A. F.

1980-01-01

Asymptotic solutions are derived for the pressure distribution in the interaction of a weak normal shock wave with a turbulent boundary layer. The undisturbed boundary layer is characterized by the law of the wall and the law of the wake for compressible flow. In the limiting case considered, for 'high' transonic speeds, the sonic line is very close to the wall. Comparisons with experiment are shown, with corrections included for the effect of longitudinal wall curvature and for the boundary-layer displacement effect in a circular pipe.

Numerical simulations of the flow in the HYPULSE expansion tube

NASA Technical Reports Server (NTRS)

Wilson, Gregory J.; Sussman, Myles A.; Bakos, Robert J.

1995-01-01

Axisymmetric numerical simulations with finite-rate chemistry are presented for two operating conditions in the HYPULSE expansion tube. The operating gas for these two cases is nitrogen and the computations are compared to experimental data. One test condition is at a total enthalpy of 15.2 MJ/Kg and a relatively low static pressure of 2 kPa. This case is characterized by a laminar boundary layer and significant chemical nonequilibrium in the acceleration gas. The second test condition is at a total enthalpy of 10.2 MJ/Kg and a static pressure of 38 kPa and is characterized by a turbulent boundary layer. For both cases, the time-varying test gas pressure predicted by the simulations is in good agreement with experimental data. The computations are also found to be in good agreement with Mirels' correlations for shock tube flow. It is shown that the nonuniformity of the test gas observed in the HYPULSE expansion tube is strongly linked to the boundary layer thickness. The turbulent flow investigated has a larger boundary layer and greater test gas nonuniformity. In order to investigate possibilities of improving expansion tube flow quality by reducing the boundary layer thickness, parametric studies showing the effect of density and turbulent transition point on the test conditions are also presented. Although an increase in the expansion tube operating pressure level would reduce the boundary layer thickness, the simulations indicate that the reduction would be less than what is predicted by flat plate boundary layer correlations.

Fuselage boundary-layer refraction of fan tones radiated from an installed turbofan aero-engine.

PubMed

Gaffney, James; McAlpine, Alan; Kingan, Michael J

2017-03-01

A distributed source model to predict fan tone noise levels of an installed turbofan aero-engine is extended to include the refraction effects caused by the fuselage boundary layer. The model is a simple representation of an installed turbofan, where fan tones are represented in terms of spinning modes radiated from a semi-infinite circular duct, and the aircraft's fuselage is represented by an infinitely long, rigid cylinder. The distributed source is a disk, formed by integrating infinitesimal volume sources located on the intake duct termination. The cylinder is located adjacent to the disk. There is uniform axial flow, aligned with the axis of the cylinder, everywhere except close to the cylinder where there is a constant thickness boundary layer. The aim is to predict the near-field acoustic pressure, and in particular, to predict the pressure on the cylindrical fuselage which is relevant to assess cabin noise. Thus no far-field approximations are included in the modelling. The effect of the boundary layer is quantified by calculating the area-averaged mean square pressure over the cylinder's surface with and without the boundary layer included in the prediction model. The sound propagation through the boundary layer is calculated by solving the Pridmore-Brown equation. Results from the theoretical method show that the boundary layer has a significant effect on the predicted sound pressure levels on the cylindrical fuselage, owing to sound radiation of fan tones from an installed turbofan aero-engine.

Summary of experimentally determined facts concerning the behavior of the boundary layer and performance of boundary layer measurements. [considering sailing flight

NASA Technical Reports Server (NTRS)

Vanness, W.

1978-01-01

A summary report of boundary layer studies is presented. Preliminary results of experimental measurements show that: (1) A very thin layer (approximately 0.4 mm) of the boundary layer seems to be accelerated; (2) the static pressure of the outer flow does not remain exactly constant through the boundary layer; and (3) an oncoming boundary layer which is already turbulent at the suction point can again become laminar behind this point without being completely sucked off.

Effects of Periodic Unsteady Wake Flow and Pressure Gradient on Boundary Layer Transition Along the Concave Surface of a Curved Plate. Part 3

NASA Technical Reports Server (NTRS)

Schobeiri, M. T.; Radke, R. E.

1996-01-01

Boundary layer transition and development on a turbomachinery blade is subjected to highly periodic unsteady turbulent flow, pressure gradient in longitudinal as well as lateral direction, and surface curvature. To study the effects of periodic unsteady wakes on the concave surface of a turbine blade, a curved plate was utilized. On the concave surface of this plate, detailed experimental investigations were carried out under zero and negative pressure gradient. The measurements were performed in an unsteady flow research facility using a rotating cascade of rods positioned upstream of the curved plate. Boundary layer measurements using a hot-wire probe were analyzed by the ensemble-averaging technique. The results presented in the temporal-spatial domain display the transition and further development of the boundary layer, specifically the ensemble-averaged velocity and turbulence intensity. As the results show, the turbulent patches generated by the wakes have different leading and trailing edge velocities and merge with the boundary layer resulting in a strong deformation and generation of a high turbulence intensity core. After the turbulent patch has totally penetrated into the boundary layer, pronounced becalmed regions were formed behind the turbulent patch and were extended far beyond the point they would occur in the corresponding undisturbed steady boundary layer.

An experimental study of the turbulent boundary layer on a transport wing in subsonic and transonic flow

NASA Technical Reports Server (NTRS)

Spaid, Frank W.; Roos, Frederick W.; Hicks, Raymond M.

1990-01-01

The upper surface boundary layer on a transport wing model was extensively surveyed with miniature yaw probes at a subsonic and a transonic cruise condition. Additional data were obtained at a second transonic test condition, for which a separated region was present at mid-semispan, aft of mid-chord. Significant variation in flow direction with distance from the surface was observed near the trailing edge except at the wing root and tip. The data collected at the transonic cruise condition show boundary layer growth associated with shock wave/boundary layer interaction, followed by recovery of the boundary layer downstream of the shock. Measurements of fluctuating surface pressure and wingtip acceleration were also obtained. The influence of flow field unsteadiness on the boundary layer data is discussed. Comparisons among the data and predictions from a variety of computational methods are presented. The computed predictions are in reasonable agreement with the experimental data in the outboard regions where 3-D effects are moderate and adverse pressure gradients are mild. In the more highly loaded mid-span region near the trailing edge, displacement thickness growth was significantly underpredicted, except when unrealistically severe adverse pressure gradients associated with inviscid calculations were used to perform boundary layer calculations.

a Fractal Permeability Model Coupling Boundary-Layer Effect for Tight Oil Reservoirs

NASA Astrophysics Data System (ADS)

Wang, Fuyong; Liu, Zhichao; Jiao, Liang; Wang, Congle; Guo, Hu

A fractal permeability model coupling non-flowing boundary-layer effect for tight oil reservoirs was proposed. Firstly, pore structures of tight formations were characterized with fractal theory. Then, with the empirical equation of boundary-layer thickness, Hagen-Poiseuille equation and fractal theory, a fractal torturous capillary tube model coupled with boundary-layer effect was developed, and verified with experimental data. Finally, the parameters influencing effective liquid permeability were quantitatively investigated. The research results show that effective liquid permeability of tight formations is not only decided by pore structures, but also affected by boundary-layer distributions, and effective liquid permeability is the function of fluid type, fluid viscosity, pressure gradient, fractal dimension, tortuosity fractal dimension, minimum pore radius and maximum pore radius. For the tight formations dominated with nanoscale pores, boundary-layer effect can significantly reduce effective liquid permeability, especially under low pressure gradient.

A review of turbulent-boundary-layer heat transfer research at Stanford, 1958-1983

NASA Technical Reports Server (NTRS)

Moffat, R. J.; Kays, W. M.

1984-01-01

For the past 25 years, there has existed in the Thermosciences Laboratory of the Mechanical Engineering Department of Stanford University a research program, primarily experimental, concerned with heat transfer through turbulent boundary layers. In the early phases of the program, the topics considered were the simple zero-pressure-gradient turbulent boundary layer with constant and with varying surface temperature, and the accelerated boundary layer. Later equilibrium boundary layers were considered along with factors affecting the boundary layer, taking into account transpired flows, flows with axial pressure gradients, transpiration, acceleration, deceleration, roughness, full-coverage film cooling, surface curvature, free convection, and mixed convection. A description is provided of the apparatus and techniques used, giving attention to the smooth plate rig, the rough plate rig, the full-coverage film cooling rig, the curvature rig, the concave wall rig, the mixed convection tunnel, and aspects of data reduction and uncertainty analysis.

User's manual for three dimensional boundary layer (BL3-D) code

NASA Technical Reports Server (NTRS)

Anderson, O. L.; Caplin, B.

1985-01-01

An assessment has been made of the applicability of a 3-D boundary layer analysis to the calculation of heat transfer, total pressure losses, and streamline flow patterns on the surface of both stationary and rotating turbine passages. In support of this effort, an analysis has been developed to calculate a general nonorthogonal surface coordinate system for arbitrary 3-D surfaces and also to calculate the boundary layer edge conditions for compressible flow using the surface Euler equations and experimental data to calibrate the method, calculations are presented for the pressure endwall, and suction surfaces of a stationary cascade and for the pressure surface of a rotating turbine blade. The results strongly indicate that the 3-D boundary layer analysis can give good predictions of the flow field, loss, and heat transfer on the pressure, suction, and endwall surface of a gas turbine passage.

Unsteady separation and vortex shedding from a laminar separation bubble over a bluff body

NASA Astrophysics Data System (ADS)

Das, S. P.; Srinivasan, U.; Arakeri, J. H.

2013-07-01

Boundary layers are subject to favorable and adverse pressure gradients because of both the temporal and spatial components of the pressure gradient. The adverse pressure gradient may cause the flow to separate. In a closed loop unsteady tunnel we have studied the initiation of separation in unsteady flow past a constriction (bluff body) in a channel. We have proposed two important scalings for the time when boundary layer separates. One is based on the local pressure gradient and the other is a convective time scale based on boundary layer parameters. The flow visualization using a dye injection technique shows the flow structure past the body. Nondimensional shedding frequency (Strouhal number) is calculated based on boundary layer and momentum thicknesses. Strouhal number based on the momentum thickness shows a close agreement with that for flat plate and circular cylinder.

Large-eddy simulations of adverse pressure gradient turbulent boundary layers

NASA Astrophysics Data System (ADS)

Bobke, Alexandra; Vinuesa, Ricardo; Örlü, Ramis; Schlatter, Philipp

2016-04-01

Adverse pressure-gradient (APG) turbulent boundary layers (TBL) are studied by performing well-resolved large-eddy simulations. The pressure gradient is imposed by defining the free-stream velocity distribution with the description of a power law. Different inflow conditions, box sizes and upper boundary conditions are tested in order to determine the final set-up. The statistics of turbulent boundary layers with two different power-law coefficients and thus magnitudes of adverse pressure gradients are then compared to zero pressure-gradient (ZPG) data. The effect of the APG on TBLs is manifested in the mean flow through a much more prominent wake region and in the Reynolds stresses through the existence of an outer peak. The pre-multiplied energy budgets show that more energy is transported from the near-wall region to farther away from the wall.

Understanding the Fundamental Roles of Momentum and Vorticity Injections in Flow Control

DTIC Science & Technology

2016-09-02

production by pitched and skewed jets in a turbulent boundary layer . AIAA Journal 30, 640–647. DISTRIBUTION A: Distribution approved for public release...adverse pressure gradient along the suction surface, which ultimately results in a separated boundary layer . Such behavior of the boundary layer can... boundary layer either directly or by utilizing free stream momentum to energize the boundary layer (Gad-el-Hak, 2000a). Directly adding momentum to the

Shock wave oscillation driven by turbulent boundary layer fluctuations

NASA Technical Reports Server (NTRS)

Plotkin, K. J.

1972-01-01

Pressure fluctuations due to the interaction of a shock wave with a turbulent boundary layer were investigated. A simple model is proposed in which the shock wave is convected from its mean position by velocity fluctuations in the turbulent boundary layer. Displacement of the shock is assumed limited by a linear restoring mechanism. Predictions of peak root mean square pressure fluctuation and spectral density are in excellent agreement with available experimental data.

In-flight boundary-layer measurements on a hollow cylinder at a Mach number of 3.0

NASA Technical Reports Server (NTRS)

Quinn, R. D.; Gong, L.

1980-01-01

Skin temperatures, shear forces, surface static pressures, boundary layer pitot pressures, and boundary layer total temperatures were measured on the external surface of a hollow cylinder that was 3.04 meters long and 0.437 meter in diameter and was mounted beneath the fuselage of the YF-12A airplane. The data were obtained at a nominal free stream Mach number of 3.0 (a local Mach number of 2.9) and at wall to recovery temperature ratios of 0.66 to 0.91. The local Reynolds number had a nominal value of 4,300,000 per meter. Heat transfer coefficients and skin friction coefficients were derived from skin temperature time histories and shear force measurements, respectively. In addition, boundary layer velocity profiles were derived from pitot pressure measurements, and a Reynolds analogy factor was obtained from the heat transfer and skin friction measurements. The measured data are compared with several boundary layer prediction methods.

Energy efficient engine, low-pressure turbine boundary layer program

NASA Technical Reports Server (NTRS)

Gardner, W. B.

1981-01-01

A study was conducted to investigate development of boundary layers under the influence of velocity distributions simulating the suction side of two state-of-the-art turbine airfoils: a forward loaded airfoil (squared-off design) and an aft loaded airfoil (aft-loaded design). These velocity distributions were simulated in a boundary layer wind tunnel. Detailed measurements of boundary layer mean velocity and turbulence intensity profiles were obtained for an inlet turbulence level of 2.4 percent and an exit Reynolds number of 800,000. Flush-mounted hot film probes identified the boundary layer transition regimes in the adverse pressure gradient regions for both velocity distributions. Wall intermittency data showed good agreement with the correlations of Dhawan and Narasimha for the intermittency factor distribution in transitional flow regimes.

F-16XL ship #1 - CAWAP boundary layer rakes and hot film on left wing

NASA Technical Reports Server (NTRS)

1996-01-01

This photo shows the boundary layer hot film and the boundary layer rakes on the left wing of NASA's single-seat F-16XL (ship #1) used for the Cranked-Arrow Wing Aerodynamic Project (CAWAP) at Dryden Flight Research Center, Edwards, California. The modified airplane features a delta 'cranked-arrow' wing with strips of tubing along the leading edge to the trailing edge to sense static on the wing and obtain pressure distribution data. The right wing receives data on pressure distribution and the left wing has three types of instrumentation - preston tubes to measure local skin friction, boundary layer rakes to measure boundary layer profiles (the layer where the air interacts with the surfaces of a moving aircraft), and hot films to determine boundary layer transition locations. The program also gathered aero data on two wing planforms for NASA's High Speed Research Program. The first flight of CAWAP occurred on November 21, 1995, and the test program ended in April 1996.

Correlation of heat transfer for the zero pressure gradient hypersonic laminar boundary layer for several gases

NASA Technical Reports Server (NTRS)

Cook, W. J.

1973-01-01

A theoretical study of heat transfer for zero pressure gradient hypersonic laminar boundary layers for various gases with particular application to the flows produced in an expansion tube facility was conducted. A correlation based on results obtained from solutions to the governing equations for five gases was formulated. Particular attention was directed toward the laminar boundary layer shock tube splitter plates in carbon dioxide flows generated by high speed shock waves. Computer analysis of the splitter plate boundary layer flow provided information that is useful in interpreting experimental data obtained in shock tube gas radiation studies.

Turbulent boundary layers subjected to multiple curvatures and pressure gradients

NASA Technical Reports Server (NTRS)

Bandyopadhyay, Promode R.; Ahmed, Anwar

1993-01-01

The effects of abruptly applied cycles of curvatures and pressure gradients on turbulent boundary layers are examined experimentally. Two two-dimensional curved test surfaces are considered: one has a sequence of concave and convex longitudinal surface curvatures and the other has a sequence of convex and concave curvatures. The choice of the curvature sequences were motivated by a desire to study the asymmetric response of turbulent boundary layers to convex and concave curvatures. The relaxation of a boundary layer from the effects of these two opposite sequences has been compared. The effect of the accompaying sequences of pressure gradient has also been examined but the effect of curvature dominates. The growth of internal layers at the curvature junctions have been studied. Measurements of the Gortler and corner vortex systems have been made. The boundary layer recovering from the sequence of concave to convex curvature has a sustained lower skin friction level than in that recovering from the sequence of convex to concave curvature. The amplification and suppression of turbulence due to the curvature sequences have also been studied.

Boundary-Layer Bypass Transition Over Large-Scale Bodies

DTIC Science & Technology

2016-12-16

shape of the streamwise velocity profile compared to the flat- plate boundary layer. The research showed that the streamwise wavenumber plays a key role...many works on the suppression of the transitional boundary layer. Most of the results in the literature are for the flat- plate boundary layer but the...behaviour of the velocity and pressure changes with the curvature. This work aims to extend the results of the flat- plate boundary layer to a Rankine

Differential analysis for the turbulent boundary layer on a compressor blade element (including boundary-layer separation)

NASA Technical Reports Server (NTRS)

Schmidt, J. F.; Todd, C. A.

1974-01-01

A two-dimensional differential analysis is developed to approximate the turbulent boundary layer on a compressor blade element with strong adverse pressure gradients, including the separated region with reverse flow. The predicted turbulent boundary layer thicknesses and velocity profiles are in good agreement with experimental data for a cascade blade, even in the separated region.

The turbulent plasmasphere boundary layer and the outer radiation belt boundary

NASA Astrophysics Data System (ADS)

Mishin, Evgeny; Sotnikov, Vladimir

2017-12-01

We report on observations of enhanced plasma turbulence and hot particle distributions in the plasmasphere boundary layer formed by reconnection-injected hot plasma jets entering the plasmasphere. The data confirm that the electron pressure peak is formed just outward of the plasmapause in the premidnight sector. Free energy for plasma wave excitation comes from diamagnetic ion currents near the inner edge of the boundary layer due to the ion pressure gradient, electron diamagnetic currents in the entry layer near the electron plasma sheet boundary, and anisotropic (sometimes ring-like) ion distributions revealed inside, and further inward of, the inner boundary. We also show that nonlinear parametric coupling between lower oblique resonance and fast magnetosonic waves significantly contributes to the VLF whistler wave spectrum in the plasmasphere boundary layer. These emissions represent a distinctive subset of substorm/storm-related VLF activity in the region devoid of substorm injected tens keV electrons and could be responsible for the alteration of the outer radiation belt boundary during (sub)storms.

Optimal disturbances in boundary layers subject to streamwise pressure gradient

NASA Technical Reports Server (NTRS)

Ashpis, David E.; Tumin, Anatoli

2003-01-01

An analysis of the optimal non-modal growth of perturbations in a boundary layer in the presence of a streamwise pressure gradient is presented. The analysis is based on PSE equations for an incompressible fluid. Examples with Falkner-Scan profiles indicate that a favorable pressure gradient decreases the non-modal growth, while an unfavorable pressure gradient leads to an increase of the amplification. It is suggested that the transient growth mechanism be utilized to choose optimal parameters of tripping elements on a low-pressure turbine (LPT) airfoil. As an example, a boundary layer flow with a streamwise pressure gradient corresponding to the pressure distribution over a LPT airfoil is considered. It is shown that there is an optimal spacing of the tripping elements and that the transient growth effect depends on the starting point.

Characteristics of Boundary Layer Transition in a Multi-Stage Low-Pressure Turbine

NASA Technical Reports Server (NTRS)

Wisler, Dave; Halstead, David E.; Okiishi, Ted

2007-01-01

An experimental investigation of boundary layer transition in a multi-stage turbine has been completed using surface-mounted hot-film sensors. Tests were carried out using the two-stage Low Speed Research Turbine of the Aerodynamics Research Laboratory of GE Aircraft Engines. Blading in this facility models current, state-of-the-art low pressure turbine configurations. The instrumentation technique involved arrays of densely-packed hot-film sensors on the surfaces of second stage rotor and nozzle blades. The arrays were located at mid-span on both the suction and pressure surfaces. Boundary layer measurements were acquired over a complete range of relevant Reynolds numbers. Data acquisition capabilities provided means for detailed data interrogation in both time and frequency domains. Data indicate that significant regions of laminar and transitional boundary layer flow exist on the rotor and nozzle suction surfaces. Evidence of relaminarization both near the leading edge of the suction surface and along much of the pressure surface was observed. Measurements also reveal the nature of the turbulent bursts occuring within and between the wake segments convecting through the blade row. The complex character of boundary layer transition resulting from flow unsteadiness due to nozzle/nozzle, rotor/nozzle, and nozzle/rotor wake interactions are elucidated using these data. These measurements underscore the need to provide turbomachinery designers with models of boundary layer transition to facilitate accurate prediction of aerodynamic loss and heat transfer.

Solution of the surface Euler equations for accurate three-dimensional boundary-layer analysis of aerodynamic configurations

NASA Technical Reports Server (NTRS)

Iyer, V.; Harris, J. E.

1987-01-01

The three-dimensional boundary-layer equations in the limit as the normal coordinate tends to infinity are called the surface Euler equations. The present paper describes an accurate method for generating edge conditions for three-dimensional boundary-layer codes using these equations. The inviscid pressure distribution is first interpolated to the boundary-layer grid. The surface Euler equations are then solved with this pressure field and a prescribed set of initial and boundary conditions to yield the velocities along the two surface coordinate directions. Results for typical wing and fuselage geometries are presented. The smoothness and accuracy of the edge conditions obtained are found to be superior to the conventional interpolation procedures.

F-16XL ship #1 wing close-up showing boundary layer detection Preston tubes

NASA Technical Reports Server (NTRS)

1995-01-01

This photo shows the boundary layer Preston tubes mounted on the left wing of NASA's single-seat F-16XL (ship #1) used for the Cranked-Arrow Wing Aerodynamic Project (CAWAP) at Dryden Flight Research Center, Edwards, California. The modified airplane features a delta 'cranked-arrow' wing with strips of tubing along the leading edge to the trailing edge to sense static on the wing and obtain pressure distribution data. The right wing receives data on pressure distribution and the left wing has three types of instrumentation - preston tubes to measure local skin friction, boundary layer rakes to measure boundary layer profiles (the layer where the air interacts with the surfaces of a moving aircraft), and hot films to determine boundary layer transition locations. The first flight of CAWAP occurred on November 21, 1995, and the test program ended in April 1996.

Some Simple Solutions to the Problem of Predicting Boundary-Layer Self-Induced Pressures

NASA Technical Reports Server (NTRS)

Bertram, Mitchel H.; Blackstock, Thomas A.

1961-01-01

Simplified theoretical approaches are shown, based on hypersonic similarity boundary-layer theory, which allow reasonably accurate estimates to be made of the surface pressures on plates on which viscous effects are important. The consideration of viscous effects includes the cases where curved surfaces, stream pressure gradients, and leadingedge bluntness are important factors.

The turbulent boundary layer on a porous plate: An experimental study of the fluid mechanics for adverse free stream pressure gradients

NASA Technical Reports Server (NTRS)

Anderson, P. S.; Kays, W. M.; Moffat, R. J.

1972-01-01

An experimental investigation of transpired turbulent boundary layers in zero and adverse pressure gradients has been carried out. Profiles of: (1) the mean velocity, (2) the three intensities of the turbulent fluctuations, and (3) the Reynolds stress were obtained by hot-wire anemometry. The friction coefficients were measured by using an integrated form of the boundary layer equation to extrapolate the measured shear stress profiles to the wall.

Stereo particle image velocimetry of nonequilibrium turbulence relaxation in a supersonic boundary layer

NASA Astrophysics Data System (ADS)

Lapsa, Andrew P.; Dahm, Werner J. A.

2011-01-01

Measurements using stereo particle image velocimetry are presented for a developing turbulent boundary layer in a wind tunnel with a Mach 2.75 free stream. As the boundary layer exits from the tunnel nozzle and moves through the wave-free test section, small initial departures from equilibrium turbulence relax, and the boundary layer develops toward the equilibrium zero-pressure-gradient form. This relaxation process is quantified by comparison of first and second order mean, fluctuation, and gradient statistics to classical inner and outer layer scalings. Simultaneous measurement of all three instantaneous velocity components enables direct assessment of the complete turbulence anisotropy tensor. Profiles of the turbulence Mach number show that, despite the M = 2.75 free stream, the incompressibility relation among spatial gradients in the velocity fluctuations applies. This result is used in constructing various estimates of the measured-dissipation rate, comparisons among which show only remarkably small differences over most of the boundary layer. The resulting measured-dissipation profiles, together with measured profiles of the turbulence kinetic energy and mean-flow gradients, enable an assessment of how the turbulence anisotropy relaxes toward its equilibrium zero-pressure-gradient state. The results suggest that the relaxation of the initially disturbed turbulence anisotropy profile toward its equilibrium zero-pressure-gradient form begins near the upper edge of the boundary layer and propagates downward through the defect layer.

Measurement of density and temperature in a hypersonic turbulent boundary layer using the electron beam fluorescence technique. Ph.D. Thesis. Final Report, 1 Oct. 1969 - 1 Sep. 1972

NASA Technical Reports Server (NTRS)

Mcronald, A. D.

1975-01-01

Mean density and temperature fluctuations were measured across the turbulent, cooled-wall boundary layer in a continuous hypersonic (Mach 9.4) wind tunnel in air, using the nitrogen fluorescence excited by a 50 kV electron beam. Data were taken at three values of the tunnel stagnation pressure, the corresponding free stream densities being equivalent to 1.2, 4.0, and 7.4 torr at room temperature, and the boundary layer thicknesses about 4.0, 4.5, and 6.0 inches. The mean temperature and density profiles were similar to those previously determined in the same facility by conventional probes (static and pitot pressure, total temperature). A static pressure variation of about 50% across the boundary layer was found, the shape of the variation changing somewhat for the three stagnation pressure levels. The quadrupole model for rotational temperature spectra gave closer agreement with the free stream isentropic level (approximately 44 K) than the dipole model.

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

NASA Technical Reports Server (NTRS)

Vijgen, P. M. H. W.; Hardin, J. D.; Yip, L. P.

1992-01-01

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.

Vortex detection through pressure measurements

NASA Astrophysics Data System (ADS)

Bhide, Aditi

Vortex Generators (VGs) are known to hinder boundary layer separation, a frequently unwanted phenomenon when it comes to external flows over aircraft wings, on-ground vehicles or internal flows within pipes, diffusers and turbomachinery. Boundary layer separation leads to loss of lift, higher drag and subsequently, energy losses. The vortices generated inhibit boundary layer separation. This thesis is an effort to discern the strength and location of these generated vortices using an array of VGs over a flat plate. Such information may be useful in the future in active control systems for streamwise vortices, which have been proposed to relaminarize turbulent boundary layers. Flow over flat plates, simulated using wind tunnel experiments, is studied for pressure variation using an array of pressure ports mounted over the plate and connected to suitable pressure sensors. Pressure coefficient and Velocity maps are generated using the data obtained from the Kirsten Wind Tunnel data acquisition system. These represent the nature of the flow field over the plate and are used to locate the vortices and determine their strength. It was found that the vortices can be detected using this method and their strength and location can be estimated.

Evaluation of Flush-Mounted, S-Duct Inlets With Large Amounts of Boundary Layer Ingestion

NASA Technical Reports Server (NTRS)

Berrier, Bobby L.; Morehouse, Melissa B.

2003-01-01

A new high Reynolds number test capability for boundary layer ingesting inlets has been developed for the NASA Langley Research Center 0.3-Meter Transonic Cryogenic Tunnel. Using this new capability, an experimental investigation of four S-duct inlet configurations with large amounts of boundary layer ingestion (nominal boundary layer thickness of about 40% of inlet height) was conducted at realistic operating conditions (high subsonic Mach numbers and full-scale Reynolds numbers). The objectives of this investigation were to 1) develop a new high Reynolds number, boundary-layer ingesting inlet test capability, 2) evaluate the performance of several boundary layer ingesting S-duct inlets, 3) provide a database for CFD tool validation, and 4) provide a baseline inlet for future inlet flow-control studies. Tests were conducted at Mach numbers from 0.25 to 0.83, Reynolds numbers (based on duct exit diameter) from 5.1 million to a fullscale value of 13.9 million, and inlet mass-flow ratios from 0.39 to 1.58 depending on Mach number. Results of this investigation indicate that inlet pressure recovery generally decreased and inlet distortion generally increased with increasing Mach number. Except at low Mach numbers, increasing inlet mass-flow increased pressure recovery and increased distortion. Increasing the amount of boundary layer ingestion (by decreasing inlet throat height and increasing inlet throat width) or ingesting a boundary layer with a distorted profile decreased pressure recovery and increased distortion. Finally, increasing Reynolds number had almost no effect on inlet distortion but increased inlet recovery by about one-half percent at a Mach number near cruise.

Experimental investigation of a two-dimensional shock-turbulent boundary layer interaction with bleed

NASA Technical Reports Server (NTRS)

Hingst, W. R.; Tanji, F. T.

1983-01-01

The two-dimensional interaction of an oblique shock wave with a turbulent boundary layer that included the effect of bleed was examined experimentally using a shock generator mounted across a supersonic wind tunnel The studies were performed at Mach numbers 2.5 and 2.0 and unit Reynolds number of approximately 2.0 x 10 to the 7th/meter. The study includes surface oil flow visualization, wall static pressure distributions and boundary layer pitot pressure profiles. In addition, the variation of the local bleed rates were measured. The results show the effect of the bleed on the boundary layer as well as the effect of the flow conditions on the local bleed rate.

Deplacement effect of the laminar boundary layer and the pressure drag

NASA Technical Reports Server (NTRS)

Gortler, H

1951-01-01

The displacement effect of the boundary layer on the outer frictionless flow is discussed for both steady and unsteady flows. The analysis is restricted to cases in which the potential flow pressure distribution remains valid for the boundary-layer calculation. Formulas are given for the dependence of the pressure drag, friction drag, and total drag of circular cylinders on the time from the start of motion for cases in which the velocity varies as a power of the time. Formulas for the locations and for the time for the appearance of the separation point are given for two dimensional bodies of arbitrary shape.

A comparison of measured and predicted test flow in an expansion tube with air and oxygen test gases

NASA Technical Reports Server (NTRS)

Aaggard, K. V.; Goad, W. K.

1975-01-01

Simultaneous time-resolved measurements of temperature, density, pitot pressure, and wall pressure in both air and O2 test gases were obtained in the Langley pilot model expansion tube. These tests show nonequilibrium chemical and vibrational relaxation significantly affect the test-flow condition. The use of an electromagnetic device to preopen the secondary diaphragm before the arrival of the primary shock wave resulted in an improvement in the agreement between the measured pitot pressure and the value inferred from measured density and interface velocity. Boundary-layer splitter plates used to reduce the wall boundary layer show that this disagreement in the measured and inferred pitot pressures is not a result of boundary-layer effects.

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

NASA Technical Reports Server (NTRS)

Benhachmi, Driss; Greber, Isaac; Hingst, Warren R.

1988-01-01

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.

Effects of forebody geometry on subsonic boundary-layer stability

NASA Technical Reports Server (NTRS)

Dodbele, Simha S.

1990-01-01

As part of an effort to develop computational techniques for design of natural laminar flow fuselages, a computational study was made of the effect of forebody geometry on laminar boundary layer stability on axisymmetric body shapes. The effects of nose radius on the stability of the incompressible laminar boundary layer was computationally investigated using linear stability theory for body length Reynolds numbers representative of small and medium-sized airplanes. The steepness of the pressure gradient and the value of the minimum pressure (both functions of fineness ratio) govern the stability of laminar flow possible on an axisymmetric body at a given Reynolds number. It was found that to keep the laminar boundary layer stable for extended lengths, it is important to have a small nose radius. However, nose shapes with extremely small nose radii produce large pressure peaks at off-design angles of attack and can produce vortices which would adversely affect transition.

A theoretical and flight test study of pressure fluctuations under a turbulent boundary layer. Part 2: Flight test study

NASA Technical Reports Server (NTRS)

Panton, R. L.; Lowery, R. L.; Reischman, M. M.

1967-01-01

The study of pressure fluctuations under a turbulent boundary layer was undertaken with the objective of extending previous work to lower frequencies. Wind tunnel and flight test measurements are invalid at low frequencies because of extraneous acoustic noises and free stream turbulence. A glider was instrumented and used as a test bed to carry microphones into a smooth flow free of acoustic noise. Hodgson had previously measured the spectrum of boundary layer noise on a glider wing. These tests showed a drop off at low frequencies that could not be reproduced in any other facility. The measurements were made on the forward fuselage of a glider where the boundary layer could develop naturally and have some length in a zero pressure gradient before the measurements were made. Two different sets of measurements were made.

Experimental investigation on the influence of boundary layer thickness on the base pressure and near-wake flow features of an axisymmetric blunt-based body

NASA Astrophysics Data System (ADS)

Mariotti, Alessandro; Buresti, Guido

2013-11-01

The influence of the thickness of the boundary layer developing over the surface of an axisymmetric bluff body upon its base pressure and near-wake flow is analyzed experimentally. The model, whose diameter-to-length ratio is d/ l = 0.175, has a forebody with an elliptical contour and a sharp-edged flat base; it is supported above a plate by means of a faired strut. The pressure distributions over the body lateral and base surfaces were obtained using numerous pressure taps, while the boundary layer profiles and the wake velocity field were measured through hot-wire anemometry. The tests were carried out at , at which the boundary layer over the lateral surface of the body becomes turbulent before reaching the base contour. Strips of emery cloth were wrapped in various positions around the body circumference in order to modify the thickness and the characteristics of the boundary layer. The results show that increasing the boundary layer thickness causes a decrease in the base suctions and a corresponding increase in the length of the mean recirculation region present behind the body. In the spectra of the velocity fluctuations measured within and aside the wake, a dominating peak becomes evident in the region downstream of the final part of the recirculation region. The relevant non-dimensional frequency decreases with increasing boundary layer thickness; however, a Strouhal number based on the wake width and the velocity defect at a suitable reference cross section downstream of the recirculation region is found to remain almost constant for the different cases.

An experimental investigation of turbulent boundary layers along curved surfaces

NASA Technical Reports Server (NTRS)

So, R. M. C.; Mellor, G. L.

1972-01-01

A curved wall tunnel was designed, and an equilibrium turbulent boundary layer was set up on the straight section preceding the curved test section. Turbulent boundary layer flows with uniform and adverse pressure distributions along convex and concave walls were investigated. Hot-wire measurements along the convex surface indicated that turbulent mixing between fluid layers was very much reduced. However, the law of the wall held and the skin friction, thus determined, correlated well with other measurements. Hot-wire measurements along the concave test wall revealed a system of longitudinal vortices inside the boundary layer and confirmed that concave curvature enhances mixing. A self-consistent set of turbulent boundary layer equations for flows along curved surfaces was derived together with a modified eddy viscosity. Solution of these equations together with the modified eddy viscosity gave results that correlated well with the present data on flows along the convex surface with arbitrary pressure distribution. However, it could only be used to predict the mean characteristics of the flow along concave walls because of the existence of the system of longitudinal vortices inside the boundary layer.

Effect of Pulsed Plasma Jets on the Recovering Boundary Layer Downstream of a Reflected Shock Interaction

NASA Astrophysics Data System (ADS)

Greene, Benton; Clemens, Noel; Magari, Patrick; Micka, Daniel; Ueckermann, Mattheus

2015-11-01

Shock-induced turbulent boundary layer separation can have many detrimental effects in supersonic inlets including flow distortion and instability, structural fatigue, poor pressure recovery, and unstart. The current study investigates the effect of pulsed plasma jets on the recovering boundary layer downstream of a reflected shock wave-boundary layer interaction. The effects of pitch and skew angle of the jet as well as the heating parameter and discharge time scale are tested using several pulsing frequencies. In addition, the effect of the plasma jets on the undisturbed boundary layer at 6 mm and 11 mm downstream of the jets is measured. A pitot-static pressure probe is used to measure the velocity profile of the boundary layer 35 mm downstream of the plasma jets, and the degree of boundary layer distortion is compared between the different models and run conditions. Additionally, the effect of each actuator configuration on the shape of the mean separated region is investigated using surface oil flow visualization. Previous studies with lower energy showed a weak effect on the downstream boundary layer. The current investigation will attempt to increase this effect using a higher-energy discharge. Funded by AFRL through and SBIR in collaboration with Creare, LLC.

The turbulent boundary layer on a porous plate: An experimental study of the heat transfer behavior with adverse pressure gradients

NASA Technical Reports Server (NTRS)

Blackwell, B. F.; Kays, W. M.; Moffat, R. J.

1972-01-01

An experimental investigation of the heat transfer behavior of the near equilibrium transpired turbulent boundary layer with adverse pressure gradient has been carried out. Stanton numbers were measured by an energy balance on electrically heated plates that form the bottom wall of the wind tunnel. Two adverse pressure gradients were studied. Two types of transpiration boundary conditions were investigated. The concept of an equilibrium thermal boundary layer was introduced. It was found that Stanton number as a function of enthalpy thickness Reynolds number is essentially unaffected by adverse pressure gradient with no transpiration. Shear stress, heat flux, and turbulent Prandtl number profiles were computed from mean temperature and velocity profiles. It was concluded that the turbulent Prandtl number is greater than unity in near the wall and decreases continuously to approximately 0.5 at the free stream.

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

NASA Technical Reports Server (NTRS)

Von Doenhoff, Albert E

1938-01-01

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.

Flight-measured base pressure coefficients for thick boundary-layer flow over an aft-facing step for Mach numbers from 0.4 to 2.5

NASA Technical Reports Server (NTRS)

Goecke, S. A.

1973-01-01

A 0.56-inch thick aft-facing step was located 52.1 feet from the leading edge of the left wing of an XB-70 airplane. A boundary-layer rake at a mirror location on the right wing was used to obtain local flow properties. Reynolds numbers were near 10 to the 8th power, resulting in a relatively thick boundary-layer. The momentum thickness ranged from slightly thinner to slightly thicker than the step height. Surface static pressures forward of the step were obtained for Mach numbers near 0.9, 1.5, 2.0, and 2.4. The data were compared with thin boundary-layer results from flight and wind-tunnel experiments and semiempirical relationships. Significant differences were found between the thick and the thin boundary-layer data.

Control of shock wave-boundary layer interactions by bleed in supersonic mixed compression inlets

NASA Technical Reports Server (NTRS)

Fukuda, M. K.; Hingst, W. G.; Reshotko, E.

1975-01-01

An experimental investigation was conducted to determine the effect of bleed on a shock wave-boundary layer interaction in an axisymmetric mixed-compression supersonic inlet. The inlet was designed for a free-stream Mach number of 2.50 with 60-percent supersonic internal area contraction. The experiment was conducted in the NASA Lewis Research Center 10-Foot Supersonic Wind Tunnel. The effects of bleed amount and bleed geometry on the boundary layer after a shock wave-boundary layer interaction were studied. The effect of bleed on the transformed form factor is such that the full realizable reduction is obtained by bleeding of a mass flow equal to about one-half of the incident boundary layer mass flow. More bleeding does not yield further reduction. Bleeding upstream or downstream of the shock-induced pressure rise is preferable to bleeding across the shock-induced pressure rise.

Evaluation of Flush-Mounted, S-Duct Inlets with Large Amounts of Boundary Layer Ingestion

NASA Technical Reports Server (NTRS)

Berrier, Bobby L.; Morehouse, Melissa B.

2003-01-01

A new high Reynolds number test capability for boundary layer ingesting inlets has been developed for the NASA Langley Research Center 0.3-Meter Transonic Cryogenic Tunnel. Using this new capability, an experimental investigation of four S-duct inlet configurations with large amounts of boundary layer ingestion (nominal boundary layer thickness of about 40% of inlet height) was conducted at realistic operating conditions (high subsonic Mach numbers and full-scale Reynolds numbers). The objectives of this investigation were to 1) provide a database for CFD tool validation on boundary layer ingesting inlets operating at realistic conditions and 2) provide a baseline inlet for future inlet flow-control studies. Tests were conducted at Mach numbers from 0.25 to 0.83, Reynolds numbers (based on duct exit diameter) from 5.1 million to a full-scale value of 13.9 million, and inlet mass-flow ratios from 0.39 to 1.58 depending on Mach number. Results of this investigation indicate that inlet pressure recovery generally decreased and inlet distortion generally increased with increasing Mach number. Except at low Mach numbers, increasing inlet mass-flow increased pressure recovery and increased distortion. Increasing the amount of boundary layer ingestion (by decreasing inlet throat height) or ingesting a boundary layer with a distorted (adverse) profile decreased pressure recovery and increased distortion. Finally, increasing Reynolds number had almost no effect on inlet distortion but increased inlet recovery by about one-half percent at a Mach number near cruise.

Detection of boundary-layer transitions in wind tunnels

NASA Technical Reports Server (NTRS)

Wood, W. R.; Somers, D. M.

1978-01-01

Accelerometer replaces stethoscope in technique for detection of laminar-to-turbulent boundary-layer transitions on wind-tunnel models. Technique allows measurements above or below atmospheric pressure because human operator is not required within tunnel. Data may be taken from accelerometer, and pressure transducer simultaneously, and delivered to systems for analysis.

Charts and Tables for Estimating the Stability of the Compressible Laminar Boundary Layer with Heat Transfer and Arbitrary Pressure Gradient

NASA Technical Reports Server (NTRS)

Tetervin, Neal

1959-01-01

The minimum critical Reynolds numbers for the similar solutions of the compressible laminar boundary layer computed by Cohen and Reshotko and also for the Falkner and Skan solutions as recomputed by Smith have been calculated by Lin's rapid approximate method for two-dimensional disturbances. These results enable the stability of the compressible laminar boundary layer with heat transfer and pressure gradient to be easily estimated after the behavior of the boundary layer has been computed by the approximate method of Cohen and Reshotko. The previously reported unusual result (NACA Technical Note 4037) that a highly cooled stagnation point flow is more unstable than a highly cooled flat-plate flow is again encountered. Moreover, this result is found to be part of the more general result that a favorable pressure gradient is destabilizing for very cool walls when the Mach number is less than that for complete stability. The minimum critical Reynolds numbers for these wall temperature ratios are, however, all larger than any value of the laminar-boundary-layer Reynolds number likely to be encountered. For Mach numbers greater than those for which complete stability occurs a favorable pressure gradient is stabilizing, even for very cool walls.

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

NASA Technical Reports Server (NTRS)

Galmes, J. M.; Pouagere, M.; Lakshminarayana, B.

1982-01-01

The Reynolds stress equation, pressure strain correlation, and dissipative terms and diffusion are discussed in relation to turbulence modelling using the Reynolds stress model. Algebraic modeling of Reynolds stresses and calculation of the boundary layer over an axial cylinder are examined with regards to the kinetic energy model for turbulence modelling. The numerical analysis of blade and hub wall boundary layers, and an experimental study of rotor blade boundary layer in an axial flow compressor rotor are discussed. The Patankar-Spalding numerical method for two dimensional boundary layers is included.

Numerical investigation of the boundary layer separation in chemical oxygen iodine laser

NASA Astrophysics Data System (ADS)

Huai, Ying; Jia, Shuqin; Wu, Kenan; Jin, Yuqi; Sang, Fengting

2017-11-01

Large eddy simulation is carried out to model the flow process in a supersonic chemical oxygen iodine laser. Unlike the common approaches relying on the tensor representation theory only, the model in the present work is an explicit anisotropy-resolving algebraic Subgrid-scale scalar flux formulation. With an accuracy in capturing the unsteady flow behaviours in the laser. Boundary layer separation initiated by the adverse pressure gradient is identified using Large Eddy Simulation. To quantify the influences of flow boundary layer on the laser performance, the fluid computations coupled with a physical optics loaded cavity model is developed. It has been found that boundary layer separation has a profound effect on the laser outputs due to the introduced shock waves. The F factor of the output beam decreases to 10% of the original one when the boundary transit into turbulence for the setup depicted in the paper. Because the pressure is always greater on the downstream of the boundary layer, there will always be a tendency of boundary separation in the laser. The results inspire designs of the laser to apply positive/passive control methods avoiding the boundary layer perturbation.

The influence of free-stream turbulence on separation of turbulent boundary layers in incompressible, two-dimensional flow

NASA Technical Reports Server (NTRS)

Potter, J. Leith; Barnett, R. Joel; Fisher, Carl E.; Koukousakis, Costas E.

1986-01-01

Experiments were conducted to determine if free-stream turbulence scale affects separation of turbulent boundary layers. In consideration of possible interrelation between scale and intensity of turbulence, the latter characteristic also was varied and its role was evaluated. Flow over a 2-dimensional airfoil in a subsonic wind tunnel was studied with the aid of hot-wire anemometry, liquid-film flow visualization, a Preston tube, and static pressure measurements. Profiles of velocity, relative turbulence intensity, and integral scale in the boundary layer were measured. Detachment boundary was determined for various angles of attack and free-stream turbulence. The free-stream turbulence intensity and scale were found to spread into the entire turbulent boundary layer, but the effect decreased as the airfoil surface was approached. When the changes in stream turbulence were such that the boundary layer velocity profiles were unchanged, detachment location was not significantly affected by the variations of intensity and scale. Pressure distribution remained the key factor in determining detachment location.

Secondary subharmonic instability of boundary layers with pressure gradient and suction

NASA Technical Reports Server (NTRS)

El-Hady, Nabil M.

1988-01-01

Three-dimensional linear secondary instability is investigated for boundary layers with pressure gradient and suction in the presence of a finite amplitude TS wave. The focus is on principal parametric resonance responsible for a strong growth of subharmonics in a low disturbance environment. Calculations are presented for the effect of pressure gradients and suction on controlling the onset and amplification of the secondary instability.

Experimental Measurements of a High Reynolds Num- ber Adverse Pressure Gradient Turbulent Boundary Layer

NASA Astrophysics Data System (ADS)

Atkinson, Callum; Amili, Omid; Stanislas, Michel; Cuvier, Christophe; Foucaut, Jean-Marc; Srinath, Sricharan; Laval, Jean-Philippe; Kaehler, Christian; Hain, Rainer; Scharnowski, Sven; Schroeder, Andreas; Geisler, Reinhard; Agocs, Janos; Roese, Anni; Willert, Christian; Klinner, Joachim; Soria, Julio

2016-11-01

The study of adverse pressure gradient turbulent boundary layers is complicated by the need to characterise both the local pressure gradient and it's upstream flow history. It is therefore necessary to measure a significant streamwise domain at a resolution sufficient to resolve the small scales features. To achieve this collaborative particle image velocimetry (PIV) measurements were performed in the large boundary layer wind-tunnel at the Laboratoire de Mecanique de Lille, including: planar measurements spanning a streamwise domain of 3.5m using 16 cameras covering 15 δ spanwise wall-normal stereo-PIV measurements, high-speed micro-PIV of the near wall region and wall shear stress; and streamwise wall-normal PIV in the viscous sub layer. Details of the measurements and preliminary results will be presented.

F-16XL ship #1 - CAWAP boundary layer hot film, left wing

NASA Technical Reports Server (NTRS)

1996-01-01

This photo shows the boundary layer hot film on the left wing of NASA's single-seat F-16XL (ship #1) used for the Cranked-Arrow Wing Aerodynamic Project (CAWAP) at Dryden Flight Research Center, Edwards, California. Hot film is used to measure temperature changes on a surface. The modified airplane features a delta 'cranked-arrow' wing with strips of tubing along the leading edge to the trailing edge to sense static on the wing and obtain pressure distribution data. The right wing receives data on pressure distribution and the left wing has three types of instrumentation - preston tubes to measure local skin friction, boundary layer rakes to measure boundary layer profiles (the layer where the air interacts with the surfaces of a moving aircraft), and hot films to determine boundary layer transition locations. The program also gathered aero data on two wing planforms for NASA's High Speed Research Program. The first flight of CAWAP occurred on November 21, 1995, and the test program ended in April 1996.

Schlieren and OH* chemiluminescence imaging of combustion in a turbulent boundary layer over a solid fuel

NASA Astrophysics Data System (ADS)

Jens, Elizabeth T.; Miller, Victor A.; Cantwell, Brian J.

2016-03-01

Combustion in a turbulent boundary layer over a solid fuel is studied using simultaneous schlieren and OH* chemiluminescence imaging. The flow configuration is representative of a hybrid rocket motor combustor. Six different hydrocarbon fuels, including both classical hybrid rocket fuels and a high regression rate fuel (paraffin wax), are burned in an undiluted oxygen free-stream at pressures ranging from atmospheric to 1524.2 kPa (221.1 psi). A detailed explanation of methods for registering the schlieren and OH* chemiluminescence images to one another is presented, and additionally, details of the routines used to extract flow features of interest (like the boundary layer height and flame location) are provided. At atmospheric pressure, the boundary layer location is consistent between all fuels; however, the flame location varies for each fuel. The flame zone appears to be smoothly distributed over the fuel surface at atmospheric pressure. At elevated pressures and correspondingly increased Dahmköhler number (but at constant Reynolds number), flame morphology is markedly different, exhibiting large rollers in a shear layer above the fuel grain and finer structures in the flame. The chemiluminescence intensity is found to be roughly proportional to the fuel burn rate at both atmospheric and elevated chamber pressures.

Stability of spatially developing boundary layers

NASA Astrophysics Data System (ADS)

Govindarajan, Rama

1993-07-01

A new formulation of the stability of boundary-layer flows in pressure gradients is presented, taking into account the spatial development of the flow. The formulation assumes that disturbance wavelength and eigenfunction vary downstream no more rapidly than the boundary-layer thickness, and includes all terms of O(1) and O(R(exp -1)) in the boundary-layer Reynolds number R. Although containing the Orr-Sommerfeld operator, the present approach does not yield the Orr-Sommerfeld equation in any rational limit. In Blasius flow, the present stability equation is consistent with that of Bertolotti et al. (1992) to terms of O(R(exp -1)). For the Falkner-Skan similarity solutions neutral boundaries are computed without the necessity of having to march in space. Results show that the effects of spatial growth are striking in flows subjected to adverse pressure gradients.

A Real-Time Method for Estimating Viscous Forebody Drag Coefficients

NASA Technical Reports Server (NTRS)

Whitmore, Stephen A.; Hurtado, Marco; Rivera, Jose; Naughton, Jonathan W.

2000-01-01

This paper develops a real-time method based on the law of the wake for estimating forebody skin-friction coefficients. The incompressible law-of-the-wake equations are numerically integrated across the boundary layer depth to develop an engineering model that relates longitudinally averaged skin-friction coefficients to local boundary layer thickness. Solutions applicable to smooth surfaces with pressure gradients and rough surfaces with negligible pressure gradients are presented. Model accuracy is evaluated by comparing model predictions with previously measured flight data. This integral law procedure is beneficial in that skin-friction coefficients can be indirectly evaluated in real-time using a single boundary layer height measurement. In this concept a reference pitot probe is inserted into the flow, well above the anticipated maximum thickness of the local boundary layer. Another probe is servomechanism-driven and floats within the boundary layer. A controller regulates the position of the floating probe. The measured servomechanism position of this second probe provides an indirect measurement of both local and longitudinally averaged skin friction. Simulation results showing the performance of the control law for a noisy boundary layer are then presented.

Acoustic Radiation From a Mach 14 Turbulent Boundary Layer

NASA Technical Reports Server (NTRS)

Zhang, Chao; Duan, Lian; Choudhari, Meelan M.

2016-01-01

Direct numerical simulations (DNS) are used to examine the turbulence statistics and the radiation field generated by a high-speed turbulent boundary layer with a nominal freestream Mach number of 14 and wall temperature of 0:18 times the recovery temperature. The flow conditions fall within the range of nozzle exit conditions of the Arnold Engineering Development Center (AEDC) Hypervelocity Tunnel No. 9 facility. The streamwise domain size is approximately 200 times the boundary-layer thickness at the inlet, with a useful range of Reynolds number corresponding to Re 450 ?? 650. Consistent with previous studies of turbulent boundary layer at high Mach numbers, the weak compressibility hypothesis for turbulent boundary layers remains applicable under this flow condition and the computational results confirm the validity of both the van Driest transformation and Morkovin's scaling. The Reynolds analogy is valid at the surface; the RMS of fluctuations in the surface pressure, wall shear stress, and heat flux is 24%, 53%, and 67% of the surface mean, respectively. The magnitude and dominant frequency of pressure fluctuations are found to vary dramatically within the inner layer (z/delta 0.< or approx. 0.08 or z+ < or approx. 50). The peak of the pre-multiplied frequency spectrum of the pressure fluctuation is f(delta)/U(sub infinity) approx. 2.1 at the surface and shifts to a lower frequency of f(delta)/U(sub infinity) approx. 0.7 in the free stream where the pressure signal is predominantly acoustic. The dominant frequency of the pressure spectrum shows a significant dependence on the freestream Mach number both at the wall and in the free stream.

Experimental Study of a Three-Dimensional Shear-Driven Turbulent Boundary Layer with Streamwise Adverse Pressure Gradient

NASA Technical Reports Server (NTRS)

Driver, David M.; Johnston, James P.

1990-01-01

The effects of a strong adverse pressure gradient on a three-dimensional turbulent boundary layer are studied in an axisymmetric spinning cylinder geometry. Velocity measurements made with a three-component laser Doppler velocimeter include all three mean flow components, all six Reynolds stress components, and all ten triple-product correlations. Reynolds stress diminishes as the flow becomes three-dimensional. Lower levels of shear stress were seen to persist under adverse pressure gradient conditions. This low level of stress was seen to roughly correlate with the magnitude of cross-flow (relative to free stream flow) for this experiment as well as most of the other experiments in the literature. Variations in pressure gradient do not appear to alter this correlation. For this reason, it is hypothesized that a three-dimensional boundary layer is more prone to separate than a two-dimensional boundary layer, although it could not be directly shown here. None of the computations performed with either a Prandtl mixing length, k-epsilon, or a Launder-Reece-Rodi full Reynolds-stress model were able to predict the reduction in Reynolds stress.

Optimal Disturbances in Boundary Layers Subject to Streamwise Pressure Gradient

NASA Technical Reports Server (NTRS)

Ashpis, David E.; Tumin, Anatoli

2003-01-01

An analysis of the non-modal growth of perturbations in a boundary layer in the presence of a streamwise pressure gradient is presented. The analysis is based on PSE equations for an incompressible fluid. Examples with Falkner- Skan profiles indicate that a favorable pressure gradient decreases the non-modal growth while an unfavorable pressure gradient leads to an increase of the amplification. It is suggested that the transient growth mechanism be utilized to choose optimal parameters of tripping elements on a low-pressure turbine (LPT) airfoil. As an example, a boundary-layer flow with a streamwise pressure gradient corresponding to the pressure distribution over a LPT airfoil is considered. It is shown that there is an optimal spacing of the tripping elements and that the transient growth effect depends on the starting point. The amplification is found to be small at the LPT s very low Reynolds numbers, but there is a possibility to enhance the transient energy growth by means of wall cooling.

Discussion of Boundary-Layer Characteristics Near the Wall of an Axial-Flow Compressor

NASA Technical Reports Server (NTRS)

Mager, Artur; Mohoney, John J; Budinger, Ray E

1952-01-01

The boundary-layer velocity profiles in the tip region of an axial-flow compressor downstream of the guide vanes and downstream of the rotor were measured by use of total-pressure and claw-type yaw probes. These velocities were resolved into two components: one along the streamline of the flow outside the boundary layer, and the other perpendicular to it. The affinity among all profiles was thus demonstrated with the boundary-layer thickness and the deflection of the boundary layer at the wall as the generalizing parameters. By use of these results and the momentum-integral equations, boundary-layer characteristics on the walls of an axial-flow compressor were qualitatively evaluated.

Numerical simulation of supersonic flow using a new analytical bleed boundary condition

NASA Technical Reports Server (NTRS)

Harloff, G. J.; Smith, G. E.

1995-01-01

A new analytical bleed boundary condition is used to compute flowfields for a strong oblique shock wave/boundary layer interaction with a baseline and three bleed rates at a freestream Mach number of 2.47 with an 8 deg shock generator. The computational results are compared to experimental Pitot pressure profiles and wall static pressures through the interaction region. An algebraic turbulence model is employed for the bleed and baseline cases, and a one equation model is also used for the baseline case where the boundary layer is separated.

Partial differential equations of 3D boundary layer and their numerical solutions in turbomachinery

NASA Astrophysics Data System (ADS)

Zhang, Guoqing; Hua, Yaonan; Wu, Chung-Hua

1991-08-01

This paper studies the 3D boundary layer equations (3DBLE) and their numerical solutions in turbomachinery: (1) the general form of 3DBLE in turbomachines with rotational and curvature effects are derived under the semiorthogonal coordinate system, in which the normal pressure gradient is not equal to zero; (2) the method of solution of the 3DBLE is discussed; (3) the 3D boundary layers on the rotating blade surface, IGV endwall, rotor endwall (with a relatively moving boundary) are numerically solved, and the predicted data correlates well with the measured data; and (4) the comparison is made between the numerical results of 3DBLE with and without normal pressure gradient.

A High-Lift Building Block Flow: Turbulent Boundary Layer Relaminarization

NASA Technical Reports Server (NTRS)

Bourassa, Corey; Thomas, Flint O.; Nelson, Robert C.

2001-01-01

A working wind tunnel test facility has been constructed at the University of Notre Dame's Hessert Center. The relaminarization test facility has been constructed in the 1.5m x 1.5m (5ft x 5 ft) atmospheric wind tunnel and generates a Re(theta)=4694 turbulent boundary layer in nominally zero-pressure gradient before it is exposed to the Case #1 pressure gradient (K approximately equal to 4.2 x 10(exp -6), which is believed to be sufficient to achieve relaminarization. Future work to be conducted will include measuring the response of the turbulent boundary layer to the favorable pressure gradients created in the test facility and documenting this response in order to understand the underlying flow physics responsible for relaminarization. It is the goal of this research to have a better understanding of accelerated turbulent boundary layers which will aid in the development of future flow diagnostic utilities to be implemented in applied aerodynamic research.

Obtaining phase velocity of turbulent boundary layer pressure fluctuations at high subsonic Mach number from wind tunnel data affected by strong background noise

NASA Astrophysics Data System (ADS)

Haxter, Stefan; Brouwer, Jens; Sesterhenn, Jörn; Spehr, Carsten

2017-08-01

Boundary layer measurements at high subsonic Mach number are evaluated in order to obtain the dominant phase velocities of boundary layer pressure fluctuations. The measurements were performed in a transonic wind tunnel which had a very strong background noise. The phase velocity was taken from phase inclination and from the convective peak in one- and two-dimensional wavenumber spectra. An approach was introduced to remove the acoustic noise from the data by applying a method based on CLEAN-SC on the two-dimensional spectra, thereby increasing the frequency range where information about the boundary layer was retrievable. A comparison with prediction models showed some discrepancies in the low-frequency range. Therefore, pressure data from a DNS calculation was used to substantiate the results of the analysis in this frequency range. Using the measured data, the DNS results and a review of the models used for comparison it was found that the phase velocity decreases at low frequencies.

STC-SAB program users manual for the turbulent boundary layer and turbulent separation prediction methods employed in the NASA Langley streamtube curvature computer program

NASA Technical Reports Server (NTRS)

Ferguson, D. R.

1972-01-01

The streamtube curvature program (STC) has been developed to predict the inviscid flow field and the pressure distribution about nacelles at transonic speeds. The effects of boundary layer are to displace the inviscid flow and effectively change the body shape. Thus, the body shape must be corrected by the displacement thickness in order to calculate the correct pressure distribution. This report describes the coupling of the Stratford and Beavers boundary layer solution with the inviscid STC analysis so that all nacelle pressure forces, friction drag, and incipient separation may be predicted. The usage of the coupled STC-SAB computer program is outlined and the program input and output are defined. Included in this manual are descriptions of the principal boundary layer tables and other revisions to the STC program. The use of the viscous option is controlled by the engineer during program input definition.

Two-layer convective heating prediction procedures and sensitivities for blunt body reentry vehicles

NASA Technical Reports Server (NTRS)

Bouslog, Stanley A.; An, Michael Y.; Wang, K. C.; Tam, Luen T.; Caram, Jose M.

1993-01-01

This paper provides a description of procedures typically used to predict convective heating rates to hypersonic reentry vehicles using the two-layer method. These procedures were used to compute the pitch-plane heating distributions to the Apollo geometry for a wind tunnel test case and for three flight cases. Both simple engineering methods and coupled inviscid/boundary layer solutions were used to predict the heating rates. The sensitivity of the heating results in the choice of metrics, pressure distributions, boundary layer edge conditions, and wall catalycity used in the heating analysis were evaluated. Streamline metrics, pressure distributions, and boundary layer edge properties were defined from perfect gas (wind tunnel case) and chemical equilibrium and nonequilibrium (flight cases) inviscid flow-field solutions. The results of this study indicated that the use of CFD-derived metrics and pressures provided better predictions of heating when compared to wind tunnel test data. The study also showed that modeling entropy layer swallowing and ionization had little effect on the heating predictions.

Flight Measurement of Wall-Pressure Fluctuations and Boundary-Layer Turbulence

NASA Technical Reports Server (NTRS)

Mull, Harold R.; Algranti, Joseph S.

1960-01-01

The results are presented for a flight test program using a fighter type jet aircraft flying at pressure altitudes of 10,000, 20,000, and 30,000 feet at Mach numbers from 0.3 to 0.8. Specially designed apparatus was used to measure and record the output of microphones and hot-wire anemometers mounted on the forward-fuselage section and wing of the airplane. Mean-velocity profiles in the boundary layers were obtained from total-pressure measurements. The ratio of the root-mean-square fluctuating wall pressure to the free-stream dynamic pressure is presented as a function of Reynolds number and Mach number. The longitudinal component of the turbulent-velocity fluctuations was measured, and the turbulence-intensity profiles are presented for the wing and forward-fuselage section. In general, the results are in agreement with wind-tunnel measurements which have been-reported in the literature. For example, the variation the square root of p(sup 2)/q times the square root of p(sup 2) is the root mean square of the wall-pressure fluctuation, and q is the free-stream dynamic pressure) with Reynolds number was found to be essentially constant for the forward-fuselage-section boundary layer, while variations at the wing station were probably unduly affected by the microphone diameter (5/8 in.), which was large compared with the boundary-layer thickness.

An experimental study of a three-dimensional shock wave/turbulent boundary-layer interaction at a hypersonic Mach number

NASA Technical Reports Server (NTRS)

Kussoy, M. I.; Horstman, K. C.; Kim, K.-S.

1991-01-01

Experimental data for a series of three-dimensional shock-wave/turbulent-boundary-layer interaction flows at Mach 8.2 are presented. The test bodies, composed of sharp fins fastened to a flat-plate test surface, were designed to generate flows with varying degrees of pressure gradient, boundary-layer separation, and turning angle. The data include surface-pressure, heat-transfer, and skin-friction distributions, as well as limited mean flowfield surveys both in the undisturbed and interaction regimes. The data were obtained for the purpose of validating computational models of these hypersonic interactions.

Documentation of Two- and Three-Dimensional Hypersonic Shock Wave/Turbulent Boundary Layer Interaction Flows

NASA Technical Reports Server (NTRS)

Kussoy, Marvin I.; Horstman, Clifford C.

1989-01-01

Experimental data for a series of two- and three-dimensional shock wave/turbulent boundary layer interaction flows at Mach 7 are presented. Test bodies, composed of simple geometric shapes, were designed to generate flows with varying degrees of pressure gradient, boundary-layer separation, and turning angle. The data include surface-pressure and heat-transfer distributions as well as limited mean-flow-field surveys in both the undisturbed and the interaction regimes. The data are presented in a convenient form for use in validating existing or future computational models of these generic hypersonic flows.

Control of shock-wave boundary-layer interactions by bleed in supersonic mixed compression inlets

NASA Technical Reports Server (NTRS)

Fukuda, M. K.; Reshotko, E.; Hingst, W. R.

1975-01-01

An experimental investigation has been conducted to determine the effect of bleed region geometry and bleed rate on shock wave-boundary layer interactions in an axisymmetric, mixed-compression inlet at a Mach number of 2.5. The full realizable reduction in transformed form factor is obtained by bleeding off about half the incident boundary layer mass flow. Bleeding upstream or downstream of the shock-induced pressure rise is preferable to bleeding across the shock-induced pressure rise. Slanted holes are more effective than normal holes. Two different bleed hole sizes were tested without detectable difference in performance.

The Turbulent Flow in Diffusers of Small Divergence Angle

NASA Technical Reports Server (NTRS)

Gourzhienko, G. A.

1947-01-01

The turbulent flow in a conical diffuser represents the type of turbulent boundary layer with positive longitudinal pressure gradient. In contrast to the boundary layer problem, however, it is not necessary that the pressure distribution along the limits of the boundary layer(along the axis of the diffuser) be given, since this distribution can be obtained from the computation. This circumstance, together with the greater simplicity of the problem as a whole, provides a useful basis for the study of the extension of the results of semiempirical theories to the case of motion with a positive pressure gradient. In the first part of the paper,formulas are derived for the computation of the velocity and.pressure distributions in the turbulent flow along, and at right angles to, the axis of a diffuser of small cone angle. The problem is solved.

An experimental study of fluctuating pressure loads beneath swept shock/boundary-layer interactions

NASA Technical Reports Server (NTRS)

Settles, Gary S.

1991-01-01

A database is established on the fluctuating pressure loads produced on aerodynamic surfaces beneath 3-D shock wave/boundary layer interactions. Such loads constitute a fundamental problem of critical concern to future supersonic and hypersonic flight vehicles. A turbulent boundary layer on a flat plate is subjected to interactions with swept planar shock waves generated by sharp fins. Fin angles from 5 to 25 deg at freestream Mach numbers between 2.5 and 4 produce a variety of interaction strengths from weak to very strong. Miniature Kulite pressure transducers mounted in the flat plate were used to measure interaction-induced wall pressure fluctuations. These data will be correlated with proposed new optical data on the fluctuations of the interaction structure, especially that of the lambda-shock system and its associated high-speed jet impingement.

Some Features of Artificially Thickened Fully Developed Turbulent Boundary Layers with Zero Pressure Gradient

NASA Technical Reports Server (NTRS)

Klebanoff, P S; Diehl, Z W

1952-01-01

Report gives an account of an investigation conducted to determine the feasibility of artificially thickening a turbulent boundary layer on a flat plate. A description is given of several methods used to thicken artificially the boundary layer. It is shown that it is possible to do substantial thickening and obtain a fully developed turbulent boundary layer, which is free from any distortions introduced by the thickening process, and, as such, is a suitable medium for fundamental research.

Viscous pressure correction in the irrotational flow outside Prandtl's boundary layer

NASA Astrophysics Data System (ADS)

Joseph, Daniel; Wang, Jing

2004-11-01

We argue that boundary layers on solid with irrotational motion outside are like a gas bubble because the shear stress vanishes at the edge of the boundary layer but the irrotational shear stress does not. This discrepancy induces a pressure correction and an additional drag which can be advertised as due to the viscous dissipation of the irrotational flow. Typically, this extra correction to the drag would be relatively small. A much more interesting implication of the extra pressure theory arises from the consideration of the effects of viscosity on the normal stress on a solid boundary which are entirely neglected in Prandtl's theory. It is very well known and easily demonstrated that as a consequence of the continuity equation the viscous normal stress must vanish on a rigid solid. It follows that all the greatly important effects of viscosity on the normal stress are buried in the pressure and the leading order effects of viscosity on the normal stress can be obtained from the viscous correction of viscous potential flow.

Intermittent Behavior of the Separated Boundary Layer along the Suction Surface of a Low Pressure Turbine Blade under Periodic Unsteady Flow Conditions

NASA Technical Reports Server (NTRS)

Oeztuerk, B; Schobeiri, M. T.; Ashpis, David E.

2005-01-01

The paper experimentally and theoretically studies the effects of periodic unsteady wake flow and aerodynamic characteristics on boundary layer development, separation and re-attachment along the suction surface of a low pressure turbine blade. The experiments were carried out at Reynolds number of 110,000 (based on suction surface length and exit velocity). For one steady and two different unsteady inlet flow conditions with the corresponding passing frequencies, intermittency behaviors were experimentally and theoretically investigated. The current investigation attempts to extend the intermittency unsteady boundary layer transition model developed in previously to the LPT cases, where separation occurs on the suction surface at a low Reynolds number. The results of the unsteady boundary layer measurements and the intermittency analysis were presented in the ensemble-averaged and contour plot forms. The analysis of the boundary layer experimental data with the flow separation, confirms the universal character of the relative intermittency function which is described by a Gausssian function.

Theory of viscous transonic flow over airfoils at high Reynolds number

NASA Technical Reports Server (NTRS)

Melnik, R. E.; Chow, R.; Mead, H. R.

1977-01-01

This paper considers viscous flows with unseparated turbulent boundary layers over two-dimensional airfoils at transonic speeds. Conventional theoretical methods are based on boundary layer formulations which do not account for the effect of the curved wake and static pressure variations across the boundary layer in the trailing edge region. In this investigation an extended viscous theory is developed that accounts for both effects. The theory is based on a rational analysis of the strong turbulent interaction at airfoil trailing edges. The method of matched asymptotic expansions is employed to develop formal series solutions of the full Reynolds equations in the limit of Reynolds numbers tending to infinity. Procedures are developed for combining the local trailing edge solution with numerical methods for solving the full potential flow and boundary layer equations. Theoretical results indicate that conventional boundary layer methods account for only about 50% of the viscous effect on lift, the remaining contribution arising from wake curvature and normal pressure gradient effects.

First results of a study on turbulent boundary layers in oscillating flow with a mean adverse pressure gradient

NASA Technical Reports Server (NTRS)

Houdeville, R.; Cousteix, J.

1979-01-01

The development of a turbulent unsteady boundary layer with a mean pressure gradient strong enough to induce separation, in order to complete the extend results obtained for the flat plate configuration is presented. The longitudinal component of the velocity is measured using constant temperature hot wire anemometer. The region where negative velocities exist is investigated with a laser Doppler velocimeter system with BRAGG cells. The boundary layer responds by forced pulsation to the perturbation of potential flow. The unsteady effects observed are very important. The average location of the zero skin friction point moves periodically at the perturbation frequency. Average velocity profiles from different instants in the cycle are compared. The existence of a logarithmic region enables a simple calculation of the maximum phase shift of the velocity in the boundary layer. An attempt of calculation by an integral method of boundary layer development is presented, up to the point where reverse flow starts appearing.

The Azimuthally Averaged Boundary Layer Structure of a Numerically Simulated Major Hurricane

DTIC Science & Technology

2015-08-14

layer in which the effects of sur- face friction are associated with significant departures from gradient wind balance. The boundary layer in the... effects of surface friction are associated with significant departures from gradient wind balance. More specifically, we follow Key Points: The...comprises a balance between three horizontal forces: Coriolis , pressure gradient, and friction. The boundary layer flow is characterized by a large Reynolds

Experimental Studies of Flow Separation of the NACA 2412 Airfoil at Low Speeds

NASA Technical Reports Server (NTRS)

Seetharam, H. C.; Rodgers, E. J.; Wentz, W. H., Jr.

1997-01-01

Wind tunnel tests have been conducted on an NACA 2412 airfoil section at Reynolds number of 2.2 x 10(exp 6) and Mach number of 0.13. Detailed measurements of flow fields associated with turbulent boundary layers have been obtained at angles of attack of 12.4 degrees, 14.4 degrees, and 16.4 degrees. Pre- and post-separated velocity and pressure survey results over the airfoil and in the associated wake are presented. Extensive force, pressure, tuft survey, hot-film survey, local skin friction, and boundary layer data are also included. Pressure distributions and separation point locations show good agreement with theory for the two layer angles of attack. Boundary layer displacement thickness, momentum thickness, and shape factor agree well with theory up to the point of separation. There is considerable disparity between extent of flow reversal in the wake as measured by pressure and hot-film probes. The difference is attributed to the intermittent nature of the flow reversal.

Base Pressure at Supersonic Speeds on Two-dimensional Airfoils and on Bodies of Revolution with and Without Fins Having Turbulent Boundary Layers

NASA Technical Reports Server (NTRS)

LOVE EUGENE S

1957-01-01

An analysis has been made of available experimental data to show the effects of most of the variables that are more predominant in determining base pressure at supersonic speeds. The analysis covers base pressures for two-dimensional airfoils and for bodies of revolution with and without stabilizing fins and is restricted to turbulent boundary layers. The present status of available experimental information is summarized as are the existing methods for predicting base pressure. A simple semiempirical method is presented for estimating base pressure. For two-dimensional bases, this method stems from an analogy established between the base-pressure phenomena and the peak pressure rise associated with the separation of the boundary layer. An analysis made for axially symmetric flow indicates that the base pressure for bodies of revolution is subject to the same analogy. Based upon the methods presented, estimations are made of such effects as Mach number, angle of attack, boattailing, fineness ratio, and fins. These estimations give fair predictions of experimental results. (author)

One-stage free-vortex aerodynamic window with pressure ratio 100 and atmospheric exhaust

NASA Astrophysics Data System (ADS)

Malkov, Victor M.; Trilis, A. V.; Savin, Andrew V.; Druzhinin, S. L.

2005-03-01

The aerodynamic windows (AW) are intended for a high power extraction from the gas laser optical cavity, where the pressure is much lower than environment pressure. The main requirements for the aerodynamic windows are to satisfy a low level of optical disturbances in a laser beam extraction channel and an air leakage absence into the optical cavity. Free vortex AW are most economic from a point of working gas consumption and the greatest pressure differential is realized on them at an exhaust to atmosphere. For ideal gas it is possible to receive as much as large pressure differential, however for real gas a pressure differential more than P>=50 is difficult to achieve. To achieve the pressure ratio 100 in free vortex single-stage AW the method of stabilizing of boundary layer was used. The gas of curtain was decelerated in the diffuser and was exhausted into the atmosphere straightly. The pressure recovery improvement was achieved by using the boundary layer blowing inside the diffuser. Only 10% of total mass flow was used for boundary layer blowing.

Cyclone separator having boundary layer turbulence control

DOEpatents

Krishna, Coimbatore R.; Milau, Julius S.

1985-01-01

A cyclone separator including boundary layer turbulence control that is operable to prevent undue build-up of particulate material at selected critical areas on the separator walls, by selectively varying the fluid pressure at those areas to maintain the momentum of the vortex, thereby preventing particulate material from inducing turbulence in the boundary layer of the vortical fluid flow through the separator.

In-Flight Boundary-Layer Transition on a Large Flat Plate at Supersonic Speeds

NASA Technical Reports Server (NTRS)

Banks, Daniel W.; Fredericks, Michael Alan; Tracy, Richard R.; Matisheck, Jason R.; Vanecek, Neal D.

2012-01-01

A flight experiment was conducted to investigate the pressure distribution, local flow conditions, and boundary-layer transition characteristics on a large flat plate in flight at supersonic speeds up to Mach 2.0. The primary objective of the test was to characterize the local flow field in preparation for future tests of a high Reynolds number natural laminar flow test article. The tests used a F-15B testbed aircraft with a bottom centerline mounted test fixture. A second objective was to determine the boundary-layer transition characteristics on the flat plate and the effectiveness of using a simplified surface coating for future laminar flow flight tests employing infrared thermography. Boundary-layer transition was captured using an onboard infrared imaging system. The infrared imagery was captured in both analog and digital formats. Surface pressures were measured with electronically scanned pressure modules connected to 60 surface-mounted pressure orifices. The local flow field was measured with five 5-hole conical probes mounted near the leading edge of the test fixture. Flow field measurements revealed the local flow characteristics including downwash, sidewash, and local Mach number. Results also indicated that the simplified surface coating did not provide sufficient insulation from the metallic structure, which likely had a substantial effect on boundary-layer transition compared with that of an adiabatic surface. Cold wall conditions were predominant during the acceleration to maximum Mach number, and warm wall conditions were evident during the subsequent deceleration. The infrared imaging system was able to capture shock wave impingement on the surface of the flat plate in addition to indicating laminar-to-turbulent boundary-layer transition.

Calculation of sidewall boundary-layer parameters from rake measurements for the Langley 0.3-meter transonic cryogenic tunnel

NASA Technical Reports Server (NTRS)

Murthy, A. V.

1987-01-01

Correction of airfoil data for sidewall boundary-layer effects requires a knowledge of the boundary-layer displacement thickness and the shape factor with the tunnel empty. To facilitate calculation of these quantities under various test conditions for the Langley 0.3 m Transonic Cryogenic Tunnel, a computer program was written. This program reads the various tunnel parameters and the boundary-layer rake total head pressure measurements directly from the Engineering Unit tapes to calculate the required sidewall boundary-layer parameters. Details of the method along with the results for a sample case are presented.

Effect of aspect ratio on sidewall boundary-layer influence in two-dimensional airfoil testing

NASA Technical Reports Server (NTRS)

Murthy, A. V.

1986-01-01

The effect of sidewall boundary layers in airfoil testing in two-dimensional wind tunnels is investigated. The non-linear crossflow velocity variation induced because of the changes in the sidewall boundary-layer thickness is represented by the flow between a wavy wall and a straight wall. Using this flow model, a correction for the sidewall boundary-layer effects is derived in terms of the undisturbed sidewall boundary-layer properties, the test Mach number and the airfoil aspect ratio. Application of the proposed correction to available experimental data showed good correlation for the shock location and pressure distribution on airfoils.

A Galilean and tensorial invariant k-epsilon model for near wall turbulence

NASA Technical Reports Server (NTRS)

Yang, Z.; Shih, T. H.

1993-01-01

A k-epsilon model is proposed for wall bounded turbulent flows. In this model, the eddy viscosity is characterized by a turbulent velocity scale and a turbulent time scale. The time scale is bounded from below by the Kolmogorov time scale. The dissipation rate equation is reformulated using this time scale and no singularity exists at the wall. A new parameter R = k/S(nu) is introduced to characterize the damping function in the eddy viscosity. This parameter is determined by local properties of both the mean and the turbulent flow fields and is free from any geometry parameter. The proposed model is then Galilean and tensorial invariant. The model constants used are the same as in the high Reynolds number Standard k-epsilon Model. Thus, the proposed model will also be suitable for flows far from the wall. Turbulent channel flows and turbulent boundary layer flows with and without pressure gradients are calculated. Comparisons with the data from direct numerical simulations and experiments show that the model predictions are excellent for turbulent channel flows and turbulent boundary layers with favorable pressure gradients, good for turbulent boundary layers with zero pressure gradients, and fair for turbulent boundary layer with adverse pressure gradients.

On the turbulent friction layer for rising pressure

NASA Technical Reports Server (NTRS)

Wieghardt, K; Tillmann, W

1951-01-01

Among the information presented are included displacement, momentum, and kinetic energy thicknesses, shearing stress distributions across boundary layer, and surface friction coefficients. The Gruschwitz method and its modifications are examined and tested. An energy theorem for the turbulent boundary layer is introduced and discussed but does not lead to a method for the prediction of the behavior of the turbulent boundary layer because relations for the shearing stress and the surface friction are lacking.

High Reynolds Number Investigation of a Flush Mounted, S-Duct Inlet With Large Amounts of Boundary Layer Ingestion

NASA Technical Reports Server (NTRS)

Berrier, Bobby L.; Carter, Melissa B.; Allan, Brian G.

2005-01-01

An experimental investigation of a flush-mounted, S-duct inlet with large amounts of boundary layer ingestion has been conducted at Reynolds numbers up to full scale. The study was conducted in the NASA Langley Research Center 0.3-Meter Transonic Cryogenic Tunnel. In addition, a supplemental computational study on one of the inlet configurations was conducted using the Navier-Stokes flow solver, OVERFLOW. Tests were conducted at Mach numbers from 0.25 to 0.83, Reynolds numbers (based on aerodynamic interface plane diameter) from 5.1 million to 13.9 million (full-scale value), and inlet mass-flow ratios from 0.29 to 1.22, depending on Mach number. Results of the study indicated that increasing Mach number, increasing boundary layer thickness (relative to inlet height) or ingesting a boundary layer with a distorted profile decreased inlet performance. At Mach numbers above 0.4, increasing inlet airflow increased inlet pressure recovery but also increased distortion. Finally, inlet distortion was found to be relatively insensitive to Reynolds number, but pressure recovery increased slightly with increasing Reynolds number.This CD-ROM supplement contains inlet data including: Boundary layer data, Duct static pressure data, performance-AIP (fan face) data, Photos, Tunnel wall P-PTO data and definitions.

Effects of Reynolds number on orifice induced pressure error

NASA Technical Reports Server (NTRS)

Plentovich, E. B.; Gloss, B. B.

1982-01-01

Data previously reported for orifice induced pressure errors are extended to the case of higher Reynolds number flows, and a remedy is presented in the form of a porous metal plug for the orifice. Test orifices with apertures 0.330, 0.660, and 1.321 cm in diam. were fabricated on a flat plate for trials in the NASA Langley wind tunnel at Mach numbers 0.40-0.72. A boundary layer survey rake was also mounted on the flat plate to allow measurement of the total boundary layer pressures at the orifices. At the high Reynolds number flows studied, the orifice induced pressure error was found to be a function of the ratio of the orifice diameter to the boundary layer thickness. The error was effectively eliminated by the insertion of a porous metal disc set flush with the orifice outside surface.

Acoustic Radiation from a Mach 14 Turbulent Boundary layer

NASA Astrophysics Data System (ADS)

Duan, Lian; Choudhari, Meelan

2014-11-01

Direct numerical simulations (DNS) are used to examine the pressure fluctuations generated by a high-speed turbulent boundary layer with a nominal freestream Mach number of 14 and wall temperature of 0.15 times the recovery temperature. The emphasis is on characterizing the acoustic radiation from the turbulent boundary layer and comparing it with previous simulations at Mach 2.5 and Mach 6 to assess the Mach-number dependence of the freestream pressure fluctuations. In particular, the numerical database is used to provide insights into the pressure disturbance spectrum and amplitude scaling with respect to the freestream Mach number as well as to understand the acoustic source mechanisms at very high Mach numbers. Such information is important for characterizing the freestream disturbance environment in conventional (i.e., noisy) hypersonic wind tunnels. Spectral characteristics of pressure fluctuations at the surface are also investigated. Supported by AFOSR and NASA Langley Research Center.

Experimental Investigation of Separated and Transitional Boundary Layers Under Low-Pressure Turbine Airfoil Conditions

NASA Technical Reports Server (NTRS)

Hultgren, Lennart S.; Volino, Ralph J.

2002-01-01

Modern low-pressure turbine airfoils are subject to increasingly stronger pressure gradients as designers impose higher loading in an effort to improve efficiency and to reduce part count. The adverse pressure gradients on the suction side of these airfoils can lead to boundary-layer separation, particularly under cruise conditions. Separation bubbles, notably those which fail to reattach, can result in a significant degradation of engine efficiency. Accurate prediction of separation and reattachment is hence crucial to improved turbine design. This requires an improved understanding of the transition flow physics. Transition may begin before or after separation, depending on the Reynolds number and other flow conditions, has a strong influence on subsequent reattachment, and may even eliminate separation. Further complicating the problem are the high free-stream turbulence levels in a real engine environment, the strong pressure gradients along the airfoils, the curvature of the airfoils, and the unsteadiness associated with wake passing from upstream stages. Because of the complicated flow situation, transition in these devices can take many paths that can coexist, vary in importance, and possibly also interact, at different locations and instances in time. The present work was carried out in an attempt to systematically sort out some of these issues. Detailed velocity measurements were made along a flat plate subject to the same nominal dimensionless pressure gradient as the suction side of a modern low-pressure turbine airfoil ('Pak-B'). The Reynolds number based on wetted plate length and nominal exit velocity, Re, was varied from 50;000 to 300; 000, covering cruise to takeoff conditions. Low, 0.2%, and high, 7%, inlet free-stream turbulence intensities were set using passive grids. These turbulence levels correspond to about 0.2% and 2.5% turbulence intensity in the test section when normalized with the exit velocity. The Reynolds number and free-stream turbulence level do not have a significant effect on the location of boundary-layer separation unless they are high enough to induce transition upstream of separation. The location and extent of the transition zone, in contrast, depend strongly on Re and TI. The beginning of reattachment closely follows the onset of transition. Under low free-stream turbulence conditions the boundary layer is laminar at separation and then begins to exhibit fluctuations in a finite frequency band in the shear layer over the separation bubble. These fluctuations are due to instability waves. The fluctuations grow in magnitude, higher harmonics are generated, and finally lead to a breakdown to turbulence. Transition begins in the shear layer, but quickly spreads to the near wall region and causes the boundary layer to reattach. The transition is rapid and the resulting turbulence contains a full range of high and low frequencies. Under high free-stream turbulence conditions, slowly growing low-frequency fluctuations are induced in the pretransitional boundary layer by the free-stream. The separation bubbles are considerably thinner than in the low TI cases, resulting in thinner boundary layers at the end of the test wall. At Re=50,000 and 100,000, the pre-transitional boundary layer separates at about the same location as in the low TI cases. Transition occurs through a bypass mode, begins upstream of the corresponding low-TI location, and proceeds in a manner similar to that of an attached boundary layer. Under high TI at Re=200,000 and 300,000, transition begins before separation. The boundary layer may separate, but if it does the separation bubble is very short and does not significantly affect the downstream development of the boundary layer. A comparison is made to previous work in a simulated cascade.

Supersonic turbulent boundary layers with periodic mechanical non-equilibrium

NASA Astrophysics Data System (ADS)

Ekoto, Isaac Wesley

Previous studies have shown that favorable pressure gradients reduce the turbulence levels and length scales in supersonic flow. Wall roughness has been shown to reduce the large-scales in wall bounded flow. Based on these previous observations new questions have been raised. The fundamental questions this dissertation addressed are: (1) What are the effects of wall topology with sharp versus blunt leading edges? and (2) Is it possible that a further reduction of turbulent scales can occur if surface roughness and favorable pressure gradients are combined? To answer these questions and to enhance the current experimental database, an experimental analysis was performed to provide high fidelity documentation of the mean and turbulent flow properties along with surface and flow visualizations of a high-speed (M = 2.86), high Reynolds number (Retheta ≈ 60,000) supersonic turbulent boundary layer distorted by curvature-induced favorable pressure gradients and large-scale ( k+s ≈ 300) uniform surface roughness. Nine models were tested at three separate locations. Three pressure gradient models strengths (a nominally zero, a weak, and a strong favorable pressure gradient) and three roughness topologies (aerodynamically smooth, square, and diamond shaped roughness elements) were used. Highly resolved planar measurements of mean and fluctuating velocity components were accomplished using particle image velocimetry. Stagnation pressure profiles were acquired with a traversing Pitot probe. Surface pressure distributions were characterized using pressure sensitive paint. Finally flow visualization was accomplished using schlieren photographs. Roughness topology had a significant effect on the boundary layer mean and turbulent properties due to shock boundary layer interactions. Favorable pressure gradients had the expected stabilizing effect on turbulent properties, but the improvements were less significant for models with surface roughness near the wall due to increased tendency towards flow separation. It was documented that proper roughness selection coupled with a sufficiently strong favorable pressure gradient produced regions of "negative" production in the transport of turbulent stress. This led to localized areas of significant turbulence stress reduction. With proper roughness selection and sufficient favorable pressure gradient strength, it is believed that localized relaminarization of the boundary layer is possible.

Boundary-layer transition on a plate subjected to simultaneous spanwise and chordwise pressure gradients

NASA Technical Reports Server (NTRS)

Boldman, D. R.; Brinich, P. F.

1974-01-01

The boundary-layer transition on a short plate was studied by means of the china-clay visual technique. The plate model was mounted in a wind tunnel so that it was subjected to small simultaneous spanwise and chordwise pressure gradients. Results of the experimental study, which was performed at three subsonic velocities, indicated that the transition pattern was appreciably curved in the spanwise direction but quite smooth and well behaved. Reasonable comparisons between predictions of transition and experiment were obtained from two finite-difference two-dimensional boundary-layer calculation methods which incorporated transition models based on the concept of a transition intermittency factor.

Characteristics of turbulence in boundary layer with zero pressure gradient

NASA Technical Reports Server (NTRS)

Klebanoff, P S

1955-01-01

The results of an experimental investigation of a turbulent boundary layer with zero pressure gradient are presented. Measurements with the hot-wire anemometer were made of turbulent energy and turbulent shear stress, probability density and flattening factor of u-fluctuation (fluctuation in x-direction), spectra of turbulent energy and shear stress, and turbulent dissipation. The importance of the region near the wall and the inadequacy of the concept of local isotropy are demonstrated. Attention is given to the energy balance and the intermittent character of the outer region of the boundary layer. Also several interesting features of the spectral distribution of the turbulent motions are discussed.

Large-Eddy Simulation of the Flat-plate Turbulent Boundary Layer at High Reynolds numbers

NASA Astrophysics Data System (ADS)

Inoue, Michio

The near-wall, subgrid-scale (SGS) model [Chung and Pullin, "Large-eddy simulation and wall-modeling of turbulent channel flow'', J. Fluid Mech. 631, 281--309 (2009)] is used to perform large-eddy simulations (LES) of the incompressible developing, smooth-wall, flat-plate turbulent boundary layer. In this model, the stretched-vortex, SGS closure is utilized in conjunction with a tailored, near-wall model designed to incorporate anisotropic vorticity scales in the presence of the wall. The composite SGS-wall model is presently incorporated into a computer code suitable for the LES of developing flat-plate boundary layers. This is then used to study several aspects of zero- and adverse-pressure gradient turbulent boundary layers. First, LES of the zero-pressure gradient turbulent boundary layer are performed at Reynolds numbers Retheta based on the free-stream velocity and the momentum thickness in the range Retheta = 103-1012. Results include the inverse skin friction coefficient, 2/Cf , velocity profiles, the shape factor H, the Karman "constant", and the Coles wake factor as functions of Re theta. Comparisons with some direct numerical simulation (DNS) and experiment are made, including turbulent intensity data from atmospheric-layer measurements at Retheta = O (106). At extremely large Retheta , the empirical Coles-Fernholz relation for skin-friction coefficient provides a reasonable representation of the LES predictions. While the present LES methodology cannot of itself probe the structure of the near-wall region, the present results show turbulence intensities that scale on the wall-friction velocity and on the Clauser length scale over almost all of the outer boundary layer. It is argued that the LES is suggestive of the asymptotic, infinite Reynolds-number limit for the smooth-wall turbulent boundary layer and different ways in which this limit can be approached are discussed. The maximum Retheta of the present simulations appears to be limited by machine precision and it is speculated, but not demonstrated, that even larger Retheta could be achieved with quad- or higher-precision arithmetic. Second, the time series velocity signals obtained from LES within the logarithmic region of the zero-pressure gradient turbulent boundary layer are used in combination with an empirical, predictive inner--outer wall model [Marusic et al., "Predictive model for wall-bounded turbulent flow'', Science 329, 193 (2010)] to calculate the statistics of the fluctuating streamwise velocity in the inner region of the zero-pressure gradient turbulent boundary layer. Results, including spectra and moments up to fourth order, are compared with equivalent predictions using experimental time series, as well as with direct experimental measurements at Reynolds numbers Retau based on the friction velocity and the boundary layer thickness, Retau = 7,300, 13,600 and 19,000. LES combined with the wall model are then used to extend the inner-layer predictions to Reynolds numbers Retau = 62,000, 100,000 and 200,000 that lie within a gap in log(Retau) space between laboratory measurements and surface-layer, atmospheric experiments. The present results support a log-like increase in the near-wall peak of the streamwise turbulence intensities with Retau and also provide a means of extending LES results at large Reynolds numbers to the near-wall region of wall-bounded turbulent flows. Finally, we apply the wall model to LES of a turbulent boundary layer subject to an adverse pressure gradient. Computed statistics are found to be consistent with recent experiments and some Reynolds number similarity is observed over a range of two orders of magnitude.

Vortex/boundary layer interactions

NASA Technical Reports Server (NTRS)

Cutler, A. D.; Bradshaw, P.

1989-01-01

Detailed and high quality measurements with hot-wires and pressure probes are presented for two different interactions between a vortex pair with common flow down and a turbulent boundary layer. The interactions studied have larger values of the vortex circulation parameter than those studied previously. The results indicate that the boundary layer under the vortex pair is thinned by lateral divergence and that boundary layer fluid is entrained into the vortex. The effect of the interaction on the vortex core (other than the inviscid effect of the image vortices behind the surface) is small.

A novel concept for subsonic inlet boundary-layer control

NASA Technical Reports Server (NTRS)

Miller, B. A.

1977-01-01

A self-bleeding method for boundary layer control is described and tested for a subsonic inlet designed to operate in the flowfield generated by high angles of attack. Naturally occurring surface static pressure gradients are used to remove the boundary layer from a separation-prone region of the inlet and to reinject it at a less critical location with a net performance gain. The results suggest that this self-bleeding method for boundary-layer control might be successfully applied to other inlets operating at extreme aerodynamic conditions.

Comparison of theoretical and experimental boundary-layer development in a Mach 2.5 mixed-compression inlet

NASA Technical Reports Server (NTRS)

Hingst, W. R.; Towne, C. E.

1974-01-01

An analytical investigation was made of the boundary layer flow in an axisymmetric Mach 2.5 mixed compression inlet, and the results were compared with experimental measurements. The inlet tests were conducted in the Lewis 10- by 10-foot supersonic wind tunnel at a unit Reynolds number of 8.2 million/m. The inlet incorporated porous bleed regions for boundary layer control, and the effect of this bleed was taken into account in the analysis. The experimental boundary layer data were analyzed by using similarity laws from which the skin friction coefficient was obtained. The boundary layer analysis included predictions of laminar and turbulent boundary layer growth, transition, and the effects of the shock boundary layer interactions. In addition, the surface static pressures were compared with those obtained from an inviscid characteristics program. The results of investigation showed that the analytical techniques gave satisfactory predictions of the boundary layer flow except in regions that were badly distorted by the terminal shock.

Increased Jet Noise Due to a "Nominally Laminar" State of Nozzle Exit Boundary Layer

NASA Technical Reports Server (NTRS)

Zaman, K. B. M. Q.

2017-01-01

A set of 2-in. diameter nozzles is used to investigate the effect of varying exit boundary layer state on the radiated noise from high-subsonic jets. It is confirmed that nozzles involving turbulent boundary layers are the quietest while nozzles involving a "nominally laminar" boundary layer are loud especially on the high-frequency side of the sound pressure level spectrum. The latter boundary layer state involves a "Blasius-like" mean velocity profile but higher turbulence intensities compared to those in the turbulent state. The higher turbulence in the initial region of the jet shear layer leads to increased high-frequency noise. The results strongly suggest that an anomaly noted with subsonic jet noise databases in the literature is due to a similar effect of differences in the initial boundary layer state.

Increased Jet Noise Due to a "Nominally Laminar" State of Nozzle Exit Boundary Layer

NASA Technical Reports Server (NTRS)

Zaman, K. B. M. Q.

2017-01-01

A set of 2-inch diameter nozzles is used to investigate the effect of varying exit boundary layer state on the radiated noise from high-subsonic jets. It is confirmed that nozzles involving turbulent boundary layers are the quietest while nozzles involving a nominally-laminar boundary layer are loud especially on the high-frequency side of the sound pressure level spectrum. The latter boundary layer state involves a Blasius-like mean velocity profile but higher turbulence intensities compared to those in the turbulent state. The higher turbulence in the initial region of the jet shear layer leads to increased high-frequency noise. The results strongly suggest that an anomaly noted with subsonic jet noise databases in the literature is due to a similar effect of differences in the initial boundary layer state.

Numerical study of Reynolds stress in compressible flows

NASA Technical Reports Server (NTRS)

Vandromme, D.; Hamin, H.

1985-01-01

A second order closure has been implemented in an implicit Navier-Stokes solver to study the behavior of the Reynolds stresses under the influence of severe pressure gradients. In the boundary layer zone, the strongly sheared character of the mean flow dominates the turbulence generation mechanisms. However, the pressure gradients also play a very important role for these processes, but at different locations within the boundary layer.

Interferometric data for a shock-wave/boundary-layer interaction

NASA Technical Reports Server (NTRS)

Dunagan, Stephen E.; Brown, James L.; Miles, John B.

1986-01-01

An experimental study of the axisymmetric shock-wave / boundary-layer strong interaction flow generated in the vicinity of a cylinder-cone intersection was conducted. The study data are useful in the documentation and understanding of compressible turbulent strong interaction flows, and are part of a more general effort to improve turbulence modeling for compressible two- and three-dimensional strong viscous/inviscid interactions. The nominal free stream Mach number was 2.85. Tunnel total pressures of 1.7 and 3.4 atm provided Reynolds number values of 18 x 10(6) and 36 x 10(6) based on model length. Three cone angles were studied giving negligible, incipient, and large scale flow separation. The initial cylinder boundary layer upstream of the interaction had a thickness of 1.0 cm. The subsonic layer of the cylinder boundary layer was quite thin, and in all cases, the shock wave penetrated a significant portion of the boundary layer. Owing to the thickness of the cylinder boundary layer, considerable structural detail was resolved for the three shock-wave / boundary-layer interaction cases considered. The primary emphasis was on the application of the holographic interferometry technique. The density field was deduced from an interferometric analysis based on the Able transform. Supporting data were obtained using a 2-D laser velocimeter, as well as mean wall pressure and oil flow measurements. The attached flow case was observed to be steady, while the separated cases exhibited shock unsteadiness. Comparisons with Navier-Stokes computations using a two-equation turbulence model are presented.

A review of quasi-coherent structures in a numerically simulated turbulent boundary layer

NASA Technical Reports Server (NTRS)

Robinson, S. K.; Kline, S. J.; Spalart, P. R.

1989-01-01

Preliminary results of a comprehensive study of the structural aspects of a numerically simulated number turbulent boundary layer are presented. A direct Navier-Stokes simulation of a flat-plate, zero pressure gradient boundary layer at Re0 = 670 was used. Most of the known nonrandom, coherent features of turbulent boundary layers are confirmed in the simulation, and several new aspects of their spatial character are reported. The spatial relationships between many of the various structures are described, forming the basis for a more complete kinematical picture of boundary layer physics than has been previously known. In particular, the importance of vortex structures of various forms to the generation of Reynolds shear stress is investigated.

A documentation of two- and three-dimensional shock-separated turbulent boundary layers

NASA Technical Reports Server (NTRS)

Brown, J. D.; Brown, J. L.; Kussoy, M. I.

1988-01-01

A shock-related separation of a turbulent boundary layer has been studied and documented. The flow was that of an axisymmetric turbulent boundary layer over a 5.02-cm-diam cylinder that was aligned with the wind tunnel axis. The boundary layer was compressed by a 30 deg half-angle conical flare, with the cone axis inclined at an angle alpha to the cylinder axis. Nominal test conditions were P sub tau equals 1.7 atm and M sub infinity equals 2.85. Measurements were confined to the upper-symmetry, phi equals 0 deg, plane. Data are presented for the cases of alpha equal to 0. 5. and 10 deg and include mean surface pressures, streamwise and normal mean velocities, kinematic turbulent stresses and kinetic energies, as well as reverse-flow intermittencies. All data are given in tabular form; pressures, streamwise velocities, turbulent shear stresses, and kinetic energies are also presented graphically.

Turbulent transport of heat and momentum in a boundary layer subject to deceleration, suction and variable wall temperature

NASA Technical Reports Server (NTRS)

Orlando, A. F.; Moffat, R. J.; Kays, W. M.

1974-01-01

The relationship between the turbulent transport of heat and momentum in an adverse pressure gradient boundary layer was studied. An experimental study was conducted of turbulent boundary layers subject to strong adverse pressure gradients with suction. Near-equilibrium flows were attained, evidenced by outer-region similarity in terms of defect temperature and defect velocity profiles. The relationship between Stanton number and enthalpy thickness was shown to be the same as for a flat plate flow both for constant wall temperature boundary conditions and for steps in wall temperature. The superposition principle used with the step-wall-temperature experimental result was shown to accurately predict the Stanton number variation for two cases of arbitrarily varying wall temperature. The Reynolds stress tensor components were measured for strong adverse pressure gradient conditions and different suction rates. Two peaks of turbulence intensity were found: one in the inner and one in the outer regions. The outer peak is shown to be displaced outward by an adverse pressure gradient and suppressed by suction.

Numerical Study of Pressure Fluctuations due to a Mach 6 Turbulent Boundary Layer

NASA Technical Reports Server (NTRS)

Duan, Lian; Choudhari, Meelan M.

2013-01-01

Direct numerical simulations (DNS) are used to examine the pressure fluctuations generated by a Mach 6 turbulent boundary layer with nominal freestream Mach number of 6 and Reynolds number of Re(sub t) approx. =. 464. The emphasis is on comparing the primarily vortical pressure signal at the wall with the acoustic freestream signal under higher Mach number conditions. Moreover, the Mach-number dependence of pressure signals is demonstrated by comparing the current results with those of a supersonic boundary layer at Mach 2.5 and Re(sub t) approx. = 510. It is found that the freestream pressure intensity exhibits a strong Mach number dependence, irrespective of whether it is normalized by the mean wall shear stress or by the mean pressure, with the normalized fluctuation amplitude being significantly larger for the Mach 6 case. Spectral analysis shows that both the wall and freestream pressure fluctuations of the Mach 6 boundary layer have enhanced energy content at high frequencies, with the peak of the premultiplied frequency spectrum of freestream pressure fluctuations being at a frequency of omega(delta)/U(sub infinity) approx. = 3.1, which is more than twice the corresponding frequency in the Mach 2.5 case. The space-time correlations indicate that the pressure-carrying eddies for the higher Mach number case are of smaller size, less elongated in the spanwise direction, and convect with higher convection speeds relative to the Mach 2.5 case. The demonstrated Mach-number dependence of the pressure field, including radiation intensity, directionality, and convection speed, is consistent with the trend exhibited in experimental data and can be qualitatively explained by the notion of "eddy Mach wave" radiation.

Receptivity of Hypersonic Boundary Layers to Acoustic and Vortical Disturbances (Invited)

NASA Technical Reports Server (NTRS)

Balakumar, P.

2015-01-01

Boundary-layer receptivity to two-dimensional acoustic and vortical disturbances for hypersonic flows over two-dimensional and axi-symmetric geometries were numerically investigated. The role of bluntness, wall cooling, and pressure gradients on the receptivity and stability were analyzed and compared with the sharp nose cases. It was found that for flows over sharp nose geometries in adiabatic wall conditions the instability waves are generated in the leading-edge region and that the boundary layer is much more receptive to slow acoustic waves as compared to the fast waves. The computations confirmed the stabilizing effect of nose bluntness and the role of the entropy layer in the delay of boundary layer transition. The receptivity coefficients in flows over blunt bodies are orders of magnitude smaller than that for the sharp cone cases. Wall cooling stabilizes the first mode strongly and destabilizes the second mode. However, the receptivity coefficients are also much smaller compared to the adiabatic case. The adverse pressure gradients increased the unstable second mode regions.

Turbulent Output-Based Anisotropic Adaptation

NASA Technical Reports Server (NTRS)

Park, Michael A.; Carlson, Jan-Renee

2010-01-01

Controlling discretization error is a remaining challenge for computational fluid dynamics simulation. Grid adaptation is applied to reduce estimated discretization error in drag or pressure integral output functions. To enable application to high O(10(exp 7)) Reynolds number turbulent flows, a hybrid approach is utilized that freezes the near-wall boundary layer grids and adapts the grid away from the no slip boundaries. The hybrid approach is not applicable to problems with under resolved initial boundary layer grids, but is a powerful technique for problems with important off-body anisotropic features. Supersonic nozzle plume, turbulent flat plate, and shock-boundary layer interaction examples are presented with comparisons to experimental measurements of pressure and velocity. Adapted grids are produced that resolve off-body features in locations that are not known a priori.

Measurement of corner separation zone response on a compression ramp to plasma actuation within the hypersonic boundary layer

NASA Astrophysics Data System (ADS)

Hedlund, Brock E.; Houpt, Alec W.; Gordeyev, Stanislav V.; Leonov, Sergey B.

2017-10-01

This study was performed to characterize the dominant frequencies present in the boundary layer uptsream of and in the corner separation zone of a compression surface in Mach 4.5 flow and to determine a control effect of transient plasma actuation on the boundary layer. Schlieren imaging was used to distinguish the corner separation zone for 20°, 25°, and 30° compression ramps mounted on flat plates. Spectra of the natural disturbances present in the boundary layer and separation zone were gathered using a high-speed Shack-Hartmann wavefront sensor and surface mounted PCBTM pressure sensors while varying flow parameters by adjusting total pressure, temperature, and ramp angle. Shallow cavity discharge plasma actuators were used as a high-frequency localized thermal forcing mechanism of the boundary layer. The plasma effect was negligible for forcing frequencies (50 kHz) below the natural dominant frequency (~55-80 kHz). High frequency perturbations that can promote the transition to turbulence were amplified when the plasma forcing frequency (100 kHz) was higher than the natural dominant frequency (~55-80 kHz). This technique can potentially be used for active control of hypersonic boundary layer transition and the supersonic flow structure on the compression surface.

Boundary-Layer-Ingesting Inlet Flow Control

NASA Technical Reports Server (NTRS)

Owens, Lewis R.; Allan, Brian G.; Gorton, Susan A.

2008-01-01

An experimental study was conducted to provide the first demonstration of an active flow control system for a flush-mounted inlet with significant boundary-layer-ingestion in transonic flow conditions. The effectiveness of the flow control in reducing the circumferential distortion at the engine fan-face location was assessed using a 2.5%-scale model of a boundary-layer-ingesting offset diffusing inlet. The inlet was flush mounted to the tunnel wall and ingested a large boundary layer with a boundary-layer-to-inlet height ratio of 35%. Different jet distribution patterns and jet mass flow rates were used in the inlet to control distortion. A vane configuration was also tested. Finally a hybrid vane/jet configuration was tested leveraging strengths of both types of devices. Measurements were made of the onset boundary layer, the duct surface static pressures, and the mass flow rates through the duct and the flow control actuators. The distortion and pressure recovery were measured at the aerodynamic interface plane. The data show that control jets and vanes reduce circumferential distortion to acceptable levels. The point-design vane configuration produced higher distortion levels at off-design settings. The hybrid vane/jet flow control configuration reduced the off-design distortion levels to acceptable ones and used less than 0.5% of the inlet mass flow to supply the jets.

A fast wind-farm boundary-layer model to investigate gravity wave effects and upstream flow deceleration

NASA Astrophysics Data System (ADS)

Allaerts, Dries; Meyers, Johan

2017-11-01

Wind farm design and control often relies on fast analytical wake models to predict turbine wake interactions and associated power losses. Essential input to these models are the inflow velocity and turbulent intensity at hub height, which come from prior measurement campaigns or wind-atlas data. Recent LES studies showed that in some situations large wind farms excite atmospheric gravity waves, which in turn affect the upstream wind conditions. In the current study, we develop a fast boundary-layer model that computes the excitation of gravity waves and the perturbation of the boundary-layer flow in response to an applied force. The core of the model is constituted by height-averaged, linearised Navier-Stokes equations for the inner and outer layer, and the effect of atmospheric gravity waves (excited by the boundary-layer displacement) is included via the pressure gradient. Coupling with analytical wake models allows us to study wind-farm wakes and upstream flow deceleration in various atmospheric conditions. Comparison with wind-farm LES results shows excellent agreement in terms of pressure and boundary-layer displacement levels. The authors acknowledge support from the European Research Council (FP7-Ideas, Grant No. 306471).

Contraction design for small low-speed wind tunnels

NASA Technical Reports Server (NTRS)

Bell, James H.; Mehta, Rabindra D.

1988-01-01

An iterative design procedure was developed for two- or three-dimensional contractions installed on small, low-speed wind tunnels. The procedure consists of first computing the potential flow field and hence the pressure distributions along the walls of a contraction of given size and shape using a three-dimensional numerical panel method. The pressure or velocity distributions are then fed into two-dimensional boundary layer codes to predict the behavior of the boundary layers along the walls. For small, low-speed contractions it is shown that the assumption of a laminar boundary layer originating from stagnation conditions at the contraction entry and remaining laminar throughout passage through the successful designs if justified. This hypothesis was confirmed by comparing the predicted boundary layer data at the contraction exit with measured data in existing wind tunnels. The measured boundary layer momentum thicknesses at the exit of four existing contractions, two of which were 3-D, were found to lie within 10 percent of the predicted values, with the predicted values generally lower. From the contraction wall shapes investigated, the one based on a fifth-order polynomial was selected for installation on a newly designed mixing layer wind tunnel.

Ground effect for V/STOL aircraft configurations and its simulation in the wind tunnel. Part 1: Introduction and theoretical studies

NASA Technical Reports Server (NTRS)

Hackett, J. E.; Praytor, E. B.

1972-01-01

Theoretical studies are made of three dimensional turbulent boundary layer behavior on fixed grounds and on moving grounds of the type used in wind tunnel tests. It is shown that, for several widely-varying STOL configurations, the ground static pressure distributions possess a remarkable degree of fore-aft symmetry about the center of lift. At low Renolds number, corresponding to small-tunnel testing, the boundary layer displacement surface reflects to a large degree the symmetry of the pressure distribution. For this reason, induced incidence at the model is small for unseparated ground flow. At high Reynolds number, the displacement thickness decrease aft of the static pressure maximum is noticeably more rapid than the corresponding rise. This is attributed to trailing-vortex-induced spanwise pumping within the boundary layer.

Local Characteristics of the Nocturnal Boundary Layer in Response to External Pressure Forcing

NASA Astrophysics Data System (ADS)

van der Linden, Steven; Baas, Peter; van Hooft, Antoon; van Hooijdonk, Ivo; Bosveld, Fred; van de Wiel, Bas

2017-04-01

Geostrophic wind speed data, derived from pressure observations, are used in combination with tower measurements to investigate the nocturnal stable boundary layer at Cabauw, The Netherlands. Since the geostrophic wind speed is not directly influenced by local nocturnal stability, it may be regarded as an external forcing parameter of the nocturnal stable boundary layer. This is in contrast to local parameters such as in situ wind speed, the Monin-Obukhov stability parameter (z/L) or the local Richardson number. To characterize the stable boundary layer, ensemble averages of clear-sky nights with similar geostrophic wind speed are formed. In this manner, the mean dynamical behavior of near-surface turbulent characteristics, and composite profiles of wind and temperature is systematically investigated. We find that the classification results in a gradual ordering of the diagnosed variables in terms of the geostrophic wind speed. In an ensemble sense the transition from the weakly stable to very stable boundary layer is more gradual than expected. Interestingly, for very weak geostrophic winds turbulent activity is found to be negligibly small while the resulting boundary cooling stays finite. Realistic numerical simulations for those cases should therefore have a a solid description of other thermodynamic processes such as soil heat conduction and radiative transfer. This prerequisite poses a challenge for Large-Eddy Simulations of weak wind nocturnal boundary layers.

An Optimization Technique for the Development of Two-Dimensional Steady Turbulent Boundary Layer Models.

DTIC Science & Technology

1982-03-01

observed coherent structure of the wall layer flow and will now be briefly described. Over the past decade, it has been well documented (see, for example...D2, and x are all arbitrary constants. Equilibrium flows have been examined experimentally for a number of years and an equilibrium boundary layer...CP93, Paper No. 27, 6. Clauser, F.H. (1954). "Turbulent Boundary Layers in Adverse Pressure Gradients", J. Aeronaut. Sci., 21, pp. 91-108. 7. Clauser

Numerical investigations of shock wave interaction with laminar boundary layer on compressor profile

NASA Astrophysics Data System (ADS)

Piotrowicz, M.; Flaszyński, P.

2016-10-01

The investigation of shockwave boundary layer interaction on suction side of transonic compressor blade is one of main objectives of TFAST project (Transition Location Effect on Shock Wave Boundary Layer Interaction). In order to look more closely into the flow structure on suction side of a profile, a design of generic test section in linear transonic wind tunnel was proposed. The experimental and numerical results of flow structure on a suction side of the compressor profile investigations are presented. The numerical simulations are carried out for EARSM (Explicit Algebraic Reynolds Stress Model) turbulence model with transition model. The result are compared with oil flow visualisation, schlieren pictures, Pressure Sensitive Paint (PSP) and static pressure.

An investigation of bleed configurations and their effect on shock wave/boundary layer interactions

NASA Technical Reports Server (NTRS)

Hamed, Awatef

1995-01-01

The design of high efficiency supersonic inlets is a complex task involving the optimization of a number of performance parameters such as pressure recovery, spillage, drag, and exit distortion profile, over the flight Mach number range. Computational techniques must be capable of accurately simulating the physics of shock/boundary layer interactions, secondary corner flows, flow separation, and bleed if they are to be useful in the design. In particular, bleed and flow separation, play an important role in inlet unstart, and the associated pressure oscillations. Numerical simulations were conducted to investigate some of the basic physical phenomena associated with bleed in oblique shock wave boundary layer interactions that affect the inlet performance.

F-16XL ship #1 CAWAP flight

NASA Technical Reports Server (NTRS)

1996-01-01

The single-seat F-16XL (ship #1) makes another run during the Cranked-Arrow Wing Aerodynamic Project (CAWAP) at Dryden Flight Research Center, Edwards, California. The modified airplane features a delta 'cranked-arrow' wing with strips of tubing along the leading edge to the trailing edge to sense static on the wing and obtain pressure distribution data. The right wing receives data on pressure distribution and the left wing has three types of instrumentation - preston tubes to measure local skin friction, boundary layer rakes to measure boundary layer profiles (the layer where the air interacts with the surfaces of a moving aircraft), and hot films to determine boundary layer transition locations. The first flight of CAWAP occurred on November 21, 1995, and the test program ended in April 1996.

Reynolds stress structures in a self-similar adverse pressure gradient turbulent boundary layer at the verge of separation.

NASA Astrophysics Data System (ADS)

Atkinson, C.; Sekimoto, A.; Jiménez, J.; Soria, J.

2018-04-01

Mean Reynolds stress profiles and instantaneous Reynolds stress structures are investigated in a self-similar adverse pressure gradient turbulent boundary layer (APG-TBL) at the verge of separation using data from direct numerical simulations. The use of a self-similar APG-TBL provides a flow domain in which the flow gradually approaches a constant non-dimensional pressure gradient, resulting in a flow in which the relative contribution of each term in the governing equations is independent of streamwise position over a domain larger than two boundary layer thickness. This allows the flow structures to undergo a development that is less dependent on the upstream flow history when compared to more rapidly decelerated boundary layers. This APG-TBL maintains an almost constant shape factor of H = 2.3 to 2.35 over a momentum thickness based Reynolds number range of Re δ 2 = 8420 to 12400. In the APG-TBL the production of turbulent kinetic energy is still mostly due to the correlation of streamwise and wall-normal fluctuations, 〈uv〉, however the contribution form the other components of the Reynolds stress tensor are no longer negligible. Statistical properties associated with the scale and location of sweeps and ejections in this APG-TBL are compared with those of a zero pressure gradient turbulent boundary layer developing from the same inlet profile, resulting in momentum thickness based range of Re δ 2 = 3400 to 3770. In the APG-TBL the peak in both the mean Reynolds stress and the production of turbulent kinetic energy move from the near wall region out to a point consistent with the displacement thickness height. This is associated with a narrower distribution of the Reynolds stress and a 1.6 times higher relative number of wall-detached negative uv structures. These structures occupy 5 times less of the boundary layer volume and show a similar reduction in their streamwise extent with respect to the boundary layer thickness. A significantly lower percentage of wall-attached structures is observed in the present case when compared with a similar investigation of a rapidly decelerating APG-TBL, suggesting that these wall-attached features could be the remanent from the lower pressure gradient domain upstream.

Experimental studies on the stability and transition of 3-dimensional boundary layers

NASA Technical Reports Server (NTRS)

Nitschke-Kowsky, P.

1987-01-01

Three-dimensional unstable boundary layers were investigated as to their characteristic instabilities, leading to turbulence. Standing cross-flow instabilities and traveling waves preceding the transition were visualized with the hydrogen bubble technique in the boundary layer above the wall of a swept cylinder. With the sublimation method and hot film technique, a model consisting of a swept flat plate with a pressure-inducing displacement body in the 1 m wind tunnel was studied. Standing waves and traveling waves in a broad frequency are observed. The boundary layer of this model is close to the assumptions of the theory.

An experimental investigation of wall boundary layer transition Reynolds numbers in an expansion tube

NASA Technical Reports Server (NTRS)

Weilmuenster, K. J.

1974-01-01

Experimental measurements of boundary-layer transition in an expansion-tube test-gas flow are presented along with radial distributions of pitot pressure. An integral method for calculating constant Reynolds number lines for an expansion-tube flow is introduced. Comparison of experimental data and constant Reynolds number calculations has shown that for given conditions, wall boundary-layer transition occurs at a constant Reynolds number in an expansion-tube flow. Operating conditions in the expansion tube were chosen so that the effects of test-gas nonequilibrium on boundary-layer transition could be studied.

Gas diffusion in and out of super-hydrophobic surface in transitional and turbulent boundary layers

NASA Astrophysics Data System (ADS)

Ling, Hangjian; Fu, Matthew; Hultmark, Marcus; Katz, Joseph

2017-11-01

The rate of gas diffusion in and out of a super-hydrophobic surface (SHS) located in boundary layers is investigated at varying Reynolds numbers and ambient pressures. The hierarchical SHS consists of nano-textured, 100 μm wide spanwise grooves. The boundary layers over the SHS under the Cassie-Baxter and Wenzel states as well as a smooth wall at same conditions are characterized by particle image velocimetry. The Reynolds number based on momentum thickness of the smooth wall, ReΘ0, ranges from 518 to 2088, covering transitional and turbulent boundary layer regimes. The mass diffusion rate is estimated by using microscopy to measure the time-evolution of plastron shape and volume. The data is used for calculating the Sherwood number based on smooth wall momentum thickness, ShΘ0. As expected, the diffusion rate increases linearly with the under- or super-saturation level, i.e., ShΘ0 is independent of ambient pressure. For the turbulent boundary layers, the data collapses onto ShΘ0 = 0.47ReΘ00.77 . For the transitional boundary layer, ShΘ0 is lower than the turbulent power law. When ShΘ0 is plotted against the friction Reynolds number (Reτ0) , both the transitional and turbulent boundary layer data collapse onto a single power law, ShΘ0 = 0.34Reτ00.913 . Results scaled based on Wenzel state momentum thickness show very similar trends. Sponsored by ONR.

Including fluid shear viscosity in a structural acoustic finite element model using a scalar fluid representation

PubMed Central

Cheng, Lei; Li, Yizeng; Grosh, Karl

2013-01-01

An approximate boundary condition is developed in this paper to model fluid shear viscosity at boundaries of coupled fluid-structure system. The effect of shear viscosity is approximated by a correction term to the inviscid boundary condition, written in terms of second order in-plane derivatives of pressure. Both thin and thick viscous boundary layer approximations are formulated; the latter subsumes the former. These approximations are used to develop a variational formation, upon which a viscous finite element method (FEM) model is based, requiring only minor modifications to the boundary integral contributions of an existing inviscid FEM model. Since this FEM formulation has only one degree of freedom for pressure, it holds a great computational advantage over the conventional viscous FEM formulation which requires discretization of the full set of linearized Navier-Stokes equations. The results from thick viscous boundary layer approximation are found to be in good agreement with the prediction from a Navier-Stokes model. When applicable, thin viscous boundary layer approximation also gives accurate results with computational simplicity compared to the thick boundary layer formulation. Direct comparison of simulation results using the boundary layer approximations and a full, linearized Navier-Stokes model are made and used to evaluate the accuracy of the approximate technique. Guidelines are given for the parameter ranges over which the accurate application of the thick and thin boundary approximations can be used for a fluid-structure interaction problem. PMID:23729844

Including fluid shear viscosity in a structural acoustic finite element model using a scalar fluid representation.

PubMed

Cheng, Lei; Li, Yizeng; Grosh, Karl

2013-08-15

An approximate boundary condition is developed in this paper to model fluid shear viscosity at boundaries of coupled fluid-structure system. The effect of shear viscosity is approximated by a correction term to the inviscid boundary condition, written in terms of second order in-plane derivatives of pressure. Both thin and thick viscous boundary layer approximations are formulated; the latter subsumes the former. These approximations are used to develop a variational formation, upon which a viscous finite element method (FEM) model is based, requiring only minor modifications to the boundary integral contributions of an existing inviscid FEM model. Since this FEM formulation has only one degree of freedom for pressure, it holds a great computational advantage over the conventional viscous FEM formulation which requires discretization of the full set of linearized Navier-Stokes equations. The results from thick viscous boundary layer approximation are found to be in good agreement with the prediction from a Navier-Stokes model. When applicable, thin viscous boundary layer approximation also gives accurate results with computational simplicity compared to the thick boundary layer formulation. Direct comparison of simulation results using the boundary layer approximations and a full, linearized Navier-Stokes model are made and used to evaluate the accuracy of the approximate technique. Guidelines are given for the parameter ranges over which the accurate application of the thick and thin boundary approximations can be used for a fluid-structure interaction problem.

Experimental study of the separating confluent boundary-layer. Volume 2: Experimental data

NASA Technical Reports Server (NTRS)

Braden, J. A.; Whipkey, R. R.; Jones, G. S.; Lilley, D. E.

1983-01-01

An experimental low speed study of the separating confluent boundary layer on a NASA GAW-1 high lift airfoil is described. The airfoil was tested in a variety of high lift configurations comprised of leading edge slat and trailing edge flap combinations. The primary test instrumentation was a two dimensional laser velocimeter (LV) system operating in a backscatter mode. Surface pressures and corresponding LV derived boundary layer profiles are given in terms of velocity components, turbulence intensities and Reynolds shear stresses as characterizing confluent boundary layer behavior up to and beyond stall. LV derived profiles and associated boundary layer parameters and those obtained from more conventional instrumentation such as pitot static transverse, Preston tube measurements and hot-wire surveys are compared.

Bypass transition in boundary layers including curvature and favorable pressure gradient effects

NASA Technical Reports Server (NTRS)

Volino, R. J.; Simon, T. W.

1991-01-01

Recent studies of 2-D boundary layers undergoing bypass transition were reviewed. Bypass transition is characterized by the sudden appearance of turbulent spots in boundary layer without first the regular, observable growth of disturbances predicted by linear stability theory. There are no standard criteria or parameters for defining bypass transition, but it is known to be the mode of transition when the flow is disturbed by perturbations of sufficient amplitude.

The Hydroacoustics of Beveled Steps and Gaps

DTIC Science & Technology

2015-11-01

2014). "Pressure fluctuations produced by forward steps immersed in a turbulent boundary layer ." Journal of Fluid Mechanics 756: 384-421. Catlett...M., W. J. Devenport and S. Glegg (2014). "Sound from boundary layer flow over steps and gaps." Journal of Sound and Vibration 333: 4170-4186...Conference, Berlin, June 2013 Catlett, MR., Devenport, W. and Glegg, S., 2014, "Sound from Boundary Layer Flow over Steps and Gaps", Journal of Sound

The structure of a three-dimensional turbulent boundary layer

NASA Technical Reports Server (NTRS)

Degani, A. T.; Smith, F. T.; Walker, J. D. A.

1993-01-01

The three-dimensional turbulent boundary layer is shown to have a self-consistent two-layer asymptotic structure in the limit of large Reynolds number. In a streamline coordinate system, the streamwise velocity distribution is similar to that in two-dimensional flows, having a defect-function form in the outer layer which is adjusted to zero at the wall through an inner wall layer. An asymptotic expansion accurate to two orders is required for the cross-stream velocity which is shown to exhibit a logarithmic form in the overlap region. The inner wall-layer flow is collateral to leading order but the influence of the pressure gradient, at large but finite Reynolds numbers, is not negligible and can cause substantial skewing of the velocity profile near the wall. Conditions under which the boundary layer achieves self-similarity and the governing set of ordinary differential equations for the outer layer are derived. The calculated solution of these equations is matched asymptotically to an inner wall-layer solution and the composite profiles so formed describe the flow throughout the entire boundary layer. The effects of Reynolds number and cross-stream pressure gradient on the crossstream velocity profile are discussed and it is shown that the location of the maximum cross-stream velocity is within the overlap region.

Vortex Generators in a Two-Dimensional, External-Compression Supersonic Inlet

NASA Technical Reports Server (NTRS)

Baydar, Ezgihan; Lu, Frank K.; Slater, John W.

2016-01-01

Vortex generators within a two-dimensional, external-compression supersonic inlet for Mach 1.6 were investigated to determine their ability to increase total pressure recovery, reduce total pressure distortion, and improve the boundary layer. The vortex generators studied included vanes and ramps. The geometric factors of the vortex generators studied included height, length, spacing, and positions upstream and downstream of the inlet terminal shock. The flow through the inlet was simulated through the computational solution of the steady-state Reynolds-averaged Navier-Stokes equations on multi-block, structured grids. The vortex generators were simulated by either gridding the geometry of the vortex generators or modeling the vortices generated by the vortex generators. The inlet performance was characterized by the inlet total pressure recovery, total pressure distortion, and incompressible shape factor of the boundary-layer at the engine face. The results suggested that downstream vanes reduced the distortion and improved the boundary layer. The height of the vortex generators had the greatest effect of the geometric factors.

Analysis and calculation by integral methods of laminar compressible boundary-layer with heat transfer and with and without pressure gradient

NASA Technical Reports Server (NTRS)

Morduchow, Morris

1955-01-01

A survey of integral methods in laminar-boundary-layer analysis is first given. A simple and sufficiently accurate method for practical purposes of calculating the properties (including stability) of the laminar compressible boundary layer in an axial pressure gradient with heat transfer at the wall is presented. For flow over a flat plate, the method is applicable for an arbitrarily prescribed distribution of temperature along the surface and for any given constant Prandtl number close to unity. For flow in a pressure gradient, the method is based on a Prandtl number of unity and a uniform wall temperature. A simple and accurate method of determining the separation point in a compressible flow with an adverse pressure gradient over a surface at a given uniform wall temperature is developed. The analysis is based on an extension of the Karman-Pohlhausen method to the momentum and the thermal energy equations in conjunction with fourth- and especially higher degree velocity and stagnation-enthalpy profiles.

Interference heating from interactions of shock waves with turbulent boundary layers at Mach 6

NASA Technical Reports Server (NTRS)

Johnson, C. B.; Kaufman, L. G., II

1974-01-01

An experimental investigation of interference heating resulting from interactions of shock waves and turbulent boundary layers was conducted. Pressure and heat-transfer distributions were measured on a flat plate in the free stream and on the wall of the test section of the Langley Mach 6 high Reynolds number tunnel for Reynolds numbers ranging from 2 million to 400 million. Various incident shock strengths were obtained by varying a wedge-shock generator angle (from 10 deg to 15 deg) and by placing a spherical-shock generator at different vertical positions above the instrumented flat plate and tunnel wall. The largest heating-rate amplification factors obtained for completely turbulent boundary layers were 22.1 for the flat plate and 11.6 for the tunnel wall experiments. Maximum heating correlated with peak pressures using a power law with a 0.85 exponent. Measured pressure distributions were compared with those calculated using turbulent free-interaction pressure rise theories, and separation lengths were compared with values calculated by using different methods.

Measurements in a Transitional Boundary Layer Under Low-Pressure Turbine Airfoil Conditions

NASA Technical Reports Server (NTRS)

Simon, Terrence W.; Qiu, Songgang; Yuan, Kebiao; Ashpis, David (Technical Monitor); Simon, Fred (Technical Monitor)

2000-01-01

This report presents the results of an experimental study of transition from laminar to turbulent flow in boundary layers or in shear layers over separation zones on a convex-curved surface which simulates the suction surface of a low-pressure turbine airfoil. Flows with various free-stream turbulence intensity (FSTI) values (0.5%, 2.5% and 10%), and various Reynolds numbers (50,000, 100,000 200,000 and 300,000) are investigated. Reynold numbers in the present study are based on suction surface length and passage exit mean velocity. Flow separation followed by transition within the separated flow region is observed for the lower-Re cases at each of the FSTI levels. At the highest Reynolds numbers and at elevated FSn, transition of the attached boundary layer begins before separation, and the separation zone is small. Transition proceeds in the shear layer over the separation bubble. For both the transitional boundary layer and the transitional shear layer, mean velocity, turbulence intensity and intermittency (the fraction of the time the flow is turbulent) distributions are presented. The present data are compared to published distribution models for bypass transition, intermittency distribution through transition, transition start position, and transition length. A model developed for transition of separated flows is shown to adequately predict the location of the beginning of transition, for these cases, and a model developed for transitional boundary layer flows seems to adequately predict the path of intermittency through transition when the transition start and end are known. These results are useful for the design of low-pressure turbine stages which are known to operate under conditions replicated by these tests.

Modeling Disturbance Dynamics in Transitional and Turbulent Boundary Layers

NASA Technical Reports Server (NTRS)

Grosch, C. E.; Gatski, T. B. (Technical Monitor)

2002-01-01

The dynamics of an ensemble of linear disturbances in boundary-layer flows at various Reynolds numbers is studied through an analysis of the transport equations for the mean disturbance kinetic energy and energy dissipation rate. Effects of adverse and favorable pressure-gradients on the disturbance dynamics are also included in the analysis. Unlike the fully turbulent regime where nonlinear phase scrambling of the fluctuations affects the flow field even in proximity to the wall, the early stage transition regime fluctuations studied here are influenced across the boundary layer by the solid boundary. In addition, the dominating dynamics in the disturbance kinetic energy equation is governed by the energy production, pressure-transport and viscous diffusion - also in contrast to the fully turbulent regime. For the disturbance dissipation rate, a dynamic balance exists between the destruction and diffusion of dissipation.

Thermocouple Rakes for Measuring Boundary Layer Flows Extremely Close to Surface

NASA Technical Reports Server (NTRS)

Hwang, Danny P.; Fralick, Gustave C.; Martin, Lisa C.; Blaha, Charles A.

2001-01-01

Of vital interest to aerodynamic researchers is precise knowledge of the flow velocity profile next to the surface. This information is needed for turbulence model development and the calculation of viscous shear force. Though many instruments can determine the flow velocity profile near the surface, none of them can make measurements closer than approximately 0.01 in. from the surface. The thermocouple boundary-layer rake can measure much closer to the surface than conventional instruments can, such as a total pressure boundary layer rake, hot wire, or hot film. By embedding the sensors (thermocouples) in the region where the velocity is equivalent to the velocity ahead of a constant thickness strut, the boundary-layer flow profile can be obtained. The present device fabricated at the NASA Glenn Research Center microsystem clean room has a heater made of platinum and thermocouples made of platinum and gold. Equal numbers of thermocouples are placed both upstream and downstream of the heater, so that the voltage generated by each pair at the same distance from the surface is indicative of the difference in temperature between the upstream and downstream thermocouple locations. This voltage differential is a function of the flow velocity, and like the conventional total pressure rake, it can provide the velocity profile. In order to measure flow extremely close to the surface, the strut is made of fused quartz with extremely low heat conductivity. A large size thermocouple boundary layer rake is shown in the following photo. The latest medium size sensors already provide smooth velocity profiles well into the boundary layer, as close as 0.0025 in. from the surface. This is about 4 times closer to the surface than the previously used total pressure rakes. This device also has the advantage of providing the flow profile of separated flow and also it is possible to measure simultaneous turbulence levels within the boundary layer.

Interaction of solar wind with the magnetopause-boundary layer and generation of magnetic impulse events

NASA Technical Reports Server (NTRS)

Lee, L. C.; Wei, C. Q.

1993-01-01

The transport of mass, momentum, energy and waves from the solar wind to the Earth's magnetosphere takes place in the magnetopause-boundary layer region. Various plasma processes that may occur in this region have been proposed and studied. In this paper, we present a brief review of the plasma processes in the dayside magnetopause-boundary layer. These processes include (1) flux transfer events at the dayside magnetopause, (2) formation of plasma vortices in the low-latitude boundary layer by the Kelvin-Helmholtz instability and coupling to the polar ionosphere, (3) the response of the magnetopause to the solar wind dynamic pressure pulses, and (4) the impulsive penetration of solar wind plasma filaments through the dayside magnetopause into the magnetospheric boundary layer. Through the coupling of the magnetopause-boundary layer to the polar ionosphere, those above processes may lead to occurrence of magnetic impulse events observed in the high-latitude stations.

Boundary layer transition detection on the X-15 vertical fin using surface-pressure-fluctuation measurements

NASA Technical Reports Server (NTRS)

Lewis, T. L.; Banner, R. D.

1971-01-01

A flush-mounted microphone on the vertical fin of an X-15 airplane was used to investigate boundary layer transition phenomenon during flights to peak altitudes of approximately 70,000 meters. The flight results were compared with those from wind tunnel studies, skin temperature measurements, and empirical prediction data. The Reynolds numbers determined for the end of transition were consistent with those obtained from wind tunnel studies. Maximum surface-pressure-fluctuation coefficients in the transition region were about an order of magnitude greater than those for fully developed turbulent flow. This was also consistent with wind tunnel data. It was also noted that the power-spectral-density estimates of the surface-pressure fluctuations were characterized by a shift in power from high frequencies to low frequencies as the boundary layer changed from turbulent to laminar flow. Large changes in power at the lowest frequencies appeared to mark the beginning of transition.

An Approach to the Constrained Design of Natural Laminar Flow Airfoils

NASA Technical Reports Server (NTRS)

Green, Bradford E.

1997-01-01

A design method has been developed by which an airfoil with a substantial amount of natural laminar flow can be designed, while maintaining other aerodynamic and geometric constraints. After obtaining the initial airfoil's pressure distribution at the design lift coefficient using an Euler solver coupled with an integral turbulent boundary layer method, the calculations from a laminar boundary layer solver are used by a stability analysis code to obtain estimates of the transition location (using N-Factors) for the starting airfoil. A new design method then calculates a target pressure distribution that will increase the laminar flow toward the desired amount. An airfoil design method is then iteratively used to design an airfoil that possesses that target pressure distribution. The new airfoil's boundary layer stability characteristics are determined, and this iterative process continues until an airfoil is designed that meets the laminar flow requirement and as many of the other constraints as possible.

An approach to the constrained design of natural laminar flow airfoils

NASA Technical Reports Server (NTRS)

Green, Bradford Earl

1995-01-01

A design method has been developed by which an airfoil with a substantial amount of natural laminar flow can be designed, while maintaining other aerodynamic and geometric constraints. After obtaining the initial airfoil's pressure distribution at the design lift coefficient using an Euler solver coupled with an integml turbulent boundary layer method, the calculations from a laminar boundary layer solver are used by a stability analysis code to obtain estimates of the transition location (using N-Factors) for the starting airfoil. A new design method then calculates a target pressure distribution that will increase the larninar flow toward the desired amounl An airfoil design method is then iteratively used to design an airfoil that possesses that target pressure distribution. The new airfoil's boundary layer stability characteristics are determined, and this iterative process continues until an airfoil is designed that meets the laminar flow requirement and as many of the other constraints as possible.

Similar solutions for the compressible laminar boundary layer with heat transfer and pressure gradient

NASA Technical Reports Server (NTRS)

Cohen, Clarence B; Reshotko, Eli

1956-01-01

Stewartson's transformation is applied to the laminar compressible boundary-layer equations and the requirement of similarity is introduced, resulting in a set of ordinary nonlinear differential equations previously quoted by Stewartson, but unsolved. The requirements of the system are Prandtl number of 1.0, linear viscosity-temperature relation across the boundary layer, an isothermal surface, and the particular distributions of free-stream velocity consistent with similar solutions. This system admits axial pressure gradients of arbitrary magnitude, heat flux normal to the surface, and arbitrary Mach numbers. The system of differential equations is transformed to integral system, with the velocity ratio as the independent variable. For this system, solutions are found by digital computation for pressure gradients varying from that causing separation to the infinitely favorable gradient and for wall temperatures from absolute zero to twice the free-stream stagnation temperature. Some solutions for separated flows are also presented.

Application of rotor mounted pressure transducers to analysis of inlet turbulence. [flow distortion in turbofan engine inlet

NASA Technical Reports Server (NTRS)

Hanson, D. B.

1976-01-01

Miniature pressure transducers installed near the leading edge of a fan blade were used to diagnose the non-uniform flow entering a subsonic tip speed turbofan on a static test stand. The pressure response of the blade to the inlet flow variations was plotted in a form which shows the space-time history of disturbances ingested by the rotor. Also, periodically sampled data values were auto- and cross-correlated as if they had been acquired from fixed hot wire anemometers at 150 equally spaced angles around the inlet. With a clean inlet and low wind, evidence of long, narrow turbulence eddies was easily found both in the boundary layer of the fan duct and outside the boundary layer. The role of the boundary layer was to follow and amplify disturbances in the outer flow. These eddies frequently moved around the inlet with a corkscrew motion as they passed through.

Results of correlations for transition location on a clean-up glove installed on an F-14 aircraft and design studies for a laminar glove for the X-29 aircraft accounting for spanwise pressure gradient

NASA Technical Reports Server (NTRS)

Goradia, S. H.; Bobbitt, P. J.; Morgan, H. L.; Ferris, J. C.; Harvey, William D.

1989-01-01

Results of correlative and design studies for transition location, laminar and turbulent boundary-layer parameters, and wake drag for forward swept and aft swept wings are presented. These studies were performed with the use of an improved integral-type boundary-layer and transition-prediction methods. Theoretical predictions were compared with flight measurements at subsonic and transonic flow conditions for the variable aft swept wing F-14 aircraft for which experimental pressure distributions, transition locations, and turbulent boundary-layer velocity profiles were measured. Flight data were available at three spanwise stations for several values of sweep, freestream unit Reynolds number, Mach numbers, and lift coefficients. Theory/experiment correlations indicate excellent agreement for both transition location and turbulent boundary-layer parameters. The results of parametric studies performed during the design of a laminar glove for the forward swept wing X-29 aircraft are also presented. These studies include the effects of a spanwise pressure gradient on transition location and wake drag for several values of freestream Reynolds numbers at a freestream Mach number of 0.9.

Experimental Evaluation of the Effect of Angle-of-attack on the External Aerodynamics and Mass Capture of a Symmetric Three-engine Air-breathing Launch Vehicle Configuration at Supersonic Speeds

NASA Technical Reports Server (NTRS)

Kim, Hyun D.; Frate, Franco C.

2001-01-01

A subscale aerodynamic model of the GTX air-breathing launch vehicle was tested at NASA Glenn Research Center's 10- by 10-Foot Supersonic Wind Tunnel from Mach 2.0 to 3.5 at various angles-of-attack. The objective of the test was to investigate the effect of angle-of-attack on inlet mass capture, inlet diverter effectiveness, and the flowfield at the cowl lip plane. The flow-through inlets were tested with and without boundary-layer diverters. Quantitative measurements such as inlet mass flow rates and pitot-pressure distributions in the cowl lip plane are presented. At a 3deg angle-of-attack, the flow rates for the top and side inlets were within 8 percent of the zero angle-of-attack value, and little distortion was evident at the cowl lip plane. Surface oil flow patterns showing the shock/boundary-layer interaction caused by the inlet spikes are shown. In addition to inlet data, vehicle forebody static pressure distributions, boundary-layer profiles, and temperature-sensitive paint images to evaluate the boundary-layer transition are presented. Three-dimensional parabolized Navier-Stokes computational fluid dynamics calculations of the forebody flowfield are presented and show good agreement with the experimental static pressure distributions and boundary-layer profiles. With the boundary-layer diverters installed, no adverse aerodynamic phenomena were found that would prevent the inlets from operating at the required angles-of-attack. We recommend that phase 2 of the test program be initiated, where inlet contraction ratio and diverter geometry variations will be tested.

Contraction design for small low-speed wind tunnels

NASA Technical Reports Server (NTRS)

Bell, James H.; Mehta, Rabindra D.

1988-01-01

An iterative design procedure was developed for 2- or 3-dimensional contractions installed on small, low speed wind tunnels. The procedure consists of first computing the potential flow field and hence the pressure distributions along the walls of a contraction of given size and shape using a 3-dimensional numerical panel method. The pressure or velocity distributions are then fed into 2-dimensional boundary layer codes to predict the behavior of the boundary layers along the walls. For small, low speed contractions, it is shown that the assumption of a laminar boundary layer originating from stagnation conditions at the contraction entry and remaining laminar throughout passage through the successful designs is justified. This hypothesis was confirmed by comparing the predicted boundary layer data at the contraction exit with measured data in existing wind tunnels. The measured boundary layer momentum thicknesses at the exit of four existing contractions, two of which were 3-D, were found to lie within 10 percent of the predicted values, with the predicted values generally lower. From the contraction wall shapes investigated, the one based on a 5th order polynomial was selected for newly designed mixing wind tunnel installation.

Assessment of Turbulent Shock-Boundary Layer Interaction Computations Using the OVERFLOW Code

NASA Technical Reports Server (NTRS)

Oliver, A. B.; Lillard, R. P.; Schwing, A. M.; Blaisdell, G> A.; Lyrintzis, A. S.

2007-01-01

The performance of two popular turbulence models, the Spalart-Allmaras model and Menter s SST model, and one relatively new model, Olsen & Coakley s Lag model, are evaluated using the OVERFLOWcode. Turbulent shock-boundary layer interaction predictions are evaluated with three different experimental datasets: a series of 2D compression ramps at Mach 2.87, a series of 2D compression ramps at Mach 2.94, and an axisymmetric coneflare at Mach 11. The experimental datasets include flows with no separation, moderate separation, and significant separation, and use several different experimental measurement techniques (including laser doppler velocimetry (LDV), pitot-probe measurement, inclined hot-wire probe measurement, preston tube skin friction measurement, and surface pressure measurement). Additionally, the OVERFLOW solutions are compared to the solutions of a second CFD code, DPLR. The predictions for weak shock-boundary layer interactions are in reasonable agreement with the experimental data. For strong shock-boundary layer interactions, all of the turbulence models overpredict the separation size and fail to predict the correct skin friction recovery distribution. In most cases, surface pressure predictions show too much upstream influence, however including the tunnel side-wall boundary layers in the computation improves the separation predictions.

Prediction of mean flow data for adiabatic 2-D compressible turbulent boundary layers

NASA Astrophysics Data System (ADS)

Motallebi, Fariborz

1995-02-01

This report presents a method for the prediction of mean flow data (i.e. , skin friction, velocity profile, and shape parameter) for adiabatic two-dimensional compressible turbulent boundary layers at zero pressure gradient. The transformed law of the wall, law of the wake, the van Driest model for the complete inner region, and a correlation between the Reynolds number based on the boundary layer integral length scale (Re(sub Delta*)) and the Reynolds number based on the boundary layer momentum thickness (Re(sub theta)) were used to predict the mean flow quantities. The results for skin friction coefficient show good agreement with a number of existing theories including those of van Driest and Huang et al. Comparison with a large number of experimental data suggests that at least for transonic and supersonic flows, the velocity profile as described by van Driest and Coles is Reynolds number dependent and should not be presumed universal. Extra information or perhaps a better physical approach to the formulation of the mean structure of compressible turbulent boundary layers, even in zero pressure gradient and adiabatic condition, is required in order to achieve complete (physical and mathematical) convergence when it is applied in any prediction methods.

F-16XL ship #1 outboard rake #7

NASA Technical Reports Server (NTRS)

1996-01-01

This photo shows the #7 outboard rake on the left wing of NASA's single-seat F-16XL (ship #1) used for the Cranked-Arrow Wing Aerodynamic Project (CAWAP) at Dryden Flight Research Center, Edwards, California. The modified airplane features a delta 'cranked-arrow' wing with strips of tubing along the leading edge to the trailing edge to sense static on the wing and obtain pressure distribution data. The right wing receives data on pressure distribution and the left wing has three types of instrumentation - preston tubes to measure local skin friction, boundary layer rakes to measure boundary layer profiles (the layer where the air interacts with the surfaces of a moving aircraft), and hot films to determine boundary layer transition locations. The first flight of CAWAP occurred on November 21, 1995, and the test program ended in April 1996.

The measurement of boundary layers on a compressor blade in cascade. Volume 2: Data tables

NASA Technical Reports Server (NTRS)

Zierke, William C.; Deutsch, Steven

1989-01-01

Measurements were made of the boundary layers and wakes about a highly loaded, double-circular-arc compressor blade in cascade. These laser Doppler velocimetry measurements have yielded a very detailed and precise data base with which to test the application of viscous computational codes to turbomachinery. In order to test the computational codes at off-design conditions, the data have been acquired at a chord Reynolds number of 500,000 and at three incidence angles. Average values and 95 percent confidence bands were tabularized for the velocity, local turbulence intensity, skewness, kurtosis, and percent backflow. Tables also exist for the blade static-pressure distributions and boundary layer velocity profiles reconstructed to account for the normal pressure gradient.

Assessment of a 3-D boundary layer code to predict heat transfer and flow losses in a turbine

NASA Technical Reports Server (NTRS)

Anderson, O. L.

1984-01-01

Zonal concepts are utilized to delineate regions of application of three-dimensional boundary layer (DBL) theory. The zonal approach requires three distinct analyses. A modified version of the 3-DBL code named TABLET is used to analyze the boundary layer flow. This modified code solves the finite difference form of the compressible 3-DBL equations in a nonorthogonal surface coordinate system which includes coriolis forces produced by coordinate rotation. These equations are solved using an efficient, implicit, fully coupled finite difference procedure. The nonorthogonal surface coordinate system is calculated using a general analysis based on the transfinite mapping of Gordon which is valid for any arbitrary surface. Experimental data is used to determine the boundary layer edge conditions. The boundary layer edge conditions are determined by integrating the boundary layer edge equations, which are the Euler equations at the edge of the boundary layer, using the known experimental wall pressure distribution. Starting solutions along the inflow boundaries are estimated by solving the appropriate limiting form of the 3-DBL equations.

The inducement of planetary boundary layer mass convergence associated with varying vorticity beneath tropospheric wind maximum

NASA Technical Reports Server (NTRS)

Johnson, D. R.

1984-01-01

The effects of the vorticity distribution are applied to study planetary boundary layer mass convergence beneath free tropospheric wind maximum. For given forcing by viscous and pressure gradient forces beneath a wind maximum, boundary layer cross stream mass transport is increased by anticyclonic vorticity on the right flank and decreased by cyclonic vorticity on the left flank. Such frictionally forced mass transport induces boundary layer mass convergence beneath the relative wind maximum. This result is related to the empirical rule that the most intense convection and severe weather frequently develop beneath the 500 mb zero relative vorticity isopleth.

Effects of boundary layer refraction and fuselage scattering on fuselage surface noise from advanced turboprop propellers

NASA Technical Reports Server (NTRS)

Mcaninch, G. L.; Rawls, J. W., Jr.

1984-01-01

An acoustic disturbance's propagation through a boundary layer is discussed with a view to the analysis of the acoustic field generated by a propfan rotor incident to the fuselage of an aircraft. Applying the parallel flow assumption, the resulting partial differential equations are reduced to an ordinary acoustic pressure differential equation by means of the Fourier transform. The methods used for the solution of this equation include those of Frobenius and of analytic continuation; both yield exact solutions in series form. Two models of the aircraft fuselage-boundary layer system are considered, in the first of which the fuselage is replaced by a flat plate and the acoustic field is assumed to be two-dimensional, while in the second the fuselage is a cylinder in a fully three-dimensional acoustic field. It is shown that the boundary layer correction improves theory-data comparisons over simple application of a pressure-doubling rule at the fuselage.

Wall Boundary Layer Measurements for the NASA Langley Transonic Dynamics Tunnel

NASA Technical Reports Server (NTRS)

Wieseman, Carol D.; Bennett, Robert M.

2007-01-01

Measurements of the boundary layer parameters in the NASA Langley Transonic Dynamics tunnel were conducted during extensive calibration activities following the facility conversion from a Freon-12 heavy-gas test medium to R-134a. Boundary-layer rakes were mounted on the wind-tunnel walls, ceiling, and floor. Measurements were made over the range of tunnel operation envelope in both heavy gas and air and without a model in the test section at three tunnel stations. Configuration variables included open and closed east sidewall wall slots, for air and R134a test media, reentry flap settings, and stagnation pressures over the full range of tunnel operation. The boundary layer thickness varied considerably for the six rakes. The thickness for the east wall was considerably larger that the other rakes and was also larger than previously reported. There generally was some reduction in thickness at supersonic Mach numbers, but the effect of stagnation pressure, and test medium were not extensive.

Flight-measured pressure characteristics of aft-facing steps in high Reynolds number flow at Mach numbers of 2.20, 2.50, and 2.80 and comparison with other data

NASA Technical Reports Server (NTRS)

Powers, S. G.

1978-01-01

The YF-12 airplane was studied to determine the pressure characteristics associated with an aft-facing step in high Reynolds number flow for nominal Mach numbers of 2.20, 2.50, and 2.80. Base pressure coefficients were obtained for three step heights. The surface static pressures ahead of and behind the step were measured for the no-step condition and for each of the step heights. A boundary layer rake was used to determine the local boundary layer conditions. The Reynolds number based on the length of flow ahead of the step was approximately 10 to the 8th power and the ratios of momentum thickness to step height ranged from 0.2 to 1.0. Base pressure coefficients were compared with other available data at similar Mach numbers and at ratios of momentum thickness to step height near 1.0. In addition, the data were compared with base pressure coefficients calculated by a semiempirical prediction method. The base pressure ratios are shown to be a function of Reynolds number based on momentum thickness. Profiles of the surface pressures ahead of and behind the step and the local boundary layer conditions are also presented.

Experimental measurements of unsteady turbulent boundary layers near separation

NASA Technical Reports Server (NTRS)

Simpson, R. L.

1982-01-01

Investigations conducted to document the behavior of turbulent boundary layers on flat surfaces that separate due to adverse pressure gradients are reported. Laser and hot wire anemometers measured turbulence and flow structure of a steady free stream separating turbulent boundary layer produced on the flow of a wind tunnel section. The effects of sinusoidal and unsteadiness of the free stream velocity on this separating turbulent boundary layer at a reduced frequency were determined. A friction gage and a thermal tuft were developed and used to measure the surface skin friction and the near wall fraction of time the flow moves downstream for several cases. Abstracts are provided of several articles which discuss the effects of the periodic free stream unsteadiness on the structure or separating turbulent boundary layers.

Effect of Sub-Boundary Layer Vortex Generations on Incident Turbulence

NASA Technical Reports Server (NTRS)

Casper, J.; Lin, J. C.; Yao, C. S.

2003-01-01

Sub-boundary layer vortex generators were tested in a wind tunnel to assess their effect on the velocity field within the wake region of a turbulent boundary layer. Both mean flow quantities and turbulence statistics were measured. Although very small relative to the boundary layer thickness, these so-called micro vortex generators were found to have a measurable effect on the power spectra and integral length scales of the turbulence at a distance many times the height of the devices themselves. In addition, the potential acoustic impact of these devices is also discussed. Measured turbulence spectra are used as input to an acoustic formulation in a manner that compares predicted sound pressure levels that result from the incident boundary-layer turbulence, with and without the vortex generators in the flow.

Exploratory Investigation of the Effects of Boundary-Layer Control on the Pressure-Recovery Characteristics of a Circular Internal-Contraction Inlet with Translating Centerbody at Mach Numbers of 2.00 and 2.35

NASA Technical Reports Server (NTRS)

Martin, Norman J.

1959-01-01

Exploratory tests of a circular internal-contraction inlet were made at Mach numbers of 2.00 and 2.35 to determine the effect of a cowl-type boundary-layer control located downstream of the inlet throat. The inlet was designed for a Mach number of 2.5. Tests were also made of the inlet modified to correspond to design Mach numbers of 2.35 and 2.25. Surveys near the minimum area section of the inlet without boundary-layer control indicated maximum averaged pressure recoveries between 0.90 and 0.92 at a free-stream Mach number, M(sub infinity), of 2.35 for the inlets. Farther downstream, after partial subsonic diffusion, a maximum pressure recovery of 0.842 was obtained with the inlet at M(sub infinity) = 2.35. The pressure recovery of the inlet was increased by 0.03 at a Mach number of 2.35 and decreased by 0.02 at a Mach number of 2.00 by the application of cowl-type boundary-layer control. Further investigation with the inlet without bleed demonstrated that an increase of angle of attack from 0 deg to 3 deg reduced the pressure recovery 0.04. The effect of Reynolds number was to increase pressure recovery 0.07 (from 0.785 to 0.855) with an increase in Reynolds number (based on inlet diameter) from 0.79 x 10(exp 6) to 3.19 x 10(exp 6).

Near-wall similarity in a pressure-driven three-dimensional turbulent boundary layer

NASA Technical Reports Server (NTRS)

Pierce, F. J.; Mcallister, J. E.

1980-01-01

Mean velocity, measured wall pressure and wall shear stress fields were made in a three dimensional pressure-driven turbulent boundary layer created by a cylinder with trailing edge placed normal to a flat plate floor. The direct force wall shear stress measurements were made with floating element direct force sensing shear meter that responded to both the magnitude and direction of the local wall shear stress. The ability of 10 near wall similarity models to describe the near wall velocity field for the measured flow under a wide range of skewing conditions and a variety of pressure gradient and wall shear vector orientations was used.

Non-Gaussian PDF Modeling of Turbulent Boundary Layer Fluctuating Pressure Excitation

NASA Technical Reports Server (NTRS)

Steinwolf, Alexander; Rizzi, Stephen A.

2003-01-01

The purpose of the study is to investigate properties of the probability density function (PDF) of turbulent boundary layer fluctuating pressures measured on the exterior of a supersonic transport aircraft. It is shown that fluctuating pressure PDFs differ from the Gaussian distribution even for surface conditions having no significant discontinuities. The PDF tails are wider and longer than those of the Gaussian model. For pressure fluctuations upstream of forward-facing step discontinuities and downstream of aft-facing step discontinuities, deviations from the Gaussian model are more significant and the PDFs become asymmetrical. Various analytical PDF distributions are used and further developed to model this behavior.

Influence of collision on the flow through in-vitro rigid models of the vocal folds

NASA Astrophysics Data System (ADS)

Deverge, M.; Pelorson, X.; Vilain, C.; Lagrée, P.-Y.; Chentouf, F.; Willems, J.; Hirschberg, A.

2003-12-01

Measurements of pressure in oscillating rigid replicas of vocal folds are presented. The pressure upstream of the replica is used as input to various theoretical approximations to predict the pressure within the glottis. As the vocal folds collide the classical quasisteady boundary layer theory fails. It appears however that for physiologically reasonable shapes of the replicas, viscous effects are more important than the influence of the flow unsteadiness due to the wall movement. A simple model based on a quasisteady Bernoulli equation corrected for viscous effect, combined with a simple boundary layer separation model does globally predict the observed pressure behavior.

Transient interaction between a reaction control jet and a hypersonic crossflow

NASA Astrophysics Data System (ADS)

Miller, Warrick A.; Medwell, Paul R.; Doolan, Con J.; Kim, Minkwan

2018-04-01

This paper presents a numerical study that focuses on the transient interaction between a reaction control jet and a hypersonic crossflow with a laminar boundary layer. The aim is to better understand the underlying physical mechanisms affecting the resulting surface pressure and control force. Implicit large-eddy simulations were performed with a round, sonic, perfect air jet issuing normal to a Mach 5 crossflow over a flat plate with a laminar boundary layer, at a jet-to-crossflow momentum ratio of 5.3 and a pressure ratio of 251. The pressure distribution induced on the flat plate is unsteady and is influenced by vortex structures that form around the jet. A horseshoe vortex structure forms upstream and consists of six vortices: two quasi-steady vortices and two co-rotating vortex pairs that periodically coalesce. Shear-layer vortices shed periodically and cause localised high pressure regions that convect downstream with constant velocity. A longitudinal counter-rotating vortex pair is present downstream of the jet and is formed from a series of trailing vortices which rotate about a common axis. Shear-layer vortex shedding causes periodic deformation of barrel and bow shocks. This changes the location of boundary layer separation which also affects the normal force on the plate.

Experimental Study of the Effects of Periodic Unsteady Wakes on Flow Separation in Low Pressure Turbines

NASA Technical Reports Server (NTRS)

Ozturk, Burak; Schobeiri, Meinhard T.

2009-01-01

The present study, which is the first of a series of investigations of low pressure turbine (LPT) boundary layer aerodynamics, is aimed at providing detailed unsteady boundary layer flow information to understand the underlying physics of the inception, onset, and extent of the separation zone. A detailed experimental study on the behavior of the separation zone on the suction surface of a highly loaded LPT-blade under periodic unsteady wake flow is presented. Experimental investigations were performed on a large-scale, high-subsonic unsteady turbine cascade research facility with an integrated wake generator and test section unit. Blade Pak B geometry was used in the cascade. The wakes were generated by continuously moving cylindrical bars device. Boundary layer investigations were performed using hot wire anemometry at Reynolds number of 110,000, based on the blade suction surface length and the exit velocity, for one steady and two unsteady inlet flow conditions, with the corresponding passing frequencies, wake velocities, and turbulence intensities. The reduced frequencies cover the entire operation range of LP-turbines. In addition to the unsteady boundary layer measurements, blade surface pressure measurements were performed at Re = 50,000, 75,000, 100,000, 110,000, and 125,000. For each Reynolds number, surface pressure measurements are carried out at one steady and two periodic unsteady inlet flow conditions. Detailed unsteady boundary layer measurement identifies the onset and extension of the separation zone as well as its behavior under unsteady wake flow. The results, presented in ensemble-averaged and contour plot forms, help to understand the physics of the separation phenomenon under periodic unsteady wake flow.

Interaction of two glancing, crossing shock waves with a turbulent boundary-layer at various Mach numbers

NASA Technical Reports Server (NTRS)

Hingst, Warren R.; Williams, Kevin E.

1991-01-01

A preliminary experimental investigation was conducted to study two crossing, glancing shock waves of equal strengths, interacting with the boundary-layer developed on a supersonic wind tunnel wall. This study was performed at several Mach numbers between 2.5 and 4.0. The shock waves were created by fins (shock generators), spanning the tunnel test section, that were set at angles varying from 4 to 12 degrees. The data acquired are wall static pressure measurements, and qualitative information in the form of oil flow and schlieren visualizations. The principle aim is two-fold. First, a fundamental understanding of the physics underlying this flow phenomena is desired. Also, a comprehensive data set is needed for computational fluid dynamic code validation. Results indicate that for small shock generator angles, the boundary-layer remains attached throughout the flow field. However, with increasing shock strengths (increasing generator angles), boundary layer separation does occur and becomes progressively more severe as the generator angles are increased further. The location of the separation, which starts well downstream of the shock crossing point, moves upstream as shock strengths are increased. At the highest generator angles, the separation appears to begin coincident with the generator leading edges and engulfs most of the area between the generators. This phenomena occurs very near the 'unstart' limit for the generators. The wall pressures at the lower generator angles are nominally consistent with the flow geometries (i.e. shock patterns) although significantly affected by the boundary-layer upstream influence. As separation occurs, the wall pressures exhibit a gradient that is mainly axial in direction in the vicinity of the separation. At the limiting conditions the wall pressure gradients are primarily in the axial direction throughout.

Interaction of viscous and inviscid instability modes in separation-bubble transition

NASA Astrophysics Data System (ADS)

Brinkerhoff, Joshua R.; Yaras, Metin I.

2011-12-01

This paper describes numerical simulations that are used to examine the interaction of viscous and inviscid instability modes in laminar-to-turbulent transition in a separation bubble. The results of a direct numerical simulation are presented in which separation of a laminar boundary-layer occurs in the presence of an adverse streamwise pressure gradient. The simulation is performed at low freestream-turbulence levels and at a flow Reynolds number and pressure distribution approximating those typically encountered on the suction side of low-pressure turbine blades in a gas-turbine engine. The simulation results reveal the development of a viscous instability upstream of the point of separation which produces streamwise-oriented vortices in the attached laminar boundary layer. These vortices remain embedded in the flow downstream of separation and are carried into the separated shear layer, where they are amplified by the local adverse pressure-gradient and contribute to the formation of coherent hairpin-like vortices. A strong interaction is observed between these vortices and the inviscid instability that typically dominates the shear layer in the separated zone. The interaction is noted to determine the spanwise extent of the vortical flow structures that periodically shed from the downstream end of the separated shear layer. The structure of the shed vortical flow structures is examined and compared with the coherent structures typically observed within turbulent boundary layers.

Theoretical face pressure and drag characteristics of forward-facing steps in supersonic turbulent boundary layers

NASA Technical Reports Server (NTRS)

Patel, D. K.; Czarnecki, K. R.

1975-01-01

A theoretical investigation of the pressure distributions and drag characteristics was made for forward facing steps in turbulent flow at supersonic speeds. An approximate solution technique proposed by Uebelhack has been modified and extended to obtain a more consistent numerical procedure. A comparison of theoretical calculations with experimental data generally indicated good agreement over the experimentally available range of ratios of step height to boundary layer thickness from 7 to 0.05.

Transitional Boundary Layers Under the Influence of High Free Stream Turbulence, Intensive Wall Cooling and High Pressure Gradients in Hot Gas Circulation. Ph.D. Thesis - Technische Hochschule, Karlsruhe, 1985

NASA Technical Reports Server (NTRS)

Rued, Klaus

1987-01-01

The requirements for fundamental experimental studies of the influence of free stream turbulence, pressure gradients and wall cooling are discussed. Under turbine-like free stream conditions, comprehensive tests of transitional boundary layers with laminar, reversing and turbulent flow increments were performed to decouple the effects of the parameters and to determine the effects during mutual interaction.

Shock Tunnel Operation and Correlation of Boundary Layer Transition on a Cone in Hypervelocity Flow

DTIC Science & Technology

2013-07-01

conditions from the ideal reflected-shock pressure to measured reservoir pressure using an isentropic expan- sion. Furthermore, the 1-D nozzle computation...does not account for boundary layer growth within the nozzle , off-design operation conditions that lead to flow nonuni- formity, or vibration...translation nonequilibrium and freezing within the nozzle , which is significant for the N2 cases. For the uncertainties that can be quantified, we have combined

Performance of the active sidewall boundary-layer removal system for the Langley 0.3-meter Transonic Cryogenic Tunnel

NASA Technical Reports Server (NTRS)

Balakrishna, S.; Kilgore, W. Allen; Murthy, A. V.

1989-01-01

A performance evaluation of an active sidewall boundary-layer removal system for the Langley 0.3-m Transonic Cryogenic Tunnel (TCT) was evaluated in 1988. This system uses a compressor and two throttling digital valves to control the boundary-layer mass flow removal from the tunnel. The compressor operates near the maximum pressure ratio for all conditions. The system uses a surge prevention and flow recirculation scheme. A microprocessor based controller is used to provide the necessary mass flow and compressor pressure ratio control. Initial tests on the system indicated problems in realizing smooth mass flow control while running the compressor at high speed and high pressure ratios. An alternate method has been conceived to realize boundary-layer mass flow control which avoids the recirculation of the compressor mass flow and operation near the compressor surge point. This scheme is based on varying the speed of the compressor for a sufficient pressure ratio to provide needed mass flow removal. The system has a mass flow removal capability of about 10 percent of test section flow at M = 0.3 and 4 percent at M = 0.8. The system performance has been evaluated in the form of the compressor map, and compressor tunnel interface characteristics covering most of the 0.3-m TCT operational envelope.

Generalized wall function and its application to compressible turbulent boundary layer over a flat plate

NASA Astrophysics Data System (ADS)

Liu, J.; Wu, S. P.

2017-04-01

Wall function boundary conditions including the effects of compressibility and heat transfer are improved for compressible turbulent boundary flows. Generalized wall function formulation at zero-pressure gradient is proposed based on coupled velocity and temperature profiles in the entire near-wall region. The parameters in the generalized wall function are well revised. The proposed boundary conditions are integrated into Navier-Stokes computational fluid dynamics code that includes the shear stress transport turbulence model. Numerical results are presented for a compressible boundary layer over a flat plate at zero-pressure gradient. Compared with experimental data, the computational results show that the generalized wall function reduces the first grid spacing in the directed normal to the wall and proves the feasibility and effectivity of the generalized wall function method.

Results of tests on a Rockwell International space shuttle orbiter (-139 configuration) 0.0175-scale model (no. 29-0) in AEDC tunnel B to determine boundary layer characteristics

NASA Technical Reports Server (NTRS)

Quan, M.

1975-01-01

Results of wind tunnel tests were conducted to determine boundary layer characteristics on the lower surface of a space shuttle orbiter. Total pressure and temperature profile data at various model stations were obtained using a movable, four-degree-of-freedom probe mechanism and static pressure taps on the model surface. During a typical run, the probe was located over a preselected model location, then driven down through the bondary layer until contact was made with the model surface.

An investigation of turbulence structure in a low-Reynolds-number incompressible turbulent boundary

NASA Technical Reports Server (NTRS)

White, B. R.; Strataridakis, C. J.

1987-01-01

An existing high turbulence intensity level (5%) atmospheric boundary-layer wind tunnel has been successfully converted to a relatively low level turbulence (0.3%) wind tunnel through extensive modification, testing, and calibration. A splitter plate was designed, built, and installed into the wind-tunnel facility to create thick, mature, two-dimensional turbulent boundary layer flow at zero pressure gradient. Single and cross hot-wire measurements show turbulent boundary layer characteristics of good quality with unusually large physical size, i.e., viscous sublayer of the order of 1 mm high. These confirm the potential ability of the tunnel to be utilized for future high-quality near-wall turbulent boundary layer measurements. It compares very favorably with many low turbulence research tunnels.

Laser interferometer/Preston tube skin-friction comparison in shock/boundary-layer interaction

NASA Technical Reports Server (NTRS)

Kim, K.-S.; Lee, Y.; Settles, G. S.

1991-01-01

An evaluation is conducted of the accuracy of the 'Preston tube' surface pitot-pressure skin friction measurement method relative to the already proven laser interferometer skin-friction meter in a swept shock wave/turbulent boundary-layer interaction. The Preston tube was used to estimate the total shear-stress distribution in a fin-generated swept shock-wave/turbulent boundary-layer interaction. The Keener-Hopkins calibration method using the isentropic relation to calculate the Preston-tube Mach number produces the best results.

Hydrodynamic structure of the boundary layers in a rotating cylindrical cavity with radial inflow

DOE Office of Scientific and Technical Information (OSTI.GOV)

Herrmann-Priesnitz, Benjamín, E-mail: bherrman@ing.uchile.cl; Torres, Diego A.; Advanced Mining Technology Center, Universidad de Chile, Av. Tupper 2007, Santiago

A flow model is formulated to investigate the hydrodynamic structure of the boundary layers of incompressible fluid in a rotating cylindrical cavity with steady radial inflow. The model considers mass and momentum transfer coupled between boundary layers and an inviscid core region. Dimensionless equations of motion are solved using integral methods and a space-marching technique. As the fluid moves radially inward, entraining boundary layers develop which can either meet or become non-entraining. Pressure and wall shear stress distributions, as well as velocity profiles predicted by the model, are compared to numerical simulations using the software OpenFOAM. Hydrodynamic structure of themore » boundary layers is governed by a Reynolds number, Re, a Rossby number, Ro, and the dimensionless radial velocity component at the periphery of the cavity, U{sub o}. Results show that boundary layers merge for Re < < 10 and Ro > > 0.1, and boundary layers become predominantly non-entraining for low Ro, low Re, and high U{sub o}. Results may contribute to improve the design of technology, such as heat exchange devices, and turbomachinery.« less

F-16XL ship #1 CAWAP flight - alpha 5 degrees, altitude 10,000 feet

NASA Technical Reports Server (NTRS)

1996-01-01

The single-seat F-16XL (ship #1) makes another run during the Cranked-Arrow Wing Aerodynamic Project (CAWAP) at Dryden Flight Research Center, Edwards, California. The modified airplane features a delta 'cranked-arrow' wing with strips of tubing along the leading edge to the trailing edge to sense static on the wing and obtain pressure distribution data. The right wing receives data on pressure distribution and the left wing has three types of instrumentation - preston tubes to measure local skin friction, boundary layer rakes to measure boundary layer profiles (the layer where the air interacts with the surfaces of a moving aircraft), and hot films to determine boundary layer transition locations. This photo shows the aircraft gathering data at an altitude of 10,000 feet, with an angle of attack of 5 degrees. The first flight of CAWAP occurred on November 21, 1995, and the test program ended in April 1996.

F-16XL ship #1 - CAWAP outboard rakes #7 and inboard rack #3

NASA Technical Reports Server (NTRS)

1996-01-01

This photo shows the #7 outboard rake and the #3 inboard rake on the left wing of NASA's single-seat F-16XL (ship #1) used for the Cranked-Arrow Wing Aerodynamic Project (CAWAP) at Dryden Flight Research Center, Edwards, California. The modified airplane features a delta 'cranked-arrow' wing with strips of tubing along the leading edge to the trailing edge to sense static on the wing and obtain pressure distribution data. The right wing receives data on pressure distribution and the left wing has three types of instrumentation - preston tubes to measure local skin friction, boundary layer rakes to measure boundary layer profiles (the layer where the air interacts with the surfaces of a moving aircraft), and hot films to determine boundary layer transition locations. The first flight of CAWAP occurred on November 21, 1995, and the test program ended in April 1996.

F-16XL ship #1 CAWAP flight

NASA Technical Reports Server (NTRS)

1996-01-01

The single-seat F-16XL (ship #1) makes another run during the Cranked-Arrow Wing Aerodynamic Project (CAWAP) at Dryden Flight Research Center, Edwards, California. The modified airplane features a delta 'cranked-arrow' wing with strips of tubing along the leading edge to the trailing edge to sense static on the wing and obtain pressure distribution data. The right wing receives data on pressure distribution and the left wing has three types of instrumentation - preston tubes to measure local skin friction, boundary layer rakes to measure boundary layer profiles (the layer where the air interacts with the surfaces of a moving aircraft), and hot films to determine boundary layer transition locations. The program also gathered aero data on two wing planforms for NASA's High Speed Research Program. The first flight of CAWAP occurred on November 21, 1995, and the test program ended in April 1996.

F-16XL ship #1 - CAWAP outboard rake #7

NASA Technical Reports Server (NTRS)

1996-01-01

This photo shows the #7 outboard rake on the left wing of NASA's single-seat F-16XL (ship #1) used for the Cranked-Arrow Wing Aerodynamic Project (CAWAP) at Dryden Flight Research Center, Edwards, California. The modified airplane features a delta 'cranked-arrow' wing with strips of tubing along the leading edge to the trailing edge to sense static on the wing and obtain pressure distribution data. The right wing receives data on pressure distribution and the left wing has three types of instrumentation - preston tubes to measure local skin friction, boundary layer rakes to measure boundary layer profiles (the layer where the air interacts with the surfaces of a moving aircraft), and hot films to determine boundary layer transition locations. The program also gathered aero data on two wing planforms for NASA's High Speed Research Program. The first flight of CAWAP occurred on November 21, 1995, and the test program ended in April 1996.

Control of Boundary Layers for Aero-optical Applications

DTIC Science & Technology

2015-06-23

range of subsonic and supersonic Mach numbers was developed and shown to correctly predict experimentally-observed reductions. Heating the wall allows...40 3.3 Extension to supersonic speeds...boundary layers at supersonic speeds. Comparing the model prediction to the experimental results, it was speculated that while the pressure effects can

Numerical investigation of an internal layer in turbulent flow over a curved hill

NASA Technical Reports Server (NTRS)

Kim, S-W.

1989-01-01

The development of an internal layer in a turbulent boundary layer flow over a curved hill is investigated numerically. The turbulence field of the boundary layer flow over the curved hill is compared with that of a turbulent flow over a symmetric airfoil (which has the same geometry as the curved hill except that the leading and trailing edge plates were removed) to study the influence of the strongly curved surface on the turbulence field. The turbulent flow equations are solved by a control-volume based finite difference method. The turbulence is described by a multiple-time-scale turbulence model supplemented with a near-wall turbulence model. Computational results for the mean flow field (pressure distributions on the walls, wall shearing stresses and mean velocity profiles), the turbulence structure (Reynolds stress and turbulent kinetic energy profiles), and the integral parameters (displacement and momentum thicknesses) compared favorably with the measured data. Computational results show that the internal layer is a strong turbulence field which is developed beneath the external boundary layer and is located very close to the wall. Development of the internal layer was more obviously observed in the Reynolds stress profiles and in the turbulent kinetic energy profiles than in the mean velocity profiles. In this regard, the internal layers is significantly different from wall-bounded simple shear layers in which the mean velocity profile characterizes the boundary layer most distinguishably. Development of such an internal layer, characterized by an intense turbulence field, is attributed to the enormous mean flow strain rate caused by the streamline curvature and the strong pressure gradient. In the turbulent flow over the curved hill, the internal layer begin to form near the forward corner of the hill, merges with the external boundary layer, and develops into a new fully turbulent boundary layer as the fluid flows in the downstream direction. For the flow over the symmetric airfoil, the boundary layer began to form from almost the same location as that of the curved hill, grew in its strength, and formed a fully turbulent boundary layer from mid-part of the airfoil and in the downstream region. Computational results also show that the detailed turbulence structure in the region very close to the wall of the curved hill is almost the same as that of the airfoil in most of the curved regions except near the leading edge. Thus the internal layer of the curved hill and the boundary layer of the airfoil were also almost the same. Development of the wall shearing stress and separation of the boundary layer at the rear end of the curved hill mostly depends on the internal layer and is only slightly influenced by the external boundary layer flow.

Analysis of in-flight boundary-layer state measurements on a subsonic transport wing in high-lift configuration

NASA Technical Reports Server (NTRS)

vanDam, C. P.; Los, S. M.; Miley, S. J.; Yip, L. P.; Banks, D. W.; Roback, V. E.; Bertelrud, A.

1995-01-01

Flight experiments on NASA Langley's B737-100 (TSRV) airplane have been conducted to document flow characteristics in order to further the understanding of high-lift flow physics, and to correlate and validate computational predictions and wind-tunnel measurements. The project is a cooperative effort involving NASA, industry, and universities. In addition to focusing on in-flight measurements, the project includes extensive application of various computational techniques, and correlation of flight data with computational results and wind-tunnel measurements. Results obtained in the most recent phase of flight experiments are analyzed and presented in this paper. In-flight measurements include surface pressure distributions, measured using flush pressure taps and pressure belts on the slats, main element, and flap elements; surface shear stresses, measured using Preston tubes; off-surface velocity distributions, measured using shear-layer rakes; aeroelastic deformations of the flap elements, measured using an optical positioning system; and boundary-layer transition phenomena, measured using hot-film anemometers and an infrared imaging system. The analysis in this paper primarily focuses on changes in the boundary-layer state that occurred on the slats, main element, and fore flap as a result of changes in flap setting and/or flight condition. Following a detailed description of the experiment, the boundary-layer state phenomenon will be discussed based on data measured during these recent flight experiments.

Non-local sub-characteristic zones of influence in unsteady interactive boundary-layers

NASA Technical Reports Server (NTRS)

Rothmayer, A. P.

1992-01-01

The properties of incompressible, unsteady, interactive, boundary layers are examined for a model hypersonic boundary layer and internal flow past humps or, equivalently, external flow past short-scaled humps. Using a linear high frequency analysis, it is shown that the domains of dependence within the viscous sublayer may be a strong function of position within the sublayer and may be strongly influenced by the pressure displacement interaction, or the prescribed displacement condition. Detailed calculations are presented for the hypersonic boundary layer. This effect is found to carry over directly to the fully viscous problem as well as the nonlinear problem. In the fully viscous problem, the non-local character of the domains of dependence manifests itself in the sub-characteristics. Potential implications of the domain of dependence structure on finite difference computations of unsteady boundary layers are briefly discussed.

Effects of Riblets on Skin Friction in High-Speed Turbulent Boundary Layers

NASA Technical Reports Server (NTRS)

Duan, Lian; Choudhari, Meelan M.

2012-01-01

Direct numerical simulations of spatially developing turbulent boundary layers over riblets are conducted to examine the effects of riblets on skin friction at supersonic speeds. Zero-pressure gradient boundary layers with an adiabatic wall, a Mach number of M1 = 2.5, and a Reynolds number based on momentum thickness of Re = 1720 are considered. Simulations are conducted for boundary-layer flows over a clean surface and symmetric V- groove riblets with nominal spacings of 20 and 40 wall units. The DNS results confirm the few existing experimental observations and show that a drag reduction of approximately 7% is achieved for riblets with proper spacing. The influence of riblets on turbulence statistics is analyzed in detail with an emphasis on identifying the differences, if any, between the drag reduction mechanisms for incompressible and high-speed boundary layers.

The behaviour of turbulence anisotropy through shock waves and expansions

NASA Technical Reports Server (NTRS)

Minh, H. H.; Kollmann, W.; Vandromme, D.

1985-01-01

A second order closure has been implemented in an implicit Navier-Stokes solver to study the behavior of the Reynolds stresses under the influence of severe pressure gradients. In the boundary layer zone, the strongly sheared character of the mean flow dominates the turbulence generation mechanisms. However, the pressure gradients play also a very important role for these processes, but at different locations within the boundary layer. This aspect may be emphasized by the analysis of turbulence anisotropy through shock waves and expansions.

Swept shock/boundary layer interaction experiments in support of CFD code validation

NASA Technical Reports Server (NTRS)

Settles, G. S.; Lee, Y.

1990-01-01

Research on the topic of shock wave/turbulent boundary layer interaction was carried out. Skin friction and surface pressure measurements in fin-induced, swept interactions were conducted, and heat transfer measurements in the same flows are planned. The skin friction data for a strong interaction case (Mach 4, fin-angles equal 16 and 20 degrees) were obtained, and their comparison with computational results was published. Surface pressure data for weak-to-strong fin interactions were also obtained.

Exact solutions of laminar-boundary-layer equations with constant property values for porous wall with variable temperature

NASA Technical Reports Server (NTRS)

Donoughe, Patrick L; Livingood, John N B

1955-01-01

Exact solution of the laminar-boundary-layer equations for wedge-type flow with constant property values are presented for transpiration-cooled surfaces with variable wall temperatures. The difference between wall and stream temperature is assumed proportional to a power of the distance from the leading edge. Solutions are given for a Prandtl number of 0.7 and ranges of pressure-gradient, cooling-air-flow, and wall-temperature-gradient parameters. Boundary-layer profiles, dimensionless boundary-layer thicknesses, and convective heat-transfer coefficients are given in both tabular and graphical form. Corresponding results for constant wall temperature and for impermeable surfaces are included for comparison purposes.

Inverse boundary-layer theory and comparison with experiment

NASA Technical Reports Server (NTRS)

Carter, J. E.

1978-01-01

Inverse boundary layer computational procedures, which permit nonsingular solutions at separation and reattachment, are presented. In the first technique, which is for incompressible flow, the displacement thickness is prescribed; in the second technique, for compressible flow, a perturbation mass flow is the prescribed condition. The pressure is deduced implicitly along with the solution in each of these techniques. Laminar and turbulent computations, which are typical of separated flow, are presented and comparisons are made with experimental data. In both inverse procedures, finite difference techniques are used along with Newton iteration. The resulting procedure is no more complicated than conventional boundary layer computations. These separated boundary layer techniques appear to be well suited for complete viscous-inviscid interaction computations.

Estimated Performance of Radial-Flow Exit Nozzles for Air in Chemical Equilibrium

NASA Technical Reports Server (NTRS)

Englert, Gerald W.; Kochendorfer, Fred D.

1959-01-01

The thrust, boundary-layer, and heat-transfer characteristics were computed for nozzles having radial flow in the divergent part. The working medium was air in chemical equilibrium, and the boundary layer was assumed to be all turbulent. Stagnation pressure was varied from 1 to 32 atmospheres, stagnation temperature from 1000 to 6000 R, and wall temperature from 1000 to 3000 R. Design pressure ratio was varied from 5 to 320, and operating pressure ratio was varied from 0.25 to 8 times the design pressure ratio. Results were generalized independent of divergence angle and were also generalized independent of stagnation pressure in the temperature range of 1000 to 3000 R. A means of determining the aerodynamically optimum wall angle is provided.

Flight Demonstration of a Shock Location Sensor Using Constant Voltage Hot-Film Anemometry

NASA Technical Reports Server (NTRS)

Moes, Timothy R.; Sarma, Garimella R.; Mangalam, Siva M.

1997-01-01

Flight tests have demonstrated the effectiveness of an array of hot-film sensors using constant voltage anemometry to determine shock position on a wing or aircraft surface at transonic speeds. Flights were conducted at the NASA Dryden Flight Research Center using the F-15B aircraft and Flight Test Fixture (FTF). A modified NACA 0021 airfoil was attached to the side of the FTF, and its upper surface was instrumented to correlate shock position with pressure and hot-film sensors. In the vicinity of the shock-induced pressure rise, test results consistently showed the presence of a minimum voltage in the hot-film anemometer outputs. Comparing these results with previous investigations indicate that hot-film anemometry can identify the location of the shock-induced boundary layer separation. The flow separation occurred slightly forward of the shock- induced pressure rise for a laminar boundary layer and slightly aft of the start of the pressure rise when the boundary layer was tripped near the airfoil leading edge. Both minimum mean output and phase reversal analyses were used to identify the shock location.

Hydraulic transport across hydrophilic and hydrophobic nanopores: Flow experiments with water and n-hexane.

PubMed

Gruener, Simon; Wallacher, Dirk; Greulich, Stefanie; Busch, Mark; Huber, Patrick

2016-01-01

We experimentally explore pressure-driven flow of water and n-hexane across nanoporous silica (Vycor glass monoliths with 7- or 10-nm pore diameters, respectively) as a function of temperature and surface functionalization (native and silanized glass surfaces). Hydraulic flow rates are measured by applying hydrostatic pressures via inert gases (argon and helium, pressurized up to 70 bar) on the upstream side in a capacitor-based membrane permeability setup. For the native, hydrophilic silica walls, the measured hydraulic permeabilities can be quantitatively accounted for by bulk fluidity provided we assume a sticking boundary layer, i.e., a negative velocity slip length of molecular dimensions. The thickness of this boundary layer is discussed with regard to previous capillarity-driven flow experiments (spontaneous imbibition) and with regard to velocity slippage at the pore walls resulting from dissolved gas. Water flow across the silanized, hydrophobic nanopores is blocked up to a hydrostatic pressure of at least 70 bar. The absence of a sticking boundary layer quantitatively accounts for an enhanced n-hexane permeability in the hydrophobic compared to the hydrophilic nanopores.

Sound Transmission through Cylindrical Shell Structures Excited by Boundary Layer Pressure Fluctuations

NASA Technical Reports Server (NTRS)

Tang, Yvette Y.; Silcox, Richard J.; Robinson, Jay H.

1996-01-01

This paper examines sound transmission into two concentric cylindrical sandwich shells subject to turbulent flow on the exterior surface of the outer shell. The interior of the shells is filled with fluid medium and there is an airgap between the shells in the annular space. The description of the pressure field is based on the cross-spectral density formulation of Corcos, Maestrello, and Efimtsov models of the turbulent boundary layer. The classical thin shell theory and the first-order shear deformation theory are applied for the inner and outer shells, respectively. Modal expansion and the Galerkin approach are used to obtain closed-form solutions for the shell displacements and the radiation and transmission pressures in the cavities including both the annular space and the interior. The average spectral density of the structural responses and the transmitted interior pressures are expressed explicitly in terms of the summation of the cross-spectral density of generalized force induced by the boundary layer turbulence. The effects of acoustic and hydrodynamic coincidences on the spectral density are observed. Numerical examples are presented to illustrate the method for both subsonic and supersonic flows.

Transitional and turbulent boundary layer with heat transfer

NASA Astrophysics Data System (ADS)

Wu, Xiaohua; Moin, Parviz

2010-08-01

We report on our direct numerical simulation of an incompressible, nominally zero-pressure-gradient flat-plate boundary layer from momentum thickness Reynolds number 80-1950. Heat transfer between the constant-temperature solid surface and the free-stream is also simulated with molecular Prandtl number Pr=1. Skin-friction coefficient and other boundary layer parameters follow the Blasius solutions prior to the onset of turbulent spots. Throughout the entire flat-plate, the ratio of Stanton number and skin-friction St/Cf deviates from the exact Reynolds analogy value of 0.5 by less than 1.5%. Mean velocity and Reynolds stresses agree with experimental data over an extended turbulent region downstream of transition. Normalized rms wall-pressure fluctuation increases gradually with the streamwise growth of the turbulent boundary layer. Wall shear stress fluctuation, τw,rms'+, on the other hand, remains constant at approximately 0.44 over the range, 800

Unsteady separation in sharp fin-induced shock wave/turbulent boundary layer interaction at Mach 5

NASA Technical Reports Server (NTRS)

Schmisseur, J. D.; Dolling, D. S.

1992-01-01

Fluctuating wall-pressure measurements are made in shock-wave/turbulent-boundary-layer interactions generated by sharp/unswept fins at angles of attack of 16, 18, 20, 22, 24, 26, and 28 degrees at Mach 5. The experiment was conducted under approximately adiabatic wall temperature conditions. The mean and rms pressure distributions can be collapsed in conical coordinates. The wall-pressure signal near separation is intermittent for all angles of attack (16-28 deg) and is qualitatively similar to that measured in unswept flows. However, the shock frequencies are higher - about 5 kHz compared to 0.5-1 kHz. Over the range of sweepbacks examined, from 25-55 deg, the spectral content of the fluctuating pressures does not change. Thus, the increase in separation-shock frequency from 1 to 5 kHz occurs at lower interaction sweepback and is not a continuous process with increasing sweepback. Power spectra at the position of maximum rms in the intermittent region for interactions in different incoming boundary layers have the same center frequency. The maximum rms in the intermittent region correlates with interaction sweepback, not with overall inviscid pressure rise.

Boundary layer, skin friction, and boattail pressure measurements from the YF-12 airplane at Mach numbers up to 3

NASA Technical Reports Server (NTRS)

Fisher, D. F.

1978-01-01

In-flight measurements of boundary layer and skin friction data were made on YF-12 airplanes for Mach numbers between 2.0 and 3.0. Boattail pressures were also obtained for Mach numbers between 0.7 and 3.0 with Reynolds numbers up to four hundred million. Boundary layer data measured along the lower fuselage centerline indicate local displacement and momentum thicknesses can be much larger than predicted. Skin friction coefficients measured at two of five lower fuselage stations were significantly less than predicted by flat plate theory. The presence of large differences between measured boattail pressure drag and values calculated by a potential flow solution indicates the presence of vortex effects on the upper boattail surface. At both subsonic and supersonic speeds, pressure drag on the longer of two boattail configurations was equal to or less than the pressure drag on the shorter configuration. At subsonic and transonic speeds, the difference in the drag coefficient was on the order of 0.0008 to 0.0010. In the supersonic cruise range, the difference in the drag coefficient was on the order of 0.002. Boattail drag coefficients are based on wing reference area.

Modeling of wave-coherent pressures in the turbulent boundary layer above water waves

NASA Technical Reports Server (NTRS)

Papadimitrakis, Yiannis ALEX.

1988-01-01

The behavior of air pressure fluctuations induced by progressive water waves generated mechanically in a laboratory tank was simulated by solving a modified Orr-Sommerfeld equation in a transformed Eulerian wave-following frame of reference. Solution is obtained by modeling the mean and wave-coherent turbulent Reynolds stresses, the behavior of which in the turbulent boundary layer above the waves was simulated using a turbulent kinetic energy-dissipation model, properly modified to account for free-surface proximity and favorable pressure gradient effects. The distribution of both the wave-coherent turbulent Reynolds stress and pressure amplitudes and their corresponding phase lags was found to agree reasonably well with available laboratory data.

In-Flight Boundary-Layer Transition of a Large Flat Plate at Supersonic Speeds

NASA Technical Reports Server (NTRS)

Banks, D. W.; Frederick, M. A.; Tracy, R. R.; Matisheck, J. R.; Vanecek, N. D.

2012-01-01

A flight experiment was conducted to investigate the pressure distribution, local-flow conditions, and boundary-layer transition characteristics on a large flat plate in flight at supersonic speeds up to Mach 2.00. The tests used a NASA testbed aircraft with a bottom centerline mounted test fixture. The primary objective of the test was to characterize the local flow field in preparation for future tests of a high Reynolds number natural laminar flow test article. A second objective was to determine the boundary-layer transition characteristics on the flat plate and the effectiveness of using a simplified surface coating. Boundary-layer transition was captured in both analog and digital formats using an onboard infrared imaging system. Surface pressures were measured on the surface of the flat plate. Flow field measurements near the leading edge of the test fixture revealed the local flow characteristics including downwash, sidewash, and local Mach number. Results also indicated that the simplified surface coating did not provide sufficient insulation from the metallic structure, which likely had a substantial effect on boundary-layer transition compared with that of an adiabatic surface. Cold wall conditions were predominant during the acceleration to maximum Mach number, and warm wall conditions were evident during the subsequent deceleration.

Direct Numerical Simulation of Hypersonic Turbulent Boundary Layer inside an Axisymmetric Nozzle

NASA Technical Reports Server (NTRS)

Huang, Junji; Zhang, Chao; Duan, Lian; Choudhari, Meelan M.

2017-01-01

As a first step toward a study of acoustic disturbance field within a conventional, hypersonic wind tunnel, direct numerical simulations (DNS) of a Mach 6 turbulent boundary layer on the inner wall of a straight axisymmetric nozzle are conducted and the results are compared with those for a flat plate. The DNS results for a nozzle radius to boundary-layer thickness ratio of 5:5 show that the turbulence statistics of the nozzle-wall boundary layer are nearly unaffected by the transverse curvature of the nozzle wall. Before the acoustic waves emanating from different parts of the nozzle surface can interfere with each other and undergo reflections from adjacent portions of the nozzle surface, the rms pressure fluctuation beyond the boundary layer edge increases toward the nozzle axis, apparently due to a focusing effect inside the axisymmetric configuration. Spectral analysis of pressure fluctuations at both the wall and the freestream indicates a similar distribution of energy content for both the nozzle and the flat plate, with the peak of the premultiplied frequency spectrum at a frequency of [(omega)(delta)]/U(sub infinity) approximately 6.0 inside the free stream and at [(omega)(delta)]/U(sub infinity) approximately 2.0 along the wall. The present results provide the basis for follow-on simulations involving reverberation effects inside the nozzle.

The dynamics and control of fluctuating pressure loads in the reattachment region of a supersonic free shear layer

NASA Technical Reports Server (NTRS)

Smits, A. J.

1990-01-01

The primary aim is to investigate the mechanisms which cause the unsteady wall-pressure fluctuations in shock wave turbulent shear layer interactions. The secondary aim is to find means to reduce the magnitude of the fluctuating pressure loads by controlling the unsteady shock motion. The particular flow proposed for study is the unsteady shock wave interaction formed in the reattachment zone of a separated supersonic flow. Similar flows are encountered in many practical situations, and they are associated with high levels of fluctuating wall pressure. Wall pressure fluctuations were measured in the reattachment region of the supersonic free shear layer. The free shear layer was formed by the separation of a Mach 2.9 turbulent boundary layer from a backward facing step. Reattachment occurred on a 20 deg ramp. By adjusting the position of the ramp, the base pressure was set equal to the freestream pressure, and the free shear layer formed in the absence of a separation shock. An array of flush-mounted, miniature, high-frequency pressure transducers was used to make multichannel measurements of the fluctuating wall pressure in the vicinity of the reattachment region. Contrary to previous observations of this flow, the reattachment region was found to be highly unsteady, and the pressure fluctuations were found to be significant. The overall behavior of the wall pressure loading is similar in scale and magnitude to the unsteadiness of the wall pressure field in compression ramp flows at the same Mach number. Rayleigh scattering was used to visualize the instantaneous shock structure in the streamwise and spanwise direction. Spanwise wrinkles on the order of half the boundary layer thickness were observed.

Optimal Control of Shock Wave Turbulent Boundary Layer Interactions Using Micro-Array Actuation

NASA Technical Reports Server (NTRS)

Anderson, Bernhard H.; Tinapple, Jon; Surber, Lewis

2006-01-01

The intent of this study on micro-array flow control is to demonstrate the viability and economy of Response Surface Methodology (RSM) to determine optimal designs of micro-array actuation for controlling the shock wave turbulent boundary layer interactions within supersonic inlets and compare these concepts to conventional bleed performance. The term micro-array refers to micro-actuator arrays which have heights of 25 to 40 percent of the undisturbed supersonic boundary layer thickness. This study covers optimal control of shock wave turbulent boundary layer interactions using standard micro-vane, tapered micro-vane, and standard micro-ramp arrays at a free stream Mach number of 2.0. The effectiveness of the three micro-array devices was tested using a shock pressure rise induced by the 10 shock generator, which was sufficiently strong as to separate the turbulent supersonic boundary layer. The overall design purpose of the micro-arrays was to alter the properties of the supersonic boundary layer by introducing a cascade of counter-rotating micro-vortices in the near wall region. In this manner, the impact of the shock wave boundary layer (SWBL) interaction on the main flow field was minimized without boundary bleed.

Factors Influencing Pitot Probe Centerline Displacement in a Turbulent Supersonic Boundary Layer

NASA Technical Reports Server (NTRS)

Grosser, Wendy I.

1997-01-01

When a total pressure probe is used for measuring flows with transverse total pressure gradients, a displacement of the effective center of the probe is observed (designated Delta). While this phenomenon is well documented in incompressible flow and supersonic laminar flow, there is insufficient information concerning supersonic turbulent flow. In this study, three NASA Lewis Research Center Supersonic Wind Tunnels (SWT's) were used to investigate pitot probe centerline displacement in supersonic turbulent boundary layers. The relationship between test conditions and pitot probe centerline displacement error was to be determined. For this investigation, ten circular probes with diameter-to-boundary layer ratios (D/delta) ranging from 0.015 to 0.256 were tested in the 10 ft x 10 ft SWT, the 15 cm x 15 cm SWT, and the 1 ft x 1 ft SWT. Reynolds numbers of 4.27 x 10(exp 6)/m, 6.00 x 10(exp 6)/in, 10.33 x 10(exp 6)/in, and 16.9 x 10(exp 6)/m were tested at nominal Mach numbers of 2.0 and 2.5. Boundary layer thicknesses for the three tunnels were approximately 200 mm, 13 mm, and 30 mm, respectively. Initial results indicate that boundary layer thickness, delta, and probe diameter, D/delta play a minimal role in pitot probe centerline offset error, Delta/D. It appears that the Mach gradient, dM/dy, is an important factor, though the exact relationship has not yet been determined. More data is needed to fill the map before a conclusion can be drawn with any certainty. This research provides valuable supersonic, turbulent boundary layer data from three supersonic wind tunnels with three very different boundary layers. It will prove a valuable stepping stone for future research into the factors influencing pitot probe centerline offset error.

Turbulent boundary layer separation over a rearward facing ramp and its control through mechanical excitation

NASA Technical Reports Server (NTRS)

Mckinzie, Daniel J., Jr.

1991-01-01

A vane oscillating about a fixed point at the inlet to a two-dimensional 20 degree rearward facing ramp has proven effective in delaying the separation of a turbulent boundary layer. Measurements of the ramp surface static pressure coefficient obtained under the condition of vane oscillation and constant inlet velocity revealed that two different effects occurred with surface distance along the ramp. In the vicinity of the oscillating vane, the pressure coefficients varied as a negative function of the vane's trailing edge rms velocity; the independent variable on which the rms velocity depends are the vane's oscillation frequency and its displacement amplitude. From a point downstream of the vane to the exit of the ramp; however, the pressure coefficient varied as a more complex function of the two independent variables. That is, it was found to vary as a function of the vane's oscillation frequency throughout the entire range of frequencies covered during the test, but over only a limited range of the trailing edge displacement amplitudes covered. More specifically, the value of the pressure coefficient was independent of increases in the vane's displacement amplitude above approximately 35 inner wall units of the boundary layer. Below this specific amplitude it varied as a function of the vane's trailing edge rms velocity. This height is close to the upper limit of the buffer layer. A parametric study was made to determine the variation of the maximum static pressure recovery as a function of the vane's oscillation frequency, for several ramp inlet velocities and a constant displacement amplitude of the vane's trailing edge. The results indicate that the phenomenon producing the optimum delay of separation may be Strouhal number dependent. Corona anemometer measurements obtained in the inner wall regions of the boundary layer for the excited case reveal a large range of unsteadiness in the local velocities. These measurements imply the existence of inflections in the profiles, which provide a mechanism for resulting inviscid flow instabilities to produce turbulence in the near wall region, thereby delaying separation of the boundary layer.

Reynolds-Stress Budgets in an Impinging Shock Wave/Boundary-Layer Interaction

NASA Technical Reports Server (NTRS)

Vyas, Manan A.; Yoder, Dennis A.; Gaitonde, Datta V.

2018-01-01

Implicit large-eddy simulation (ILES) of a shock wave/boundary-layer interaction (SBLI) was performed. Comparisons with experimental data showed a sensitivity of the current prediction to the modeling of the sidewalls. This was found to be common among various computational studies in the literature where periodic boundary conditions were used in the spanwise direction, as was the case in the present work. Thus, although the experiment was quasi-two-dimensional, the present simulation was determined to be two-dimensional. Quantities present in the exact equation of the Reynolds-stress transport, i.e., production, molecular diffusion, turbulent transport, pressure diffusion, pressure strain, dissipation, and turbulent mass flux were calculated. Reynolds-stress budgets were compared with past large-eddy simulation and direct numerical simulation datasets in the undisturbed portion of the turbulent boundary layer to validate the current approach. The budgets in SBLI showed the growth in the production term for the primary normal stress and energy transfer mechanism was led by the pressure strain term in the secondary normal stresses. The pressure diffusion term, commonly assumed as negligible by turbulence model developers, was shown to be small but non-zero in the normal stress budgets, however it played a key role in the primary shear stress budget.

A Marine Boundary Layer Water Vapor Climatology Derived from Microwave and Near-Infrared Imagery

NASA Astrophysics Data System (ADS)

Millan Valle, L. F.; Lebsock, M. D.; Teixeira, J.

2017-12-01

The synergy of the collocated Advanced Microwave Scanning Radiometer (AMSR) and the Moderate Resolution Imaging Spectroradiometer (MODIS) provides daily global estimates of partial marine planetary boundary layer water vapor. AMSR microwave radiometry provides the total column water vapor, while MODIS near-infrared imagery provides the water vapor above the cloud layers. The difference between the two gives the vapor between the surface and the cloud top, which may be interpreted as the boundary layer water vapor. Comparisons against radiosondes, and GPS-Radio occultation data demonstrate the robustness of these boundary layer water vapor estimates. We exploit the 14 years of AMSR-MODIS synergy to investigate the spatial, seasonal, and inter-annual variations of the boundary layer water vapor. Last, it is shown that the measured AMSR-MODIS partial boundary layer water vapor can be generally prescribed using sea surface temperature, cloud top pressure and the lifting condensation level. The multi-sensor nature of the analysis demonstrates that there exists more information on boundary layer water vapor structure in the satellite observing system than is commonly assumed when considering the capabilities of single instruments. 2017 California Institute of Technology. U.S. Government sponsorship acknowledged.

Unsteady pressure loads in a generic high speed engine model

NASA Technical Reports Server (NTRS)

Parrott, Tony L.; Jones, Michael G.; Thurlow, Ernie M.

1992-01-01

Unsteady pressure loads were measured along the top interior wall of a generic high-speed engine (GHSE) model undergoing performance tests in the combustion-Heated Scramjet Test Facility at the Langley Research Center. Flow to the model inlet was simulated at 72000 ft and a flight Mach number of 4. The inlet Mach number was 3.5 with a total temperature and pressure of 1640 R and 92 psia. The unsteady pressure loads were measured with 5 piezoresistive gages, recessed into the wall 4 to 12 gage diameters to reduce incident heat flux to the diaphragms, and distributed from the inlet to the combustor. Contributors to the unsteady pressure loads included boundary layer turbulence, combustion noise, and transients generated by unstart loads. Typical turbulent boundary layer rms pressures in the inlet ranged from 133 dB in the inlet to 181 dB in the combustor over the frequency range from 0 to 5 kHz. Downstream of the inlet exist, combustion noise was shown to dominate boundary layer turbulence noise at increased heat release rates. Noise levels in the isolator section increased by 15 dB when the fuel-air ratio was increased from 0.37 to 0.57 of the stoichiometric ratio. Transient pressure disturbances associated with engine unstarts were measured in the inlet and have an upstream propagation speed of about 7 ft/sec and pressure jumps of at least 3 psia.

Delay of Turbulent Boundary Layer Detachment by Mechanical Excitation: Application to Rearward-facing Ramp

NASA Technical Reports Server (NTRS)

McKinzie, Daniel J., Jr.

1996-01-01

A vane oscillating about a fixed point at the inlet to a two-dimensional 20 deg rearward-facing ramp proved effective in delaying the detachment of a turbulent boundary layer. Flow-field, surface static pressure, and smoke-wire flow visualization measurements were made. Surface pressure coefficient distributions revealed that two different effects occurred with axial distance along the ramp surface. The surface pressure coefficient varied as a complex function of the vane oscillation frequency and its trailing edge displacement amplitude; that is, it varied as a function of the vane oscillation frequency throughout the entire range of frequencies covered during the test, but it varied over only a limited range of the trailing edge displacement amplitudes covered.The complexity of these findings prompted a detailed investigation, the results of which revealed a combination of phenomena that explain qualitatively how the mechanically generated, periodic, sinusoidal perturbing signal produced by the oscillating vane reacts with the fluid flow to delay the detachment of a turbulent boundary layer experiencing transitory detachment.

A Comparison of Active and Passive Methods for Control of Hypersonic Boundary Layers on Airbreathing Configurations

NASA Technical Reports Server (NTRS)

Berry, Scott A.; Nowak, Robert J.

2003-01-01

Active and passive methods for control of hypersonic boundary layers have been experimentally examined in NASA Langley Research Center wind tunnels on a Hyper-X model. Several configurations for forcing transition using passive discrete roughness elements and active mass addition, or blowing, methods were compared in two hypersonic facilities, the 20-Inch Mach 6 Air and the 31-Inch Mach 10 Air tunnels. Heat transfer distributions, obtained via phosphor thermography, shock system details, and surface streamline patterns were measured on a 0.333-scale model of the Hyper-X forebody. The comparisons between the active and passive methods for boundary layer control were conducted at test conditions that nearly match the nominal Mach 7 flight trajectory of an angle-of-attack of 2-deg and length Reynolds number of 5.6 million. For the passive roughness examination, the primary parametric variation was a range of trip heights within the calculated boundary layer thickness for several trip concepts. The prior passive roughness study resulted in a swept ramp configuration being selected for the Mach 7 flight vehicle that was scaled to be roughly 0.6 of the calculated boundary layer thickness. For the active jet blowing study, the blowing manifold pressure was systematically varied for each configuration, while monitoring the mass flow, to determine the jet penetration height with schlieren and transition movement with the phosphor system for comparison to the passive results. All the blowing concepts tested were adequate for providing transition onset near the trip location with manifold stagnation pressures on the order of 40 times the model static pressure or higher.

Scaling of heat transfer augmentation due to mechanical distortions in hypervelocity boundary layers

NASA Astrophysics Data System (ADS)

Flaherty, W.; Austin, J. M.

2013-10-01

We examine the response of hypervelocity boundary layers to global mechanical distortions due to concave surface curvature. Surface heat transfer and visual boundary layer thickness data are obtained for a suite of models with different concave surface geometries. Results are compared to predictions using existing approximate methods. Near the leading edge, good agreement is observed, but at larger pressure gradients, predictions diverge significantly from the experimental data. Up to a factor of five underprediction is reported in regions with greatest distortion. Curve fits to the experimental data are compared with surface equations. We demonstrate that reasonable estimates of the laminar heat flux augmentation may be obtained as a function of the local turning angle for all model geometries, even at the conditions of greatest distortion. This scaling may be explained by the application of Lees similarity. As a means of introducing additional local distortions, vortex generators are used to impose streamwise structures into the boundary layer. The response of the large scale vortices to an adverse pressure gradient is investigated. Surface streak evolution is visualized over the different surface geometries using fast response pressure sensitive paint. For a flat plate baseline case, heat transfer augmentation at similar levels to turbulent flow is measured. For the concave geometries, increases in heat transfer by factors up to 2.6 are measured over the laminar values. The scaling of heat transfer with turning angle that is identified for the laminar boundary layer response is found to be robust even in the presence of the imposed vortex structures.

Compressible turbulence measurements in a supersonic boundary layer including favorable pressure gradient effects

NASA Astrophysics Data System (ADS)

Miller, Raymond S.

1994-12-01

The effect of a favorable pressure gradient on the turbulent flow structure in a Mach 2.9 boundary layer (Re/m approximately equal to 1.5 x 10(exp 7)) is investigated experimentally. Conventional flow and hot film measurements of turbulent fluctuation properties have been made upstream of and along an expansion ramp. Upstream measurements were taken in a zero pressure gradient boundary layer 44 cm from the nozzle throat in a 6.35 cm square test section. Measurements are obtained in the boundary layer, above the expansion ramp, 71.5 cm from the nozzle throat. Mean flow and turbulent flow characteristics are measured in all three dimensions. Comparisons are made between data obtained using single and multiple-overheat cross-wire anemometry as well as conventional mean flow probes. Conventional flow measurements were taken using a Pitot probe and a 10 degree cone static probe. Flow visualization was conducted via imaging techniques (Schlieren and shadowgraph photographs). Results suggest that compressibility effects, as seen through the density fluctuations in the Reynolds shear stress, are roughly 10% relative to the mean velocity and are large relative to the velocity fluctuations. This is also observed in the total Reynolds shear stress; compressibility accounts for 50-75% of the total shear. This is particularly true in the favorable pressure gradient region, where though the peak fluctuation intensities are diminished, the streamwise component of the mean flow is larger, hence the contribution of the compressibility term is significant in the Reynolds shear.

Flow regimes in a trapped vortex cell

NASA Astrophysics Data System (ADS)

Lasagna, D.; Iuso, G.

2016-03-01

This paper presents results of an experimental investigation on the flow in a trapped vortex cell, embedded into a flat plate, and interacting with a zero-pressure-gradient boundary layer. The objective of the work is to describe the flow features and elucidate some of the governing physical mechanisms, in the light of recent investigations on flow separation control using vortex cells. Hot-wire velocity measurements of the shear layer bounding the cell and of the boundary layers upstream and downstream are reported, together with spectral and correlation analyses of wall-pressure fluctuation measurements. Smoke flow visualisations provide qualitative insight into some relevant features of the internal flow, namely a large-scale flow unsteadiness and possible mechanisms driving the rotation of the vortex core. Results are presented for two very different regimes: a low-Reynolds-number case where the incoming boundary layer is laminar and its momentum thickness is small compared to the cell opening, and a moderately high-Reynolds-number case, where the incoming boundary layer is turbulent and the ratio between the momentum thickness and the opening length is significantly larger than in the first case. Implications of the present findings to flow control applications of trapped vortex cells are also discussed.

Wall turbulence control

NASA Technical Reports Server (NTRS)

Wilkinson, Stephen P.; Lindemann, A. Margrethe; Beeler, George B.; Mcginley, Catherine B.; Goodman, Wesley L.; Balasubramanian, R.

1986-01-01

A variety of wall turbulence control devices which were experimentally investigated are discussed; these include devices for burst control, alteration of outer flow structures, large eddy substitution, increased heat transfer efficiency, and reduction of wall pressure fluctuations. Control of pre-burst flow was demonstrated with a single, traveling surface depression which is phase-locked to elements of the burst production process. Another approach to wall turbulence control is to interfere with the outer layer coherent structures. A device in the outer part of a boundary layer was shown to suppress turbulence and reduce drag by opposing both the mean and unsteady vorticity in the boundary layer. Large eddy substitution is a method in which streamline curvature is introduced into the boundary layer in the form of streamwise vortices. Riblets, which were already shown to reduce turbulent drag, were also shown to exhibit superior heat transfer characteristics. Heat transfer efficiency as measured by the Reynolds Analogy Factor was shown to be as much as 36 percent greater than a smooth flat plate in a turbulent boundary layer. Large Eddy Break-Up (LEBU) which are also known to reduce turbulent drag were shown to reduce turbulent wall pressure fluctuation.

The Effects of Blade Count on Boundary Layer Development in a Low-Pressure Turbine

NASA Technical Reports Server (NTRS)

Dorney, Daniel J.; Flitan, Horia C.; Ashpis, David E.; Solomon, William J.

2000-01-01

Experimental data from jet-engine tests have indicated that turbine efficiencies at takeoff can be as much as two points higher than those at cruise conditions. Recent studies have shown that Reynolds number effects contribute to the lower efficiencies at cruise conditions. In the current study numerical simulations have been performed to study the boundary layer development in a two-stage low-pressure turbine, and to evaluate the models available for low Reynolds number flows in turbomachinery. In a previous study using the same geometry the predicted time-averaged boundary layer quantities showed excellent agreement with the experimental data, but the predicted unsteady results showed only fair agreement with the experimental data. It was surmised that the blade count approximation used in the numerical simulations generated more unsteadiness than was observed in the experiments. In this study a more accurate blade approximation has been used to model the turbine, and the method of post-processing the boundary layer information has been modified to more closely resemble the process used in the experiments. The predicted results show improved agreement with the unsteady experimental data.

High Reynolds Number Investigation of a Flush-Mounted, S-Duct Inlet With Large Amounts of Boundary Layer Ingestion

NASA Technical Reports Server (NTRS)

Berrier, Bobby L.; Carter, Melissa B.; Allan, Brian G.

2005-01-01

An experimental investigation of a flush-mounted, S-duct inlet with large amounts of boundary layer ingestion has been conducted at Reynolds numbers up to full scale. The study was conducted in the NASA Langley Research Center 0.3-Meter Transonic Cryogenic Tunnel. In addition, a supplemental computational study on one of the inlet configurations was conducted using the Navier-Stokes flow solver, OVERFLOW. Tests were conducted at Mach numbers from 0.25 to 0.83, Reynolds numbers (based on aerodynamic interface plane diameter) from 5.1 million to 13.9 million (full-scale value), and inlet mass-flow ratios from 0.29 to 1.22, depending on Mach number. Results of the study indicated that increasing Mach number, increasing boundary layer thickness (relative to inlet height) or ingesting a boundary layer with a distorted profile decreased inlet performance. At Mach numbers above 0.4, increasing inlet airflow increased inlet pressure recovery but also increased distortion. Finally, inlet distortion was found to be relatively insensitive to Reynolds number, but pressure recovery increased slightly with increasing Reynolds number.

Derivation of Zagarola-Smits scaling in zero-pressure-gradient turbulent boundary layers

NASA Astrophysics Data System (ADS)

Wei, Tie; Maciel, Yvan

2018-01-01

This Rapid Communication derives the Zagarola-Smits scaling directly from the governing equations for zero-pressure-gradient turbulent boundary layers (ZPG TBLs). It has long been observed that the scaling of the mean streamwise velocity in turbulent boundary layer flows differs in the near surface region and in the outer layer. In the inner region of small-velocity-defect boundary layers, it is generally accepted that the proper velocity scale is the friction velocity, uτ, and the proper length scale is the viscous length scale, ν /uτ . In the outer region, the most generally used length scale is the boundary layer thickness, δ . However, there is no consensus on velocity scales in the outer layer. Zagarola and Smits [ASME Paper No. FEDSM98-4950 (1998)] proposed a velocity scale, U ZS=(δ1/δ ) U∞ , where δ1 is the displacement thickness and U∞ is the freestream velocity. However, there are some concerns about Zagarola-Smits scaling due to the lack of a theoretical base. In this paper, the Zagarola-Smits scaling is derived directly from a combination of integral, similarity, and order-of-magnitude analysis of the mean continuity equation. The analysis also reveals that V∞, the mean wall-normal velocity at the edge of the boundary layer, is a proper scale for the mean wall-normal velocity V . Extending the analysis to the streamwise mean momentum equation, we find that the Reynolds shear stress in ZPG TBLs scales as U∞V∞ in the outer region. This paper also provides a detailed analysis of the mass and mean momentum balance in the outer region of ZPG TBLs.

A perspective on coherent structures and conceptual models for turbulent boundary layer physics

NASA Technical Reports Server (NTRS)

Robinson, Stephen K.

1990-01-01

Direct numerical simulations of turbulent boundary layers have been analyzed to develop a unified conceptual model for the kinematics of coherent motions in low Reynolds number canonical turbulent boundary layers. All classes of coherent motions are considered in the model, including low-speed streaks, ejections and sweeps, vortical structures, near-wall and outer-region shear layers, sublayer pockets, and large-scale outer-region eddies. The model reflects the conclusions from the study of the simulated boundary layer that vortical structures are directly associated with the production of turbulent shear stresses, entrainment, dissipation of turbulence kinetic energy, and the fluctuating pressure field. These results, when viewed from the perspective of the large body of published work on the subject of coherent motions, confirm that vortical structures may be considered the central dynamic element in the maintenance of turbulence in the canonical boundary layer. Vortical structures serve as a framework on which to construct a unified picture of boundary layer structure, providing a means to relate the many known structural elements in a consistent way.

An Innovative Flow-Measuring Device: Thermocouple Boundary Layer Rake

NASA Technical Reports Server (NTRS)

Hwang, Danny P.; Fralick, Gustave C.; Martin, Lisa C.; Wrbanek, John D.; Blaha, Charles A.

2001-01-01

An innovative flow-measuring device, a thermocouple boundary layer rake, was developed. The sensor detects the flow by using a thin-film thermocouple (TC) array to measure the temperature difference across a heater strip. The heater and TC arrays are microfabricated on a constant-thickness quartz strut with low heat conductivity. The device can measure the velocity profile well into the boundary layer, about 65 gm from the surface, which is almost four times closer to the surface than has been possible with the previously used total pressure tube.

Turbulent boundary-layer velocity profiles on a nonadiabatic at Mach number 6.5

NASA Technical Reports Server (NTRS)

Keener, E. R.; Hopkins, E. J.

1972-01-01

Velocity profiles were obtained from pitot-pressure and total-temperature measurements within a turbulent boundary layer on a large sharp-edged flat plate. Momentum-thickness Reynolds number ranged from 2590 to 8860 and wall-to-adiabatic-wall temperature ratios ranged from 0.3 to 0.5. Measurements were made both with and without boundary layer trips. Five methods are evaluated for correlating the measured velocity profiles with the incompressible law-of-the-wall and the velocity defect law. The mixing-length generalization of Van Driest gives the best correlation.

Introduction to boundary-layer theory. [viscous friction loss calculation for turbine blade design

NASA Technical Reports Server (NTRS)

Mcnally, W. D.

1973-01-01

The pressure ratio across a turbine provides a certain amount of ideal energy that is available to the turbine for producing work. The portion of the ideal energy that is not converted to work is considered to be a loss. One of the more important and difficult aspects of turbine design is the prediction of the losses. The primary cause of losses is the boundary layer that develops on the blade and end wall surfaces. Boundary-layer theory is used to calculate the parameters needed to estimate viscous (friction) losses.

Wall Turbulence with Designer Properties: Identification, Characterization and Manipulation of Energy Pathways

DTIC Science & Technology

2016-02-26

zero-pressure- gradient boundary layer to develop over a flat plate . As shown in figure 6.1, the flat plate contains an insert to allow for a thin...B. J. ‘Triadic scale interactions in a turbulent boundary layer ’ J. Fluid Mech., 767, R4 (2015). 6. Luhar, M., Sharma, A. S. & McKeon, B. J. ‘A... boundary layer ’, Paper H22.00003, 68th Meeting of the American Physical Society Division of Fluid Dynamics, Boston, MA, Nov., 2015.  Duvvuri

Unified aeroacoustics analysis for high speed turboprop aerodynamics and noise. Volume 5: Propagation of propeller tone noise through a fuselage boundary layer

NASA Technical Reports Server (NTRS)

Magliozzi, B.; Hanson, D. B.

1991-01-01

An analysis of tone noise propagation through a boundary layer and fuselage scattering effects was derived. This analysis is a three dimensional and the complete wave field is solved by matching analytical expressions for the incident and scattered waves in the outer flow to a numerical solution in the boundary layer flow. The outer wave field is constructed analytically from an incident wave appropriate to the source and a scattered wave in the standard Hankel function form. For the incident wave, an existing function - domain propeller noise radiation theory is used. In the boundary layer region, the wave equation is solved by numerical methods. The theoretical analysis is embodied in a computer program which allows the calculation of correction factors for the fuselage scattering and boundary layer refraction effects. The effects are dependent on boundary layer profile, flight speed, and frequency. Corrections can be derived for any point on the fuselage, including those on the opposite side from the source. The theory was verified using limited cases and by comparing calculations with available measurements from JetStar tests of model prop-fans. For the JetStar model scale, the boundary layer refraction effects produce moderate fuselage pressure reinforcements aft of and near the plane of rotation and significant attenuation forward of the plane of rotation at high flight speeds. At lower flight speeds, the calculated boundary layer effects result in moderate amplification over the fuselage area of interest. Apparent amplification forward of the plane of rotation is a result of effective changes in the source directivity due to boundary layer refraction effects. Full scale effects are calculated to be moderate, providing fuselage pressure amplification of about 5 dB at the peak noise location. Evaluation using available noise measurements was made under high-speed, high-altitude flight conditions. Comparisons of calculations made of free field noise, using a current frequency-domain propeller noise prediction method, and fuselage effects using this new procedure show good agreement with fuselage measurements over a wide range of flight speeds and frequencies. Correction factors for the JetStar measurements made on the fuselage are provided in an Appendix.

Investigation of blown boundary layers with an improved wall jet system

NASA Technical Reports Server (NTRS)

Saripalli, K. R.; Simpson, R. L.

1980-01-01

Measurements were made in a two dimensional incompressible wall jet submerged under a thick upstream boundary layer with a zero pressure gradient and an adverse pressure gradient. The measurements included mean velocity and Reynolds stresses profiles, skin friction, and turbulence spectra. The measurements were confined to practical ratios (less than 2) of the jet velocity to the free stream velocity. The wall jet used in the experiments had an asymmetric velocity profile with a relatively higher concentration of momentum away from the wall. An asymmetric jet velocity profile has distinct advantages over a uniform jet velocity profile, especially in the control of separation. Predictions were made using Irwin's (1974) method for blown boundary layers. The predictions clearly show the difference in flow development between an asymmetric jet velocity profile and a uniform jet velocity profile.

Base Heating Sensitivity Study for a 4-Cluster Rocket Motor Configuration in Supersonic Freestream

NASA Technical Reports Server (NTRS)

Mehta, Manish; Canabal, Francisco; Tashakkor, Scott B.; Smith, Sheldon D.

2011-01-01

In support of launch vehicle base heating and pressure prediction efforts using the Loci-CHEM Navier-Stokes computational fluid dynamics solver, 35 numerical simulations of the NASA TND-1093 wind tunnel test have been modeled and analyzed. This test article is composed of four JP-4/LOX 500 lbf rocket motors exhausting into a Mach 2 - 3.5 wind tunnel at various ambient pressure conditions. These water-cooled motors are attached to a base plate of a standard missile forebody. We explore the base heating profiles for fully coupled finite-rate chemistry simulations, one-way coupled RAMP (Reacting And Multiphase Program using Method of Characteristics)-BLIMPJ (Boundary Layer Integral Matrix Program - Jet Version) derived solutions and variable and constant specific heat ratio frozen flow simulations. Variations in turbulence models, temperature boundary conditions and thermodynamic properties of the plume have been investigated at two ambient pressure conditions: 255 lb/sq ft (simulated low altitude) and 35 lb/sq ft (simulated high altitude). It is observed that the convective base heat flux and base temperature are most sensitive to the nozzle inner wall thermal boundary layer profile which is dependent on the wall temperature, boundary layer s specific energy and chemical reactions. Recovery shock dynamics and afterburning significantly influences convective base heating. Turbulence models and external nozzle wall thermal boundary layer profiles show less sensitivity to base heating characteristics. Base heating rates are validated for the highest fidelity solutions which show an agreement within +/-10% with respect to test data.

Fuselage Boundary Layer Ingestion Propulsion Applied to a Thin Haul Commuter Aircraft for Optimal Efficiency

NASA Technical Reports Server (NTRS)

Mikic, Gregor Veble; Stoll, Alex; Bevirt, JoeBen; Grah, Rok; Moore, Mark D.

2016-01-01

Theoretical and numerical aspects of aerodynamic efficiency of propulsion systems are studied. Focus is on types of propulsion that closely couples to the aerodynamics of the complete vehicle. We discuss the effects of local flow fields, which are affected both by conservative flow acceleration as well as total pressure losses, on the efficiency of boundary layer immersed propulsion devices. We introduce the concept of a boundary layer retardation turbine that helps reduce skin friction over the fuselage. We numerically investigate efficiency gains offered by boundary layer and wake interacting devices. We discuss the results in terms of a total energy consumption framework and show that efficiency gains offered depend on all the elements of the propulsion system.

High-Reynolds-number turbulent-boundary-layer wall pressure fluctuations with skin-friction reduction by air injection.

PubMed

Winkel, Eric S; Elbing, Brian R; Ceccio, Steven L; Perlin, Marc; Dowling, David R

2008-05-01

The hydrodynamic pressure fluctuations that occur on the solid surface beneath a turbulent boundary layer are a common source of flow noise. This paper reports multipoint surface pressure fluctuation measurements in water beneath a high-Reynolds-number turbulent boundary layer with wall injection of air to reduce skin-friction drag. The experiments were conducted in the U.S. Navy's Large Cavitation Channel on a 12.9-m-long, 3.05-m-wide hydrodynamically smooth flat plate at freestream speeds up to 20 ms and downstream-distance-based Reynolds numbers exceeding 200 x 10(6). Air was injected from one of two spanwise slots through flush-mounted porous stainless steel frits (approximately 40 microm mean pore diameter) at volume flow rates from 17.8 to 142.5 l/s per meter span. The two injectors were located 1.32 and 9.78 m from the model's leading edge and spanned the center 87% of the test model. Surface pressure measurements were made with 16 flush-mounted transducers in an "L-shaped" array located 10.7 m from the plate's leading edge. When compared to no-injection conditions, the observed wall-pressure variance was reduced by as much as 87% with air injection. In addition, air injection altered the inferred convection speed of pressure fluctuation sources and the streamwise coherence of pressure fluctuations.

Analysis of Numerical Simulation Database for Pressure Fluctuations Induced by High-Speed Turbulent Boundary Layers

NASA Technical Reports Server (NTRS)

Duan, Lian; Choudhari, Meelan M.

2014-01-01

Direct numerical simulations (DNS) of Mach 6 turbulent boundary layer with nominal freestream Mach number of 6 and Reynolds number of Re(sub T) approximately 460 are conducted at two wall temperatures (Tw/Tr = 0.25, 0.76) to investigate the generated pressure fluctuations and their dependence on wall temperature. Simulations indicate that the influence of wall temperature on pressure fluctuations is largely limited to the near-wall region, with the characteristics of wall-pressure fluctuations showing a strong temperature dependence. Wall temperature has little influence on the propagation speed of the freestream pressure signal. The freestream radiation intensity compares well between wall-temperature cases when normalized by the local wall shear; the propagation speed of the freestream pressure signal and the orientation of the radiation wave front show little dependence on the wall temperature.

Similar solutions for viscous hypersonic flow over a slender three-fourths-power body of revolution

NASA Technical Reports Server (NTRS)

Lin, Chin-Shun

1987-01-01

For hypersonic flow with a shock wave, there is a similar solution consistent throughout the viscous and inviscid layers along a very slender three-fourths-power body of revolution The strong pressure interaction problem can then be treated by the method of similarity. Numerical calculations are performed in the viscous region with the edge pressure distribution known from the inviscid similar solutions. The compressible laminar boundary-layer equations are transformed into a system of ordinary differential equations. The resulting two-point boundary value problem is then solved by the Runge-Kutta method with a modified Newton's method for the corresponding boundary conditions. The effects of wall temperature, mass bleeding, and body transverse curvature are investigated. The induced pressure, displacement thickness, skin friction, and heat transfer due to the previously mentioned parameters are estimated and analyzed.

Experimental Investigation of the Effects of Viscosity on the Drag and Base Pressure of Bodies of Revolution at a Mach Number 1.5

NASA Technical Reports Server (NTRS)

Chapman, Dean R; Perkins, Edward W

1951-01-01

Models were tested to evaluate effects of Reynolds number for both laminar and turbulent boundary layers. Principal geometric variables investigated were afterbody shape and length-diameter ratio. Force tests and base-pressure measurements were made. Schlieren photographs were used to analyze the effects of viscosity on flow separation and shock-wave configuration and to verify the condition of the boundary layer as deduced from the force tests. The results are discussed and compared with theoretical calculations.

Investigation at supersonic and subsonic Mach numbers of auxiliary inlets supplying secondary air flow to ejector exhaust nozzles

NASA Technical Reports Server (NTRS)

Hearth, Donald P; Cubbison, Robert W

1956-01-01

The results indicated increases in auxiliary-inlet pressure recovery with increases in scoop height relative to the boundary-layer thickness. The pressure recovery increased at about the same rate as theoretically predicted for an inlet in a boundary layer having a one-seventh power profile, but was only about 0.68 to 0.75 of the theoretically obtainable values. Under some operating conditions, flow from the primary jet was exhausted through the auxiliary inlet. This phenomenon could be predicted from the ejector pumping characteristics.

Role of boundary layer diffusion in vapor deposition growth of chalcogenide nanosheets: the case of GeS.

PubMed

Li, Chun; Huang, Liang; Snigdha, Gayatri Pongur; Yu, Yifei; Cao, Linyou

2012-10-23

We report a synthesis of single-crystalline two-dimensional GeS nanosheets using vapor deposition processes and show that the growth behavior of the nanosheet is substantially different from those of other nanomaterials and thin films grown by vapor depositions. The nanosheet growth is subject to strong influences of the diffusion of source materials through the boundary layer of gas flows. This boundary layer diffusion is found to be the rate-determining step of the growth under typical experimental conditions, evidenced by a substantial dependence of the nanosheet's size on diffusion fluxes. We also find that high-quality GeS nanosheets can grow only in the diffusion-limited regime, as the crystalline quality substantially deteriorates when the rate-determining step is changed away from the boundary layer diffusion. We establish a simple model to analyze the diffusion dynamics in experiments. Our analysis uncovers an intuitive correlation of diffusion flux with the partial pressure of source materials, the flow rate of carrier gas, and the total pressure in the synthetic setup. The observed significant role of boundary layer diffusions in the growth is unique for nanosheets. It may be correlated with the high growth rate of GeS nanosheets, ~3-5 μm/min, which is 1 order of magnitude higher than other nanomaterials (such as nanowires) and thin films. This fundamental understanding of the effect of boundary layer diffusions may generally apply to other chalcogenide nanosheets that can grow rapidly. It can provide useful guidance for the development of general paradigms to control the synthesis of nanosheets.

Transient Growth Analysis of Compressible Boundary Layers with Parabolized Stability Equations

NASA Technical Reports Server (NTRS)

Paredes, Pedro; Choudhari, Meelan M.; Li, Fei; Chang, Chau-Lyan

2016-01-01

The linear form of parabolized linear stability equations (PSE) is used in a variational approach to extend the previous body of results for the optimal, non-modal disturbance growth in boundary layer flows. This methodology includes the non-parallel effects associated with the spatial development of boundary layer flows. As noted in literature, the optimal initial disturbances correspond to steady counter-rotating stream-wise vortices, which subsequently lead to the formation of stream-wise-elongated structures, i.e., streaks, via a lift-up effect. The parameter space for optimal growth is extended to the hypersonic Mach number regime without any high enthalpy effects, and the effect of wall cooling is studied with particular emphasis on the role of the initial disturbance location and the value of the span-wise wavenumber that leads to the maximum energy growth up to a specified location. Unlike previous predictions that used a basic state obtained from a self-similar solution to the boundary layer equations, mean flow solutions based on the full Navier-Stokes (NS) equations are used in select cases to help account for the viscous-inviscid interaction near the leading edge of the plate and also for the weak shock wave emanating from that region. These differences in the base flow lead to an increasing reduction with Mach number in the magnitude of optimal growth relative to the predictions based on self-similar mean-flow approximation. Finally, the maximum optimal energy gain for the favorable pressure gradient boundary layer near a planar stagnation point is found to be substantially weaker than that in a zero pressure gradient Blasius boundary layer.

Determination of the Pressure Drag of Airfoils by Integration of Surface Pressures

NASA Technical Reports Server (NTRS)

Phillips, William H.

1990-01-01

A study was conducted of the causes of pressure drag of subsonic airfoils. In a previous paper by the author, the pressure drag is obtained by calculating the total drag from the momentum defect in the boundary layer at the trailing edge and subtracting the friction drag obtained from integration of surface friction along the chord. Herein, the pressure drag is obtained by integrating the streamwise components of surface pressure around the airfoil. Studies were made to verify the accuracy of the integration procedure. The values of pressure drag were much smaller than those obtained by the previous method. This lack of agreement is attributed to the difficulty of calculating boundary layer conditions in the vicinity of the trailing edge and to the extreme sensitivity of the circulation and lift to the trailing edge conditions. The results of these studies are compared with those of previous investigations.

An interpretation of radiosonde errors in the atmospheric boundary layer

Treesearch

Bernadette H. Connell; David R. Miller

1995-01-01

The authors review sources of error in radiosonde measurements in the atmospheric boundary layer and analyze errors of two radiosonde models manufactured by Atmospheric Instrumentation Research, Inc. The authors focus on temperature and humidity lag errors and wind errors. Errors in measurement of azimuth and elevation angles and pressure over short time intervals and...

Earth's core-mantle boundary - Results of experiments at high pressures and temperatures

NASA Technical Reports Server (NTRS)

Knittle, Elise; Jeanloz, Raymond

1991-01-01

Laboratory experiments document that liquid iron reacts chemically with silicates at high pressures (above 2.4 x 10 to the 10th Pa) and temperatures. In particular, (Mg,Fe)SiO3 perovskite, the most abundant mineral of earth's lower mantle, is expected to react with liquid iron to produce metallic alloys (FeO and FeSi) and nonmetallic silicates (SiO2 stishovite and MgSiO3 perovskite) at the pressures of the core-mantle boundary, 14 x 10 to the 10th Pa. The experimental observations, in conjunction with seismological data, suggest that the lowermost 200 to 300 km of earth's mantle, the D-double-prime layer, may be an extremely heterogeneous region as a result of chemical reactions between the silicate mantle and the liquid iron alloy of earth's core. The combined thermal-chemical-electrical boundary layer resulting from such reactions offers a plausible explanation for the complex behavior of seismic waves near the core-mantle boundary and could influence earth's magnetic field observed at the surface.

Experimental and Computational Evaluation of Flush-Mounted, S-Duct Inlets

NASA Technical Reports Server (NTRS)

Berrier, Bobby L.; Allan, Brian G.

2004-01-01

A new high Reynolds number test capability for boundary layer ingesting inlets has been developed for the NASA Langley Research Center 0.3-Meter Transonic Cryogenic Tunnel. Using this new capability. an experimental investigation of four S-duct inlet configurations was conducted. A computational study of one of the inlets was also conducted using a Navier-Stokes solver. The objectives of this investigation were to: 1) develop a new high Reynolds number inlet test capability for flush-mounted inlets; 2) provide a database for CFD tool validation; 3) evaluate the performance of S-duct inlets with large amounts of boundary layer ingestion; and 4) provide a baseline inlet for future inlet flow-control studies. Tests were conducted at Mach numbers from 0.25 to 0.83. Reynolds numbers (based on duct exit diameter) from 5.1 million to a full-scale value of 13.9 million, and inlet mass-flow ratios from 0.39 to 1.58 depending on Mach number. Results of the experimental study indicate that inlet pressure recovery generally decreased and inlet distortion generally increased with increasing Mach number. Except at low Mach numbers, increasing inlet mass-flow increased pressure recovery and increased distortion. Increasing the amount of boundary layer ingestion or ingesting a boundary layer with a distorted profile decreased pressure recovery and increased distortion. Finally, increasing Reynolds number had almost no effect on inlet distortion but increased inlet recovery by about one-half percent at a Mach number near cruise. The computational results captured the inlet pressure recovery and distortion trends with Mach number and inlet mass-flow well: the reversal of the pressure recovery trend with increasing inlet mass-flow at low and high Mach numbers was predicted by CFD. However, CFD results were generally more pessimistic (larger losses) than measured experimentally.

Effect of Reynolds Number and Periodic Unsteady Wake Flow Condition on Boundary Layer Development, Separation, and Intermittency Behavior Along the Suction Surface of a Low Pressure Turbine Blade

NASA Technical Reports Server (NTRS)

Schobeiri, M. T.; Ozturk, B.; Ashpis, David E.

2007-01-01

The paper experimentally studies the effects of periodic unsteady wake flow and different Reynolds numbers on boundary layer development, separation and re-attachment along the suction surface of a low pressure turbine blade. The experimental investigations were performed on a large scale, subsonic unsteady turbine cascade research facility at Turbomachinery Performance and Flow Research Laboratory (TPFL) of Texas A&M University. The experiments were carried out at Reynolds numbers of 110,000 and 150,000 (based on suction surface length and exit velocity). One steady and two different unsteady inlet flow conditions with the corresponding passing frequencies, wake velocities, and turbulence intensities were investigated. The reduced frequencies chosen cover the operating range of LP turbines. In addition to the unsteady boundary layer measurements, surface pressure measurements were performed. The inception, onset, and the extent of the separation bubble information collected from the pressure measurements were compared with the hot wire measurements. The results presented in ensemble-averaged, and the contour plot forms help to understand the physics of the separation phenomenon under periodic unsteady wake flow and different Reynolds number. It was found that the suction surface displayed a strong separation bubble for these three different reduced frequencies. For each condition, the locations defining the separation bubble were determined carefully analyzing and examining the pressure and mean velocity profile data. The location of the boundary layer separation was dependent of the Reynolds number. It is observed that starting point of the separation bubble and the re-attachment point move further downstream by increasing Reynolds number from 110,000 to 150,000. Also, the size of the separation bubble is smaller when compared to that for Re=110,000.

Effect of Reynolds Number and Periodic Unsteady Wake Flow Condition on Boundary Layer Development, Separation, and Re-attachment along the Suction Surface of a Low Pressure Turbine Blade

NASA Technical Reports Server (NTRS)

Ozturk, B.; Schobeiri, M. T.; Ashpis, David E.

2005-01-01

The paper experimentally studies the effects of periodic unsteady wake flow and different Reynolds numbers on boundary layer development, separation and re-attachment along the suction surface of a low pressure turbine blade. The experimental investigations were performed on a large scale, subsonic unsteady turbine cascade research facility at Turbomachinery Performance and Flow Research Laboratory (TPFL) of Texas A&M University. The experiments were carried out at Reynolds numbers of 110,000 and 150,000 (based on suction surface length and exit velocity). One steady and two different unsteady inlet flow conditions with the corresponding passing frequencies, wake velocities, and turbulence intensities were investigated. The reduced frequencies chosen cover the operating range of LP turbines. In addition to the unsteady boundary layer measurements, surface pressure measurements were performed. The inception, onset, and the extent of the separation bubble information collected from the pressure measurements were compared with the hot wire measurements. The results presented in ensemble-averaged, and the contour plot forms help to understand the physics of the separation phenomenon under periodic unsteady wake flow and different Reynolds number. It was found that the suction surface displayed a strong separation bubble for these three different reduced frequencies. For each condition, the locations defining the separation bubble were determined carefully analyzing and examining the pressure and mean velocity profile data. The location of the boundary layer separation was dependent of the Reynolds number. It is observed that starting point of the separation bubble and the re-attachment point move further downstream by increasing Reynolds number from 110,000 to 150,000. Also, the size of the separation bubble is smaller when compared to that for Re=110,000.

An experimental study of the compressor rotor blade boundary layer

NASA Technical Reports Server (NTRS)

Pouagare, M.; Lakshminarayana, B.; Galmes, J. M.

1984-01-01

The three-dimensional turbulent boundary layer developing on a rotor blade of an axial flow compressor was measured using a miniature 'x' configuration hot-wire probe. The measurements were carried out at nine radial locations on both surfaces of the blade at various chordwise locations. The data derived includes streamwise and radial mean velocities and turbulence intensities. The validity of conventional velocity profiles such as the 'power law profile' for the streamwise profile, and Mager and Eichelbrenner's for the radial profile, is examined. A modification to Mager's crossflow profile is proposed. Away from the blade tip, the streamwise component of the blade boundary layer seems to be mainly influenced by the streamwise pressure gradient. Near the tip of the blade, the behavior of the blade boundary layer is affected by the tip leakage flow and the annulus wall boundary layer. The 'tangential blockage' due to the blade boundary layer is derived from the data. The profile losses are found to be less than that of an equivalent cascade, except in the tip region of the blade.

Modeling of the heat transfer in bypass transitional boundary-layer flows

NASA Technical Reports Server (NTRS)

Simon, Frederick F.; Stephens, Craig A.

1991-01-01

A low Reynolds number k-epsilon turbulence model and conditioned momentum, energy and turbulence equations were used to predict bypass transition heat transfer on a flat plate in a high-disturbance environment with zero pressure gradient. The use of conditioned equations was demonstrated to be an improvement over the use of the global-time-averaged equations for the calculation of velocity profiles and turbulence intensity profiles in the transition region of a boundary layer. The approach of conditioned equations is extended to include heat transfer and a modeling of transition events is used to predict transition onset and the extent of transition on a flat plate. The events, which describe the boundary layer at the leading edge, result in boundary-layer regions consisting of: (1) the laminar, (2) pseudolaminar, (3) transitional, and (4) turbulent boundary layers. The modeled transition events were incorporated into the TEXSTAN 2-D boundary-layer code which is used to numerically predict the heat transfer. The numerical predictions in general compared well with the experimental data and revealed areas where additional experimental information is needed.

Investigation of boundary layer and turbulence characteristics inside the passages of an axial flow inducer

NASA Technical Reports Server (NTRS)

Anand, A.; Gorton, C.; Lakshminarayana, B.; Yamaoka, H.

1973-01-01

A study of the boundary layer and turbulence characteristics inside the passages of an axial flow inducer is reported. The first part deals with the analytical and experimental investigation of the boundary layer characteristics in a four bladed flat plate inducer passage operated with no throttle. An approximate analysis for the prediction of radial and chordwise velocity profiles across the passage is carried out. The momentum integral technique is used to predict the gross properties of the boundary layer. Equations are given for the exact analysis of the turbulent boundary layer characteristics using the turbulent field method. Detailed measurement of boundary layer profiles, limiting streamline angle and skin friction stress on the rotating blade is also reported. Part two of this report deals with the prediction of the flow as well as blade static pressure measurements in a three bladed inducer with cambered blades operated at a flow coefficient of 0.065. In addition, the mean velocity and turbulence measurements carried out inside the passage using a rotating triaxial probe is reported.

F-16XL ship #1 CAWAP flight - alpha 15 degrees, altitude 5,000 feet

NASA Technical Reports Server (NTRS)

1996-01-01

The single-seat F-16XL (ship #1) makes another run during the Cranked-Arrow Wing Aerodynamic Project (CAWAP) at Dryden Flight Research Center, Edwards, California. The modified airplane features a delta 'cranked-arrow' wing with strips of tubing along the leading edge to the trailing edge to sense static on the wing and obtain pressure distribution data. The right wing receives data on pressure distribution and the left wing has three types of instrumentation - preston tubes to measure local skin friction, boundary layer rakes to measure boundary layer profiles (the layer where the air interacts with the surfaces of a moving aircraft), and hot films to determine boundary layer transition locations. This photo shows the aircraft gathering data at an altitude of 5000 feet, with an angle of attack of 15 degrees. The program also gathered aero data on two wing planforms for NASA's High Speed Research Program. The first flight of CAWAP occurred on November 21, 1995, and the test program ended in April 1996.

Flow analysis for the nacelle of an advanced ducted propeller at high angle-of-attack and at cruise with boundary layer control

NASA Technical Reports Server (NTRS)

Hwang, D. P.; Boldman, D. R.; Hughes, C. E.

1994-01-01

An axisymmetric panel code and a three dimensional Navier-Stokes code (used as an inviscid Euler code) were verified for low speed, high angle of attack flow conditions. A three dimensional Navier-Stokes code (used as an inviscid code), and an axisymmetric Navier-Stokes code (used as both viscous and inviscid code) were also assessed for high Mach number cruise conditions. The boundary layer calculations were made by using the results from the panel code or Euler calculation. The panel method can predict the internal surface pressure distributions very well if no shock exists. However, only Euler and Navier-Stokes calculations can provide a good prediction of the surface static pressure distribution including the pressure rise across the shock. Because of the high CPU time required for a three dimensional Navier-Stokes calculation, only the axisymmetric Navier-Stokes calculation was considered at cruise conditions. The use of suction and tangential blowing boundary layer control to eliminate the flow separation on the internal surface was demonstrated for low free stream Mach number and high angle of attack cases. The calculation also shows that transition from laminar flow to turbulent flow on the external cowl surface can be delayed by using suction boundary layer control at cruise flow conditions. The results were compared with experimental data where possible.

Characterization of the Boundary Layer on Full-Scale Bluefin Tuna

NASA Astrophysics Data System (ADS)

Amaral, Brian; Cipolla, Kimberly; Henoch, Charles

2014-11-01

The physics that enable tuna to cross large expanses of ocean while feeding and avoiding predators is not presently understood, and could involve complex control of turbulent boundary layer transition and drag reduction. Typical swimming speeds of Bluefin tuna are 1-2 m/s, but can be higher during strong accelerations. The goal of this work is to experimentally determine the approximate lateral location at which transition to turbulence occurs on the tuna for various speeds. The question is whether laminar flow or an advanced propulsion mechanism (or both) allows them to swim at high speeds. Uncertainties include the surface roughness of the skin, local favorable and adverse pressure gradients, and discontinuities such as the open mouth or juncture at the fins. Historically, much of the fluid mechanics work in the area of fish locomotion has focused on vortex shedding issues rather than the boundary layer. Here, the focus is obtaining information on the boundary layer characteristics of a rigid tuna model. A full scale model of a Pacific Bluefin tuna was fabricated using a mold made from an actual deceased tuna, preserving the surface features and details of the appendages. The model was instrumented with 32 wall pressure sensors and experiments performed in a tow tank. Results from flow visualization, drag and wall pressure measurements over a range of speeds and varying angles of attack will be presented.

Pressure and heating-rate distributions on a corrugated surface in a supersonic turbulent boundary layer

NASA Technical Reports Server (NTRS)

Sawyer, J. W.

1977-01-01

Drag and heating rates on wavy surfaces typical of current corrugated plate designs for thermal protection systems were determined experimentally. Pressure-distribution, heating-rate, and oil-flow tests were conducted in the Langley Unitary Plan wind tunnel at Mach numbers of 2.4 and 4.5 with the corrugated surface exposed to both thick and thin turbulent boundary layers. Tests were conducted with the corrugations at cross-flow angles from 0 deg to 90 deg to the flow. Results show that for cross-flow angles of 30 deg or less, the pressure drag coefficients are less than the local flat-plate skin-friction coefficients and are not significantly affected by Mach number, Reynolds number, or boundary-layer thickness over the ranges investigated. For cross-flow angles greater than 30 deg, the drag coefficients increase significantly with cross-flow angle and moderately with Reynolds number. Increasing the Mach number causes a significant reduction in the pressure drag. The average and peak heating penalties due to the corrugated surface are small for cross-flow angles of 10 deg or less but are significantly higher for the larger cross-flow angles.

Effects of local and global mechanical distortions to hypervelocity boundary layers

NASA Astrophysics Data System (ADS)

Flaherty, William P.

The response of hypervelocity boundary layers to global mechanical distortions due to concave surface curvature is examined. Surface heat transfer, visual boundary layer thickness, and pressure sensitive paint (PSP) data are obtained for a suite of models with different concave surface geometries. Results are compared to predictions using existing approximate methods. Near the leading edge, good agreement is observed, but at larger pressure gradients, predictions diverge significantly from the experimental data. Up to a factor of five underprediction is reported in regions with greatest distortion. Curve fits to the experimental data are compared with surface equations. It is demonstrated that reasonable estimates of the laminar heat flux augmentation may be obtained as a function of the local turning angle for all model geometries, even at the conditions of greatest distortion. As a means of introducing additional local distortions, vortex generators are used to impose streamwise structures into the boundary layer. The response of the large scale vortical structures to an adverse pressure gradient is investigated. For a flat plate baseline case, heat transfer augmentation at similar levels to turbulent flow is measured. For the concave geometries, increases in heat transfer by factors up to 2.6 are measured over the laminar values, though for higher turning angle cases, a relaxation to below undisturbed values is reported at turning angles between 10 and 15 degrees. The scaling of heat transfer with turning angle that is identified for the laminar boundary layer response is found to be robust even in the presence of the imposed vortex structures. PSP measurements indicated that natural streaks form over concave models even when imposed vorticity is present. Correlations found between the heat transfer and natural streak formation are discussed and indicate possible vortex interactions.

Boundary Layer Control for Hypersonic Airbreathing Vehicles

NASA Technical Reports Server (NTRS)

Berry, Scott A.; Nowak, Robert J.; Horvath, Thomas J.

2004-01-01

Active and passive methods for tripping hypersonic boundary layers have been examined in NASA Langley Research Center wind tunnels using a Hyper-X model. This investigation assessed several concepts for forcing transition, including passive discrete roughness elements and active mass addition (or blowing), in the 20-Inch Mach 6 Air and the 31-Inch Mach 10 Air Tunnels. Heat transfer distributions obtained via phosphor thermography, shock system details, and surface streamline patterns were measured on a 0.333-scale model of the Hyper-X forebody. The comparisons between the active and passive methods for boundary layer control were conducted at test conditions that nearly match the Hyper-X nominal Mach 7 flight test-point of an angle-of-attack of 2-deg and length Reynolds number of 5.6 million. For passive roughness, the primary parametric variation was a range of trip heights within the calculated boundary layer thickness for several trip concepts. The passive roughness study resulted in a swept ramp configuration, scaled to be roughly 0.6 of the calculated boundary layer thickness, being selected for the Mach 7 flight vehicle. For the active blowing study, the manifold pressure was systematically varied (while monitoring the mass flow) for each configuration to determine the jet penetration height, with schlieren, and transition movement, with the phosphor system, for comparison to the passive results. All the blowing concepts tested, which included various rows of sonic orifices (holes), two- and three-dimensional slots, and random porosity, provided transition onset near the trip location with manifold stagnation pressures on the order of 40 times the model surface static pressure, which is adequate to ensure sonic jets. The present results indicate that the jet penetration height for blowing was roughly half the height required with passive roughness elements for an equivalent amount of transition movement.

High-Area-Ratio Rocket Nozzle at High Combustion Chamber Pressure: Experimental and Analytical Validation

NASA Technical Reports Server (NTRS)

Jankovsky, Robert S.; Smith, Timothy D.; Pavli, Albert J.

1999-01-01

Experimental data were obtained on an optimally contoured nozzle with an area ratio of 1025:1 and on a truncated version of this nozzle with an area ratio of 440:1. The nozzles were tested with gaseous hydrogen and liquid oxygen propellants at combustion chamber pressures of 1800 to 2400 psia and mixture ratios of 3.89 to 6.15. This report compares the experimental performance, heat transfer, and boundary layer total pressure measurements with theoretical predictions of the current Joint Army, Navy, NASA, Air Force (JANNAF) developed methodology. This methodology makes use of the Two-Dimensional Kinetics (TDK) nozzle performance code. Comparisons of the TDK-predicted performance to experimentally attained thrust performance indicated that both the vacuum thrust coefficient and the vacuum specific impulse values were approximately 2.0-percent higher than the turbulent prediction for the 1025:1 configurations, and approximately 0.25-percent higher than the turbulent prediction for the 440:1 configuration. Nozzle wall temperatures were measured on the outside of a thin-walled heat sink nozzle during the test fittings. Nozzle heat fluxes were calculated front the time histories of these temperatures and compared with predictions made with the TDK code. The heat flux values were overpredicted for all cases. The results range from nearly 100 percent at an area ratio of 50 to only approximately 3 percent at an area ratio of 975. Values of the integral of the heat flux as a function of nozzle surface area were also calculated. Comparisons of the experiment with analyses of the heat flux and the heat rate per axial length also show that the experimental values were lower than the predicted value. Three boundary layer rakes mounted on the nozzle exit were used for boundary layer measurements. This arrangement allowed total pressure measurements to be obtained at 14 different distances from the nozzle wall. A comparison of boundary layer total pressure profiles and analytical predictions show good agreement for the first 0.5 in. from the nozzle wall; but the further into the core flow that measurements were taken, the more that TDK overpredicted the boundary layer thickness.

Direct Numerical Simulation of Acoustic Noise Generation from the Nozzle Wall of a Hypersonic Wind Tunnel

NASA Astrophysics Data System (ADS)

Huang, Junji; Duan, Lian; Choudhari, Meelan; Missouri Univ of Sci; Tech Team; NASA Langley Research Center Team

2017-11-01

Direct numerical simulations (DNS) are used to examine the acoustic noise generation from the turbulent boundary layer on the nozzle wall of a Mach 6 Ludwieg Tube. The emphasis is on characterizing the freestream acoustic pressure disturbances radiated from the nozzle-wall turbulent boundary layer and comparing it with acoustic noise generated from a single, flat wall in an unconfined setting at a similar freestream Mach number to assess the effects of noise reverberation. In particular, the numerical database is used to provide insights into the pressure disturbance spectrum and amplitude scaling with respect to the boundary-layer parameters as well as to understand the acoustic source mechanisms. Such information is important for characterizing the freestream disturbance environment in conventional (i.e., noisy) hypersonic wind tunnels. Air Force Office of Scientific Research Award No. FA9550-14-1-0170.

The Compressible Laminar Boundary Layer with Heat Transfer and Arbitrary Pressure Gradient

NASA Technical Reports Server (NTRS)

Cohen, Clarence B; Reshotko, Eli

1956-01-01

An approximate method for the calculation of the compressible laminar boundary layer with heat transfer and arbitrary pressure gradient, based on Thwaites' correlation concept, is presented. With the definition of dimensionless shear and heat-transfer parameters and an assumed correlation of these parameters in terms of a momentum parameter, a complete system of relations for calculating skin friction and heat transfer results. Knowledge of velocity or temperature profiles is not necessary in using this calculation method. When the method is applied to a convergent-divergent, axially symmetric rocket nozzle, it shows that high rates of heat transfer are obtained at the initial stagnation point and at the throat of the nozzle. Also indicated are negative displacement thicknesses in the convergent portion of the nozzle; these occur because of the high density within the lower portions of the cooled boundary layer. (author)

F-8 supercritical wing flight pressure, Boundary layer, and wake measurements and comparisons with wind tunnel data

NASA Technical Reports Server (NTRS)

Montoya, L. C.; Banner, R. D.

1977-01-01

Data for speeds from Mach 0.50 to Mach 0.99 are presented for configurations with and without fuselage area-rule additions, with and without leading-edge vortex generators, and with and without boundary-layer trips on the wing. The wing pressure coefficients are tabulated. Comparisons between the airplane and model data show that higher second velocity peaks occurred on the airplane wing than on the model wing. The differences were attributed to wind tunnel wall interference effects that caused too much rear camber to be designed into the wing. Optimum flow conditions on the outboard wing section occurred at Mach 0.98 at an angle of attack near 4 deg. The measured differences in section drag with and without boundary-layer trips on the wing suggested that a region of laminar flow existed on the outboard wing without trips.

Flight-determined characteristics of an air intake system on an F-111A airplane

NASA Technical Reports Server (NTRS)

Hughes, D. L.; Johnson, H. J.

1972-01-01

Flow phenomena of the F-111A air intake system were investigated over a large range of Mach number, altitude, and angle of attack. Boundary-layer variations are shown for the fuselage splitter plate and inlet entrance stations. Inlet performance is shown in terms of pressure recovery, airflow, mass-flow ratio, turbulence factor, distortion factor, and power spectral density. The fuselage boundary layer was found to be not completely removed from the upper portion of the splitter plate at all Mach numbers investigated. Inlet boundary-layer ingestion started at approximately Mach 1.6 near the translating spike and cone. Pressure-recovery distribution at the compressor face showed increasing distortion with increasing angle of attack and increasing Mach number. The time-averaged distortion-factor value approached 1300, which is near the distortion tolerance of the engine at Mach numbers above 2.1.

Investigation of Turbulent Boundary-Layer Separation Using Laser Velocimetry

NASA Technical Reports Server (NTRS)

Modarress, D.; Johnson, D. A.

1979-01-01

Boundary-layer measurements realized by laser velocimetry are presented for a Much 2.9, two-dimensional, shock-wave/turbulent boundary-layer interaction containing an extensive region of separated flow. Mean velocity and turbulent intensity profiles were obtained from upstream of the interaction zone to downstream of the mean reattachment point. The superiority of the laser velocimeter technique over pressure sensors in turbulent separated flows is demonstrated by a comparison of the laser velocimeter data with results obtained from local pilot and static pressure measurements for the same flow conditions. The locations of the mean separation and reattachment points as deduced from the mean velocity measurements are compared to oil-now visualization results. Representative velocity probability density functions obtained in the separated now region are also presented. Critical to the success of this investigation were: the use of Bragg cell frequency shifting and artificial seeding of the now with submicron light-scattering particles.

Boundary-layer effects in droplet splashing

NASA Astrophysics Data System (ADS)

Riboux, Guillaume; Gordillo, Jose Manuel

2017-11-01

A drop falling onto a solid substrate will disintegrate into smaller parts when its impact velocity exceeds the so called critical velocity for splashing. Under these circumstances, the very thin liquid sheet ejected tangentially to the solid after the drop touches the substrate, lifts off as a consequence of the aerodynamic forces exerted on it and finally breaks into smaller droplets, violently ejected radially outwards, provoking the splash. Here, the tangential deceleration experienced by the fluid entering the thin liquid sheet is investigated making use of boundary layer theory. The velocity component tangent to the solid, computed using potential flow theory provides the far field boundary condition as well as the pressure gradient for the boundary layer equations. The structure of the flow permits to find a self similar solution of the boundary layer equations. This solution is then used to calculate the boundary layer thickness at the root of the lamella as well as the shear stress at the wall. The splash model presented in, which is slightly modified to account for the results obtained from the boundary layer analysis, provides a very good agreement between the measurements and the predicted values of the critical velocity for the splash.

Calculation of turbulent boundary layers with heat transfer and pressure gradient utilizing a compressibility transformation. Part 3: Computer program manual

NASA Technical Reports Server (NTRS)

Schneider, J.; Boccio, J.

1972-01-01

A computer program is described capable of determining the properties of a compressible turbulent boundary layer with pressure gradient and heat transfer. The program treats the two-dimensional problem assuming perfect gas and Crocco integral energy solution. A compressibility transformation is applied to the equation for the conservation of mass and momentum, which relates this flow to a low speed constant property flow with simultaneous mass transfer and pressure gradient. The resulting system of describing equations consists of eight ordinary differential equations which are solved numerically. For Part 1, see N72-12226; for Part 2, see N72-15264.

Rotor blade boundary layer measurement hardware feasibility demonstration

NASA Technical Reports Server (NTRS)

Clark, D. R.; Lawton, T. D.

1972-01-01

A traverse mechanism which allows the measurement of the three dimensional boundary layers on a helicopter rotor blade has been built and tested on a full scale rotor to full scale conditions producing centrifugal accelerations in excess of 400 g and Mach numbers of 0.6 and above. Boundary layer velocity profiles have been measured over a range of rotor speeds and blade collective pitch angles. A pressure scanning switch and transducer were also tested on the full scale rotor and found to be insensitive to centrifugal effects within the normal main rotor operating range. The demonstration of the capability to measure boundary layer behavior on helicopter rotor blades represents the first step toward obtaining, in the rotating system, data of a quality comparable to that already existing for flows in the fixed system.

Large Eddy Simulations of a Bottom Boundary Layer Under a Shallow Geostrophic Front

NASA Astrophysics Data System (ADS)

Bateman, S. P.; Simeonov, J.; Calantoni, J.

2017-12-01

The unstratified surf zone and the stratified shelf waters are often separated by dynamic fronts that can strongly impact the character of the Ekman bottom boundary layer. Here, we use large eddy simulations to study the turbulent bottom boundary layer associated with a geostrophic current on a stratified shelf of uniform depth. The simulations are initialized with a spatially uniform vertical shear that is in geostrophic balance with a pressure gradient due to a linear horizontal temperature variation. Superposed on the temperature front is a stable vertical temperature gradient. As turbulence develops near the bottom, the turbulence-induced mixing gradually erodes the initial uniform temperature stratification and a well-mixed layer grows in height until the turbulence becomes fully developed. The simulations provide the spatial distribution of the turbulent dissipation and the Reynolds stresses in the fully developed boundary layer. We vary the initial linear stratification and investigate its effect on the height of the bottom boundary layer and the turbulence statistics. The results are compared to previous models and simulations of stratified bottom Ekman layers.

Numerical Investigation of Vortex Generator Flow Control for External-Compression Supersonic Inlets

NASA Astrophysics Data System (ADS)

Baydar, Ezgihan

Vortex generators (VGs) within external-compression supersonic inlets for Mach 1.6 were investigated to determine their ability to increase total pressure recovery and reduce total pressure distortion. Ramp and vane-type VGs were studied. The geometric factors of interest included height, length, spacing, angle-of-incidence, and positions upstream and downstream of the inlet terminal shock. The flow through the inlet was simulated numerically through the solution of the steady-state, Reynolds-averaged Navier-Stokes equations on multi-block, structured grids using the Wind-US flow solver. The inlet performance was characterized by the inlet total pressure recovery and the radial and circumferential total pressure distortion indices at the engine face. Previous research of downstream VGs in the low-boom supersonic inlet demonstrated improvement in radial distortion up to 24% while my work on external-compression supersonic inlets improved radial distortion up to 86%, which is significant. The design of experiments and statistical analysis methods were applied to quantify the effect of the geometric factors of VGs and search for optimal VG arrays. From the analysis, VG angle-of-incidence and VG height were the most influential factors in increasing total pressure recovery and reducing distortion. The study on the two-dimensional external-compression inlet determined which passive flow control devices, such as counter-rotating vanes or ramps, reduce high distortion levels and improve the health of the boundary layer, relative to the baseline. Downstream vanes demonstrate up to 21% improvement in boundary layer health and 86% improvement in radial distortion. Upstream vanes demonstrated up to 3% improvement in boundary layer health and 9% improvement in radial distortion. Ramps showed no improvement in boundary layer health and radial distortion. Micro-VGs were preferred for their reduced viscous drag and improvement in total pressure recovery at the AIP. Although traditional VGs energize the flow with stronger vortex structures compared to micro-VGs, the AIP is affected with overwhelming amounts of reduced and enhanced flow regions. In summary, vanes are exceptional in reducing radial distortion and improving the health of the boundary layer compared to the ramps. In the study of the STEX inlet, vane-type vortex generators were the preferred devices for boundary layer flow control. In the supersonic diffuser, co-rotating vane arrays and counter-rotating vane arrays did not show improvement. In the subsonic diffuser, co-rotating vane arrays with negative angles-of-incidence and counter-rotating vane arrays were exceptional in reducing radial distortion and improving total pressure recovery. Downstream co-rotating vanes demonstrated up to 41% improvement in radial distortion whereas downstream counter-rotating vanes demonstrated up to 73% improvement. For downstream counter-rotating vanes, a polynomial trend between VG height and radial distortion indicate that increasing VG height improves inlet distortion. In summary, downstream vanes are exceptional in improving total pressure recovery compared to upstream vanes.

Experimental analysis of the boundary layer transition with zero and positive pressure gradient

NASA Technical Reports Server (NTRS)

Arnal, D.; Jullen, J. C.; Michel, R.

1980-01-01

The influence of a positive pressure gradient on the boundary layer transition is studied. The mean velocity and turbulence profiles of four cases are examined. As the intensity of the pressure gradient is increased, the Reynolds number of the transition onset and the length of the transition region are reduced. The Tollmein-Schlichting waves disturb the laminar regime; the amplification of these waves is in good agreement with the stability theory. The three dimensional deformation of the waves leads finally to the appearance of turbulence. In the case of zero pressure gradient, the properties of the turbulent spots are studied by conditional sampling of the hot-wire signal; in the case of positive pressure gradient, the turbulence appears in a progressive manner and the turbulent spots are much more difficult to characterize.

Predictions of Separated and Transitional Boundary Layers Under Low-Pressure Turbine Airfoil Conditions Using an Intermittency Transport Equation

NASA Technical Reports Server (NTRS)

Suzen, Y. Bora; Huang, P. G.; Hultgren, Lennart S.; Ashpis, David E.

2001-01-01

A new transport equation for the intermittency factor was proposed to predict separated and transitional boundary layers under low-pressure turbine airfoil conditions. The intermittent behavior of the transitional flows is taken into account and incorporated into computations by modifying the eddy viscosity, mu(sub t), with the intermittency factor, gamma. Turbulent quantities are predicted by using Menter's two-equation turbulence model (SST). The intermittency factor is obtained from a transport equation model, which not only can reproduce the experimentally observed streamwise variation of the intermittency in the transition zone, but also can provide a realistic cross-stream variation of the intermittency profile. In this paper, the intermittency model is used to predict a recent separated and transitional boundary layer experiment under low pressure turbine airfoil conditions. The experiment provides detailed measurements of velocity, turbulent kinetic energy and intermittency profiles for a number of Reynolds numbers and freestream turbulent intensity conditions and is suitable for validation purposes. Detailed comparisons of computational results with experimental data are presented and good agreements between the experiments and predictions are obtained.

Predictions of Separated and Transitional Boundary Layers Under Low-Pressure Turbine Airfoil Conditions Using an Intermittency Transport Equation

NASA Technical Reports Server (NTRS)

Suzen, Y. B.; Huang, P. G.; Hultgren, Lennart S.; Ashpis, David E.

2003-01-01

A new transport equation for the intermittency factor was proposed to predict separated and transitional boundary layers under low-pressure turbine airfoil conditions. The intermittent behavior of the transitional flows is taken into account and incorporated into computations by modifying the eddy viscosity, t , with the intermittency factor, y. Turbulent quantities are predicted by using Menter s two-equation turbulence model (SST). The intermittency factor is obtained from a transport equation model, which not only can reproduce the experimentally observed streamwise variation of the intermittency in the transition zone, but also can provide a realistic cross-stream variation of the intermittency profile. In this paper, the intermittency model is used to predict a recent separated and transitional boundary layer experiment under low pressure turbine airfoil conditions. The experiment provides detailed measurements of velocity, turbulent kinetic energy and intermittency profiles for a number of Reynolds numbers and freestream turbulent intensity conditions and is suitable for validation purposes. Detailed comparisons of computational results with experimental data are presented and good agreements between the experiments and predictions are obtained.

Theoretical Calculation of Viscous-Inviscid Transonic Flows.

DTIC Science & Technology

1980-08-01

Taylor Naval Ship Research and Development Center Aviation and Surface Effects Department / (See reverse side) Bethesda, Maryland 20084 ! CONTROLLING...Interactions ... .......... ... 46 18 -ffect of Boundary Layer on Blade Surface Pressures in a Transonic Fan Rotor Tip Section Cascade...complicated by the viscous effect . The strong viscous-inviscid interaction caused by the shock wave thickens the boundary layer rapidly, and the flow eventually

An experimental investigation of shock wave-turbulent boundary layer interactions with and without boundary layer suction: A data summary report

NASA Technical Reports Server (NTRS)

Sun, C. C.; Childs, M. E.

1977-01-01

Tabulated data from a series of experimental studies of the interaction of a shock wave with a turbulent boundary layer in axisymmetric flow configurations is presented. The studies were conducted at the walls of circular wind tunnels and on the cylindrical centerbody of an annular flow channel. Detailed pitot pressure profiles and wall static pressure profiles upstream of, within and downstream of the interaction region are given. Results are presented for flows at nominal freestream Mach Numbers of 2, 3 and 4. For studies at the tunnel sidewalls, the shock waves were produced by conical shock generators mounted on the centerline of the wind tunnel at zero angle of attack. The annular ring generator was used to produce the shock wave at the centerbody of the annular flow channel. The effects of boundary layer bleed were examined in the investigation. Both bleed rate and bleed location were studied. Most of the bleed studies were conducted with bleed holes drilled normal to the wall surface but the effects of slot suction were also examined. A summary of the principal results and conclusions is given.

Enhanced viscous flow drag reduction using acoustic excitation

NASA Technical Reports Server (NTRS)

Nagel, R. T.

1988-01-01

Large eddy break up devices (LEBUs) constitute a promising method of obtaining drag reduction in a turbulent boundary layer. Enhancement of the LEBU effectiveness by exciting its trailing edge with acoustic waves phase locked to the large scale structure influencing the momentum transfer to the wall is sought. An initial estimate of the required sound pressure level for an effective pulse was obtained by considering the magnitude of the pressure perturbations at the near wake of a thin plate in inviscid flow. Detailed skin friction measurments were obtained in the flow region downstream of a LEBU excited with acoustic waves. The data are compared with skin friction measurements of a simply manipulated flow, without acoustic excitation and with a plain flow configuration. The properties and the scales of motion in the flow regime downstream of the acoustically excited LEBU are studied. A parametric study based upon the characteristics of the acoustic input was pursued in addition to the careful mapping of the drag reduction phenomenon within the acoustically manipulated boundary layer. This study of boundary layer manipulation has lead to improved skin friction drag reduction and further understanding of the turbulent boundary layer.

Vehicle cabin cooling system for capturing and exhausting heated boundary layer air from inner surfaces of solar heated windows

DOEpatents

Farrington, Robert B.; Anderson, Ren

2001-01-01

The cabin cooling system includes a cooling duct positioned proximate and above upper edges of one or more windows of a vehicle to exhaust hot air as the air is heated by inner surfaces of the windows and forms thin boundary layers of heated air adjacent the heated windows. The cabin cooling system includes at least one fan to draw the hot air into the cooling duct at a flow rate that captures the hot air in the boundary layer without capturing a significant portion of the cooler cabin interior air and to discharge the hot air at a point outside the vehicle cabin, such as the vehicle trunk. In a preferred embodiment, the cooling duct has a cross-sectional area that gradually increases from a distal point to a proximal point to the fan inlet to develop a substantially uniform pressure drop along the length of the cooling duct. Correspondingly, this cross-sectional configuration develops a uniform suction pressure and uniform flow rate at the upper edge of the window to capture the hot air in the boundary layer adjacent each window.

On the dynamics of the flow in the vicinity of micro-scale coatings composed by organized elements

NASA Astrophysics Data System (ADS)

Doosttalab, Ali; Bocanegra Evans, Humberto; Gorumlu, Serdar; Aksak, Burak; Chamorro, Leonardo P.; Castillo, Luciano

2017-11-01

A set of high-resolution PIV experiments were carried out in a refractive index-matched facility under zero pressure gradient turbulent boundary layer to investigate the flow dynamics around two customized coatings composed of uniformly distributed fibers of different geometry. The two type of fibers shared a cylindrical shape and height y+ < 1 however, one of those had diverging tip similar to that of a shark skin. Results evidence an inter-layer acting between the viscous-dominated flow within the pillars canopy (where Re 1) and the inertia dominated flow in the boundary layer. Using averaged 2D N-S equations, it is possible to show that the inter-layer wall shear stress is τoiw+ = [∂U+/∂y+ - < uv+ > ] - [Pw+h+(y+/h+ - 1) + ] , with first term in the RHS representing the wall shear stress and the second term indicating the inter-layer form drag. A wall-normal Reynolds stress exist which depends on the pressure difference across the boundary layer and at the wall, = +( - < P >) / ρ . This reveals a basic mechanism where the flow is modulated by unsteady blowing and suction at the interface.

An experimental study of the effects of bodyside compression on forward swept sidewall compression inlets ingesting a turbulent boundary layer

NASA Technical Reports Server (NTRS)

Rodi, Patrick E.

1993-01-01

Forward swept sidewall compression inlets have been tested in the Mach 4 Blowdown Facility at the NASA Langley Research Center to study the effects of bodyside compression surfaces on inlet performance in the presence of an incoming turbulent boundary layer. The measurements include mass flow capture and mean surface pressure distributions obtained during simulated combustion pressure increases downstream of the inlet. The kerosene-lampblack surface tracer technique has been used to obtain patterns of the local wall shear stress direction. Inlet performance is evaluated using starting and unstarting characteristics, mass capture, mean surface pressure distributions and permissible back pressure limits. The results indicate that inlet performance can be improved with selected bodyside compression surfaces placed between the inlet sidewalls.

Performance of a Blunt-lip Side Inlet with Ramp Bleed, Bypass, and a Long Constant-area Duct Ahead of the Engine : Mach Numbers 0.66 and 1.5 to 2.1

NASA Technical Reports Server (NTRS)

Allen, John L

1956-01-01

Unsteady shock-induced separation of the ramp boundary layer was reduced and stabilized more effectively by external perforations than by external or internal slots. At Mach 2.0 peak total-pressure recovery was increased from 0.802 to 0.89 and stable mass-flow range was increased 185 percent over that for the solid ramp. Peak pressure recovery occurred just before instability. The 7 and one-third-diameter duct ahead of the engine reduced large total-pressure distortions but was not as successful for small distortions as obtained with throat bleed. By removing boundary-layer air the bypass nearly recovered the total-pressure loss due to the long duct.

Modelling Unsteady Wall Pressures Beneath Turbulent Boundary Layers

NASA Technical Reports Server (NTRS)

Ahn, B-K.; Graham, W. R.; Rizzi, S. A.

2004-01-01

As a structural entity of turbulence, hairpin vortices are believed to play a major role in developing and sustaining the turbulence process in the near wall region of turbulent boundary layers and may be regarded as the simplest conceptual model that can account for the essential features of the wall pressure fluctuations. In this work we focus on fully developed typical hairpin vortices and estimate the associated surface pressure distributions and their corresponding spectra. On the basis of the attached eddy model, we develop a representation of the overall surface pressure spectra in terms of the eddy size distribution. Instantaneous wavenumber spectra and spatial correlations are readily derivable from this representation. The model is validated by comparison of predicted wavenumber spectra and cross-correlations with existing emperical models and experimental data.

Receptivity of the Boundary Layer to Vibrations of the Wing Surface

NASA Astrophysics Data System (ADS)

Bernots, Tomass; Ruban, Anatoly; Pryce, David; Laminar Flow Control UK Group Team

2014-11-01

In this work we study generation of Tollmien-Schlichting (T-S) waves in the boundary layer due to elastic vibrations of the wing surface. The flow is investigated based on the asymptotic analysis of the Navier-Stokes equations at large values of the Reynolds number. It is assumed that in the spectrum of the wing vibrations there is a harmonic which comes in resonance with the T-S wave on the lower branch of the stability curve. It was found that the vibrations of the wing surface produce pressure perturbations in the flow outside the boundary layer which can be calculated with the help of the piston theory. As the pressure perturbations penetrate into the boundary layer, a Stokes layer forms on the wing surface which appears to be influenced significantly by the compressibility of the flow, and is incapable of producing the T-S waves. The situation changes when the Stokes layer encounters an roughness; near which the flow is described using the triple-deck theory. The solution of the triple-deck problem can be found in an analytic form. Our main concern is with the flow behaviour downstream of the roughness and, in particular, with the amplitude of the generated Tollmien-Schlichting waves. This research was performed in the Laminar Flow Control Centre (LFC-UK) at Imperial College London. The centre is supported by EPSRC, Airbus UK and EADS Innovation Works.

Measurements in Transitional Boundary Layers Under High Free-Stream Turbulence and Strong Acceleration Conditions

NASA Technical Reports Server (NTRS)

Volino, Ralph J.; Simon, Terrence W.

1995-01-01

Measurements from transitional, heated boundary layers along a concave-curved test wall are presented and discussed. A boundary layer subject to low free-stream turbulence intensity (FSTI), which contains stationary streamwise (Gortler) vortices, is documented. The low FSTI measurements are followed by measurements in boundary layers subject to high (initially 8%) free-stream turbulence intensity and moderate to strong streamwise acceleration. Conditions were chosen to simulate those present on the downstream half of the pressure side of a gas turbine airfoil. Mean flow characteristics as well as turbulence statistics, including the turbulent shear stress, turbulent heat flux, and turbulent Prandtl number, are documented. A technique called "octant analysis" is introduced and applied to several cases from the literature as well as to data from the present study. Spectral analysis was applied to describe the effects of turbulence scales of different sizes during transition. To the authors'knowledge, this is the first detailed documentation of boundary layer transition under such high free-stream turbulence conditions.

Control of unsteadiness of a shock wave/turbulent boundary layer interaction by using a pulsed-plasma-jet actuator

NASA Astrophysics Data System (ADS)

Narayanaswamy, Venkateswaran; Raja, Laxminarayan L.; Clemens, Noel T.

2012-07-01

A pulsed-plasma jet actuator is used to control the unsteady motion of the separation shock of a shock wave/boundary layer interaction formed by a compression ramp in a Mach 3 flow. The actuator is based on a plasma-generated synthetic jet and is configured as an array of three jets that can be injected normal to the cross-flow, pitched, or pitched and skewed. The typical peak jet exit velocity of the actuators is about 300 m/s and the pulsing frequencies are a few kilohertz. A study of the interaction between the pulsed-plasma jets and the shock/boundary layer interaction was performed in a time-resolved manner using 10 kHz schlieren imaging. When the actuator, pulsed at StL ≈ 0.04 (f = 2 kHz), was injected into the upstream boundary layer, the separation shock responded to the plasma jet by executing a rapid upstream motion followed by a gradual downstream recovery motion. Schlieren movies of the interaction showed that the separation shock unsteadiness was locked to the pulsing frequency of the actuator, with amplitude of about one boundary layer thickness. Wall-pressure measurements made under the intermittent region showed about a 30% decrease in the overall magnitude of the pressure fluctuations in the low-frequency band associated with unsteady large-scale motion of the separated flow. Furthermore, by increasing the pulsing frequency to 3.3 kHz, the amplitude of the separation shock oscillation was reduced to less than half the boundary layer thickness. Investigation into the effect of the actuator location on the shock wave/boundary layer interaction (SWBLI) showed qualitatively and quantitatively that the actuator placed upstream of the separation shock caused significant modification to the SWBLI unsteadiness, whereas injection from inside the separation bubble did not cause a noticeable effect.

Wall boundary layer development near the tip region of an IGV of an axial flow compressor

NASA Technical Reports Server (NTRS)

Lakshminarayana, B.; Sitaram, N.

1983-01-01

The annulus wall boundary layer inside the blade passage of the inlet guide vane (IGV) passage of a low-speed axial compressor stage was measured with a miniature five-hole probe. The three-dimensional velocity and pressure fields were measured at various axial and tangential locations. Limiting streamline angles and static pressures were also measured on the casing of the IGV passage. Strong secondary vorticity was developed. The data were analyzed and correlated with the existing velocity profile correlations. The end wall losses were also derived from these data.

Experimental Investigation of Three-Dimensional Shock Wave Turbulent Boundary Layer Interaction: An Exploratory Study of Blunt Fin-Induced Flows.

DTIC Science & Technology

1980-03-01

distributions could be obtained. The pressure tappings were sampled using two computer controlled 48 port Model 48J4 Scanivalves equipped with Druck ...the boundary layer becomes turbulent, the upstream in- fluence drops to between 2 and 3D . 3.2 Pressure Distributions Off the Plane of Symmetry 3.2.1...upstream influence varies between 0.3 cm (0.12") and 7.6 cm (3.0"), a ratio of about 25, yet in terms of D , Iu lies between 2 and 3D . The figure shows

Experiments with an Airfoil Model on which the Boundary Layers are Controlled Without the Use of Supplementary Equipment

NASA Technical Reports Server (NTRS)

Abbott, I H

1931-01-01

This report describes test made in the Variable Density Wind Tunnel of the NACA to determine the possibility of controlling the boundary layer on the upper surface of an airfoil by use of the low pressure existing near the leading edge. The low pressure was used to induce flow through slots in the upper surface of the wing. The tests showed that the angle of attack for maximum lift was increased at the expense of a reduction in the maximum lift coefficient and an increase in the drag coefficient.

Properties of the Equatorial Magnetotail Flanks ˜50-200 RE Downtail

NASA Astrophysics Data System (ADS)

Artemyev, A. V.; Angelopoulos, V.; Runov, A.; Wang, C.-P.; Zelenyi, L. M.

2017-12-01

In space, thin boundaries separating plasmas with different properties serve as a free energy source for various plasma instabilities and determine the global dynamics of large-scale systems. In planetary magnetopauses and shock waves, classical examples of such boundaries, the magnetic field makes a significant contribution to the pressure balance and plasma dynamics. The configuration and properties of such boundaries have been well investigated and modeled. However, much less is known about boundaries that form between demagnetized plasmas where the magnetic field is not important for pressure balance. The most accessible example of such a plasma boundary is the equatorial boundary layer of the Earth's distant magnetotail. Rather, limited measurements since its first encounter in the late 1970s by the International Sun-Earth Explorer-3 spacecraft revealed the basic properties of this boundary, but its statistical properties and structure have not been studied to date. In this study, we use Geotail and Acceleration, Reconnection, Turbulence and Electrodynamics of the Moon's Interaction with the Sun (ARTEMIS) missions to investigate the equatorial boundary layer from lunar orbit (˜55 Earth radii, RE, downtail) to as far downtail as ˜200 RE. Although the magnetic field has almost no effect on the structure of the boundary layer, the layer separates well the hot, rarefied plasma sheet from dense cold magnetosheath plasmas. We suggest that the most important role in plasma separation is played by polarization electric fields, which modify the efficiency of magnetosheath ion penetration into the plasma sheet. We also show that the total energies (bulk flow plus thermal) of plasma sheet ions and magnetosheath ions are very similar; that is, magnetosheath ion thermalization (e.g., via ion scattering by magnetic field fluctuations) is sufficient to produce hot plasma sheet ions without any additional acceleration.

An experimental investigation of the effect of boundary layer refraction on the noise from a high-speed propeller

NASA Technical Reports Server (NTRS)

Dittmar, J. H.; Burns, R. J.; Leciejewski, D. J.

1984-01-01

Models of supersonic propellers were previously tested for acoustics in the Lewis 8- by 6-Foot Wind Tunnel using pressure transducers mounted in the tunnel ceiling. The boundary layer on the tunnel ceiling is believed to refract some of the propeller noise away from the measurement transducers. Measurements were made on a plate installed in the wind tunnel which had a thinner boundary layer than the ceiling boundary layer. The plate was installed in two locations for comparison with tunnel ceiling noise data and with fuselage data taken on the NASA Dryden Jetstar airplane. Analysis of the data indicates that the refraction increases with: increasing boundary layer thickness; increasing free stream Mach number; increasing frequency; and decreasing sound radiation angle (toward the inlet axis). At aft radiation angles greater than about 100 deg there was little or no refraction. Comparisons with the airplane data indicated that not only is the boundary layer thickness important but also the shape of the velocity profile. Comparisons with an existing two-dimensional theory, using an idealized shear layer to approximate the boundary layer, showed that the theory and data had the same trends. Analysis of the data taken in the tunnel at two different distances from the propeller indicates a decay with distance in the wind tunnel at high Mach numbers but the decay at low Mach numbers is not as clear.

Flowfield dynamics in blunt fin-induced shock wave/turbulent boundary layer interactions

NASA Technical Reports Server (NTRS)

Dolling, David S.; Brusniak, Leon

1994-01-01

Fluctuating wall pressure measurements have been made on centerline upstream of a blunt fin in a Mach 5 turbulent boundary layer. By examining the ensemble averaged wall pressure distributions for different separation shock foot positions, it has been shown that local fluctuating wall pressure measurements are due to a distinct pressure distribution, Rho(sub i), which undergoes a stretching and flattening effect as its upstream boundary translates aperiodically between the upstream influence and separation lines. The locations of the maxima and minima in the wall pressure standard deviation can be accurately predicted using this distribution, providing quantitative confirmation of the model. This model also explains the observed cross-correlations and ensemble average measurements within the interaction. Using the Rho(sub i) model, wall pressure signals from under the separated flow region were used to reproduce the position-time history of the separation shock foot. Further, the negative time delay peak in the cross-correlation between the predicted and actual shock foot histories suggests that the separated region fluctuations precede shock foot motion. The unsteady behavior of the primary horseshoe vortex and its relation to the unsteady separation shock are described.

Comparison of predicted and measured low-speed performance of two 51 centimeter-diameter inlets at incidence angle

NASA Technical Reports Server (NTRS)

Albers, J. A.

1973-01-01

Theoretical and experimental internal flow characteristics of two 51-cm-diameter inlets are compared. Theoretical flow characteristics along the inlet surface were obtained from an axisymmetric potential flow and boundary layer analysis. The experimental data were obtained from low-speed tests of a high-bypass-ratio turbofan engine simulator. Comparisons between calculated internal surface pressure distributions and experimental data are presented for a free-system velocity of 45 m/sec and for incidence angles from 0 deg to 50 deg. Analysis of boundary layer separation on the inlet lip at incidence angle is the major emphasis of this report. Theoretical boundary layer shape factors, skin friction coefficients, and velocity profiles in the boundary layer are presented, along with the location of the transition region. Theoretical and experimental separation locations are also discussed.

Numerical modeling of the transitional boundary layer over a flat plate

NASA Astrophysics Data System (ADS)

Ivanov, Dimitry; Chorny, Andrei

2015-11-01

Our example is connected with fundamental research on understanding how an initially laminar boundary layer becomes turbulent. We have chosen the flow over a flat plate as a prototype for boundary-layer flows around bodies. Special attention was paid to the near-wall region in order to capture all levels of the boundary layer. In this study, the numerical software package OpenFOAM has been used in order to solve the flow field. The results were used in a comparative study with data obtained from Large Eddy Simulation (LES). The composite SGS-wall model is presently incorporated into a computer code suitable for the LES of developing flat-plate boundary layers. Presently this model is extended to the LES of the zero-pressure gradient, flat-plate turbulent boundary layer. In current study the time discretization is based on a second order Crank-Nicolson/Adams-Bashforth method. LES solver using Smagorinsky and the one-equation LES turbulence models. The transition models significantly improve the prediction of the onset location compared to the fully turbulent models.LES methods appear to be the most promising new tool for the design and analysis of flow devices including transition regions of the turbulent flow.

DOE Office of Scientific and Technical Information (OSTI.GOV)

Newsom, R. K.; Sivaraman, C.; Shippert, T. R.

Wind speed and direction, together with pressure, temperature, and relative humidity, are the most fundamental atmospheric state parameters. Accurate measurement of these parameters is crucial for numerical weather prediction. Vertically resolved wind measurements in the atmospheric boundary layer are particularly important for modeling pollutant and aerosol transport. Raw data from a scanning coherent Doppler lidar system can be processed to generate accurate height-resolved measurements of wind speed and direction in the atmospheric boundary layer.

Simulation of the planetary boundary layer with the UCLA general circulation model

NASA Technical Reports Server (NTRS)

Suarez, M. J.; Arakawa, A.; Randall, D. A.

1981-01-01

A planetary boundary layer (PBL) model is presented which employs a mixed layer entrainment formulation to describe the mass exchange between the mixed layer with the upper, laminar atmosphere. A modified coordinate system couples the mixed layer model with large scale and sub-grid scale processes of a general circulation model. The vertical coordinate is configured as a sigma coordinate with the lower boundary, the top of the PBL, and the prescribed pressure level near the tropopause expressed as coordinate surfaces. The entrainment mass flux is parameterized by assuming the dissipation rate of turbulent kinetic energy to be proportional to the positive part of the generation by convection or mechanical production. The results of a simulation of July are presented for the entire globe.

Flow field predictions for a slab delta wing at incidence

NASA Technical Reports Server (NTRS)

Conti, R. J.; Thomas, P. D.; Chou, Y. S.

1972-01-01

Theoretical results are presented for the structure of the hypersonic flow field of a blunt slab delta wing at moderately high angle of attack. Special attention is devoted to the interaction between the boundary layer and the inviscid entropy layer. The results are compared with experimental data. The three-dimensional inviscid flow is computed numerically by a marching finite difference method. Attention is concentrated on the windward side of the delta wing, where detailed comparisons are made with the data for shock shape and surface pressure distributions. Surface streamlines are generated, and used in the boundary layer analysis. The three-dimensional laminar boundary layer is computed numerically using a specially-developed technique based on small cross-flow in streamline coordinates. In the rear sections of the wing the boundary layer decreases drastically in the spanwise direction, so that it is still submerged in the entropy layer at the centerline, but surpasses it near the leading edge. Predicted heat transfer distributions are compared with experimental data.

Pressure estimation from single-snapshot tomographic PIV in a turbulent boundary layer

NASA Astrophysics Data System (ADS)

Schneiders, Jan F. G.; Pröbsting, Stefan; Dwight, Richard P.; van Oudheusden, Bas W.; Scarano, Fulvio

2016-04-01

A method is proposed to determine the instantaneous pressure field from a single tomographic PIV velocity snapshot and is applied to a flat-plate turbulent boundary layer. The main concept behind the single-snapshot pressure evaluation method is to approximate the flow acceleration using the vorticity transport equation. The vorticity field calculated from the measured instantaneous velocity is advanced over a single integration time step using the vortex-in-cell (VIC) technique to update the vorticity field, after which the temporal derivative and material derivative of velocity are evaluated. The pressure in the measurement volume is subsequently evaluated by solving a Poisson equation. The procedure is validated considering data from a turbulent boundary layer experiment, obtained with time-resolved tomographic PIV at 10 kHz, where an independent surface pressure fluctuation measurement is made by a microphone. The cross-correlation coefficient of the surface pressure fluctuations calculated by the single-snapshot pressure method with respect to the microphone measurements is calculated and compared to that obtained using time-resolved pressure-from-PIV, which is regarded as benchmark. The single-snapshot procedure returns a cross-correlation comparable to the best result obtained by time-resolved PIV, which uses a nine-point time kernel. When the kernel of the time-resolved approach is reduced to three measurements, the single-snapshot method yields approximately 30 % higher correlation. Use of the method should be cautioned when the contributions to fluctuating pressure from outside the measurement volume are significant. The study illustrates the potential for simplifying the hardware configurations (e.g. high-speed PIV or dual PIV) required to determine instantaneous pressure from tomographic PIV.

Active Brownian particles near straight or curved walls: Pressure and boundary layers

NASA Astrophysics Data System (ADS)

Duzgun, Ayhan; Selinger, Jonathan V.

2018-03-01

Unlike equilibrium systems, active matter is not governed by the conventional laws of thermodynamics. Through a series of analytic calculations and Langevin dynamics simulations, we explore how systems cross over from equilibrium to active behavior as the activity is increased. In particular, we calculate the profiles of density and orientational order near straight or circular walls and show the characteristic width of the boundary layers. We find a simple relationship between the enhancements of density and pressure near a wall. Based on these results, we determine how the pressure depends on wall curvature and hence make approximate analytic predictions for the motion of curved tracers, as well as the rectification of active particles around small openings in confined geometries.

Analytical and experimental evaluation of a 3-D hypersonic fixed-geometry, swept, mixed compression inlet

NASA Technical Reports Server (NTRS)

Agnone, Anthony M.

1987-01-01

The performance of a fixed-geometry, swept, mixed compression hypersonic inlet is presented. The experimental evaluation was conducted for a Mach number of 6.0 and for several angles of attack. The measured surface pressures and pitot pressure surveys at the inlet throat are compared to computations using a three-dimensional Euler code and an integral boundary layer theory. Unique features of the intake design, including the boundary layer control, insure a high inlet performance. The experimental data show the inlet has a high mass averaged total pressure recovery, a high mass capture and nearly uniform flow diffusion. The swept inlet exhibits excellent starting characteristics, and high flow stability at angle of attack.

Influence of a heated leading edge on boundary layer growth, stability, and transition

DOE Office of Scientific and Technical Information (OSTI.GOV)

Landrum, D.B.; Macha, J.M.

1987-06-01

This paper presents the results of a combined theoretical and experimental study of the influence of a heated leading edge on the growth, stability, and transition of a two-dimensional boundary layer. The findings are directly applicable to aircraft wings and nacelles that use surface heating for anti-icing protection. The potential effects of the non-adiabatic condition are particularly important for laminar-flow sections where even small perturbations can result in significantly degraded aerodynamic performance. The results of the study give new insight to the fundamental coupling between streamwise pressure gradient and surface heat flux in laminar and transitional boundary layers. 13 references.

Influence of a heated leading edge on boundary layer growth, stability, and transition

DOE Office of Scientific and Technical Information (OSTI.GOV)

Landrum, D.B.; Macha, J.M.

1987-01-01

This paper presents the results of a combined theoretical and experimental study of the influence of a heated leading edge on the growth, stability, and transition of a two-dimensional boundary layer. The findings are directly applicable to aircraft wings and nacelles that use surface heating for anti-icing protection. The potential effects of the non-adiabatic condition are particularly important for laminar-flow sections where even small perturbations can result in significantly degraded aerodynamic performance. The results of the study give new insight to the fundamental coupling between streamwise pressure gradient and surface heat flux in laminar and transitional boundary layers.

Effects of Hybrid Flow Control on a Normal Shock Boundary-Layer Interaction

NASA Technical Reports Server (NTRS)

Hirt, Stefanie M.; Vyas, Manan A.

2013-01-01

Hybrid flow control, a combination of micro-ramps and steady micro-jets, was experimentally investigated in the 15x15 cm Supersonic Wind Tunnel at the NASA Glenn Research Center. A central composite design of experiments method, was used to develop response surfaces for boundary-layer thickness and reversed-flow thickness, with factor variables of inter-ramp spacing, ramp height and chord length, and flow injection ratio. Boundary-layer measurements and wall static pressure data were used to understand flow separation characteristics. A limited number of profiles were measured in the corners of the tunnel to aid in understanding the three-dimensional characteristics of the flowfield.

Boundary layer separation on isolated boattail nozzles. M.S. Thesis; [conducted in the Langley 16-foot transonic wind tunnel

NASA Technical Reports Server (NTRS)

Abeyounis, W. K.

1977-01-01

The phenomenon of separated flow on a series of circular-arc afterbodies was investigated using the Langley 16-foot transonic tunnel at free-stream Mach numbers from 0.40 to 0.95 at 0 deg angle of attack. Both high-pressure air and solid circular cylinders with a diameter equal to the nozzle exit diameter were used to simulate jet exhausts. A detailed data base of boundary layer separation locations was obtained using oil-flow techniques. The results indicate that boundary layer separation is most extensive on steep boattails at high Mach numbers.

Pressure fluctuations beneath instability wavepackets and turbulent spots in a hypersonic boundary layer

DOE PAGES

Casper, Katya M.; Beresh, Steven J.; Schneider, Steven P.

2014-09-09

To investigate the pressure-fluctuation field beneath turbulent spots in a hypersonic boundary layer, a study was conducted on the nozzle wall of the Boeing/AFOSR Mach-6 Quiet Tunnel. Controlled disturbances were created by pulsed-glow perturbations based on the electrical breakdown of air. Under quiet-flow conditions, the nozzle-wall boundary layer remains laminar and grows very thick over the long nozzle length. This allows the development of large disturbances that can be well-resolved with high-frequency pressure transducers. A disturbance first grows into a second-mode instability wavepacket that is concentrated near its own centreline. Weaker disturbances are seen spreading from the centre. The wavesmore » grow and become nonlinear before breaking down to turbulence. The breakdown begins in the core of the packets where the wave amplitudes are largest. Second-mode waves are still evident in front of and behind the breakdown point and can be seen propagating in the spanwise direction. The turbulent core grows downstream, resulting in a spot with a classical arrowhead shape. Behind the spot, a low-pressure calmed region develops. However, the spot is not merely a localized patch of turbulence; instability waves remain an integral part. Limited measurements of naturally occurring disturbances show many similar characteristics. From the controlled disturbance measurements, the convection velocity, spanwise spreading angle, and typical pressure-fluctuation field were obtained.« less

Analytical method for predicting the pressure distribution about a nacelle at transonic speeds

NASA Technical Reports Server (NTRS)

Keith, J. S.; Ferguson, D. R.; Merkle, C. L.; Heck, P. H.; Lahti, D. J.

1973-01-01

The formulation and development of a computer analysis for the calculation of streamlines and pressure distributions around two-dimensional (planar and axisymmetric) isolated nacelles at transonic speeds are described. The computerized flow field analysis is designed to predict the transonic flow around long and short high-bypass-ratio fan duct nacelles with inlet flows and with exhaust flows having appropriate aerothermodynamic properties. The flow field boundaries are located as far upstream and downstream as necessary to obtain minimum disturbances at the boundary. The far-field lateral flow field boundary is analytically defined to exactly represent free-flight conditions or solid wind tunnel wall effects. The inviscid solution technique is based on a Streamtube Curvature Analysis. The computer program utilizes an automatic grid refinement procedure and solves the flow field equations with a matrix relaxation technique. The boundary layer displacement effects and the onset of turbulent separation are included, based on the compressible turbulent boundary layer solution method of Stratford and Beavers and on the turbulent separation prediction method of Stratford.

A comparison of disturbance levels measured in hypersonic tunnels using a hot-wire anemometer and a pitot pressure probe.

NASA Technical Reports Server (NTRS)

Stainback, P. C.; Wagner, R. D.

1972-01-01

Disturbance levels were measured in the test section of a Mach 5 blowdown jet using a constant-current, hot-wire anemometer and a pressure transducer. The disturbance levels, measured by the two instruments and normalized by local mean values, agreed within about 30%, with the pitot data higher than the hot-wire data. The rms disturbance levels measured with the hot-wire anemometer and converted to pitot pressures using a quasi-steady flow analysis, were about two-thirds the levels measured with the pitot probe. The variation of the normalized rms disturbance levels with stagnation pressure indicated that transition occurred in the boundary layer on the nozzle wall and influenced the outputs of the instruments located at the exit of the nozzle when the total pressure was about 35 N/sq cm. Below this pressure the disturbance levels decreased markedly. At higher pressures the disturbances were predominantly aerodynamic noise generated by the turbulent boundary layer on the nozzle wall.

Fluctuating pressures measured beneath a high-temperature, turbulent boundary layer on a flat plate at Mach number of 5

NASA Technical Reports Server (NTRS)

Parrott, Tony L.; Jones, Michael G.; Albertson, Cindy W.

1989-01-01

Fluctuating pressures were measured beneath a Mach 5, turbulent boundary layer on a flat plate with an array of piezoresistive sensors. The data were obtained with a digital signal acquisition system during a test run of 4 seconds. Data sampling rate was such that frequency analysis up to 62.5 kHz could be performed. To assess in situ frequency response of the sensors, a specially designed waveguide calibration system was employed to measure transfer functions of all sensors and related instrumentation. Pressure time histories were approximated well by a Gaussian prohibiting distribution. Pressure spectra were very repeatable over the array span of 76 mm. Total rms pressures ranged from 0.0017 to 0.0046 of the freestream dynamic pressure. Streamwise, space-time correlations exhibited expected decaying behavior of a turbulence generated pressure field. Average convection speed was 0.87 of freestream velocity. The trendless behavior with sensor separation indicated possible systematic errors.

Flow-field measurements in the windward surface shock layer of space shuttle orbiter configurations at Mach number 8. [wind tunnel tests of scale models

NASA Technical Reports Server (NTRS)

Martindale, W. R.; Carter, L. D.

1975-01-01

Pitot pressure and total-temperature measurements were made in the windward surface shock layer of two 0.0175-scale space shuttle orbiter models at simulated re-entry conditions. Corresponding surface static pressure measurements were also made. Flow properties at the edge of the model boundary layer were derived from these measurements and compared with values calculated using conventional methods.

New Models for Velocity/Pressure-Gradient Correlations in Turbulent Boundary Layers

NASA Astrophysics Data System (ADS)

Poroseva, Svetlana; Murman, Scott

2014-11-01

To improve the performance of Reynolds-Averaged Navier-Stokes (RANS) turbulence models, one has to improve the accuracy of models for three physical processes: turbulent diffusion, interaction of turbulent pressure and velocity fluctuation fields, and dissipative processes. The accuracy of modeling the turbulent diffusion depends on the order of a statistical closure chosen as a basis for a RANS model. When the Gram-Charlier series expansions for the velocity correlations are used to close the set of RANS equations, no assumption on Gaussian turbulence is invoked and no unknown model coefficients are introduced into the modeled equations. In such a way, this closure procedure reduces the modeling uncertainty of fourth-order RANS (FORANS) closures. Experimental and direct numerical simulation data confirmed the validity of using the Gram-Charlier series expansions in various flows including boundary layers. We will address modeling the velocity/pressure-gradient correlations. New linear models will be introduced for the second- and higher-order correlations applicable to two-dimensional incompressible wall-bounded flows. Results of models' validation with DNS data in a channel flow and in a zero-pressure gradient boundary layer over a flat plate will be demonstrated. A part of the material is based upon work supported by NASA under award NNX12AJ61A.

Planar Reflection of Gaseous Detonations

NASA Astrophysics Data System (ADS)

Damazo, Jason Scott

Pipes containing flammable gaseous mixtures may be subjected to internal detonation. When the detonation normally impinges on a closed end, a reflected shock wave is created to bring the flow back to rest. This study built on the work of Karnesky (2010) and examined deformation of thin-walled stainless steel tubes subjected to internal reflected gaseous detonations. A ripple pattern was observed in the tube wall for certain fill pressures, and a criterion was developed that predicted when the ripple pattern would form. A two-dimensional finite element analysis was performed using Johnson-Cook material properties; the pressure loading created by reflected gaseous detonations was accounted for with a previously developed pressure model. The residual plastic strain between experiments and computations was in good agreement. During the examination of detonation-driven deformation, discrepancies were discovered in our understanding of reflected gaseous detonation behavior. Previous models did not accurately describe the nature of the reflected shock wave, which motivated further experiments in a detonation tube with optical access. Pressure sensors and schlieren images were used to examine reflected shock behavior, and it was determined that the discrepancies were related to the reaction zone thickness extant behind the detonation front. During these experiments reflected shock bifurcation did not appear to occur, but the unfocused visualization system made certainty impossible. This prompted construction of a focused schlieren system that investigated possible shock wave-boundary layer interaction, and heat-flux gauges analyzed the boundary layer behind the detonation front. Using these data with an analytical boundary layer solution, it was determined that the strong thermal boundary layer present behind the detonation front inhibits the development of reflected shock wave bifurcation.

Unsteady flow through in-vitro models of the glottis

NASA Astrophysics Data System (ADS)

Hofmans, G. C. J.; Groot, G.; Ranucci, M.; Graziani, G.; Hirschberg, A.

2003-03-01

The unsteady two-dimensional flow through fixed rigid in vitro models of the glottis is studied in some detail to validate a more accurate model based on the prediction of boundary-layer separation. The study is restricted to the flow phenomena occurring within the glottis and does not include effects of vocal-fold movement on the flow. Pressure measurements have been carried out for a transient flow through a rigid scale model of the glottis. The rigid model with a fixed geometry driven by an unsteady pressure is used in order to achieve a high accuracy in the specification of the geometry of the glottis. The experimental study is focused on flow phenomena as they might occur in the glottis, such as the asymmetry of the flow due to the Coanda effect and the transition to turbulent flow. It was found that both effects need a relatively long time to establish themselves and are therefore unlikely to occur during the production of normal voiced speech when the glottis closes completely during part of the oscillation cycle. It is shown that when the flow is still laminar and symmetric the prediction of the boundary-layer model and the measurement of the pressure drop from the throat of the glottis to the exit of the glottis agree within 40%. Results of the boundary-layer model are compared with a two-dimensional vortex-blob method for viscous flow. The difference between the results of the simpiflied boundary-layer model and the experimental results is explained by an additional pressure difference between the separation point and the far field within the jet downstream of the separation point. The influence of the movement of the vocal folds on our conclusions is still unclear.

Assessment of fluctuating pressure gradient using acceleration spectra in near wall flows

NASA Astrophysics Data System (ADS)

Cadel, Daniel; Lowe, K. Todd

2015-11-01

Separation of contributions to the fluctuating acceleration from pressure gradient fluctuations and viscous shear fluctuations in the frequency domain is examined in a turbulent boundary layer. Past work leveraging turbulent accelerations for pressure gradient measurements has neglected the viscous shear term from the momentum equation--an invalid assumption in the case of near wall flows. The present study seeks to account for the influence of the viscous shear term and spectrally reject its contribution, which is thought to be concentrated at higher frequencies. Spectra of velocity and acceleration fluctuations in a flat plate, zero pressure gradient turbulent boundary layer at a momentum thickness Reynolds number of 7500 are measured using a spatially resolving three-component laser Doppler velocimeter. This canonical case data is applied for validation of the spectral approach for future application in more complex aerodynamic flows.

F-16XL ship #1 CAWAP flight - alpha 10 degrees, beta -5 degrees, altitude 10,000 feet

NASA Technical Reports Server (NTRS)

1996-01-01

The single-seat F-16XL (ship #1) makes another run during the Cranked-Arrow Wing Aerodynamic Project (CAWAP) at Dryden Flight Research Center, Edwards, California. The modified airplane features a delta 'cranked-arrow' wing with strips of tubing along the leading edge to the trailing edge to sense static on the wing and obtain pressure distribution data. The right wing receives data on pressure distribution and the left wing has three types of instrumentation - preston tubes to measure local skin friction, boundary layer rakes to measure boundary layer profiles (the layer where the air interacts with the surfaces of a moving aircraft), and hot films to determine boundary layer transition locations. This photo shows the aircraft gathering data at an altitude of 10,000 feet, with an angle of attack of 10 degrees and a sideslip angle of -5 degrees. The program also gathered aero data on two wing planforms for NASA's High Speed Research Program. The first flight of CAWAP occurred on November 21, 1995, and the test program ended in April 1996.

F-16XL ship #1 crew

NASA Technical Reports Server (NTRS)

1995-01-01

November 27, 1995 Photograph of the F-16XL Ship #1 Cranked-Arrow Wing Aerodynamic Project (CAWAP) Test Team; from left to right, Ron Wilcox; Operations Engineer, Art Cope; Aircraft Mechanic, Dave Fisher; Chief Project Engineer, Dick Denman; Aircraft Mechanic, Bob Garcia; A/C Crew Chief, Susan Ligon; Aircraft Mechanic, Rodger Tarango; Mobile Operations Facility (MOF) Staff, Jerry Cousins; Aircraft Mechanic, Bruce Gallmeyer; MOF Staff, and Mike Reardon; Aircraft Mechanic/Helper. The modified airplane features a delta 'cranked-arrow' wing with strips of tubing along the leading edge to the trailing edge to sense static on the wing and obtain pressure distribution data. The right wing receives data on pressure distribution and the left wing has three types of instrumentation - preston tubes to measure local skin friction, boundary layer rakes to measure boundary layer profiles (the layer where the air interacts with the surfaces of a moving aircraft), and hot films to determine boundary layer transition locations. The first flight of CAWAP occurred at NASA's Dryden Flight Research Center, Edwards, California, on November 21, 1995, and the test program ended in April 1996.

F-16XL ship #1 CAWAP flight - alpha 21 degrees, altitude 17,500 feet

NASA Technical Reports Server (NTRS)

1996-01-01

The single-seat F-16XL (ship #1) makes another run during the Cranked-Arrow Wing Aerodynamic Project (CAWAP) at Dryden Flight Research Center, Edwards, California. The modified airplane features a delta 'cranked-arrow' wing with strips of tubing along the leading edge to the trailing edge to sense static on the wing and obtain pressure distribution data. The right wing receives data on pressure distribution and the left wing (visible here) has three types of instrumentation - preston tubes to measure local skin friction, boundary layer rakes to measure boundary layer profiles (the layer where the air interacts with the surfaces of a moving aircraft), and hot films to determine boundary layer transition locations. This photo shows the aircraft gathering data at an altitude of 17,500 feet, with an angle of attack of 21 degrees The program also gathered aero data on two wing planforms for NASA's High Speed Research Program. The first flight of CAWAP occurred on November 21, 1995, and the test program ended in April 1996.

Surface pressure fluctuations in hypersonic turbulent boundary layers

NASA Technical Reports Server (NTRS)

Raman, K. R.

1974-01-01

The surface pressure fluctuations on a flat plate model at hypersonic Mach numbers of 5.2, 7.4 and 10.4 with an attached turbulent boundary layer were measured using flush mounted small piezoelectric sensors. A high frequency resolution of the pressure field was achieved using specially designed small piezoelectric sensors that had a good frequency response well above 300 KHz. The RMS pressures and non-dimensional energy spectra for all above Mach numbers are presented. The convective velocities, obtained from space time correlation considerations are equal to 0.7 U sub infinity. The results indicate the RMS pressures vary from 5 to 25 percent of the mean static pressures. The ratios of RMS pressure to dynamic pressure are less than the universally accepted subsonic value of 6 x 10/3. The ratio decreases in value as the Mach number or the dynamic pressure is increased. The ratio of RMS pressure to wall shear for Mach number 7.4 satisfies one smaller than or equal to p/tau sub w smaller than or equal to three.

Stability of the Boundary Layer and the Spot

NASA Technical Reports Server (NTRS)

Wygnanski, I.

2007-01-01

The similarity among turbulent spots observed in various transition experiments, and the rate in which they contaminate the surrounding laminar boundary layer is only cursory. The shape of the spot depends on the Reynolds number of the surrounding boundary layer and on the pressure gradient to which it and the surrounding laminar flow are exposed. The propagation speeds of the spot boundaries depend, in addition, on the location from which the spot originated and do not simply scale with the local free stream velocity. The understanding of the manner in which the turbulent manner in which the turbulent spot destabilizes the surrounding, vortical fluid is a key to the understanding of the transition process. We therefore turned to detailed observations near the spot boundaries in general and near the spanwise tip of the spot in particular.

Experimental Evaluation of an Isolated Synthetic Jet IN Crossflow

NASA Technical Reports Server (NTRS)

Schaeffler, Norman W.; Jenkins, Luther N.; Hepner, Timothy E.

2007-01-01

The second case for this workshop builds upon the isolated synthetic jet of Case 1 by adding a crossflow, with no streamwise pressure gradient, for the developing jet to interact with. Formally, Case 2 examines the interaction of a single, isolated, synthetic jet and a fully turbulent zero-pressure gradient boundary layer. The resulting flow has many of the characteristics that need to be modeled with fidelity if the results of the calculations are to serve as the basis for research and design with active flow control devices. These include the turbulence in the boundary layer, the time-evolution of the large vortical structure emanating from the jet orifice and its subsequent interaction with and distortion by the boundary layer turbulence, and the effect of the suction cycle on the boundary layer flow. In a synthetic jet, the flow through the orifice and out into the outer flowfield alternates between an exhaust and a suction cycle, driven by the contraction and expansion of a cavity internal to the actuator. In the present experiment, the volume changes in the internal cavity are accomplished by replacing one of the rigid walls of the cavity, the wall opposite the orifice exit, with a deformable wall. This flexible wall is driven by a bottom-mounted moveable piston. The piston is driven electro-mechanically. The synthetic jet issues into the external flow through a circular orifice. In the present experiment, this orifice has a diameter of 0.250 inches (6.35 mm). The flow is conceptually similar to that documented in Schaeffler [1]. To document the flow, several measurement techniques were utilized. The upstream boundary conditions (in-flow conditions), and several key phase-averaged velocity profiles were measured with a 3-component laser-Doppler velocimetry system. Phase-averaged velocity field measurements were made with both stereo digital particle image velocimetry and 2-D digital particle image velocimetry as the primary measurement system. Surface pressure measurements were made utilizing an electronically scanned pressure system.

An experimental study of the sources of fluctuating pressure loads beneath swept shock/boundary-layer interactions

NASA Technical Reports Server (NTRS)

Settles, G. S.; Garg, S.

1993-01-01

An experimental research program providing basic knowledge and establishing a database on the fluctuating pressure loads produced on aerodynamic surfaces beneath three dimensional shock wave/boundary layer interactions is described. Such loads constitute a fundamental problem of critical concern to future supersonic and hypersonic flight vehicles. A turbulent boundary layer on a flat plate is subjected to interactions with swept planar shock waves generated by sharp fins at angle of attack. Fin angles from 10 to 20 deg at freestream Mach numbers of 3 and 4 produce a variety of interaction strengths from weak to very strong. Miniature Kulite pressure transducers flush-mounted in the flat plate are used to measure interaction-induced wall pressure fluctuations. The distributions of properties of the pressure fluctuations, such as their ring levels, amplitude distributions, and power spectra, are also determined. Measurements were made for the first time in the aft regions of these interactions, revealing fluctuating pressure levels as high as 160 dB. These fluctuations are dominated by low frequency (0-5 kHz) signals. The maximum ring levels in the interactions show an increasing trend with increasing interaction strength. On the other hand, the maximum ring levels in the forward portion of the interactions decrease linearly with increasing interaction sweep back. These ring pressure distributions and spectra are correlated with the features of the interaction flowfield. The unsteadiness of the off-surface flowfield is studied using a new, non-intrusive technique based on the shadow graph method. The results indicate that the entire lambda-shock structure generated by the interaction undergoes relatively low-frequency oscillations. Some regions where particularly strong fluctuations are generated were identified. Fluctuating pressure measurements are also made along the line of symmetry of an axisymmetric jet impinging upon a flat plate at an angle. This flow was chosen as a simple analog to the impinging jet region found in the rear portion of the shock wave/boundary layer interactions under study. It is found that a sharp peak in ring pressure level exists at or near the mean stagnation point. It is suggested that the phenomena responsible for this peak may be active in the swept interactions as well, and may cause the extremely high fluctuating pressures observed in the impinging jet region in the present experimental program.

Improved two-equation k-omega turbulence models for aerodynamic flows

NASA Technical Reports Server (NTRS)

Menter, Florian R.

1992-01-01

Two new versions of the k-omega two-equation turbulence model will be presented. The new Baseline (BSL) model is designed to give results similar to those of the original k-omega model of Wilcox, but without its strong dependency on arbitrary freestream values. The BSL model is identical to the Wilcox model in the inner 50 percent of the boundary-layer but changes gradually to the high Reynolds number Jones-Launder k-epsilon model (in a k-omega formulation) towards the boundary-layer edge. The new model is also virtually identical to the Jones-Lauder model for free shear layers. The second version of the model is called Shear-Stress Transport (SST) model. It is based on the BSL model, but has the additional ability to account for the transport of the principal shear stress in adverse pressure gradient boundary-layers. The model is based on Bradshaw's assumption that the principal shear stress is proportional to the turbulent kinetic energy, which is introduced into the definition of the eddy-viscosity. Both models are tested for a large number of different flowfields. The results of the BSL model are similar to those of the original k-omega model, but without the undesirable freestream dependency. The predictions of the SST model are also independent of the freestream values and show excellent agreement with experimental data for adverse pressure gradient boundary-layer flows.

Effects of Sweep Angle on the Boundary-Layer Stability Characteristics of an Untapered Wing at Low Speeds

NASA Technical Reports Server (NTRS)

Boltz, Frederick W.; Kenyon, George C.; Allen, Clyde Q.

1960-01-01

An investigation was conducted in the Ames 12-Foot Low-Turbulence Pressure Tunnel to determine the effects of sweep on the boundary-layer stability characteristics of an untapered variable-sweep wing having an NACA 64(2)A015 section normal to the leading edge. Pressure distribution and transition were measured on the wing at low speeds at sweep angles of 0, 10, 20, 30, 40, and 50 deg. and at angles of attack from -3 to 3 deg. The investigation also included flow-visualization studies on the surface at sweep angles from 0 to 50 deg. and total pressure surveys in the boundary layer at a sweep angle of 30 deg. for angles of attack from -12 to 0 deg. It was found that sweep caused premature transition on the wing under certain conditions. This effect resulted from the formation of vortices in the boundary layer when a critical combination of sweep angle, pressure gradient, and stream Reynolds number was attained. A useful parameter in indicating the combined effect of these flow variables on vortex formation and on beginning transition is the crossflow Reynolds number. The critical values of crossflow Reynolds number for vortex formation found in this investigation range from about 135 to 190 and are in good agreement with those reported in previous investigations. The values of crossflow Reynolds number for beginning transitions were found to be between 190 and 260. For each condition (i.e., development of vortices and initiation of transition at a given location) the lower values in the specified ranges were obtained with a light coating of flow-visualization material on the surface. A method is presented for the rapid computation of crossflow Reynolds number on any swept surface for which the pressure distribution is known. From calculations based on this method, it was found that the maximum values of crossflow Reynolds number are attained under conditions of a strong pressure gradient and at a sweep angle of about 50 deg. Due to the primary dependence on pressure gradient, effects of sweep in causing premature transition are generally first encountered on the lower surfaces of wings operating at positive angles of attack.

Robust boundary treatment for open-channel flows in divergence-free incompressible SPH

NASA Astrophysics Data System (ADS)

Pahar, Gourabananda; Dhar, Anirban

2017-03-01

A robust Incompressible Smoothed Particle Hydrodynamics (ISPH) framework is developed to simulate specified inflow and outflow boundary conditions for open-channel flow. Being purely divergence-free, the framework offers smoothed and structured pressure distribution. An implicit treatment of Pressure Poison Equation and Dirichlet boundary condition is applied on free-surface to minimize error in velocity-divergence. Beyond inflow and outflow threshold, multiple layers of dummy particles are created according to specified boundary condition. Inflow boundary acts as a soluble wave-maker. Fluid particles beyond outflow threshold are removed and replaced with dummy particles with specified boundary velocity. The framework is validated against different cases of open channel flow with different boundary conditions. The model can efficiently capture flow evolution and vortex generation for random geometry and variable boundary conditions.

Experimental plasma studies

NASA Technical Reports Server (NTRS)

Dunn, M. G.

1972-01-01

The rate coefficients for the reactions C(+) + e(-) + e(-) yields C + e(-) and CO(+) + e(-) yields C + O were measured over the electron temperature range of approximately 1500 deg K to 7000 deg K. The measurements were performed in CO that had expanded from equilibrium reservoir conditions of 7060 deg K at 17.3 atm pressure and from 6260 deg K at 10.0 atm pressure. Two RAM flight probes were used to measure electron density and electron temperature in the expanding flow of a shock tunnel. Experiments were performed in the inviscid flow with both probes and in the nozzle-wall boundary layer with the constant bias-voltage probe. The distributions of electron density and electron temperature were independently measured using voltage-swept thin-wire probes. Thin-wire Langmuir probes were also used to measure the electron-density and electron-temperature distributions in the boundary layer of a sharp flat plate located on the nozzle centerline. Admittance measurements were performed with the RAM C and RAM C-C S-band antennas in the presence of an ionized boundary layer.

Aircraft Boundary-layer Measurements in the Gulf of Tehuantepec

NASA Astrophysics Data System (ADS)

Friehe, Carl; Melville, W. K.

2005-11-01

Airborne flux, meteorological, and wave measurements were made from the NSF/NCAR EC130Q aircraft in the Gulf of Tehuantepec under strong boundary-layer gap winds up to 25 m/sec at 33 m height. Statistics of flux estimates were obtained from multiple 33-m tracks flown under reasonably stationary and homogeneous conditions. Flux divergence was obtained from stack patterns flown at various distances from shore. Tracks flown at 33 m between the stacks provided the pressure gradient and advection terms in the momentum balance. Near shore, flux divergence was important and approximately balanced by the pressure gradient and advective terms; off-shore (400 km), divergence was small and again approximately in balance with the other two terms. Data from dropsondes and the Scanning Aerosol Backscatter LIDAR (SABL) revealed that the internal boundary layer initially thins off-shore as the gap wind field spreads horizontally, and then thickens due to turbulent mixing and possible hydraulic effects. Supported by NSF Division of Ocean Sciences.

Experimental Study of Unsteady Separation in a Laminar Boundary Layer

NASA Astrophysics Data System (ADS)

Bonacci, Andrew; Lang, Amy; Wahidi, Redha; Santos, Leo

2016-11-01

Separation, caused by an adverse pressure gradient, can be a major problem to aircraft. Reversing flow occurs in separated regions and an investigation of how this backflow forms is of interest due to the fact that this could be used as a means of initiating flow control. Specifically, backflow can bristle shark scales which may be linked to a passive, flow actuated separation control mechanism. An experiment was conducted in a water tunnel to replicate separation, with a focus on the reversing flow development near the wall within a laminar boundary layer. Using a rotating cylinder, an adverse pressure gradient was induced creating a separated region over a flat plate. In this experiment the boundary layer grows to sizes great enough that the scale of the flow is increased, making it more measurable to DPIV. In the future, this research can be utilized to better understand flow control mechanisms such as those enabled by shark skin. Funding from Army Research Office and NSF REU site Grant EEC 1358991 is greatly appreciated.

Modelling the transitional boundary layer

NASA Technical Reports Server (NTRS)

Narasimha, R.

1990-01-01

Recent developments in the modelling of the transition zone in the boundary layer are reviewed (the zone being defined as extending from the station where intermittency begins to depart from zero to that where it is nearly unity). The value of using a new non-dimensional spot formation rate parameter, and the importance of allowing for so-called subtransitions within the transition zone, are both stressed. Models do reasonably well in constant pressure 2-dimensional flows, but in the presence of strong pressure gradients further improvements are needed. The linear combination approach works surprisingly well in most cases, but would not be so successful in situations where a purely laminar boundary layer would separate but a transitional one would not. Intermittency-weighted eddy viscosity methods do not predict peak surface parameters well without the introduction of an overshooting transition function whose connection with the spot theory of transition is obscure. Suggestions are made for further work that now appears necessary for developing improved models of the transition zone.

Exploratory Calibration of Adjustable-Protrusion Surface-Obstacle (APSO) Skin Friction Vector Gage

NASA Technical Reports Server (NTRS)

Hakkinen, Raimo J.; Neubauer, Jeremy S.; Hamory, Philip J.; Bui, Trong T.; Noffz, Gregory K.; Young, Ron (Technical Monitor)

2003-01-01

The design of an adjustable-protrusion surface-obstacle (APSO) skin friction vector gage is presented. Results from exploratory calibrations conducted in laminar and turbulent boundary layers at the Washington University Low-Speed Wind Tunnel and for turbulent boundary layers at speeds up to Mach 2 on the ceiling of the NASA Glenn Research Center 8- X 6-ft Supersonic Wind Tunnel are also discussed. The adjustable-height gage was designed to yield both the magnitude and direction of the surface shear stress vector and to measure the local static pressure distribution. Results from the NASA test show good correlation for subsonic and low supersonic conditions covering several orders of magnitude in terms of the adopted similarity variables. Recommendations for future work in this area consist of identifying the physical parameters responsible for the disagreement between the university and NASA data sets, developing a compressibility correction specific to the APSO geometry, and examining the effect that static pressure distribution and skewed boundary layers have on the results from the APSO.

Transonic Shock-Wave/Boundary-Layer Interactions on an Oscillating Airfoil

NASA Technical Reports Server (NTRS)

Davis, Sanford S.; Malcolm, Gerald N.

1980-01-01

Unsteady aerodynamic loads were measured on an oscillating NACA 64A010 airfoil In the NASA Ames 11 by 11 ft Transonic Wind Tunnel. Data are presented to show the effect of the unsteady shock-wave/boundary-layer interaction on the fundamental frequency lift, moment, and pressure distributions. The data show that weak shock waves induce an unsteady pressure distribution that can be predicted quite well, while stronger shock waves cause complex frequency-dependent distributions due to flow separation. An experimental test of the principles of linearity and superposition showed that they hold for weak shock waves while flows with stronger shock waves cannot be superimposed.

Turbulence stress measurements in a nonadiabatic hypersonic boundary layer

NASA Technical Reports Server (NTRS)

Mikulla, V.; Horstman, C. C.

1975-01-01

Turbulent shear stress and direct turbulent total heat-flux measurements have been made across a nonadiabatic, zero pressure gradient, hypersonic boundary layer by using specially designed hot-wire probes free of strain-gauging and wire oscillation. Heat-flux measurements were in reasonably good agreement with values obtained by integrating the energy equation using measured profiles of velocity and temperature. The shear-stress values deduced from the measurements, by assuming zero correlation of velocity and pressure fluctuations, were lower than the values obtained by integrating the momentum equation. Statistical properties of the cross-correlations are similar to corresponding incompressible measurements at approximately the same momentum-thickness Reynolds number.

The vertical structure of the circulation and dynamics in Hudson Shelf Valley

USGS Publications Warehouse

Lentz, Steven J.; Butman, Bradford; Harris, Courtney K.

2014-01-01

Hudson Shelf Valley is a 20–30 m deep, 5–10 km wide v-shaped submarine valley that extends across the Middle Atlantic Bight continental shelf. The valley provides a conduit for cross-shelf exchange via along-valley currents of 0.5 m s−1 or more. Current profile, pressure, and density observations collected during the winter of 1999–2000 are used to examine the vertical structure and dynamics of the flow. Near-bottom along-valley currents having times scales of a few days are driven by cross-shelf pressure gradients setup by wind stresses, with eastward (westward) winds driving onshore (offshore) flow within the valley. The along-valley momentum balance in the bottom boundary layer is predominantly between the pressure gradient and bottom stress because the valley bathymetry limits current veering. Above the bottom boundary layer, the flow veers toward an along-shelf (cross-valley) orientation and a geostrophic balance with some contribution from the wind stress (surface Ekman layer). The vertical structure and strength of the along-valley current depends on the magnitude and direction of the wind stress. During offshore flows driven by westward winds, the near-bottom stratification within the valley increases resulting in a thinner bottom boundary layer and weaker offshore currents. Conversely, during onshore flows driven by eastward winds the near-bottom stratification decreases resulting in a thicker bottom boundary layer and stronger onshore currents. Consequently, for wind stress magnitudes exceeding 0.1 N m−2, onshore along-valley transport associated with eastward wind stress exceeds the offshore transport associated with westward wind stress of the same magnitude.

Detailed Study of Moderately Swept Impinging Oblique Shock/Boundary-Layer Interaction in Mach 2.3 Flow

NASA Astrophysics Data System (ADS)

Threadgill, James; Doerhmann, Adam; Little, Jesse

2017-11-01

A detailed experimental investigation of an impinging oblique Shock/Boundary Layer Interaction (SBLI) with 30° sweep in Mach 2.3 flow has been conducted. Despite its non-dimensional form, this canonical SBLI configuration has attracted little attention and remains poorly understood. Using a 12 .5° shock generator mounted in the freestream over a turbulent boundary layer, the interaction has been characterized with oil flow visualization, fast-response pressure transducers, and particle image velocimetry. Velocity vectors are used to extract the 3D interaction structure. These data are compared to wall pressure measurements and surface skin-friction streamlines. A local collapse of data normal to separation indicates a swept equivalence to Free Interaction Theory, albeit at a lower angle of sweep than imposed by the shock generator. Conditions at reattachment align with the imposed shock. Low-frequency shock motion near separation is observed, analogous to unswept SBLIs, with significant correlations that indicate spanwise traveling ripples in the shock foot. However, the magnitude of wall-pressure unsteadiness in this location is lower and shifted to higher frequencies than observed in equivalent unswept SBLI counterparts. Supported by the Air Force Office of Scientific Research (FA9550-15-1-0430).

In-flight Compressible Turbulent Boundary Layer Measurements on a Hollow Cylinder at a Mach Number of 3.0

NASA Technical Reports Server (NTRS)

Quinn, R. D.; Gong, L.

1978-01-01

Skin temperatures, shearing forces, surface static pressures, and boundary layer pitot pressures and total temperatures were measured on a hollow cylinder 3.04 meters long and 0.437 meter in diameter mounted beneath the fuselage of the YF-12A airplane. The data were obtained at a nominal free stream Mach number of 3.0 and at wall-to-recovery temperature ratios of 0.66 to 0.91. The free stream Reynolds number had a minimal value of 4.2 million per meter. Heat transfer coefficients and skin friction coefficients were derived from skin temperature time histories and shear force measurements, respectively. Boundary layer velocity profiles were derived from pitot pressure measurements, and a Reynolds analogy factor of 1.11 was obtained from the measured heat transfer and skin friction data. The skin friction coefficients predicted by the theory of van Driest were in excellent agreement with the measurements. Theoretical heat transfer coefficients, in the form of Stanton numbers calculated by using a modified Reynolds analogy between skin friction and heat transfer, were compared with measured values. The measured velocity profiles were compared to Coles' incompressible law-of-the-wall profile.

Advanced laser diagnostics for diamond deposition research

DOE Office of Scientific and Technical Information (OSTI.GOV)

Kruger, C.H.; Owano, T.G.; Wahl, E.H.

Chemical Vapor Deposition (CVD) using thermal plasmas is attractive for diamond synthesis applications due to the inherently high reactant densities and throughput, but the associated high gas-phase collision rates in the boundary layer above the substrate produce steep thermal and species gradients which can drive the complex plasma chemistry away from optimal conditions. To understand and control these environments, accurate measurements of temperature and species concentrations within the reacting boundary layer are needed. This is challenging in atmospheric pressure reactors due to the highly luminous environment, steep thermal and species gradients, and small spatial scales. The applicability of degenerate four-wavemore » mixing (DFWM) as a spectroscopic probe of atmospheric pressure reacting plasmas has been investigated. This powerful, nonlinear technique has been applied to the measurement of temperature and radical species concentrations in the boundary layer of a diamond growth substrate immersed in a flowing atmospheric pressure plasma. In-situ measurements of CH and C{sub 2} radicals have been performed to determine spatially resolved profiles of vibrational temperature, rotational temperature, and species concentration. Results of these measurements are compared with the predictions of a detailed numerical simulation.« less

Calculation of three dimensional viscous flows in annular cascades using parabolized Navier-Stokes equations

NASA Astrophysics Data System (ADS)

Lawerenz, M.

Numerical algorithms for describing the endwall boundary layers and secondary flows in high turning turbine cascades are described. Partially-parabolic methods which cover three-dimensional viscous flow effects are outlined. Introduction of tip-clearance models and modifications of no-slip conditions without the use of wall functions expand the range of application and improve accuracy. Simultaneous computation of the profile boundary layers by refinement of the mesh size in the circumferential direction makes it possible to describe the boundary layer interaction in the corners formed by the bladings and the endwalls. The partially-parabolic method means that the streamwise elliptic coupling is well represented by the given pressure field and that separation does not occur, but it is not possible to describe the separation of the endwall boundary layer near the leading edge and the horse-shoe vortex there properly.

A general method for calculating three-dimensional compressible laminar and turbulent boundary layers on arbitrary wings

NASA Technical Reports Server (NTRS)

Cebeci, T.; Kaups, K.; Ramsey, J. A.

1977-01-01

The method described utilizes a nonorthogonal coordinate system for boundary-layer calculations. It includes a geometry program that represents the wing analytically, and a velocity program that computes the external velocity components from a given experimental pressure distribution when the external velocity distribution is not computed theoretically. The boundary layer method is general, however, and can also be used for an external velocity distribution computed theoretically. Several test cases were computed by this method and the results were checked with other numerical calculations and with experiments when available. A typical computation time (CPU) on an IBM 370/165 computer for one surface of a wing which roughly consist of 30 spanwise stations and 25 streamwise stations, with 30 points across the boundary layer is less than 30 seconds for an incompressible flow and a little more for a compressible flow.

Flight-measured laminar boundary-layer transition phenomena including stability theory analysis

NASA Technical Reports Server (NTRS)

Obara, C. J.; Holmes, B. J.

1985-01-01

Flight experiments were conducted on a single-engine turboprop aircraft fitted with a 92-in-chord, 3-ft-span natural laminar flow glove at glove section lift coefficients from 0.15 to 1.10. The boundary-layer transition measurement methods used included sublimating chemicals and surface hot-film sensors. Transition occurred downstream of the minimum pressure point. Hot-film sensors provided a well-defined indication of laminar, laminar-separation, transitional, and turbulent boundary layers. Theoretical calculations of the boundary-layer parameters provided close agreement between the predicted laminar-separation point and the measured transition location. Tollmien-Schlichting (T-S) wave growth n-factors between 15 and 17 were calculated at the predicted point of laminar separation. These results suggest that for many practical airplane cruise conditions, laminar separation (as opposed to T-S instability) is the major cause of transition in predominantly two-dimensional flows.

Shock wave boundary layer interaction on suction side of compressor profile in single passage test section

NASA Astrophysics Data System (ADS)

Flaszynski, Pawel; Doerffer, Piotr; Szwaba, Ryszard; Kaczynski, Piotr; Piotrowicz, Michal

2015-11-01

The shock wave boundary layer interaction on the suction side of transonic compressor blade is one of the main objectives of TFAST project (Transition Location Effect on Shock Wave Boundary Layer Interaction). In order to investigate the flow structure on the suction side of a profile, a design of a generic test section in linear transonic wind tunnel was proposed. The experimental and numerical results for the flow structure investigations are shown for the flow conditions as the existing ones on the suction side of the compressor profile. Near the sidewalls the suction slots are applied for the corner flow structure control. It allows to control the Axial Velocity Density Ratio (AVDR), important parameter for compressor cascade investigations. Numerical results for Explicit Algebraic Reynolds Stress Model with transition modeling are compared with oil flow visualization, schlieren and Pressure Sensitive Paint. Boundary layer transition location is detected by Temperature Sensitive Paint.

Towards industrial-strength Navier-Stokes codes

NASA Technical Reports Server (NTRS)

Jou, Wen-Huei; Wigton, Laurence B.; Allmaras, Steven R.

1992-01-01

In this paper we discuss our experiences with Navier-Stokes (NS) codes using central differencing (CD) and scalar artificial dissipation (SAD). The NS-CDSAD codes have been developed by several researchers. Our results confirm that for typical commercial transport wing and wing/body configurations flying at transonic conditions with all turbulent boundary layers, NS-CDSAD codes, when used with the Johnson-King turbulence model, are capable of computing pressure distributions in excellent agreement with experimental data. However, results are not as good when laminar boundary layers are present. Exhaustive 2-D grid refinement studies supported by detailed analysis suggest that the numerical errors associated with SAD severely contaminate the solution in the laminar portion of the boundary layer. It is left as a challenge to the CFD community to find and fix the problems with Navier-Stokes codes and to produce a NS code which converges reliably and properly captures the laminar portion of the boundary layer on a reasonable grid.

Orbiter Gap Filler Bending Model for Re-entry

NASA Technical Reports Server (NTRS)

Campbell, Charles H.

2007-01-01

Pressure loads on a protruding gap filler during an Orbiter reentry are investigated to evaluate the likelihood of extraction due to pressure loads, and to ascertain how much bending will be induced by re-entry pressure loads. Oblique shock wave theory is utilized to develop a representation of the pressure loads induced on a gap filler for the ISSHVFW trajectory, representative of a heavy weight ISS return. A free body diagram is utilized to react the forces induced by the pressure forces. Preliminary results developed using these methods demonstrate that pressure loads, alone, are not likely causes of gap filler extraction during reentry. Assessment of the amount a gap filler will bend over is presented. Implications of gap filler bending during re-entry include possible mitigation of early boundary layer transition concerns, uncertainty in ground based measurement of protruding gap fillers from historical Orbiter flight history, and uncertainty in the use of Orbiter gap fillers for boundary layer prediction calibration. Authors will be added to the author list as appropriate.

Transient Growth Theory Prediction of Optimal Placing of Passive and Active Flow Control Devices for Separation Delay in LPT Airfoils

NASA Technical Reports Server (NTRS)

Tumin, Anatoli; Ashpis, David E.

2003-01-01

An analysis of the non-modal growth of perturbations in a boundary layer in the presence of a streamwise pressure gradient is presented. The analysis is based on PSE equations for an incompressible fluid. Examples with Falkner-Skan profiles indicate that a favorable pressure gradient decreases the non-modal growth while an unfavorable pressure gradient leads to an increase of the amplification. It is suggested that the transient growth mechanism be utilized to choose optimal parameters of tripping elements on a low-pressure turbine (LPT) airfoil. As an example, a boundary layer flow with a streamwise pressure gradient corresponding to the pressure distribution over a LPT airfoil is considered. It is shown that there is an optimal spacing of the tripping elements and that the transient growth effect depends on the starting point. At very low Reynolds numbers, there is a possibility to enhance the transient energy growth by means of wall cooling.

Investigation of low-speed turbulent separated flow around airfoils

NASA Technical Reports Server (NTRS)

Wadcock, Alan J.

1987-01-01

Described is a low-speed wind tunnel experiment to measure the flowfield around a two-dimensional airfoil operating close to maximum lift. Boundary layer separation occurs on the upper surface at x/c=0.85. A three-component laser velocimeter, coupled with a computer-controlled data acquisition system, was used to obtain three orthogonal mean velocity components and three components of the Reynolds stress tensor in both the boundary layer and wake of the airfoil. Pressure distributions on the airfoil, skin friction distribution on the upper surface of the airfoil, and integral properties of the airfoil boudary layer are also documented. In addition to these near-field flow properties, static pressure distributions, both upstream and downstream from the airfoil and on the walls of the wind tunnel, are also presented.

Flow unsteadiness effects on boundary layers

NASA Technical Reports Server (NTRS)

Murthy, Sreedhara V.

1989-01-01

The development of boundary layers at high subsonic speeds in the presence of either mass flux fluctuations or acoustic disturbances (the two most important parameters in the unsteadiness environment affecting the aerodynamics of a flight vehicle) was investigated. A high quality database for generating detailed information concerning free-stream flow unsteadiness effects on boundary layer growth and transition in high subsonic and transonic speeds is described. The database will be generated with a two-pronged approach: (1) from a detailed review of existing literature on research and wind tunnel calibration database, and (2) from detailed tests in the Boundary Layer Apparatus for Subsonic and Transonic flow Affected by Noise Environment (BLASTANE). Special instrumentation, including hot wire anemometry, the buried wire gage technique, and laser velocimetry were used to obtain skin friction and turbulent shear stress data along the entire boundary layer for various free stream noise levels, turbulence content, and pressure gradients. This database will be useful for improving the correction methodology of applying wind tunnel test data to flight predictions and will be helpful for making improvements in turbulence modeling laws.

Explicit finite-volume time-marching calculations of total temperature distributions in turbulent flow

NASA Technical Reports Server (NTRS)

Nicholson, Stephen; Moore, Joan G.; Moore, John

1987-01-01

A method was developed which calculates two-dimensional, transonic, viscous flow in ducts. The finite volume, time-marching formulation is used to obtain steady flow solutions of the Reynolds-averaged form of the Navier-Stokes equations. The entire calculation is performed in the physical domain. This paper investigates the introduction of a new formulation of the energy equation which gives improved transient behavior as the calculation converges. The effect of variable Prandtl number on the temperature distribution through the boundary layer is also investigated. A turbulent boundary layer in an adverse pressure gradient (M = 0.55) is used to demonstrate the improved transient temperature distribution obtained when the new formulation of the energy equation is used. A flat plate turbulent boundary layer with a supersonic free-stream Mach number of 2.8 is used to investigate the effect of Prandtl number on the distribution of properties through the boundary layer. The computed total temperature distribution and recovery factor agree well with the measurements when a variable Prandtl number is used through the boundary layer.

Explicit finite-volume time-marching calculations of total temperature distributions in turbulent flow

NASA Technical Reports Server (NTRS)

Nicholson, Stephen; Moore, Joan G.; Moore, John

1986-01-01

A method was developed which calculates two-dimensional, transonic, viscous flow in ducts. The finite volume, time-marching formulation is used to obtain steady flow solutions of the Reynolds-averaged form of the Navier-Stokes equations. The entire calculation is performed in the physical domain. This paper investigates the introduction of a new formulation of the energy equation which gives improved transient behavior as the calculation converges. The effect of variable Prandtl number on the temperature distribution through the boundary layer is also investigated. A turbulent boundary layer in an adverse pressure gradient (M = 0.55) is used to demonstrate the improved transient temperature distribution obtained when the new formulation of the energy equation is used. A flat plate turbulent boundary layer with a supersonic free-stream Mach number of 2.8 is used to investigate the effect of Prandtl number on the distribution of properties through the boundary layer. The computed total temperature distribution and recovery factor agree well with the measurements when a variable Prandtl number is used through the boundary layer.

An Experimental Investigation of Forced Mixing of a Turbulent Boundary Layer in an Annular Diffuser. Ph.D. Thesis - Ohio State Univ.; [for boundary layer control

NASA Technical Reports Server (NTRS)

Shaw, R. J.

1979-01-01

The forced mixing process of a turbulent boundary layer in an axisymmetric annular diffuser using conventional wing-like vortex generators was studied. Flow field measurements were made at four axial locations downstream of the vortex generators. At each axial location, a total of 25 equally spaced profiles were measured behind three consecutive vortex generators which formed two pairs of vortex generators. Hot film anemometry probes measured the boundary layer turbulence structure at the same locations where pressure measurements were made. Both single and cross film probes were used. The diffuser turbulence data was teken only for a nominal inlet Mach number of 0.3. Three vortex generator configurations were tested. The differences between configurations involved changes in size and relative vortex generator positions. All three vortex generator configurations tested provided increases in diffuser performance. Distinct differences in the boundary layer integral properties and skin friction levels were noted between configurations. The axial turbulence intensity and Reynolds stress profiles measured displayed similarities in trends but differences in levels for the three configurations.

The curved kinetic boundary layer of active matter.

PubMed

Yan, Wen; Brady, John F

2018-01-03

A body submerged in active matter feels the swim pressure through a kinetic accumulation boundary layer on its surface. The boundary layer results from a balance between translational diffusion and advective swimming and occurs on the microscopic length scale . Here , D T is the Brownian translational diffusivity, τ R is the reorientation time and l = U 0 τ R is the swimmer's run length, with U 0 the swim speed [Yan and Brady, J. Fluid. Mech., 2015, 785, R1]. In this work we analyze the swim pressure on arbitrary shaped bodies by including the effect of local shape curvature in the kinetic boundary layer. When δ ≪ L and l ≪ L, where L is the body size, the leading order effects of curvature on the swim pressure are found analytically to scale as J S λδ 2 /L, where J S is twice the (non-dimensional) mean curvature. Particle-tracking simulations and direct solutions to the Smoluchowski equation governing the probability distribution of the active particles show that λδ 2 /L is a universal scaling parameter not limited to the regime δ, l ≪ L. The net force exerted on the body by the swimmers is found to scale as F net /(n ∞ k s T s L 2 ) = f(λδ 2 /L), where f(x) is a dimensionless function that is quadratic when x ≪ 1 and linear when x ∼ 1. Here, k s T s = ζU 0 2 τ R /6 defines the 'activity' of the swimmers, with ζ the drag coefficient, and n ∞ is the uniform number density of swimmers far from the body. We discuss the connection of this boundary layer to continuum mechanical descriptions of active matter and briefly present how to include hydrodynamics into this purely kinetic study.

Instability waves and transition in adverse-pressure-gradient boundary layers

NASA Astrophysics Data System (ADS)

Bose, Rikhi; Zaki, Tamer A.; Durbin, Paul A.

2018-05-01

Transition to turbulence in incompressible adverse-pressure-gradient (APG) boundary layers is investigated by direct numerical simulations. Purely two-dimensional instability waves develop on the inflectional base velocity profile. When the boundary layer is perturbed by isotropic turbulence from the free stream, streamwise elongated streaks form and may interact with the instability waves. Subsequent mechanisms that trigger transition depend on the intensity of the free-stream disturbances. All evidence from the present simulations suggest that the growth rate of instability waves is sufficiently high to couple with the streaks. Under very low levels of free-stream turbulence (˜0.1 % ), transition onset is highly sensitive to the inlet disturbance spectrum and is accelerated if the spectrum contains frequency-wave-number combinations that are commensurate with the instability waves. Transition onset and completion in this regime is characterized by formation and breakdown of Λ vortices, but they are more sporadic than in natural transition. Beneath free-stream turbulence with higher intensity (1-2 % ), bypass transition mechanisms are dominant, but instability waves are still the most dominant disturbances in wall-normal and spanwise perturbation spectra. Most of the breakdowns were by disturbances with critical layers close to the wall, corresponding to inner modes. On the other hand, the propensity of an outer mode to occur increases with the free-stream turbulence level. Higher intensity free-stream disturbances induce strong streaks that favorably distort the boundary layer and suppress the growth of instability waves. But the upward displacement of high amplitude streaks brings them to the outer edge of the boundary layer and exposes them to ambient turbulence. Consequently, high-amplitude streaks exhibit an outer-mode secondary instability.

A spectrally accurate boundary-layer code for infinite swept wings

NASA Technical Reports Server (NTRS)

Pruett, C. David

1994-01-01

This report documents the development, validation, and application of a spectrally accurate boundary-layer code, WINGBL2, which has been designed specifically for use in stability analyses of swept-wing configurations. Currently, we consider only the quasi-three-dimensional case of an infinitely long wing of constant cross section. The effects of streamwise curvature, streamwise pressure gradient, and wall suction and/or blowing are taken into account in the governing equations and boundary conditions. The boundary-layer equations are formulated both for the attachment-line flow and for the evolving boundary layer. The boundary-layer equations are solved by marching in the direction perpendicular to the leading edge, for which high-order (up to fifth) backward differencing techniques are used. In the wall-normal direction, a spectral collocation method, based upon Chebyshev polynomial approximations, is exploited. The accuracy, efficiency, and user-friendliness of WINGBL2 make it well suited for applications to linear stability theory, parabolized stability equation methodology, direct numerical simulation, and large-eddy simulation. The method is validated against existing schemes for three test cases, including incompressible swept Hiemenz flow and Mach 2.4 flow over an airfoil swept at 70 deg to the free stream.

Rotor boundary layer development with inlet guide vane (IGV) wake impingement

NASA Astrophysics Data System (ADS)

Jia, Lichao; Zou, Tengda; Zhu, Yiding; Lee, Cunbiao

2018-04-01

This paper examines the transition process in a boundary layer on a rotor blade under the impingement of an inlet guide vane wake. The effects of wake strengths and the reduced frequency on the unsteady boundary layer development on a low-speed axial compressor were investigated using particle image velocimetry. The measurements were carried out at two reduced frequencies (fr = fIGVS0/U2i, fr = 1.35, and fr = 0.675) with the Reynolds number, based on the blade chord and the isentropic inlet velocity, being 97 500. At fr = 1.35, the flow separated at the trailing edge when the wake strength was weak. However, the separation was almost totally suppressed as the wake strength increased. For the stronger wake, both the wake's high turbulence and the negative jet behavior of the wake dominated the interaction between the unsteady wake and the separated boundary layer on the suction surface of the airfoil. The boundary layer displacement thickened first due to the negative jet effect. Then, as the disturbances developed underneath the wake, the boundary layer thickness reduced gradually. The high disturbance region convected downstream at a fraction of the free-stream velocity and spread in the streamwise direction. The separation on the suction surface was suppressed until the next wake's arrival. Because of the long recovery time at fr = 0.675, the boundary layer thickened gradually as the wake convected further downstream and finally separated due to the adverse pressure gradient. The different boundary layer states in turn affected the development of disturbances.

Flow-around modes for a rhomboid wing with a stall vortex in the shock layer

NASA Astrophysics Data System (ADS)

Zubin, M. A.; Maximov, F. A.; Ostapenko, N. A.

2017-12-01

The results of theoretical and experimental investigation of an asymmetrical hypersonic flow around a V-shaped wing with the opening angle larger than π on the modes with attached shockwaves on forward edges, when the stall flow is implemented on the leeward wing cantilever behind the kink point of the cross contour. In this case, a vortex of nonviscous nature is formed in which the velocities on the sphere exceeding the speed of sound and resulting in the occurrence of pressure shocks with an intensity sufficient for the separation of the turbulent boundary layer take place in the reverse flow according to the calculations within the framework of the ideal gas. It is experimentally established that a separation boundary layer can exist in the reverse flow, and its structure is subject to the laws inherent to the reverse flow in the separation region of the turbulent boundary layer arising in the supersonic conic flow under the action of a shockwave incident to the boundary layer.

Design philosophy of long range LFC transports with advanced supercritical LFC airfoils. [laminar flow control

NASA Technical Reports Server (NTRS)

Pfenninger, Werner; Vemuru, Chandra S.

1988-01-01

The achievement of 70 percent laminar flow using modest boundary layer suction on the wings, empennage, nacelles, and struts of long-range LFC transports, combined with larger wing spans and lower span loadings, could make possible an unrefuelled range halfway around the world up to near sonic cruise speeds with large payloads. It is shown that supercritical LFC airfoils with undercut front and rear lower surfaces, an upper surface static pressure coefficient distribution with an extensive low supersonic flat rooftop, a far upstream supersonic pressure minimum, and a steep subsonic rear pressure rise with suction or a slotted cruise flap could alleviate sweep-induced crossflow and attachment-line boundary-layer instability. Wing-mounted superfans can reduce fuel consumption and engine tone noise.

A Experimental Study of Fluctuating Pressure Loads Beneath Swept Shock Wave/boundary Layer Interactions

NASA Astrophysics Data System (ADS)

Garg, Sanjay

An experimental research program providing basic knowledge and establishing a database on the fluctuating pressure loads produced on aerodynamic surfaces beneath three-dimensional shock wave/boundary layer interactions is described. Such loads constitute a fundamental problem of critical concern to future supersonic and hypersonic flight vehicles. A turbulent boundary layer on a flat plate is subjected to interactions with swept planar shock waves generated by sharp fins. Fin angles from 10 ^circ to 20^circ at freestream Mach numbers of 3 and 4 produce a variety of interaction strengths from weak to very strong. Miniature pressure transducers flush-mounted in the flat plate have been used to measure interaction-induced wall pressure fluctuations. The distributions of properties of the pressure fluctuations, such as their rms level, amplitude distribution and power spectra, are also determined. Measurements have been made for the first time in the aft regions of these interactions, revealing fluctuating pressure levels as high as 155 dB, which places them in the category of significant aeroacoustic load generators. The fluctuations near the foot of the fin are dominated by low frequency (0-5 kHz) components, and are caused by a previously unrecognized random motion of the primary attachment line. This phenomenon is probably intimately linked to the unsteadiness of the separation shock at the start of the interaction. The characteristics of the pressure fluctuations are explained in light of the features of the interaction flowfield. In particular, physical mechanisms responsible for the generation of high levels of surface pressure fluctuations are proposed based on the results of the study. The unsteadiness of the flowfield of the surface is also examined via a novel, non-intrusive optical technique. Results show that the entire shock structure generated by the interaction undergoes relatively low-frequency oscillations.

Drag of two-dimensional small-amplitude symmetric and asymmetric wavy walls in turbulent boundary layers

NASA Technical Reports Server (NTRS)

Lin, J. C.; Walsh, M. J.; Balasubramanian, R.

1984-01-01

Included are results of an experimental investigation of low-speed turbulent flow over multiple two-dimensional transverse rigid wavy surfaces having a wavelength on the order of the boundary-layer thickness. Data include surface pressure and total drag measurements on symmetric and asymmetric wall waves under a low-speed turbulent boundary-layer flow. Several asymmetric wave configurations exhibited drag levels below the equivalent symmetric (sine) wave. The experimental results compare favorably with numerical predictions from a Reynolds-averaged Navier-Stokes spectral code. The reported results are of particular interest for the estimation of drag, the minimization of fabrication waviness effects, and the study of wind-wave interactions.

Turbulence measurements in high Reynolds number boundary layers

NASA Astrophysics Data System (ADS)

Vallikivi, Margit; Smits, Alexander

2013-11-01

Measurements are conducted in zero pressure gradient turbulent boundary layers for Reynolds numbers from Reθ = 9,000 to 225,000. The experiments were performed in the High Reynolds number Test Facility (HRTF) at Princeton University, which uses compressed air as the working fluid. Nano-Scale Thermal Anemometry Probes (NSTAPs) are used to acquire data with very high spatial and temporal precision. These new data are used to study the scaling behavior of the streamwise velocity fluctuations in the boundary layer and make comparisons with the scaling of other wall-bounded turbulent flows. Supported under ONR Grant N00014-09-1-0263 (program manager Ron Joslin) and NSF Grant CBET-1064257 (program manager Henning Winter).

Unsteady boundary-layer injection

NASA Technical Reports Server (NTRS)

Telionis, D. P.; Jones, G. S.

1981-01-01

The boundary-layer equations for two-dimensional incompressible flow are integrated numerically for the flow over a flat plate and a Howarth body. Injection is introduced either impulsively or periodically along a narrow strip. Results indicate that injection perpendicular to the wall is transmitted instantly across the boundary layer and has little effect on the velocity profile parallel to the wall. The effect is a little more noticeable for flows with adverse pressure gradients. Injection parallel to the wall results in fuller velocity profiles. Parallel and oscillatory injection appears to influence the mean. The amplitude of oscillation decreases with distance from the injection strip but further downstream it increases again in a manner reminiscent of an unstable process.

Thermal evolution of the high-pressure ice layers beneath a buried ocean within Titan and Ganymede

NASA Astrophysics Data System (ADS)

Choblet, G.; Tobie, G.

2015-12-01

Deep interiors of massive icy satellites such as Titan and Ganymede probably harbor a buried ocean above high-pressure (HP) ice layers. The nature and location of the lower interface of the ocean is ultimately controlled by the amount of heat transferred through the surface ice Ih layer but it also involves equilibration of heat and melt transfer in the HP ices. While the Rayleigh number associated to such HP ice layers is most probably supercritical, classical subsolidus convection might not be a viable mechanism as the radial temperature gradient in the cold boundary layer is likely to exceed the slope of the melting curve. A significant part of the heat transfer could be achieved via the mass flux of warm liquid through this cold boundary layer up to the global phase boundary, a phenomenon sometimes referred to as heat-pipe mechanism. We present 3D spherical simulations of thermal convection in these HP ice layers that address for the first time this complex interplay. First, scaling relationships are proposed to describe this configuration for a large range of Rayleigh numbers and solidus curves. We then focus on a more realistic set-up where a prescribed basal heat flux is considered in a plausible range for the thermal history of Ganymede or Titan together with the expected viscosity law for HP ices.

Active flow control of subsonic flow in an adverse pressure gradient using synthetic jets and passive micro flow control devices

NASA Astrophysics Data System (ADS)

Denn, Michael E.

Several recent studies have shown the advantages of active and/or passive flow control devices for boundary layer flow modification. Many current and future proposed air vehicles have very short or offset diffusers in order to save vehicle weight and create more optimal vehicle/engine integration. Such short coupled diffusers generally result in boundary layer separation and loss of pressure recovery which reduces engine performance and in some cases may cause engine stall. Deployment of flow control devices can alleviate this problem to a large extent; however, almost all active flow control devices have some energy penalty associated with their inclusion. One potential low penalty approach for enhancing the diffuser performance is to combine the passive flow control elements such as micro-ramps with active flow control devices such as synthetic jets to achieve higher control authority. The goal of this dissertation is twofold. The first objective is to assess the ability of CFD with URANS turbulence models to accurately capture the effects of the synthetic jets and micro-ramps on boundary layer flow. This is accomplished by performing numerical simulations replicating several experimental test cases conducted at Georgia Institute of Technology under the NASA funded Inlet Flow Control and Prediction Technologies Program, and comparing the simulation results with experimental data. The second objective is to run an expanded CFD matrix of numerical simulations by varying various geometric and other flow control parameters of micro-ramps and synthetic jets to determine how passive and active control devices interact with each other in increasing and/or decreasing the control authority and determine their influence on modification of boundary layer flow. The boundary layer shape factor is used as a figure of merit for determining the boundary layer flow quality/modification and its tendency towards separation. It is found by a large number of numerical experiments and the analysis of simulation data that a flow control device's influence on boundary layer quality is a function of three factors: (1) the strength of the longitudinal vortex emanating from the flow control device or devices, (2) the height of the vortex core above the surface and, when a synthetic jet is present, (3) the momentum added to the boundary layer flow.

Dynamics of the outgoing turbulent boundary layer in a Mach 5 unswept compression ramp interaction

NASA Technical Reports Server (NTRS)

Gramann, Richard A.; Dolling, David S.

1990-01-01

Wall pressure fluctuations have been measured under the unsteady separation shock and on the ramp face in an unswept Mach 5 compression ramp interaction. The freestream Reynolds number was 51.0 x 10 to the 6th/m, and the incoming turbulent boundary layer developed on the tunnel floor under approximately adiabatic wall temperature conditions. Standard data-acquisition methods, as well as real-time and posttest conditional sampling techniques were used. The results show that the mean and rms pressure levels are strong functions of separation shock position. At all stations on the ramp, from the corner to where the pressure reaches the theoretical inviscid value, the pressure signals have two dominant components: a low frequency component characteristic of the global unsteadiness, which correlates with the separation shock motion, and a higher frequency component associated with turbulence. The former is the major contributor to the overall signal variance.

Analysis of the separated boundary layer flow on the surface and in the wake of blunt trailing edge airfoils

NASA Technical Reports Server (NTRS)

Goradia, S. H.; Mehta, J. M.; Shrewsbury, G. S.

1977-01-01

The viscous flow phenomena associated with sharp and blunt trailing edge airfoils were investigated. Experimental measurements were obtained for a 17 percent thick, high performance GAW-1 airfoil. Experimental measurements consist of velocity and static pressure profiles which were obtained by the use of forward and reverse total pressure probes and disc type static pressure probes over the surface and in the wake of sharp and blunt trailing edge airfoils. Measurements of the upper surface boundary layer were obtained in both the attached and separated flow regions. In addition, static pressure data were acquired, and skin friction on the airfoil upper surface was measured with a specially constructed device. Comparison of the viscous flow data with data previously obtained elsewhere indicates reasonable agreement in the attached flow region. In the separated flow region, considerable differences exist between these two sets of measurements.

Diffusion in porous layers with memory

NASA Astrophysics Data System (ADS)

Caputo, Michele; Plastino, Wolfango

2004-07-01

The process of diffusion of fluid in porous media and biological membranes has usually been modelled with Darcy's constitutive equation, which states that the flux is proportional to the pressure gradient. However, when the permeability of the matrix changes during the process, solution of the equations governing the diffusion presents severe analytical difficulties because the variation of permeability is not known a priori. A diverse formulation of the constitutive law of diffusion is therefore needed and many authors have studied this problem using various methods and solutions. In this paper Darcy's constitutive equation is modified with the introduction of a memory formalism. We have also modified the second constitutive equation of diffusion which relates the density variations in the fluid to the pressure, introducing rheology in the fluid represented by memory formalisms operating on pressure variations as well as on density variations. The memory formalisms are then specified as derivatives of fractional order, solving the problem in the case of a porous layer when constant pressures are applied to its sides. For technical reasons many studies of diffusion are devoted to the flux rather than to the pressure; in this work we shall devote our attention to studying the pressure and compute the Green's function of the pressure in the layer when a constant pressure is applied to the boundary (Case A) for which we have found closed-form formulae. The described problem has already been considered for a half space (Caputo 2000); however, the results for a half space are mostly qualitative since in most practical problems the diffusion occurs in layers. The solution is also readily extended to the case when a periodic pressure is applied to one of the boundary planes while on the other the pressure is constant (Case B) which mimics the effect of the tides on sea coasts. In this case we have found a skin effect for the flux which limits the flux to a surface layer whose thickness decreases with increasing frequency. Regarding the effect of pressure due to tidal waters on the coast, it has been observed that when the medium is sand and the fluid is water, for a sinusoidal pressure of 2 × 104 Pa and a period of 24 hr at one of the boundaries and zero pressure at the other boundary, the flux is sinusoidal with the same period and amplitude decaying exponentially with distance to become negligible at a distance of a few hundred metres. A brief discussion is given concerning the mode of determination of the parameters of memory formalisms governing the diffusion using the observed pressure at several frequencies. We shall also see that, as in the classic case of pure Darcy's law behaviour, the equation governing the flux resulting in the diffusion through porous media with memory is the same as that governing the pressure.

Experimental Investigation of Soil and Atmospheric Conditions on the Momentum, Mass, and Thermal Boundary Layers Above the Land Atmosphere Interface

NASA Astrophysics Data System (ADS)

Trautz, A.; Smits, K. M.; Illangasekare, T. H.; Schulte, P.

2014-12-01

The purpose of this study is to investigate the impacts of soil conditions (i.e. soil type, saturation) and atmospheric forcings (i.e. velocity, temperature, relative humidity) on the momentum, mass, and temperature boundary layers. The atmospheric conditions tested represent those typically found in semi-arid and arid climates and the soil conditions simulate the three stages of evaporation. The data generated will help identify the importance of different soil conditions and atmospheric forcings with respect to land-atmospheric interactions which will have direct implications on future numerical studies investigating the effects of turbulent air flow on evaporation. The experimental datasets generated for this study were performed using a unique climate controlled closed-circuit wind tunnel/porous media facility located at the Center for Experimental Study of Subsurface Environmental Processes (CESEP) at the Colorado School of Mines. The test apparatus consisting of a 7.3 m long porous media tank and wind tunnel, were outfitted with a sensor network to carefully measure wind velocity, air and soil temperature, relative humidity, soil moisture, and soil air pressure. Boundary layer measurements were made between the heights of 2 and 500 mm above the soil tank under constant conditions (i.e. wind velocity, temperature, relative humidity). The soil conditions (e.g. soil type, soil moisture) were varied between datasets to analyze their impact on the boundary layers. Experimental results show that the momentum boundary layer is very sensitive to the applied atmospheric conditions and soil conditions to a much less extent. Increases in velocity above porous media leads to momentum boundary layer thinning and closely reflect classical flat plate theory. The mass and thermal boundary layers are directly dependent on both atmospheric and soil conditions. Air pressure within the soil is independent of atmospheric temperature and relative humidity - wind velocity and soil moisture effects were observed. This data provides important insight into future work of accurately modeling the exchange processes associated with evaporation under various turbulent atmospheric conditions.

Numerical calculation of boundary layers and wake characteristics of high-speed trains with different lengths

PubMed Central

Zhou, Dan; Niu, Jiqiang

2017-01-01

Trains with different numbers of cars running in the open air were simulated using the delayed detached-eddy simulation (DDES). The numbers of cars included in the simulation are 3, 4, 5 and 8. The aim of this study was to investigate how train length influences the boundary layer, the wake flow, the surface pressure, the aerodynamic drag and the friction drag. To certify the accuracy of the mesh and methods, the drag coefficients from numerical simulation of trains with 3 cars were compared with those from the wind tunnel test, and agreement was obtained. The results show that the boundary layer is thicker and the wake vortices are less symmetric as the train length increases. As a result, train length greatly affects pressure. The upper surface pressure of the tail car reduced by 2.9%, the side surface pressure of the tail car reduced by 8.3% and the underneath surface pressure of the tail car reduced by 19.7% in trains that included 3 cars to those including 8 cars. In addition, train length also has a significant effect on the friction drag coefficient and the drag coefficient. The friction drag coefficient of each car in a configuration decreases along the length of the train. In a comparison between trains consisting of 3 cars to those consisting of 8 cars, the friction drag coefficient of the tail car reduced by 8.6% and the drag coefficient of the tail car reduced by 3.7%. PMID:29261758

The lagRST Model: A Turbulence Model for Non-Equilibrium Flows

NASA Technical Reports Server (NTRS)

Lillard, Randolph P.; Oliver, A. Brandon; Olsen, Michael E.; Blaisdell, Gregory A.; Lyrintzis, Anastasios S.

2011-01-01

This study presents a new class of turbulence model designed for wall bounded, high Reynolds number flows with separation. The model addresses deficiencies seen in the modeling of nonequilibrium turbulent flows. These flows generally have variable adverse pressure gradients which cause the turbulent quantities to react at a finite rate to changes in the mean flow quantities. This "lag" in the response of the turbulent quantities can t be modeled by most standard turbulence models, which are designed to model equilibrium turbulent boundary layers. The model presented uses a standard 2-equation model as the baseline for turbulent equilibrium calculations, but adds transport equations to account directly for non-equilibrium effects in the Reynolds Stress Tensor (RST) that are seen in large pressure gradients involving shock waves and separation. Comparisons are made to several standard turbulence modeling validation cases, including an incompressible boundary layer (both neutral and adverse pressure gradients), an incompressible mixing layer and a transonic bump flow. In addition, a hypersonic Shock Wave Turbulent Boundary Layer Interaction with separation is assessed along with a transonic capsule flow. Results show a substantial improvement over the baseline models for transonic separated flows. The results are mixed for the SWTBLI flows assessed. Separation predictions are not as good as the baseline models, but the over prediction of the peak heat flux downstream of the reattachment shock that plagues many models is reduced.

Stability and Drag Reduction in a Boundary Layer with Microbubbles.

DTIC Science & Technology

1988-02-01

order accurate. .’ Since the numerical methods are not the object of this % dissertation, we decline from including the finite difference equations...previous appendix must be the special case of zero pressure gradient. Some entries of the matrices of the block tridiagonal system will be different ...of the wall mean velocity gradient was observed to be associated with the migration of the bubbles away from the boundary layer. The time scale of the

Hearing Protection - Needs, Technologies and Performance (Protection de l’ouie - besoins, technologies et resultats)

DTIC Science & Technology

2010-11-01

Human (AHAAH) which takes into account the whole signal transmission from the free sound field to the cochlear structures. Auditory Hazard Units...airflow around the aircraft canopy and the front structure of the aircraft (boundary layer flow noise), and the other is from internally generated noise...from the pressurisation and cockpit conditioning systems. The boundary layer flow noise is dependent upon the dynamic pressures on the aircraft and

HIFiRE-5 Boundary Layer Transition and HIFiRE-1 Shock Boundary Layer Interaction

DTIC Science & Technology

2015-10-01

pressure calcium-fluoride window and a new wind tunnel model consisting of a steel forebody and a PEEK plastic shell. This model contained a large...wind tunnel data. These models were then applied to the supersonic 2 Approved for public release; distribution unlimited. ascent phase of the...the tuned RANS models predicted upstream influence well. These results lend confidence to the approach of calibrating RANS models with wind tunnel

Investigations of Compression Shocks and Boundary Layers in Gases Moving at High Speed

NASA Technical Reports Server (NTRS)

Ackeret, J.; Feldmann, F.; Rott, N.

1947-01-01

The mutual influences of compression shocks and friction boundary layers were investigated by means of high speed wind tunnels.Schlieren optics provided a clear picture of the flow phenomena and were used for determining the location of the compression shocks, measurement of shock angles, and also for Mach angles. Pressure measurement and humidity measurements were also taken into consideration.Results along with a mathematical model are described.

The effects of micro-vortex generators on normal shock wave/boundary layer interactions

NASA Astrophysics Data System (ADS)

Herges, Thomas G.

Shock wave/boundary-layer interactions (SWBLIs) are complex flow phenomena that are important in the design and performance of internal supersonic and transonic flow fields such as engine inlets. This investigation was undertaken to study the effects of passive flow control devices on normal shock wave/boundary layer interactions in an effort to gain insight into the physics that govern these complex interactions. The work concentrates on analyzing the effects of vortex generators (VGs) as a flow control method by contributing a greater understanding of the flowfield generated by these devices and characterizing their effects on the SWBLI. The vortex generators are utilized with the goal of improving boundary layer health (i.e., reducing/increasing the boundary-layer incompressible shape factor/skin friction coefficient) through a SWBLI, increasing pressure recovery, and reducing flow distortion at the aerodynamic interface plane while adding minimal drag to the system. The investigation encompasses experiments in both small-scale and large-scale inlet testing, allowing multiple test beds for improving the characterization and understanding of vortex generators. Small-scale facility experiments implemented instantaneous schlieren photography, surface oil-flow visualization, pressure-sensitive paint, and particle image velocimetry to characterize the effects of an array of microramps on a normal shock wave/boundary-layer interaction. These diagnostics measured the time-averaged and instantaneous flow organization in the vicinity of the microramps and SWBLI. The results reveal that a microramp produces a complex vortex structure in its wake with two primary counter-rotating vortices surrounded by a train of Kelvin- Helmholtz (K-H) vortices. A streamwise velocity deficit is observed in the region of the primary vortices in addition to an induced upwash/downwash which persists through the normal shock with reduced strength. The microramp flow control also increased the spanwise-averaged skin-friction coefficient and reduced the spanwise-averaged incompressible shape factor, thereby improving the health of the boundary layer. The velocity in the near-wall region appears to be the best indicator of microramp effectiveness at controlling SWBLIs. Continued analysis of additional micro-vortex generator designs in the small-scale facility revealed reduced separation within a subsonic diffuser downstream of the normal shock wave/boundary layer interaction. The resulting attached flow within the diffuser from the micro-vortex generator control devices reduces shock wave position and pressure RMS fluctuations within the diffuser along with increased pressure recovery through the shock and at the entrance of the diffuser. The largest effect was observed by the micro-vortex generators that produce the strongest streamwise vortices. High-speed pressure measurements also indicated that the vortex generators shift the energy of the pressure fluctuations to higher frequencies. Implementation of micro-vortex generators into a large-scale, supersonic, axisymmetric, relaxed-compression inlet have been investigated with the use of a unique and novel flow-visualization measurement system designed and successfully used for the analysis of both upstream micro-VGs (MVGs) and downstream VGs utilizing surface oil-flow visualization and pressure-sensitive paint measurements. The inlet centerbody and downstream diffuser vortex-generator regions were imaged during wind-tunnel testing internally through the inlet cowl with the diagnostic system attached to the cowl. Surface-flow visualization revealed separated regions along the inlet centerbody for large mass-flow rates without vortex generators. Upstream vortex generators did reduce separation in the subsonic diffuser, and a unique perspective of the flowfield produced by the downstream vortex generators was obtained. In addition, pressure distributions on the inlet centerbody and vortex generators were measured with pressure-sensitive paint. At low mass-flow ratios the onset of buzz occurs in the large-scale low-boom inlet. Inlet buzz and how it is affected by vortex generators was characterized using shock tracking through high-speed schlieren imaging and pressure fluctuation measurements. The analysis revealed a dominant low frequency oscillation at 21.0 Hz for the single-stream inlet, corresponding with the duration of one buzz cycle. Pressure oscillations prior to the onset of buzz were not detected, leaving the location where the shock wave triggers large separation on the compression spike as the best indicator for the onset of buzz. The driving mechanism for a buzz cycle has been confirmed as the rate of depressurization and repressurization of the inlet as the buzz cycle fluctuates between an effectively unstarted (blocked) inlet and supercritical operation (choked flow), respectively. High-frequency shock position oscillations/pulsations (spike buzz) were also observed throughout portions of the inlet buzz cycle. The primary effect of the VGs was to trigger buzz at a higher mass-flow ratio.

Observations of the Early Evening Boundary-Layer Transition Using a Small Unmanned Aerial System

NASA Astrophysics Data System (ADS)

Bonin, Timothy; Chilson, Phillip; Zielke, Brett; Fedorovich, Evgeni

2013-01-01

The evolution of the lower portion of the planetary boundary layer is investigated using the Small Multifunction Research and Teaching Sonde (SMARTSonde), an unmanned aerial vehicle developed at the University of Oklahoma. The study focuses on the lowest 200 m of the atmosphere, where the most noticeable thermodynamic changes occur during the day. Between October 2010 and February 2011, a series of flights was conducted during the evening hours on several days to examine the vertical structure of the lower boundary layer. Data from a nearby Oklahoma Mesonet tower was used to supplement the vertical profiles of temperature, humidity, and pressure, which were collected approximately every 30 min, starting 2 h before sunset and continuing until dusk. From the profiles, sensible and latent heat fluxes were estimated. These fluxes were used to diagnose the portion of the boundary layer that was most affected by the early evening transition. During the transition period, a shallow cool and moist layer near the ground was formed, and as the evening progressed the cooling affected an increasingly shallower layer just above the surface.

Three Dimensional Viscous Flow Field in an Axial Flow Turbine Nozzle Passage

NASA Technical Reports Server (NTRS)

Ristic, D.; Lakshminarayana, B.

1997-01-01

The objective of this investigation is experimental and computational study of three dimensional viscous flow field in the nozzle passage of an axial flow turbine stage. The nozzle passage flow field has been measured using a two sensor hot-wire probe at various axial and radial stations. In addition, two component LDV measurements at one axial station (x/c(sum m) = 0.56) were performed to measure the velocity field. Static pressure measurements and flow visualization, using a fluorescent oil technique, were also performed to obtain the location of transition and the endwall limiting streamlines. A three dimensional boundary layer code, with a simple intermittency transition model, was used to predict the viscous layers along the blade and endwall surfaces. The boundary layers on the blade surface were found to be very thin and mostly laminar, except on the suction surface downstream of 70% axial chord. Strong radial pressure gradient, especially close to the suction surface, induces strong cross flow components in the trailing edge regions of the blade. On the end-walls the boundary layers were much thicker, especially near the suction corner of the casing surface, caused by secondary flow. The secondary flow region near the suction-casing surface corner indicates the presence of the passage vortex detached from the blade surface. The corner vortex is found to be very weak. The presence of a closely spaced rotor downstream (20% of the nozzle vane chord) introduces unsteadiness in the blade passage. The measured instantaneous velocity signal was filtered using FFT square window to remove the periodic unsteadiness introduced by the downstream rotor and fans. The filtering decreased the free stream turbulence level from 2.1% to 0.9% but had no influence on the computed turbulence length scale. The computation of the three dimensional boundary layers is found to be accurate on the nozzle passage blade surfaces, away from the end-walls and the secondary flow region. On the nozzle passage endwall surfaces the presence of strong pressure gradients and secondary flow limit the validity of the boundary layer code.

Diamond anvil cell for spectroscopic investigation of materials at high temperature, high pressure and shear

DOEpatents

Westerfield, Curtis L.; Morris, John S.; Agnew, Stephen F.

1997-01-01

Diamond anvil cell for spectroscopic investigation of materials at high temperature, high pressure and shear. A cell is described which, in combination with Fourier transform IR spectroscopy, permits the spectroscopic investigation of boundary layers under conditions of high temperature, high pressure and shear.

Global stability analysis of axisymmetric boundary layer over a circular cylinder

NASA Astrophysics Data System (ADS)

Bhoraniya, Ramesh; Vinod, Narayanan

2018-05-01

This paper presents a linear global stability analysis of the incompressible axisymmetric boundary layer on a circular cylinder. The base flow is parallel to the axis of the cylinder at inflow boundary. The pressure gradient is zero in the streamwise direction. The base flow velocity profile is fully non-parallel and non-similar in nature. The boundary layer grows continuously in the spatial directions. Linearized Navier-Stokes (LNS) equations are derived for the disturbance flow quantities in the cylindrical polar coordinates. The LNS equations along with homogeneous boundary conditions forms a generalized eigenvalues problem. Since the base flow is axisymmetric, the disturbances are periodic in azimuthal direction. Chebyshev spectral collocation method and Arnoldi's iterative algorithm is used for the solution of the general eigenvalues problem. The global temporal modes are computed for the range of Reynolds numbers and different azimuthal wave numbers. The largest imaginary part of the computed eigenmodes is negative, and hence, the flow is temporally stable. The spatial structure of the eigenmodes shows that the disturbance amplitudes grow in size and magnitude while they are moving towards downstream. The global modes of axisymmetric boundary layer are more stable than that of 2D flat-plate boundary layer at low Reynolds number. However, at higher Reynolds number they approach 2D flat-plate boundary layer. Thus, the damping effect of transverse curvature is significant at low Reynolds number. The wave-like nature of the disturbance amplitudes is found in the streamwise direction for the least stable eigenmodes.

Aerodynamic pressures and heating rates on surfaces between split elevons at Mach 6.6

NASA Technical Reports Server (NTRS)

Hunt, L. Roane

1988-01-01

An aerothermal study was performed in the Langley 8-Foot High Temperature Tunnel at Mach number 6.6 to define the pressures and heating rates on the surfaces between split elevons similar to those used on the Space Shuttle. Tests were performed with both laminar and turbulent boundary layers on the wing surface upstream of the elevons. The flow in the chordwise gap between the elevons was characterized by flow separation at the gap entrance and flow reattachment at a depth into the gap inversely proportional to the gap width. The gap pressure and heating rate increased significantly with decrease of elevon gap width, and the maximum gap heating rate was proportional to the maximum gap pressure. Correlation of the present results indicate that the gap heating was directly proportional to the elevon windward surface pressure and was not dependent upon whether the boundary layer on the windward elevon surface was laminar or turbulent.

Fuel decomposition and boundary-layer combustion processes of hybrid rocket motors

NASA Technical Reports Server (NTRS)

Chiaverini, Martin J.; Harting, George C.; Lu, Yeu-Cherng; Kuo, Kenneth K.; Serin, Nadir; Johnson, David K.

1995-01-01

Using a high-pressure, two-dimensional hybrid motor, an experimental investigation was conducted on fundamental processes involved in hybrid rocket combustion. HTPB (Hydroxyl-terminated Polybutadiene) fuel cross-linked with diisocyanate was burned with GOX under various operating conditions. Large-amplitude pressure oscillations were encountered in earlier test runs. After identifying the source of instability and decoupling the GOX feed-line system and combustion chamber, the pressure oscillations were drastically reduced from +/-20% of the localized mean pressure to an acceptable range of +/-1.5% Embedded fine-wire thermocouples indicated that the surface temperature of the burning fuel was around 1000 K depending upon axial locations and operating conditions. Also, except near the leading-edge region, the subsurface thermal wave profiles in the upstream locations are thicker than those in the downstream locations since the solid-fuel regression rate, in general, increases with distance along the fuel slab. The recovered solid fuel slabs in the laminar portion of the boundary layer exhibited smooth surfaces, indicating the existence of a liquid melt layer on the burning fuel surface in the upstream region. After the transition section, which displayed distinct transverse striations, the surface roughness pattern became quite random and very pronounced in the downstream turbulent boundary-layer region. Both real-time X-ray radiography and ultrasonic pulse-echo techniques were used to determine the instantaneous web thickness burned and instantaneous solid-fuel regression rates over certain portions of the fuel slabs. Globally averaged and axially dependent but time-averaged regression rates were also obtained and presented.

Numerical simulation of adverse-pressure-gradient boundary layer with or without roughness

NASA Astrophysics Data System (ADS)

Mottaghian, Pouya; Yuan, Junlin; Piomelli, Ugo

2014-11-01

Large-eddy and direct numerical simulations are carried out on flat-plate boundary layer over smooth and rough surfaces, with adverse pressure gradient.The deceleration is achieved by imposing a wall-normal freestream velocity profile, and is strong enough to cause separation at the wall. The Reynolds number based on momentum thickness and freestream velocity at inlet is 600. Numerical sandgrain roughness is applied based on an immersed boundary method, yielding a flow that is transitionally rough. The turbulence intensity increases before separation, and reaches a higher value for the rough case, indicating stronger mixing. Roughness also causes higher momentum deficit near the wall, leading to earlier separation. This is consistent with previous observation made on rough-wall flow separation over a ramp. In both cases, the turbulent kinetic energy peaks inside the shear layer above the detachment region, with higher values in the rough case; it then decreases approaching the reattachment region. Near the wall inside the separation bubble, the near-zero turbulent intensity indicates that the turbulent structures are lifted up in the separation region. Compared with the smooth case, the shear layer is farther from the wall and the reattachment length is longer on the rough wall.

Practical application of computer programs for supersonic combustion

NASA Technical Reports Server (NTRS)

Groves, F. R., Jr.

1972-01-01

Experimental data were interpreted using two supersonic combustion computer programs. The P1 program is based on a conventional boundary layer treatment of the mixing of concentric gas streams and complete combustion chemistry. The H1 program is based on a modified boundary layer approach which accounts for radial pressure gradients in the flow and also incorporates a finite rate chemistry calculation. The objective of the investigation was to compare the experimental data with theoretical predictions of the two programs with special emphasis on the prediction of radial pressure gradients by the H1 program. A test of the H1 program was also desired through comparison with the experimental data and with the P1 program.

The measurement of shear stress and total heat flux in a nonadiabatic turbulent hypersonic boundary layer

NASA Technical Reports Server (NTRS)

Mikulla, V.; Horstman, C. C.

1975-01-01

Turbulent shear stress and direct turbulent total heat-flux measurements have been made across a nonadiabatic, zero pressure gradient, hypersonic boundary layer by using specially designed hot-wire probes free of strain-gauging and wire oscillation. Heat-flux measurements were in reasonably good agreement with values obtained by integrating the energy equation using measured profiles of velocity and temperature. The shear-stress values deduced from the measurements, by assuming zero correlation of velocity and pressure fluctuations, were lower than the values obtained by integrating the momentum equation. Statistical properties of the cross-correlations are similar to corresponding incompressible measurements at approximately the same momentum-thickness Reynolds number.

A three-dimensional turbulent separated flow and related mesurements

NASA Technical Reports Server (NTRS)

Pierce, F. J.

1985-01-01

The applicability of and the limits on the applicability of 11 near wall similarity laws characterizing three-dimensional turbulent boundary layer flows were determined. A direct force sensing local wall shear stress meter was used in both pressure-driven and shear-driven three-dimensional turbulent boundary layers, together with extensive mean velocity field and wall pressure field data. This resulted in a relatively large number of graphical comparisons of the predictive ability of 10 of these 11 similarity models relative to measured data over a wide range of flow conditions. Documentation of a complex, separated three-dimensional turbulent flow as a standard test case for evaluating the predictive ability of numerical codes solving such flows is presented.

An Investigation into the Mechanics of Windblown Dust Entrainment from Nickel Slag Surfaces Resembling Armoured Desert Pavements

NASA Astrophysics Data System (ADS)

Sanderson, Robert Steven

The purpose of this thesis is to investigate the dynamics of PM 10 emission from a nickel slag stockpile that closely resembles a desert pavement in physical characteristics. In the field, it was observed that slag surfaces develop by natural processes into a well-armoured surface over some period of time. The surface then consists of two distinct layers; a surficial armour layer containing only non-erodible gravel and cobble-sized clasts, and an underlying dust-laden layer, which contains a wide size range of slag particles, from clay-sized to cobble-sized. This surficial armour layer protects the underlying fines from wind entrainment, at least under typical wind conditions; however, particle emissions still do occur under high wind speeds. The dynamics of particle entrainment from within these surfaces are investigated herein. It is shown that the dynamics of the boundary layer flow over these lag surfaces are influenced by the inherent roughness and permeability of the surficial armour layer, such that the flow resembles those observed over and within vegetation canopies, and those associated with permeable gravel-bed river channels. Restriction of air flow within the permeable surface produces a high-pressure zone within the pore spaces, resulting in a Kelvin-Helmholtz shear instability, which triggers coherent motions in the form of repeating burst-sweep cycles. Using Laser Doppler Anemometry (LDA), it is demonstrated that the lower boundary layer is characterized by both Q4 sweeping motions and Q2 bursting motions, while the upper boundary layer is dominated by Q2 bursts. Pore air motions within the slag material were measured using buried pressure ports. It is shown that the mean pressure gradient which forms within the slag material results in net upward displacement of air, or wind pumping. However, this net upward motion is a result of rapid oscillatory motions which are directly driven by coherent boundary layer motions. It is also demonstrated that these coherent motions are able to penetrate at least 4 cm through the surficial armour layer, thereby transporting turbulent kinetic energy (TKE) downward to the dust-laden sub-surface layer. This represents a mechanism of momentum transfer that is able to reach the erodible material, while the wind pumping effect represents a mechanism for particle exhaustion.

Experimental and theoretical investigation of three-dimensional turbulent boundary layers and turbulence characteristics inside an axial flow inducer passage. Final Report. Ph.D. Thesis, Jun. 1971

NASA Technical Reports Server (NTRS)

Anand, A. K.; Lakshminarayana, B.

1977-01-01

Analytical and experimental investigations of the characteristics of three dimensional turbulent boundary layers in a rotating helical passage of an inducer rotor are reported. Expressions are developed for the velocity profiles in the inner layer, where the viscous effects dominate, in the outer layer, where the viscous effects are small, and in the interference layer, where the end walls influence the flow. The prediction of boundary layer growth is based on the momentum integral technique. The equations derived are general enough to be valid for all turbomachinery rotors with arbitrary pressure gradients. The experimental investigations are carried out in a flat plate inducer 3 feet in diameter. The mean velocity profiles, turbulence intensities and shear stresses, wall shear stress, and limiting streamline angles are measured at various radial and chordwise locations by using rotating probes. The measurements are in general agreement with the predictions. The radial flows are well represented by an expression which includes the effect of stagger angle and radial pressure gradient. The radial flows in the rotor channel are higher than those on a single blade. The collateral region exists only very near the blade surface. The radial component of turbulence intensity is higher than the streamwise component because of the effect of rotation.

High-pressure flame visualization of autoignition and flashback phenomena with liquid-fuel spray

NASA Technical Reports Server (NTRS)

Marek, C. J.; Baker, C. E.

1983-01-01

A study was undertaken to determine the effect of boundary layers on autoignition and flashback for premixed Jet-A fuel in a unique high-pressure windowed test facility. A plate was placed in the center of the fuel-air stream to establish a boundary layer. Four experimental configurations were tested: a 24.5-cm-long plate with either a pointed leading edge, a rounded edge or an edge with a 0.317-cm step, or the duct without the plate. Experiments at an equivalence ratio ranging from 0.4 to 0.9 were performed at pressures to 2500 kPa (25 atm.) at temperatures of 600, 645, and 700 K and velocities to 115 meters per second. Flame shapes were observed during flashback and autoignition using high speed cinematography. Flashback and autoignition limits were determined.

Generalized Wall Function for Complex Turbulent Flows

NASA Technical Reports Server (NTRS)

Shih, Tsan-Hsing; Povinelli, Louis A.; Liu, Nan-Suey; Chen, Kuo-Huey

2000-01-01

A generalized wall function was proposed by Shih et al., (1999). It accounts the effect of pressure gradients on the flow near the wall. Theory shows that the effect of pressure gradients on the flow in the inertial sublayer is very significant and the standard wall function should be replaced by a generalized wall function. Since the theory is also valid for boundary layer flows toward separation, the generalized wall function may be applied to complex turbulent flows with acceleration, deceleration, separation and recirculation. This paper is to verify the generalized wall function with numerical simulations for boundary layer flows with various adverse and favorable pressure gradients, including flows about to separate. Furthermore, a general procedure of implementation of the generalized wall function for National Combustion Code (NCC) is described, it can be applied to both structured and unstructured CFD codes.

An experimental/computational study of sharp fin induced shock wave/turbulent boundary layer interactions at Mach 5 - Experimental results

NASA Technical Reports Server (NTRS)

Rodi, Patrick E.; Dolling, David S.

1992-01-01

A combined experimental/computational study has been performed of sharp fin induced shock wave/turbulent boundary layer interactions at Mach 5. The current paper focuses on the experiments and analysis of the results. The experimental data include mean surface heat transfer, mean surface pressure distributions and surface flow visualization for fin angles of attack of 6, 8, 10, 12, 14 and 16-degrees at Mach 5 under a moderately cooled wall condition. Comparisons between the results and correlations developed earlier show that Scuderi's correlation for the upstream influence angle (recast in a conical form) is superior to other such correlations in predicting the current results, that normal Mach number based correlations for peak pressure heat transfer are adequate and that the initial heat transfer peak can be predicted using pressure-interaction theory.

Aerodynamic characteristics of NACA RM-10 missile in 8- by 6-foot supersonic wind tunnel at Mach numbers from 1.49 to 1.98 I : presentation and analysis of pressure measurements (stabilizing fins removed)

NASA Technical Reports Server (NTRS)

Luidens, Roger W; Simon, Paul C

1950-01-01

Experimental investigation of flow about a slender body of revolution (NACA RM-10 missile) aligned and inclined to a supersonic stream was conducted at Mach numbers from 1.49 to 1.98 at a Reynolds number of approximately 30,000,000. Boundary-layer measurements at zero angle of attack are correlated with subsonic formulations for predicting boundary-layer thickness and profile. Comparison of pressure coefficients predicted by theory with experimental values showed close agreement at zero angle of attack and angle of attack except over the aft leeward side of body. At angle of attack, pitot pressure measurements in plane of model base indicated a pair of symmetrically disposed vortices on leeward side of body.

Simulation of hypersonic shock wave - laminar boundary layer interactions

NASA Astrophysics Data System (ADS)

Kianvashrad, N.; Knight, D.

2017-06-01

The capability of the Navier-Stokes equations with a perfect gas model for simulation of hypersonic shock wave - laminar boundary layer interactions is assessed. The configuration is a hollow cylinder flare. The experimental data were obtained by Calspan-University of Buffalo (CUBRC) for total enthalpies ranging from 5.07 to 21.85 MJ/kg. Comparison of the computed and experimental surface pressure and heat transfer is performed and the computed §ow¦eld structure is analyzed.

The Influence of Hydrofoil Oscillation on Boundary Layer Transition and Cavitation Noise.

DTIC Science & Technology

1981-04-01

p., and V. are fluid density, reference free-stream AVERAGED static pressure and reference free-stream velocity, respec- ORSPL3. Lair tively. S3.1...of cavitation on a propeller in both uni- ly, the location of boundary layer transition with the foil in form and nonuniform flow. He concluded that...the presence of oscillation must be determined either theoretically or ex- sheet and bubble cavitation in nonuniform flow can be perimentally. Thirdly

Boundary layer simulator improvement

NASA Technical Reports Server (NTRS)

Praharaj, S. C.; Schmitz, C.; Frost, C.; Engel, C. D.; Fuller, C. E.; Bender, R. L.; Pond, J.

1984-01-01

High chamber pressure expander cycles proposed for orbit transfer vehicles depend primarily on the heat energy transmitted from the combustion products through the thrust wall chamber wall. The heat transfer to the nozzle wall is affected by such variables as wall roughness, relamarization, and the presence of particles in the flow. Motor performance loss for these nozzles with thick boundary layers is inaccurate using the existing procedure coded BLIMPJ. Modifications and innovations to the code are examined. Updated routines are listed.

Application of the Program Profile for the Design of Low-Speed, Low- Observable Configuration Airfoils

DTIC Science & Technology

1992-12-01

112 61 . Airfoil T503 - t/c = 3.79% .... ........... .. 113 62. Airfoil T503 Leading-Edge - t/c = 3.79% ..... ... 114 63. Pressure...points on C unit circle, 6 slope of airfoil surface near trailing edge 61 boundary-layer displacement thickness 62 boundary-layer momentum thickness 63...equivalent thickness NACA 4-digit airfoils . 4 II. Theory Potential-Flow Design Method This section will overview the basic theory used in PROFILE. Eppler

Mars boundary layer simulations - Comparison with Viking lander and entry observations

NASA Technical Reports Server (NTRS)

Haberle, R. M.; Houben, H. C.

1991-01-01

Diurnal variations of wind and temperature in the lower Martian atmosphere are simulated with a boundary layer model that includes radiative heating in a dusty CO2 atmosphere, turbulence generated by convection and/or shear stresses, a surface heat budget, and time varying pressure forces due to sloping terrain. Model results for early northern summer are compared with Viking lander observations to determine the model's strengths and weaknesses, and suitability as an engineering model.

Conical similarity of shock/boundary layer interactions generated by swept fins

NASA Technical Reports Server (NTRS)

Lu, F. K.; Settles, G. S.

1983-01-01

A parametric experimental study has been made of the class of 3D shock wave/turbulent boundary layer interactions generated by swept-leading-edge fins. The fin sweepback angles ranged from 0 to 65 deg at angles of attack of 5, 9, and 15 deg. Two equilibrium 2D turbulent boundary layers with a free-stream Mach number of 2.95 and a Reynolds number of 6.3 x 10 to the 7th/m were used as incoming flow conditions. All the resulting interactions were found to possess conical symmetry of surface pressures and skin friction lines beyond an initial inception zone. Further, these interactions revealed a simple similarity based on inviscid shock strength irrespective of fin sweepback or angle of attack.

Boundary-layer measurements on a transonic low-aspect ratio wing

NASA Technical Reports Server (NTRS)

Keener, Earl R.

1985-01-01

Tabulations and plots are presented of boundary-layer velocity and flow-direction surveys from wind-tunnel tests of a large-scale (0.90 m semi-span) model of the NASA/Lockheed Wing C. This wing is a generic, transonic, supercritical, highly three-dimensional, low-aspect-ratio configuration designed with the use of a three-dimensional, transonic full-potential-flow wing code (FLO22). Tests were conducted at the design angle of attack of 5 deg over a Mach number range from 0.25 to 0.96 and a Reynolds number range of 3.4x10 to the 6th power. Wing pressures were measured at five span stations, and boundary-layer surveys were measured at the midspan station. The data are presented without analysis.

Experiments with a wing from which the boundary layer is removed by pressure or suction

NASA Technical Reports Server (NTRS)

Wieland, K

1928-01-01

With an unsymmetrical wing and a rotating Magnus cylinder, the lift is produced by the superposition of parallel and circulatory flows. An explanation of the circulatory flow is furnished by the boundary-layer theory of Prandtl and the consequent vortex formation. According to this explanation, it must evidently be possible to increase the circulation either by increasing the size of the stronger (lower) vortex or by decreasing the size of the weaker (upper) vortex. In this sense, according to Professor H. Zickendraht, we have a new type of wing from which the boundary layer is removed by forcing air out or sucking it in through openings in the upper surface of the wing near its trailing edge.

Leading-edge effects on boundary-layer receptivity

NASA Technical Reports Server (NTRS)

Gatski, Thomas B.; Kerschen, Edward J.

1990-01-01

Numerical calculations are presented for the incompressible flow over a parabolic cylinder. The computational domain extends from a region upstream of the body downstream to the region where the Blasius boundary-layer solution holds. A steady mean flow solution is computed and the results for the scaled surface vorticity, surface pressure and displacement thickness are compared to previous studies. The unsteady problem is then formulated as a perturbation solution starting with and evolving from the mean flow. The response to irrotational time harmonic pulsation of the free-stream is examined. Results for the initial development of the velocity profile and displacement thickness are presented. These calculations will be extended to later times to investigate the initiation of instability waves within the boundary-layer.

Analysis of the interaction of a weak normal shock wave with a turbulent boundary layer

NASA Technical Reports Server (NTRS)

Melnik, R. E.; Grossman, B.

1974-01-01

The method of matched asymptotic expansions is used to analyze the interaction of a normal shock wave with an unseparated turbulent boundary layer on a flat surface at transonic speeds. The theory leads to a three-layer description of the interaction in the double limit of Reynolds number approaching infinity and Mach number approaching unity. The interaction involves an outer, inviscid rotational layer, a constant shear-stress wall layer, and a blending region between them. The pressure distribution is obtained from a numerical solution of the outer-layer equations by a mixed-flow relaxation procedure. An analytic solution for the skin friction is determined from the inner-layer equations. The significance of the mathematical model is discussed with reference to existing experimental data.

Numerical Experiments in Error Control for Sound Propagation Using a Damping Layer Boundary Treatment

NASA Technical Reports Server (NTRS)

Goodrich, John W.

2017-01-01

This paper presents results from numerical experiments for controlling the error caused by a damping layer boundary treatment when simulating the propagation of an acoustic signal from a continuous pressure source. The computations are with the 2D Linearized Euler Equations (LEE) for both a uniform mean flow and a steady parallel jet. The numerical experiments are with algorithms that are third, fifth, seventh and ninth order accurate in space and time. The numerical domain is enclosed in a damping layer boundary treatment. The damping is implemented in a time accurate manner, with simple polynomial damping profiles of second, fourth, sixth and eighth power. At the outer boundaries of the damping layer the propagating solution is uniformly set to zero. The complete boundary treatment is remarkably simple and intrinsically independant from the dimension of the spatial domain. The reported results show the relative effect on the error from the boundary treatment by varying the damping layer width, damping profile power, damping amplitude, propagtion time, grid resolution and algorithm order. The issue that is being addressed is not the accuracy of the numerical solution when compared to a mathematical solution, but the effect of the complete boundary treatment on the numerical solution, and to what degree the error in the numerical solution from the complete boundary treatment can be controlled. We report maximum relative absolute errors from just the boundary treatment that range from O[10-2] to O[10-7].

Deformation of leaky-dielectric fluid globules under strong electric fields: Boundary layers and jets at large Reynolds numbers

NASA Astrophysics Data System (ADS)

Schnitzer, Ory; Frankel, Itzchak; Yariv, Ehud

2013-11-01

In Taylor's theory of electrohydrodynamic drop deformation (Proc. R. Soc. Lond. A, vol. 291, 1966, pp. 159-166), inertia is neglected at the outset, resulting in fluid velocity that scales as the square of the applied-field magnitude. For large drops, with increasing field strength the Reynolds number predicted by this scaling may actually become large, suggesting the need for a complementary large-Reynolds-number investigation. Balancing viscous stresses and electrical shear forces in this limit reveals a different velocity scaling, with the 4/3-power of the applied-field magnitude. We focus here on the flow over a gas bubble. It is essentially confined to two boundary layers propagating from the poles to the equator, where they collide to form a radial jet. At leading order in the Capillary number, the bubble deforms due to (i) Maxwell stresses; (ii) the hydrodynamic boundary-layer pressure associated with centripetal acceleration; and (iii) the intense pressure distribution acting over the narrow equatorial deflection zone, appearing as a concentrated load. Remarkably, the unique flow topology and associated scalings allow to obtain a closed-form expression for this deformation through application of integral mass and momentum balances. On the bubble scale, the concentrated pressure load is manifested in the appearance of a non-smooth equatorial dimple.

Analysis of Transition-Sensitized Turbulent Transport Equations

NASA Technical Reports Server (NTRS)

Rumsey, Christopher L.; Thacker, William D.; Gatski, Thomas B.; Grosch, Chester E,

2005-01-01

The dynamics of an ensemble of linear disturbances in boundary-layer flows at various Reynolds numbers is studied through an analysis of the transport equations for the mean disturbance kinetic energy and energy dissipation rate. Effects of adverse and favorable pressure-gradients on the disturbance dynamics are also included in the analysis Unlike the fully turbulent regime where nonlinear phase scrambling of the fluctuations affects the flow field even in proximity to the wall, the early stage transition regime fluctuations studied here are influenced cross the boundary layer by the solid boundary. The dominating dynamics in the disturbance kinetic energy and dissipation rate equations are described. These results are then used to formulate transition-sensitized turbulent transport equations, which are solved in a two-step process and applied to zero-pressure-gradient flow over a flat plate. Computed results are in good agreement with experimental data.

Implicit Large-Eddy Simulations of Zero-Pressure Gradient, Turbulent Boundary Layer

NASA Technical Reports Server (NTRS)

Sekhar, Susheel; Mansour, Nagi N.

2015-01-01

A set of direct simulations of zero-pressure gradient, turbulent boundary layer flows are conducted using various span widths (62-630 wall units), to document their influence on the generated turbulence. The FDL3DI code that solves compressible Navier-Stokes equations using high-order compact-difference scheme and filter, with the standard recycling/rescaling method of turbulence generation, is used. Results are analyzed at two different Re values (500 and 1,400), and compared with spectral DNS data. They show that a minimum span width is required for the mere initiation of numerical turbulence. Narrower domains ((is) less than 100 w.u.) result in relaminarization. Wider spans ((is) greater than 600 w.u.) are required for the turbulent statistics to match reference DNS. The upper-wall boundary condition for this setup spawns marginal deviations in the mean velocity and Reynolds stress profiles, particularly in the buffer region.

Propagation of propeller tone noise through a fuselage boundary layer

NASA Technical Reports Server (NTRS)

Hanson, D. B.; Magliozzi, B.

1984-01-01

In earlier experimental and analytical studies, it was found that the boundary layer on an aircraft could provide significant shielding from propeller noise at typical transport airplane cruise Mach numbers. In this paper a new three-dimensional theory is described that treats the combined effects of refraction and scattering by the fuselage and boundary layer. The complete wave field is solved by matching analytical expressions for the incident and scattered waves in the outer flow to a numerical solution in the boundary layer flow. The model for the incident waves is a near-field frequency-domain propeller source theory developed previously for free field studies. Calculations for an advanced turboprop (Prop-Fan) model flight test at 0.8 Mach number show a much smaller than expected pressure amplification at the noise directivity peak, strong boundary layer shielding in the forward quadrant, and shadowing around the fuselage. Results are presented showing the difference between fuselage surface and free-space noise predictions as a function of frequency and Mach number. Comparison of calculated and measured effects obtained in a Prop-Fan model flight test show good agreement, particularly near and aft of the plane of rotation at high cruise Mach number.

Flow field survey near the rotational plane of an advanced design propeller on a JetStar airplane

NASA Technical Reports Server (NTRS)

Walsh, K. R.

1985-01-01

An investigation was conducted to obtain upper fuselage surface static pressures and boundary layer velocity profiles below the centerline of an advanced design propeller. This investigation documents the upper fuselage velocity flow field in support of the in-flight acoustic tests conducted on a JetStar airplane. Initial results of the boundary layer survey show evidence of an unusual flow disturbance, which is attributed to the two windshield wiper assemblies on the aircraft. The assemblies were removed, eliminating the disturbances from the flow field. This report presents boundary layer velocity profiles at altitudes of 6096 and 9144 m (20,000 and 30,000 ft) and Mach numbers from 0.6 to 0.8, and it investigated the effects of windshield wiper assemblies on these profiles. Because of the unconventional velocity profiles that were obtained with the assemblies mounted, classical boundary layer parameters, such as momentum and displacement thicknesses, are not presented. The effects of flight test variables (Mach number and angles of attack and sideslip) and an advanced design propeller on boundary layer profiles - with the wiper assemblies mounted and removed - are presented.

A hypersonic aeroheating calculation method based on inviscid outer edge of boundary layer parameters

NASA Astrophysics Data System (ADS)

Meng, ZhuXuan; Fan, Hu; Peng, Ke; Zhang, WeiHua; Yang, HuiXin

2016-12-01

This article presents a rapid and accurate aeroheating calculation method for hypersonic vehicles. The main innovation is combining accurate of numerical method with efficient of engineering method, which makes aeroheating simulation more precise and faster. Based on the Prandtl boundary layer theory, the entire flow field is divided into inviscid and viscid flow at the outer edge of the boundary layer. The parameters at the outer edge of the boundary layer are numerically calculated from assuming inviscid flow. The thermodynamic parameters of constant-volume specific heat, constant-pressure specific heat and the specific heat ratio are calculated, the streamlines on the vehicle surface are derived and the heat flux is then obtained. The results of the double cone show that at the 0° and 10° angle of attack, the method of aeroheating calculation based on inviscid outer edge of boundary layer parameters reproduces the experimental data better than the engineering method. Also the proposed simulation results of the flight vehicle reproduce the viscid numerical results well. Hence, this method provides a promising way to overcome the high cost of numerical calculation and improves the precision.

Transitional and turbulent flat-plate boundary layers with heat transfer

NASA Astrophysics Data System (ADS)

Wu, Xiaohua; Moin, Parviz

2010-11-01

We report on our direct numerical simulation of two incompressible, nominally zero-pressure-gradient flat-plate boundary layers from momentum thickness Reynolds number 80 to 1950. Heat transfer between the constant-temperature solid surface and the free-stream is also simulated with molecular Prandtl number=1. Throughout the entire flat-plate, the ratio of Stanton number and skin-friction St/Cfdeviates from the exact Reynolds analogy value of 0.5 by less than 1.5%. Turbulent Prandtl number t peaks at the wall. Preponderance of hairpin vortices is observed in both the transitional and turbulent regions of the boundary layers. In particular, the internal structure of merged turbulent spots is hairpin forest; the internal structure of infant turbulent spots is hairpin packet. Numerous hairpin vortices are readily detected in both the near-wall and outer regions of the boundary layers up to momentum thickness Reynolds number 1950. This suggests that the hairpin vortices in the turbulent region are not simply the aged hairpin forests convected from the upstream transitional region. Temperature iso-surfaces in the companion thermal boundary layers are found to be a useful tracer in identifying hairpin vortex structures.

Interaction of aerodynamic noise with laminar boundary layers in supersonic wind tunnels

NASA Technical Reports Server (NTRS)

Schopper, M. R.

1984-01-01

The interaction between incoming aerodynamic noise and the supersonic laminar boundary layer is studied. The noise field is modeled as a Mach wave radiation field consisting of discrete waves emanating from coherent turbulent entities moving downstream within the supersonic turbulent boundary layer. The individual disturbances are likened to miniature sonic booms and the laminar boundary layer is staffed by the waves as the sources move downstream. The mean, autocorrelation, and power spectral density of the field are expressed in terms of the wave shapes and their average arrival rates. Some consideration is given to the possible appreciable thickness of the weak shock fronts. The emphasis in the interaction analysis is on the behavior of the shocklets in the noise field. The shocklets are shown to be focused by the laminar boundary layer in its outer region. Borrowing wave propagation terminology, this region is termed the caustic region. Using scaling laws from sonic boom work, focus factors at the caustic are estimated to vary from 2 to 6 for incoming shocklet strengths of 1 to .01 percent of the free stream pressure level. The situation regarding experimental evidence of the caustic region is reviewed.

Effects of surface roughness on an adverse-pressure-gradient separating turbulent boundary layer

NASA Astrophysics Data System (ADS)

Wu, Wen; Piomelli, Ugo; Turbulence Simulation; Modelling Laboratory Team

2017-11-01

Separating turbulent boundary layers over smooth and rough flat plates are investigated by large-eddy simulations. A suction-blowing velocity distribution is imposed at the top boundary to produce an adverse-to-favourable pressure gradient and a closed separation bubble. Sandgrain roughness in the fully-rough regime is modelled by an immersed boundary method. In the rough-wall case, streamline detachment occurs earlier and the separation region is substantially larger due to the momentum deficit caused by the roughness. The adverse pressure gradient decreases the form drag and causes a thin reversed-flow region below the roughness crest, so that Cf = 0 does not coincide with the detachment of the flow from the surface. The wake regions behind roughness elements affect the intermittency of the near-wall flow, so that upstream of the detachment point the flow can be reversed half of the time, but its average is positive. The separated shear layer exhibits higher turbulent kinetic energy (TKE); the growth of the TKE there begins earlier relative to the separation point, and the peak TKE occurs close to the separation point. The momentum deficit caused by the roughness, again, plays a critical role in these changes. The authors acknowledge the support from Hydro-Québec and the NSERC Collaborative Research & Development program (CRDPJ 418786-11). The simulations were performed at CAC Queen't site. UP also thanks the support of Canada Research Chair Program.

Experiments on a smooth wall hypersonic boundary layer at Mach 6

NASA Astrophysics Data System (ADS)

Neeb, Dominik; Saile, Dominik; Gülhan, Ali

2018-04-01

The turbulent boundary layer along the surface of high-speed vehicles drives shear stress and heat flux. Although essential to the vehicle design, the understanding of compressible turbulent boundary layers at high Mach numbers is limited due to the lack of available data. This is particularly true if the surface is rough, which is typically the case for all technical surfaces. To validate a methodological approach, as initial step, smooth wall experiments were performed. A hypersonic turbulent boundary layer at Ma = 6 (Ma_e=5.4) along a 7{}° sharp cone model at low Reynolds numbers Re_{θ } ≈ 3000 was characterized. The mean velocities in the boundary layer were acquired by means of Pitot pressure and particle image velocimetry (PIV) measurements. Furthermore, the PIV data were used to extract turbulent intensities along the profile. The mean velocities in the boundary layer agree with numerical data, independent of the measurement technique. Based on the profile data, three different approaches to extract the skin friction velocity were applied and show favorable comparison to literature and numerical data. The extracted values were used for inner and outer scaling of the van Driest transformed velocity profiles which are in good agreement to incompressible theoretical data. Morkovin scaled turbulent intensities show ambiguous results compared to literature data which may be influenced by inflow turbulence level, particle lag and other measurement uncertainties.

In-Flight Capability for Evaluating Skin-Friction Gages and Other Near-Wall Flow Sensors

NASA Technical Reports Server (NTRS)

Bui, Trong T.; Pipitone, Brett J.; Krake, Keith L.; Richwine, Dave (Technical Monitor)

2003-01-01

An 8-in.-square boundary-layer sensor panel has been developed for in-flight evaluation of skin-friction gages and other near-wall flow sensors on the NASA Dryden Flight Research Center F-15B/Flight Test Fixture (FTF). Instrumentation on the sensor panel includes a boundary-layer rake, temperature sensors, static pressure taps, and a Preston tube. Space is also available for skin-friction gages or other near-wall flow sensors. Pretest analysis of previous F-15B/FTF flight data has identified flight conditions suitable for evaluating skin-friction gages. At subsonic Mach numbers, the boundary layer over the sensor panel closely approximates the two-dimensional (2D), law-of-the-wall turbulent boundary layer, and skin-friction estimates from the Preston tube and the rake (using the Clauser plot method) can be used to evaluate skin-friction gages. At supersonic Mach numbers, the boundary layer over the sensor panel becomes complex, and other means of measuring skin friction are needed to evaluate the accuracy of new skin-friction gages. Results from the flight test of a new rubber-damped skin-friction gage confirm that at subsonic Mach numbers, nearly 2D, law-of-the-wall turbulent boundary layers exist over the sensor panel. Sensor panel data also show that this new skin-friction gage prototype does not work in flight.

Visualization of the structural response of a hypersonic turbulent boundary layer to convex curvature

NASA Astrophysics Data System (ADS)

Humble, R. A.; Peltier, S. J.; Bowersox, R. D. W.

2012-10-01

The effects of convex curvature on the outer structure of a Mach 4.9 turbulent boundary layer (Reθ = 4.7 × 104) are investigated using condensate Rayleigh scattering and analyzed using spatial correlations, intermittency, and fractal theory. It is found that the post-expansion boundary layer structure morphology appears subtle, but certain features exhibit a more obvious response. The large-scale flow structures survive the initial expansion, appearing to maintain the same physical size. However, due to the nature of the expansion fan, a differential acceleration effect takes place across the flow structures, causing them to be reoriented, leaning farther away from the wall. The onset of intermittency moves closer towards the boundary layer edge and the region of intermittent flow decreases. It is likely that this reflects the less frequent penetration of outer irrotational fluid into the boundary layer, consistent with a boundary layer that is losing its ability to entrain freestream fluid. The fractal dimension of the turbulent/nonturbulent interface decreases with increasing favorable pressure gradient, indicating that the interface's irregularity decreases. Because fractal scale similarity does not encompass the largest scales, this suggests that the change in fractal dimension is due to the action of the smaller-scales, consistent with the idea that the small-scale flow structures are quenched during the expansion in response to bulk dilatation.

Sustained drag reduction in a turbulent flow using a low-temperature Leidenfrost surface

PubMed Central

Saranadhi, Dhananjai; Chen, Dayong; Kleingartner, Justin A.; Srinivasan, Siddarth; Cohen, Robert E.; McKinley, Gareth H.

2016-01-01

Skin friction drag contributes a major portion of the total drag for small and large water vehicles at high Reynolds number (Re). One emerging approach to reducing drag is to use superhydrophobic surfaces to promote slip boundary conditions. However, the air layer or “plastron” trapped on submerged superhydrophobic surfaces often diminishes quickly under hydrostatic pressure and/or turbulent pressure fluctuations. We use active heating on a superhydrophobic surface to establish a stable vapor layer or “Leidenfrost” state at a relatively low superheat temperature. The continuous film of water vapor lubricates the interface, and the resulting slip boundary condition leads to skin friction drag reduction on the inner rotor of a custom Taylor-Couette apparatus. We find that skin friction can be reduced by 80 to 90% relative to an unheated superhydrophobic surface for Re in the range 26,100 ≤ Re ≤ 52,000. We derive a boundary layer and slip theory to describe the hydrodynamics in the system and show that the plastron thickness is h = 44 ± 11 μm, in agreement with expectations for a Leidenfrost surface. PMID:27757417

On the aeroacoustic and flow structures developed on a flat plate with a serrated sawtooth trailing edge

NASA Astrophysics Data System (ADS)

Chong, Tze Pei; Vathylakis, Alexandros

2015-10-01

Results of an experimental study on turbulent flow over a flat plate with a serrated sawtooth trailing edge are presented in this paper. After tripping the boundary layer to become turbulent, the broadband noise sources at the sawtooth serrated trailing edge is studied by several experimental techniques. Broadband noise reduction by the serrated sawtooth trailing edge can be realistically achieved in the flat plate configuration. The variations of wall pressure power spectral density and the spanwise coherence (which relates to the spanwise correlation length) in a sawtooth trailing edge play a minor role in the mechanisms underpinning the reduction of self noise radiation. Conditional-averaging technique was applied in the boundary layer data where a pair of pressure-driven oblique vortical structures near the sawtooth side edges is identified. In the current flat plate configuration, the interaction between the vortical structures and the local turbulent boundary layer results in a redistribution of the momentum transport and turbulent shear stress near the sawtooth side edges as well as the sawtooth tip, thus affecting the efficiency of self noise radiation.

Circulation Control Model Experimental Database for CFD Validation

NASA Technical Reports Server (NTRS)

Paschal, Keith B.; Neuhart, Danny H.; Beeler, George B.; Allan, Brian G.

2012-01-01

A 2D circulation control wing was tested in the Basic Aerodynamic Research Tunnel at the NASA Langley Research Center. A traditional circulation control wing employs tangential blowing along the span over a trailing-edge Coanda surface for the purpose of lift augmentation. This model has been tested extensively at the Georgia Tech Research Institute for the purpose of performance documentation at various blowing rates. The current study seeks to expand on the previous work by documenting additional flow-field data needed for validation of computational fluid dynamics. Two jet momentum coefficients were tested during this entry: 0.047 and 0.114. Boundary-layer transition was investigated and turbulent boundary layers were established on both the upper and lower surfaces of the model. Chordwise and spanwise pressure measurements were made, and tunnel sidewall pressure footprints were documented. Laser Doppler Velocimetry measurements were made on both the upper and lower surface of the model at two chordwise locations (x/c = 0.8 and 0.9) to document the state of the boundary layers near the spanwise blowing slot.

Improvements to the George/Castillo Boundary Layer Theory

NASA Astrophysics Data System (ADS)

Wosnik, Martin; George, William K.; Castillo, Luciano

2000-11-01

George and Castillo (1997)(George WK and Castillo L (1997) Appl.Mech.Rev.), 50, 12/1, 689-729. presented a new theory for Zero Pressure Gradient Turbulent Boundary Layers based on an application of Near-Asymptotics to scaling laws derived from equilibrium similarity to the Reynolds-averaged equations. The resulting overlap velocity profiles retained a dependence on local Reynolds number, the parameters for which had to satisfy the following constraint equation: ln \\varepsilon fracdγd ln δ^+ = fracdln [C_o/C_i] d ln δ^+ where γ is the power exponent, Co and Ci are the coefficients in inner and outer variables respectively. GC considered only the first term in an asymptotic expansion of the exact solution, but higher order terms can be considered with no increase in the number of unknowns. The improved theory is tested against new experimental Zero Pressure Gradient Turbulent Boundary Layer data of Smith (1994), Oesterlund (1999) and Johansson and Castillo (2000). For the friction law, the first order term is sufficient, but for Co and γ the higher order terms improve the fit to the velocity profiles significantly.

Characterization of structural response to hypersonic boundary-layer transition

DOE PAGES

Riley, Zachary B.; Deshmukh, Rohit; Miller, Brent A.; ...

2016-05-24

The inherent relationship between boundary-layer stability, aerodynamic heating, and surface conditions makes the potential for interaction between the structural response and boundary-layer transition an important and challenging area of study in high-speed flows. This paper phenomenologically explores this interaction using a fundamental two-dimensional aerothermoelastic model under the assumption of an aluminum panel with simple supports. Specifically, an existing model is extended to examine the impact of transition onset location, transition length, and transitional overshoot in heat flux and fluctuating pressure on the structural response of surface panels. Transitional flow conditions are found to yield significantly increased thermal gradients, and theymore » can result in higher maximum panel temperatures compared to turbulent flow. Results indicate that overshoot in heat flux and fluctuating pressure reduces the flutter onset time and increases the strain energy accumulated in the panel. Furthermore, overshoot occurring near the midchord can yield average temperatures and peak displacements exceeding those experienced by the panel subject to turbulent flow. Lastly, these results suggest that fully turbulent flow does not always conservatively predict the thermo-structural response of surface panels.« less

Correlation of transonic-cone Preston-tube data and skin friction. [characterizing the flow quality of a transonic wind tunnel

NASA Technical Reports Server (NTRS)

Reed, T. D.

1981-01-01

The distribution of Preston tube pressures within turbulent boundary layers along the surface of a sharp-nosed, ten degree cone was correlated with theoretical values of turbulent skin friction for freestream Mach numbers less than one. The mini-basic computer code, the Wu and Lock computer code, and the STAN-5 computer code were used to analyze the data and to solve the boundary layer conservation equations. The skin friction which results from using Preston tube pressures in the correlation equation, has a rms error of 1.125 percent. It was found that the effective center of the probe is not a constant but increases as the surface distance increases. For a specified unit Reynolds number, the effective center of the probe decreases as the Mach number increases. The variation of the fluid (air) properties across the face of the probe may be neglected for subsonic flows. The possible transverse errors caused by the use of the concept of a virtual origin for the turbulent boundary layer were investigated and found to be negligible.

Flight test results of riblets at supersonic speeds

NASA Technical Reports Server (NTRS)

Zuniga, Fanny A.; Anderson, Bianca T.; Bertelrud, Arild

1992-01-01

A flight experiment to test and evaluate the skin friction drag characteristics of a riblet surface in turbulent flow at supersonic speeds was conducted at NASA Dryden. Riblets of groove sizes 0.0030 and 0.0013 in. were mounted on the F-104G flight test fixture. The test surfaces were surveyed with boundary layer rakes and pressure orifices to examine the boundary layer profiles and pressure distributions of the flow. Skin friction reductions caused by the riblet surface were reported based on measured differences of momentum thickness between the smooth and riblet surfaces obtained from the boundary layer data. Flight test results for the 0.0030 in. riblet show skin friction reductions of 4 to 8 % for Mach numbers ranging from 1.2 to 1.6 and Reynolds numbers ranging from 2 to 3.4 million per unit foot. The results from the 0.0013 in. riblets show skin friction reductions of 4 to 15 % for Mach 1.2 to 1.4 and Reynolds numbers ranging from 3.6 to 6 million per unit foot.

Effects of cavity size on the control of transonic internal flow around a biconvex circular arc airfoil

NASA Astrophysics Data System (ADS)

Rahman, M. Mostaqur; Hasan, A. B. M. Toufique; Rabbi, M. S.

2017-06-01

In transonic flow conditions, self-sustained shock wave oscillation on biconvex airfoils is initiated by the complex shock wave boundary layer interaction which is frequently observed in several modern internal aeronautical applications such as inturbine cascades, compressor blades, butterfly valves, fans, nozzles, diffusers and so on. Shock wave boundary layer interaction often generates serious problems such as unsteady boundary layer separation, self-excited shock waveoscillation with large pressure fluctuations, buffeting excitations, aeroacoustic noise, nonsynchronous vibration, high cycle fatigue failure and intense drag rise. Recently, the control of the self-excited shock oscillation around an airfoil using passive control techniques is getting intense interest. Among the passive means, control using open cavity has found promising. In this study, the effect of cavity size on the control of self-sustained shock oscillation was investigated numerically. The present computations are validated with available experimental results. The results showed that the average root mean square (RMS) of pressure oscillation around the airfoil with open cavity has reduced significantly when compared to airfoil without cavity (clean airfoil).

Effect of Blowing on Boundary Layer of Scarf Inlet

NASA Technical Reports Server (NTRS)

Gerhold, Carl H.; Clark, Lorenzo R.

2004-01-01

When aircraft operate in stationary or low speed conditions, airflow into the engine accelerates around the inlet lip and pockets of turbulence that cause noise and vibration can be ingested. This problem has been encountered with engines equipped with the scarf inlet, both in full scale and in model tests, where the noise produced during the static test makes it difficult to assess the noise reduction performance of the scarf inlet. NASA Langley researchers have implemented boundary layer control in an attempt to reduce the influence of the flow nonuniformity in a 12-in. diameter model of a high bypass fan engine mounted in an anechoic chamber. Static pressures and boundary layer profiles were measured in the inlet and far field acoustic measurements were made to assess the effectiveness of the blowing treatment. The blowing system was found to lack the authority to overcome the inlet distortions. Methods to improve the implementation of boundary layer control to reduce inlet distortion are discussed.

Analytical and Experimental Evaluation of the Heat Transfer Distribution over the Surfaces of Turbine Vanes

NASA Technical Reports Server (NTRS)

Hylton, L. D.; Mihelc, M. S.; Turner, E. R.; Nealy, D. A.; York, R. E.

1983-01-01

Three airfoil data sets were selected for use in evaluating currently available analytical models for predicting airfoil surface heat transfer distributions in a 2-D flow field. Two additional airfoils, representative of highly loaded, low solidity airfoils currently being designed, were selected for cascade testing at simulated engine conditions. Some 2-D analytical methods were examined and a version of the STAN5 boundary layer code was chosen for modification. The final form of the method utilized a time dependent, transonic inviscid cascade code coupled to a modified version of the STAN5 boundary layer code featuring zero order turbulence modeling. The boundary layer code is structured to accommodate a full spectrum of empirical correlations addressing the coupled influences of pressure gradient, airfoil curvature, and free-stream turbulence on airfoil surface heat transfer distribution and boundary layer transitional behavior. Comparison of pedictions made with the model to the data base indicates a significant improvement in predictive capability.

An improved viscid/inviscid interaction procedure for transonic flow over airfoils

NASA Technical Reports Server (NTRS)

Melnik, R. E.; Chow, R. R.; Mead, H. R.; Jameson, A.

1985-01-01

A new interacting boundary layer approach for computing the viscous transonic flow over airfoils is described. The theory includes a complete treatment of viscous interaction effects induced by the wake and accounts for normal pressure gradient effects across the boundary layer near trailing edges. The method is based on systematic expansions of the full Reynolds equation of turbulent flow in the limit of Reynolds numbers, Reynolds infinity. Procedures are developed for incorporating the local trailing edge solution into the numerical solution of the coupled full potential and integral boundary layer equations. Although the theory is strictly applicable to airfoils with cusped or nearly cusped trailing edges and to turbulent boundary layers that remain fully attached to the airfoil surface, the method was successfully applied to more general airfoils and to flows with small separation zones. Comparisons of theoretical solutions with wind tunnel data indicate the present method can accurately predict the section characteristics of airfoils including the absolute levels of drag.

Analytical and experimental evaluation of the heat transfer distribution over the surfaces of turbine vanes

NASA Astrophysics Data System (ADS)

Hylton, L. D.; Mihelc, M. S.; Turner, E. R.; Nealy, D. A.; York, R. E.

1983-05-01

Three airfoil data sets were selected for use in evaluating currently available analytical models for predicting airfoil surface heat transfer distributions in a 2-D flow field. Two additional airfoils, representative of highly loaded, low solidity airfoils currently being designed, were selected for cascade testing at simulated engine conditions. Some 2-D analytical methods were examined and a version of the STAN5 boundary layer code was chosen for modification. The final form of the method utilized a time dependent, transonic inviscid cascade code coupled to a modified version of the STAN5 boundary layer code featuring zero order turbulence modeling. The boundary layer code is structured to accommodate a full spectrum of empirical correlations addressing the coupled influences of pressure gradient, airfoil curvature, and free-stream turbulence on airfoil surface heat transfer distribution and boundary layer transitional behavior. Comparison of pedictions made with the model to the data base indicates a significant improvement in predictive capability.

Study of flow control by localized volume heating in hypersonic boundary layers

NASA Astrophysics Data System (ADS)

Keller, M. A.; Kloker, M. J.; Kirilovskiy, S. V.; Polivanov, P. A.; Sidorenko, A. A.; Maslov, A. A.

2014-12-01

Boundary-layer flow control is a prerequisite for a safe and efficient operation of future hypersonic transport systems. Here, the influence of an electric discharge—modeled by a heat-source term in the energy equation—on laminar boundary-layer flows over a flat plate with zero pressure gradient at Mach 3, 5, and 7 is investigated numerically. The aim was to appraise the potential of electro-gasdynamic devices for an application as turbulence generators in the super- and hypersonic flow regime. The results with localized heat-source elements in boundary layers are compared to cases with roughness elements serving as classical passive trips. The numerical simulations are performed using the commercial code ANSYS FLUENT (by ITAM) and the high-order finite-difference DNS code NS3D (by IAG), the latter allowing for the detailed analysis of laminar flow instability. For the investigated setups with steady heating, transition to turbulence is not observed, due to the Reynolds-number lowering effect of heating.

Thermodynamic evaluation of transonic compressor rotors using the finite volume approach

NASA Technical Reports Server (NTRS)

Nicholson, S.; Moore, J.

1986-01-01

A method was developed which calculates two-dimensional, transonic, viscous flow in ducts. The finite volume, time marching formulation is used to obtain steady flow solutions of the Reynolds-averaged form of the Navier Stokes equations. The entire calculation is performed in the physical domain. The method is currently limited to the calculation of attached flows. The features of the current method can be summarized as follows. Control volumes are chosen so that smoothing of flow properties, typically required for stability, is now needed. Different time steps are used in the different governing equations to improve the convergence speed of the viscous calculations. A new pressure interpolation scheme is introduced which improves the shock capturing ability of the method. A multi-volume method for pressure changes in the boundary layer allows calculations which use very long and thin control volumes. A special discretization technique is also used to stabilize these calculations. A special formulation of the energy equation is used to provide improved transient behavior of solutions which use the full energy equation. The method is then compared with a wide variety of test cases. The freestream Mach numbers range from 0.075 to 2.8 in the calculations. Transonic viscous flow in a converging diverging nozzle is calculated with the method; the Mach number upstream of the shock is approximately 1.25. The agreement between the calculated and measured shock strength and total pressure losses is good. Essentially incompressible turbulent boundary layer flow in a adverse pressure gradient is calculated and the computed distribution of mean velocity and shear stress are in good agreement with the measurements. At the other end of the Mach number range, a flat plate turbulent boundary layer with a freestream Mach number of 2.8 is calculated using the full energy equation; the computed total temperature distribution and recovery factor agree well with the measurements when a variable Prandtl number is used through the boundary layer.

Using Temperature Sensitive Paint Technology

NASA Technical Reports Server (NTRS)

Hamner, M. P.; Popernack, T. G., Jr.; Owens, L. R.; Wahls, R. A.

2002-01-01

New facilities and test techniques afford research aerodynamicists many opportunities to investigate complex aerodynamic phenomena. For example, NASA Langley Research Center's National Transonic Facility (NTF) can hold Mach number, Reynolds number, dynamic pressure, stagnation temperature and stagnation pressure constant during testing. This is important because the wing twist associated with model construction may mask important Reynolds number effects associated with the flight vehicle. Beyond this, the NTF's ability to vary Reynolds number allows for important research into the study of boundary layer transition. The capabilities of facilities such as the NTF coupled with test techniques such as temperature sensitive paint yield data that can be applied not only to vehicle design but also to validation of computational methods. Development of Luminescent Paint Technology for acquiring pressure and temperature measurements began in the mid-1980s. While pressure sensitive luminescent paints (PSP) were being developed to acquire data for aerodynamic performance and loads, temperature sensitive luminescent paints (TSP) have been used for a much broader range of applications. For example, TSP has been used to acquire surface temperature data to determine the heating due to rotating parts in various types of mechanical systems. It has been used to determine the heating pattern(s) on circuit boards. And, it has been used in boundary layer analysis and applied to the validation of full-scale flight performance predictions. That is, data acquired on the same model can be used to develop trends from off design to full scale flight Reynolds number, e.g. to show the progression of boundary layer transition. A discussion of issues related to successfully setting-up TSP tests and using TSP systems for boundary layer studies is included in this paper, as well as results from a variety of TSP tests. TSP images included in this paper are all grey-scale so that similar to pictures from sublimating chemical tests areas of laminar flow appear "lighter," or white, and areas of turbulent flow appear "darker."

Diamond anvil cell for spectroscopic investigation of materials at high temperature, high pressure and shear

DOEpatents

Westerfield, C.L.; Morris, J.S.; Agnew, S.F.

1997-01-14

Diamond anvil cell is described for spectroscopic investigation of materials at high temperature, high pressure and shear. A cell is described which, in combination with Fourier transform IR spectroscopy, permits the spectroscopic investigation of boundary layers under conditions of high temperature, high pressure and shear. 4 figs.

Boundary-Layer-Ingesting Inlet Flow Control

NASA Technical Reports Server (NTRS)

Owens, Lewis R.; Allan, Brian G.; Gorton, Susan A.

2006-01-01

This paper gives an overview of a research study conducted in support of the small-scale demonstration of an active flow control system for a boundary-layer-ingesting (BLI) inlet. The effectiveness of active flow control in reducing engine inlet circumferential distortion was assessed using a 2.5% scale model of a 35% boundary-layer-ingesting flush-mounted, offset, diffusing inlet. This experiment was conducted in the NASA Langley 0.3-meter Transonic Cryogenic Tunnel at flight Mach numbers with a model inlet specifically designed for this type of testing. High mass flow actuators controlled the flow through distributed control jets providing the active flow control. A vortex generator point design configuration was also tested for comparison purposes and to provide a means to examine a hybrid vortex generator and control jets configuration. Measurements were made of the onset boundary layer, the duct surface static pressures, and the mass flow through the duct and the actuators. The distortion and pressure recovery were determined by 40 total pressure measurements on 8 rake arms each separated by 45 degrees and were located at the aerodynamic interface plane. The test matrix was limited to a maximum free-stream Mach number of 0.85 with scaled mass flows through the inlet for that condition. The data show that the flow control jets alone can reduce circumferential distortion (DPCP(sub avg)) from 0.055 to about 0.015 using about 2.5% of inlet mass flow. The vortex generators also reduced the circumferential distortion from 0.055 to 0.010 near the inlet mass flow design point. Lower inlet mass flow settings with the vortex generator configuration produced higher distortion levels that were reduced to acceptable levels using a hybrid vortex generator/control jets configuration that required less than 1% of the inlet mass flow.

Boundary-Layer-Ingesting Inlet Flow Control

NASA Technical Reports Server (NTRS)

Owens, Lewis R.; Allan, Brian G.; Gorton, Susan A.

2006-01-01

This paper gives an overview of a research study conducted in support of the small-scale demonstration of an active flow control system for a boundary-layer-ingesting (BLI) inlet. The effectiveness of active flow control in reducing engine inlet circumferential distortion was assessed using a 2.5% scale model of a 35% boundary-layer-ingesting flush-mounted, offset, diffusing inlet. This experiment was conducted in the NASA Langley 0.3-meter Transonic Cryogenic Tunnel at flight Mach numbers with a model inlet specifically designed for this type of testing. High mass flow actuators controlled the flow through distributed control jets providing the active flow control. A vortex generator point design configuration was also tested for comparison purposes and to provide a means to examine a hybrid vortex generator and control jets configuration. Measurements were made of the onset boundary layer, the duct surface static pressures, and the mass flow through the duct and the actuators. The distortion and pressure recovery were determined by 40 total pressure measurements on 8 rake arms each separated by 45 degrees and were located at the aerodynamic interface plane. The test matrix was limited to a maximum free-stream Mach number of 0.85 with scaled mass flows through the inlet for that condition. The data show that the flow control jets alone can reduce circumferential distortion (DPCPavg) from 0.055 to about 0.015 using about 2.5% of inlet mass flow. The vortex generators also reduced the circumferential distortion from 0.055 to 0.010 near the inlet mass flow design point. Lower inlet mass flow settings with the vortex generator configuration produced higher distortion levels that were reduced to acceptable levels using a hybrid vortex generator/control jets configuration that required less than 1% of the inlet mass flow.

Hybrid finite-difference/lattice Boltzmann simulations of microchannel and nanochannel acoustic streaming driven by surface acoustic waves

NASA Astrophysics Data System (ADS)

Tan, Ming K.; Yeo, Leslie Y.

2018-04-01

A two-dimensional hybrid numerical method that allows full coupling of the elastic motion in a piezoelectric solid (modeled using a finite-difference time-domain technique) with the resultant compressional flow in a fluid (simulated using a lattice Boltzmann scheme) is developed to study the acoustic streaming that arises in both microchannels and nanochannels under surface acoustic wave (SAW) excitation. In addition to verifying the model through a comparison of the simulations with results from experimental and numerical studies of microchannel and nanochannel flows driven by both standing and traveling SAWs in the literature, we highlight salient features of the flow field that arise and discuss the underlying mechanisms responsible for the flow. In microchannels, boundary layer streaming is the dominant mechanism when the channel height is below the sound wavelength in the liquid, whereas Eckart streaming—arising as a consequence of the attenuation of the sound wave in the liquid—dominates in the form of periodic vortices for larger channel heights. The absence of Eckart streaming and the overlapping of boundary layers in nanochannels with heights below the boundary layer thickness, on the other hand, give rise to a time-averaged dynamic acoustic pressure that results in an inertial-dominant flow, which paradoxically possesses a parabolic-like velocity profile resembling pressure-driven laminar flow. In contrast, if the nanochannel were to be filled instead with air, the significantly lower fluid density leads to a considerable reduction in the dynamic acoustic pressure and hence inertial forcing such that boundary layer streaming once again dominates, asymptotically imposing a slip condition along the channel surface that results in a negative pluglike velocity profile.

Design and analysis of axial aspirated compressor stages

NASA Astrophysics Data System (ADS)

Merchant, Ali A.

The pressure ratio of axial compressor stages can be significantly increased by controlling the development of blade and endwall boundary layers in regions of adverse pressure gradient by means of boundary layer suction. This concept is validated and demonstrated through the design and analysis of two unique aspirated compressor stages: a low-speed stage with a design pressure ratio of 1.6 at a tip speed of 750 ft/s, and a high-speed stage with a design pressure ratio of 3.5 at a tip speed of 1500 ft/s. The aspirated compressor stages were designed using a new procedure which is a synthesis of low speed and high speed blade design techniques combined with a flexible inverse design method which enabled precise independent control over the shape of the blade suction and pressure surfaces. Integration of the boundary layer suction calculation into the overall design process is an essential ingredient of the new procedure. The blade design system consists of two axisymmetric through-flow codes coupled with a quasi three-dimensional viscous cascade plane code with inverse design capability. Validation of the completed designs were carried out with three-dimensional Euler and Navier-Stokes calculations. A single spanwise slot on the blade suction surface is used to bleed the boundary layer. The suction mass flow requirement for the low-speed and high-speed stages are 1% and 4% of the inlet mass flow, respectively. Additional suction between 1-2% is also required on the compressor endwalls near shock impingement locations. The rotor is modeled with a tip shroud to eliminate tip clearance effects and to discharge the suction flow radially from the flowpath. Three-dimensional viscous evaluation of the designs showed good agreement with the quasi three-dimensional design intent, except in the endwall regions. The suction requirements predicted by the quasi three-dimensional calculation were confirmed by the three-dimensional viscous calculations. The three-dimensional viscous analysis predicted a peak pressure ratio of 1.59 at an isentropic efficiency of 89% for the low-speed stage, and a peak pressure ratio of 3.68 at an isentropic efficiency of 94% for the high-speed rotor. (Copies available exclusively from MIT Libraries, Rm. 14-0551, Cambridge, MA 02139-4307. Ph. 617-253-5668; Fax 617-253-1690.)

Computational and experimental investigation of two-dimensional scramjet inlets and hypersonic flow over a sharp flat plate

NASA Astrophysics Data System (ADS)

Messitt, Donald G.

1999-11-01

The WIND code was employed to compute the hypersonic flow in the shock wave boundary layer merged region near the leading edge of a sharp flat plate. Solutions were obtained at Mach numbers from 9.86 to 15.0 and free stream Reynolds numbers of 3,467 to 346,700 in-1 (1.365 · 105 to 1.365 · 107 m-1) for perfect gas conditions. The numerical results indicated a merged shock wave and viscous layer near the leading edge. The merged region grew in size with increasing free stream Mach number, proportional to Minfinity 2/Reinfinity. Profiles of the static pressure in the merged region indicated a strong normal pressure gradient (∂p/∂y). The normal pressure gradient has been neglected in previous analyses which used the boundary layer equations. The shock wave near the leading edge was thick, as has been experimentally observed. Computed shock wave locations and surface pressures agreed well within experimental error for values of the rarefaction parameter, chi/M infinity2 < 0.3. A preliminary analysis using kinetic theory indicated that rarefied flow effects became important above this value. In particular, the WIND solution agreed well in the transition region between the merged flow, which was predicted well by the theory of Li and Nagamatsu, and the downstream region where the strong interaction theory applied. Additional computations with the NPARC code, WIND's predecessor, demonstrated the ability of the code to compute hypersonic inlet flows at free stream Mach numbers up to 20. Good qualitative agreement with measured pressure data indicated that the code captured the important physical features of the shock wave - boundary layer interactions. The computed surface and pitot pressures fell within the combined experimental and numerical error bounds for most points. The calculations demonstrated the need for extremely fine grids when computing hypersonic interaction flows.

An integral wall model for Large Eddy Simulation (iWMLES) and applications to developing boundary layers over smooth and rough plates

NASA Astrophysics Data System (ADS)

Yang, Xiang; Sadique, Jasim; Mittal, Rajat; Meneveau, Charles

2014-11-01

A new wall model for Large-Eddy-Simulations is proposed. It is based on an integral boundary layer method that assumes a functional form for the local mean velocity profile. The method, iWMLES, evaluates required unsteady and advective terms in the vertically integrated boundary layer equations analytically. The assumed profile contains a viscous or roughness sublayer, and a logarithmic layer with an additional linear term accounting for inertial and pressure gradient effects. The iWMLES method is tested in the context of a finite difference LES code. Test cases include developing turbulent boundary layers on a smooth flat plate at various Reynolds numbers, over flat plates with unresolved roughness, and a sample application to boundary layer flow over a plate that includes resolved roughness elements. The elements are truncated cones acting as idealized barnacle-like roughness elements that often occur in biofouling of marine surfaces. Comparisons with data show that iWMLES provides accurate predictions of near-wall velocity profiles in LES while, similarly to equilibrium wall models, its cost remains independent of Reynolds number and is thus significantly lower compared to standard zonal or hybrid wall models. This work is funded by ONR Grant N00014-12-1-0582 (Dr. R. Joslin, program manager).

Hybrid Large Eddy Simulation / Reynolds Averaged Navier-Stokes Modeling in Directed Energy Applications

NASA Astrophysics Data System (ADS)

Zilberter, Ilya Alexandrovich

In this work, a hybrid Large Eddy Simulation / Reynolds-Averaged Navier Stokes (LES/RANS) turbulence model is applied to simulate two flows relevant to directed energy applications. The flow solver blends the Menter Baseline turbulence closure near solid boundaries with a Lenormand-type subgrid model in the free-stream with a blending function that employs the ratio of estimated inner and outer turbulent length scales. A Mach 2.2 mixing nozzle/diffuser system representative of a gas laser is simulated under a range of exit pressures to assess the ability of the model to predict the dynamics of the shock train. The simulation captures the location of the shock train responsible for pressure recovery but under-predicts the rate of pressure increase. Predicted turbulence production at the wall is found to be highly sensitive to the behavior of the RANS turbulence model. A Mach 2.3, high-Reynolds number, three-dimensional cavity flow is also simulated in order to compute the wavefront aberrations of an optical beam passing thorough the cavity. The cavity geometry is modeled using an immersed boundary method, and an auxiliary flat plate simulation is performed to replicate the effects of the wind-tunnel boundary layer on the computed optical path difference. Pressure spectra extracted on the cavity walls agree with empirical predictions based on Rossiter's formula. Proper orthogonal modes of the wavefront aberrations in a beam originating from the cavity center agree well with experimental data despite uncertainty about in flow turbulence levels and boundary layer thicknesses over the wind tunnel window. Dynamic mode decomposition of a planar wavefront spanning the cavity reveals that wavefront distortions are driven by shear layer oscillations at the Rossiter frequencies; these disturbances create eddy shocklets that propagate into the free-stream, creating additional optical wavefront distortion.

Off-wall boundary conditions for turbulent flows obtained from buffer-layer minimal flow units

NASA Astrophysics Data System (ADS)

Garcia-Mayoral, Ricardo; Pierce, Brian; Wallace, James

2012-11-01

There is strong evidence that the transport processes in the buffer region of wall-bounded turbulence are common across various flow configurations, even in the embryonic turbulence in transition (Park et al., Phys. Fl. 24). We use this premise to develop off-wall boundary conditions for turbulent simulations. Boundary conditions are constructed from DNS databases using periodic minimal flow units and reduced order modeling. The DNS data was taken from a channel at Reτ = 400 and a zero-pressure gradient transitional boundary layer (Sayadi et al., submitted to J . FluidMech .) . Both types of boundary conditions were first tested on a DNS of the core of the channel flow with the aim of extending their application to LES and to spatially evolving flows. 2012 CTR Summer Program.

The three-dimensional turbulent boundary layer near a plane of symmetry

NASA Technical Reports Server (NTRS)

Degani, A. T.; Smith, F. T.; Walker, J. D. A.

1992-01-01

The asymptotic structure of the three-dimensional turbulent boundary layer near a plane of symmetry is considered in the limit of large Reynolds number. A self-consistent two-layer structure is shown to exist wherein the streamwise velocity is brought to rest through an outer defect layer and an inner wall layer in a manner similar to that in two-dimensional boundary layers. The cross-stream velocity distribution is more complex and two terms in the asymptotic expansion are required to yield a complete profile which is shown to exhibit a logarithmic region. The flow in the inner wall layer is demonstrated to be collateral to leading order; pressure-gradient effects are formally of higher order but can cause the velocity profile to skew substantially near the wall at the large but finite Reynolds numbers encountered in practice. The governing set of ordinary differential equations describing a self-similar flow is derived. The calculated numerical solutions of these equations are matched asymptotically to an inner wall-layer solution and the results show trends that are consistent with experimental observations.

Slippage and boundary layer probed in an almost ideal gas by a nanomechanical oscillator.

PubMed

Defoort, M; Lulla, K J; Crozes, T; Maillet, O; Bourgeois, O; Collin, E

2014-09-26

We measure the interaction between ⁴He gas at 4.2 K and a high-quality nanoelectromechanical string device for its first three symmetric modes (resonating at 2.2, 6.7, and 11 MHz with quality factor Q>0.1×10⁶) over almost 6 orders of magnitude in pressure. This fluid can be viewed as the best experimental implementation of an almost ideal monoatomic and inert gas of which properties are tabulated. The experiment ranges from high pressure where the flow is of laminar Stokes-type presenting slippage down to very low pressures where the flow is molecular. In the molecular regime, when the mean-free path is of the order of the distance between the suspended nanomechanical probe and the bottom of the trench, we resolve for the first time the signature of the boundary (Knudsen) layer onto the measured dissipation. Our results are discussed in the framework of the most recent theories investigating boundary effects in fluids (both analytic approaches and direct simulation Monte Carlo methods).

Experimental Investigation of 'Transonic Resonance' with Convergent-Divergent Nozzles

NASA Technical Reports Server (NTRS)

Zaman, K. B. M. Q.; Dahl, M. D.; Bencic, T. J.; Zaman, Khairul (Technical Monitor)

2001-01-01

Convergent-divergent nozzles, when run at pressure ratios lower than the design value, often undergo a flow resonance accompanied by the emission of acoustic tones. The phenomenon, different in characteristics from conventional 'screech' tones, has been studied experimentally. Unlike screech, the frequency increases with increasing supply pressure. There is a 'staging' behavior; 'odd harmonic' stages resonate at lower pressures while the fundamental occurs in a range of higher pressures corresponding to a fully expanded Mach number (M(sub j)) around unity. The frequency (f(sub N)) variation with M(sub j) depends on the half angle-of-divergence (theta) of the nozzle. At smaller theta, the slope of f(sub N) versus M(sub j) curve becomes steeper. The resonance involves standing waves and is driven by unsteady shock/boundary layer interaction. The distance between the foot of the shock and the nozzle exit imposes the lengthscale (L'). The fundamental corresponds to a quarterwave resonance, the next stage at a lower supply pressure corresponds to a three-quarter-wave resonance, and so on. The principal trends in the frequency variation are explained simply from the characteristic variation of the length-scale L'. Based on the data, correlation equations are provided for the prediction of f(sub N). A striking feature is that tripping of the boundary layer near the nozzle's throat tends to suppress the resonance. In a practical nozzle a tendency for the occurrence of the phenomenon is thought to be a source of 'internal noise'; thus, there is a potential for noise benefit simply by appropriate boundary layer tripping near the nozzle's throat.

Computation of three-dimensional compressible boundary layers to fourth-order accuracy on wings and fuselages

NASA Technical Reports Server (NTRS)

Iyer, Venkit

1990-01-01

A solution method, fourth-order accurate in the body-normal direction and second-order accurate in the stream surface directions, to solve the compressible 3-D boundary layer equations is presented. The transformation used, the discretization details, and the solution procedure are described. Ten validation cases of varying complexity are presented and results of calculation given. The results range from subsonic flow to supersonic flow and involve 2-D or 3-D geometries. Applications to laminar flow past wing and fuselage-type bodies are discussed. An interface procedure is used to solve the surface Euler equations with the inviscid flow pressure field as the input to assure accurate boundary conditions at the boundary layer edge. Complete details of the computer program used and information necessary to run each of the test cases are given in the Appendix.

Application of dynamic slip wall modeling to a turbine nozzle guide vane

NASA Astrophysics Data System (ADS)

Bose, Sanjeeb; Talnikar, Chaitanya; Blonigan, Patrick; Wang, Qiqi

2015-11-01

Resolution of near-wall turbulent structures is computational prohibitive necessitating the need for wall-modeled large-eddy simulation approaches. Standard wall models are often based on assumptions of equilibrium boundary layers, which do not necessarily account for the dissimilarity of the momentum and thermal boundary layers. We investigate the use of the dynamic slip wall boundary condition (Bose and Moin, 2014) for the prediction of surface heat transfer on a turbine nozzle guide vane (Arts and de Rouvroit, 1992). The heat transfer coefficient is well predicted by the slip wall model, including capturing the transition to turbulence. The sensitivity of the heat transfer coefficient to the incident turbulence intensity will additionally be discussed. Lastly, the behavior of the thermal and momentum slip lengths will be contrasted between regions where the strong Reynolds analogy is invalid (near transition on the suction side) and an isothermal, zero pressure gradient flat plate boundary layer (Wu and Moin, 2010).

Numerical Simulation of Transition in Hypersonic Boundary Layers

DTIC Science & Technology

2011-02-01

sile domes. AGARD Report CP 493. Advisory Group for Aerospace Research and Development. 273 Horvath, T. 2002 Boundary layer transition on slender...reference skin-friction coefficient cp , cv Specific heats at constant pressure and volume, respectively cph Phase speed in propagation direction e...y)) 73 and two-dimensional (W = 0): u = U (y) + u′ , (4.9a) v = v′ , (4.9b) w = w′ , (4.9c) p = 1 + p′ , (4.9d) T = T (y) + T ′ , (4.9e) ρ = 1 T (y

Influence of Surface Roughness on the Second Order Transport of Turbulence in Non-Equilibrium Boundary Layers

DTIC Science & Technology

2006-10-08

FINAL REPORT to Air Force Office of Scientific Research (AFOSR) Project Title Influence of Surface Roughness on the Second Order Transport of...large amount of research has been performed to quantify the effects of Mach number, roughness, and wall curvature on turbulent boundary layers. However...18 a) b) c) Figure 3: a) A. D. Smith high pressure storage tank. b) Morin B series actuator controlling Virgo Engineers Trunion Mounted Ball Valve. c

Experimental investigation of localized disturbances in the straight wing boundary layer, generated by finite surface vibrations

NASA Astrophysics Data System (ADS)

Kozlov, V. V.; Katasonov, M. M.; Pavlenko, A. M.

2017-10-01

Downstream development of artificial disturbances were investigated experimentally using hot-wire constant temperature anemometry. It is shown that vibrations with high-amplitude of a three-dimensional surface lead to formation of two types of perturbations in the straight wing boundary layer: streamwise oriented localized structures and wave packets. The amplitude of streamwise structure is decay downstream. The wave packets amplitude grows in adverse pressure gradient area. The flow separation is exponentially intensified of the wave packet amplitude.

Thermoplasticity of coupled bodies in the case of stress-dependent heat transfer

NASA Technical Reports Server (NTRS)

Kilikovskaya, O. A.

1987-01-01

The problem of the thermal stresses in coupled deformable bodies is formulated for the case where the heat-transfer coefficient at the common boundary depends on the stress-strain state of the bodies (e.g., is a function of the normal pressure at the common boundary). Several one-dimensional problems are solved in this formulation. Among these problems is the determination of the thermal stresses in an n-layer plate and in a two-layer cylinder.

Progress in modeling hypersonic turbulent boundary layers

NASA Technical Reports Server (NTRS)

Zeman, Otto

1993-01-01

A good knowledge of the turbulence structure, wall heat transfer, and friction in turbulent boundary layers (TBL) at high speeds is required for the design of hypersonic air breathing airplanes and reentry space vehicles. This work reports on recent progress in the modeling of high speed TBL flows. The specific research goal described here is the development of a second order closure model for zero pressure gradient TBL's for the range of Mach numbers up to hypersonic speeds with arbitrary wall cooling requirements.

Boundary Layer Flow Over a Moving Wavy Surface

NASA Astrophysics Data System (ADS)

Hendin, Gali; Toledo, Yaron

2016-04-01

Boundary Layer Flow Over a Moving Wavy Surface Gali Hendin(1), Yaron Toledo(1) January 13, 2016 (1)School of Mechanical Engineering, Tel-Aviv University, Israel Understanding the boundary layer flow over surface gravity waves is of great importance as various atmosphere-ocean processes are essentially coupled through these waves. Nevertheless, there are still significant gaps in our understanding of this complex flow behaviour. The present work investigates the fundamentals of the boundary layer air flow over progressive, small-amplitude waves. It aims to extend the well-known Blasius solution for a boundary layer over a flat plate to one over a moving wavy surface. The current analysis pro- claims the importance of the small curvature and the time-dependency as second order effects, with a meaningful impact on the similarity pattern in the first order. The air flow over the ocean surface is modelled using an outer, inviscid half-infinite flow, overlaying the viscous boundary layer above the wavy surface. The assumption of a uniform flow in the outer layer, used in former studies, is now replaced with a precise analytical solution of the potential flow over a moving wavy surface with a known celerity, wavelength and amplitude. This results in a conceptual change from former models as it shows that the pressure variations within the boundary layer cannot be neglected. In the boundary layer, time-dependent Navier-Stokes equations are formulated in a curvilinear, orthogonal coordinate system. The formulation is done in an elaborate way that presents additional, formerly neglected first-order effects, resulting from the time-varying coordinate system. The suggested time-dependent curvilinear orthogonal coordinate system introduces a platform that can also support the formulation of turbulent problems for any surface shape. In order to produce a self-similar Blasius-type solution, a small wave-steepness is assumed and a perturbation method is applied. Consequently, a novel self-similar solution is obtained from the first order set of equations. A second order solution is also obtained, stressing the role of small curvature on the boundary layer flow. The proposed model and solution for the boundary layer problem overlaying a moving wavy surface can also be used as a base flow for stability problems that can develop in a boundary layer, including phases of transitional states.

Lubrication theory for a random fibrous medium

NASA Astrophysics Data System (ADS)

Mirbod, Parisa; Andreopoulos, Yiannis; Weinbaum, Sheldon

2007-11-01

In the classical theory for a slipper bearing one examines the relative motion of an inclined planar surface and a horizontal planar surface. The solution for the pressure distribution and lift force are independent of which boundary is moving and there is an optimum tilt k=h1/h2=2.2 for maximum lift. This symmetry is lost if the intervening space is filled with a soft porous fibrous material. In this paper the generalized Reynolds equation derived in Feng and Weinbaum (2000) J. Fluid Mech. 422:281 is extended to treat a random fiber matrix satisfying the widely used Carman-Kozeny equation. We show that the solutions are strikingly different depending on whether a) the inclined upper boundary moves or b) the upper boundary is stationary and the horizontal lower boundary moves beneath it. The behavior depends critically on the value of the dimensionless fiber interaction layer thickness α=H/√Kp . In a) the pressure and lift force increase as 2̂ and asymptotically approach a limiting behavior for large values of α since the fluid is pushed forward by the tilt of the upper boundary. In b) the pressure and lift force decay as &-2circ; since the fiber interaction layer thickness decreases and the amount of fluid dragged though the fluid gap decreases as α increases and vanishes for α>> 1.

Surface-Induced Near-Field Scaling in the Knudsen Layer of a Rarefied Gas

NASA Astrophysics Data System (ADS)

Gazizulin, R. R.; Maillet, O.; Zhou, X.; Cid, A. Maldonado; Bourgeois, O.; Collin, E.

2018-01-01

We report on experiments performed within the Knudsen boundary layer of a low-pressure gas. The noninvasive probe we use is a suspended nanoelectromechanical string, which interacts with He 4 gas at cryogenic temperatures. When the pressure P is decreased, a reduction of the damping force below molecular friction ∝P had been first reported in Phys. Rev. Lett. 113, 136101 (2014), 10.1103/PhysRevLett.113.136101 and never reproduced since. We demonstrate that this effect is independent of geometry, but dependent on temperature. Within the framework of kinetic theory, this reduction is interpreted as a rarefaction phenomenon, carried through the boundary layer by a deviation from the usual Maxwell-Boltzmann equilibrium distribution induced by surface scattering. Adsorbed atoms are shown to play a key role in the process, which explains why room temperature data fail to reproduce it.

Transient integral boundary layer method to calculate the translesional pressure drop and the fractional flow reserve in myocardial bridges.

PubMed

Bernhard, Stefan; Möhlenkamp, Stefan; Tilgner, Andreas

2006-06-21

The pressure drop-flow relations in myocardial bridges and the assessment of vascular heart disease via fractional flow reserve (FFR) have motivated many researchers the last decades. The aim of this study is to simulate several clinical conditions present in myocardial bridges to determine the flow reserve and consequently the clinical relevance of the disease. From a fluid mechanical point of view the pathophysiological situation in myocardial bridges involves fluid flow in a time dependent flow geometry, caused by contracting cardiac muscles overlying an intramural segment of the coronary artery. These flows mostly involve flow separation and secondary motions, which are difficult to calculate and analyse. Because a three dimensional simulation of the haemodynamic conditions in myocardial bridges in a network of coronary arteries is time-consuming, we present a boundary layer model for the calculation of the pressure drop and flow separation. The approach is based on the assumption that the flow can be sufficiently well described by the interaction of an inviscid core and a viscous boundary layer. Under the assumption that the idealised flow through a constriction is given by near-equilibrium velocity profiles of the Falkner-Skan-Cooke (FSC) family, the evolution of the boundary layer is obtained by the simultaneous solution of the Falkner-Skan equation and the transient von-Kármán integral momentum equation. The model was used to investigate the relative importance of several physical parameters present in myocardial bridges. Results have been obtained for steady and unsteady flow through vessels with 0 - 85% diameter stenosis. We compare two clinical relevant cases of a myocardial bridge in the middle segment of the left anterior descending coronary artery (LAD). The pressure derived FFR of fixed and dynamic lesions has shown that the flow is less affected in the dynamic case, because the distal pressure partially recovers during re-opening of the vessel in diastole. We have further calculated the wall shear stress (WSS) distributions in addition to the location and length of the flow reversal zones in dependence on the severity of the disease. The described boundary layer method can be used to simulate frictional forces and wall shear stresses in the entrance region of vessels. Earlier models are supplemented by the viscous effects in a quasi three-dimensional vessel geometry with a prescribed wall motion. The results indicate that the translesional pressure drop and the mean FFR compares favourably to clinical findings in the literature. We have further shown that the mean FFR under the assumption of Hagen-Poiseuille flow is overestimated in developing flow conditions.

Transient integral boundary layer method to calculate the translesional pressure drop and the fractional flow reserve in myocardial bridges

PubMed Central

Bernhard, Stefan; Möhlenkamp, Stefan; Tilgner, Andreas

2006-01-01

Background The pressure drop – flow relations in myocardial bridges and the assessment of vascular heart disease via fractional flow reserve (FFR) have motivated many researchers the last decades. The aim of this study is to simulate several clinical conditions present in myocardial bridges to determine the flow reserve and consequently the clinical relevance of the disease. From a fluid mechanical point of view the pathophysiological situation in myocardial bridges involves fluid flow in a time dependent flow geometry, caused by contracting cardiac muscles overlying an intramural segment of the coronary artery. These flows mostly involve flow separation and secondary motions, which are difficult to calculate and analyse. Methods Because a three dimensional simulation of the haemodynamic conditions in myocardial bridges in a network of coronary arteries is time-consuming, we present a boundary layer model for the calculation of the pressure drop and flow separation. The approach is based on the assumption that the flow can be sufficiently well described by the interaction of an inviscid core and a viscous boundary layer. Under the assumption that the idealised flow through a constriction is given by near-equilibrium velocity profiles of the Falkner-Skan-Cooke (FSC) family, the evolution of the boundary layer is obtained by the simultaneous solution of the Falkner-Skan equation and the transient von-Kármán integral momentum equation. Results The model was used to investigate the relative importance of several physical parameters present in myocardial bridges. Results have been obtained for steady and unsteady flow through vessels with 0 – 85% diameter stenosis. We compare two clinical relevant cases of a myocardial bridge in the middle segment of the left anterior descending coronary artery (LAD). The pressure derived FFR of fixed and dynamic lesions has shown that the flow is less affected in the dynamic case, because the distal pressure partially recovers during re-opening of the vessel in diastole. We have further calculated the wall shear stress (WSS) distributions in addition to the location and length of the flow reversal zones in dependence on the severity of the disease. Conclusion The described boundary layer method can be used to simulate frictional forces and wall shear stresses in the entrance region of vessels. Earlier models are supplemented by the viscous effects in a quasi three-dimensional vessel geometry with a prescribed wall motion. The results indicate that the translesional pressure drop and the mean FFR compares favourably to clinical findings in the literature. We have further shown that the mean FFR under the assumption of Hagen-Poiseuille flow is overestimated in developing flow conditions. PMID:16790065

High-Subsonic Performance Characteristics and Boundary-Layer Investigations of a 12 deg 10-Inch-Inlet-Diameter Conical Diffuser

DTIC Science & Technology

1950-05-11

available condition supersonic flow was obtained as far as K.5 inches downstream from the diffueer inlet with a maximum Mach number of M % 1.5...Boundary—layer total-pressure measurements were made with the rake shown in figure k. The tubes varied in size from 0.030-Inch outside diameter...at the wall to 0.050—inch outside diameter farther out. A static-pressure tube was mounted on the rake to measure static pressures at the same

Shear-induced surface alignment of polymer dispersed liquid crystal microdroplets on the boundary layer

NASA Technical Reports Server (NTRS)

Parmar, D. S.; Singh, J. J.

1993-01-01

Polymer dispersed liquid crystal thin films have been deposited on a glass substrate, utilizing the processes of polymerization and solvent evaporation induced phase separation. Liquid crystal microdroplets trapped on the upper surface of the thin film respond to the shear stress due to air or gas flow on the surface layer. Response to an applied step shear stress input on the surface layer has been measured by measuring the time response of the transmitted light intensity. Initial results on the measurements of the light transmission as a function of the air flow differential pressure indicate that these systems offer features suitable for boundary layer and gas flow sensors.

Experimental Study of Boundary Layer Behavior in a Simulated Low Pressure Turbine. Degree awarded by the University of Toledo, May 1998

NASA Technical Reports Server (NTRS)

Shyne, Rickey J.

1998-01-01

A detailed investigation of the flow physics occurring on the suction side of a simulated Low Pressure Turbine (LPT) blade was performed. A contoured upper wall was designed to simulate the pressure distribution of an actual LPT blade onto a flat plate. The experiments were carried out at Reynolds numbers of 100,000 and 250,000 with three levels of freestream turbulence. Freestream turbulence levels ranging from 0.8% to 3% was used in this experiment. Smoke-wire flow visualization data was used to confirm that the boundary layer was separated and formed a bubble. Hot-wires (single and x-wire) and surface mounted hot-film gases and static pressure taps were used to map the flowfield. The transition process over the separated flow region is observed to be similar to a laminar free shear layer flow with the formation of a large coherent eddy structure. For each condition, the locations defining the separation bubble were determined by careful examination of pressure and mean velocity profile data. Transition onset location and length determined from intermittency profiles decrease as freestream turbulence levels increase. Additionally, the length and height of the laminar separation bubbles were observed to be inversely proportional to the levels of freestream turbulence.

An experimental and computational investigation of the flow field about a transonic airfoil in supercritical flow with turbulent boundary-layer separation

NASA Technical Reports Server (NTRS)

Rubesin, M. W.; Okuno, A. F.; Levy, L. L., Jr.; Mcdevitt, J. B.; Seegmiller, H. L.

1976-01-01

A combined experimental and computational research program is described for testing and guiding turbulence modeling within regions of separation induced by shock waves incident in turbulent boundary layers. Specifically, studies are made of the separated flow the rear portion of an 18%-thick circular-arc airfoil at zero angle of attack in high Reynolds number supercritical flow. The measurements include distributions of surface static pressure and local skin friction. The instruments employed include highfrequency response pressure cells and a large array of surface hot-wire skin-friction gages. Computations at the experimental flow conditions are made using time-dependent solutions of ensemble-averaged Navier-Stokes equations, plus additional equations for the turbulence modeling.

Persistent Structures in the Turbulent Boundary Layer

NASA Technical Reports Server (NTRS)

Palumbo, Dan; Chabalko, Chris

2005-01-01

Persistent structures in the turbulent boundary layer are located and analyzed. The data are taken from flight experiments on large commercial aircraft. An interval correlation technique is introduced which is able to locate the structures. The Morlet continuous wavelet is shown to not only locates persistent structures but has the added benefit that the pressure data are decomposed in time and frequency. To better understand how power is apportioned among these structures, a discrete Coiflet wavelet is used to decompose the pressure data into orthogonal frequency bands. Results indicate that some structures persist a great deal longer in the TBL than would be expected. These structure contain significant power and may be a primary source of vibration energy in the airframe.

Supersonic compressor

DOE Office of Scientific and Technical Information (OSTI.GOV)

Roberts, II, William Byron; Lawlor, Shawn P.; Breidenthal, Robert E.

A supersonic compressor including a rotor to deliver a gas at supersonic conditions to a diffuser. The diffuser includes a plurality of aerodynamic ducts that have converging and diverging portions, for deceleration of gas to subsonic conditions and then for expansion of subsonic gas, to change kinetic energy of the gas to static pressure. The aerodynamic ducts include vortex generating structures for controlling boundary layer, and structures for changing the effective contraction ratio to enable starting even when the aerodynamic ducts are designed for high pressure ratios, and structures for boundary layer control. In an embodiment, aerodynamic ducts are providedmore » having an aspect ratio of in excess of two to one, when viewed in cross-section orthogonal to flow direction at an entrance to the aerodynamic duct.« less

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

NASA Technical Reports Server (NTRS)

Adamson, T. C., Jr.; Liou, M. S.; Messiter, A. F.

1980-01-01

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.

Synthetic Vortex Generator Jets Used to Control Separation on Low-Pressure Turbine Airfoils

NASA Technical Reports Server (NTRS)

Ashpis, David E.; Volino, Ralph J.

2005-01-01

Low-pressure turbine (LPT) airfoils are subject to increasingly stronger pressure gradients as designers impose higher loading in an effort to improve efficiency and lower cost by reducing the number of airfoils in an engine. When the adverse pressure gradient on the suction side of these airfoils becomes strong enough, the boundary layer will separate. Separation bubbles, particularly those that fail to reattach, can result in a significant loss of lift and a subsequent degradation of engine efficiency. The problem is particularly relevant in aircraft engines. Airfoils optimized to produce maximum power under takeoff conditions may still experience boundary layer separation at cruise conditions because of the thinner air and lower Reynolds numbers at altitude. Component efficiency can drop significantly between takeoff and cruise conditions. The decrease is about 2 percent in large commercial transport engines, and it could be as large as 7 percent in smaller engines operating at higher altitudes. Therefore, it is very beneficial to eliminate, or at least reduce, the separation bubble.

Measurement and Empirical Correlation of Transpiration-Cooling Parameters on a 25 degree Cone in a Turbulent Boundary Layer in Both Free Flight and a Hot-Gas Jet

NASA Technical Reports Server (NTRS)

Walton, Thomas E., Jr.; Rashis, Bernard

1961-01-01

Transpiration-cooling parameters are presented for a turbulent boundary layer on a cone configuration with a total angle of 250 which was tested in both free flight and in an ethylene-heated high-temperature jet at a Mach number of 2.0. The flight-tested cone was flown to a maximum Mach number of 4.08 and the jet tests were conducted at stagnation temperatures ranging from 937 R to 1,850 R. In general, the experimental heat transfer was in good agreement with the theoretical values. Inclusion of the ratio of local stream temperature to wall temperature in the nondimensional flow rate parameter enabled good correlation of both sets of transpiration data. The measured pressure at the forward station coincided with the theoretical pressure over a sharp cone; however, the measured pressure increased with distance from the nose tip.

DNS of a non-equilibrium adverse pressure gradient turbulent boundary layer

NASA Astrophysics Data System (ADS)

Gungor, Taygun R.; Gungor, Ayse G.; Maciel, Yvan; Simens, Mark P.

2017-11-01

A new direct numerical simulation (DNS) dataset of a non-equilibrium adverse pressure gradient (APG) turbulent boundary layer (TBL) that evolves from a zero-pressure-gradient (ZPG) TBL to a TBL which is very close to separation at Reθ is around 8200 is presented. There are two simulations running together in the DNS computational setup. The APG TBL spans Reθ = 1476 - 8276 . Mean velocity results do not satisfy the log law as the defect in the velocity increases. The production and the Reynolds stress peak are observed around y /δ* = 1 after the flow is evolved up to a certain point. The new dataset is compared with other datasets in terms of mean values, Reynolds stresses and turbulent kinetic energy budgets and using this comparison scaling study is performed. Funded by in part by ITU-AYP and NSERC of Canada.

Effects of bleed-hole geometry and plenum pressure on three-dimensional shock-wave/boundary-layer/bleed interactions

NASA Technical Reports Server (NTRS)

Chyu, Wei J.; Rimlinger, Mark J.; Shih, Tom I.-P.

1993-01-01

A numerical study was performed to investigate 3D shock-wave/boundary-layer interactions on a flat plate with bleed through one or more circular holes that vent into a plenum. This study was focused on how bleed-hole geometry and pressure ratio across bleed holes affect the bleed rate and the physics of the flow in the vicinity of the holes. The aspects of the bleed-hole geometry investigated include angle of bleed hole and the number of bleed holes. The plenum/freestream pressure ratios investigated range from 0.3 to 1.7. This study is based on the ensemble-averaged, 'full compressible' Navier-Stokes (N-S) equations closed by the Baldwin-Lomax algebraic turbulence model. Solutions to the ensemble-averaged N-S equations were obtained by an implicit finite-volume method using the partially-split, two-factored algorithm of Steger on an overlapping Chimera grid.

Effect of Reynolds number, turbulence level and periodic wake flow on heat transfer on low pressure turbine blades.

PubMed

Suslov, D; Schulz, A; Wittig, S

2001-05-01

The development of effective cooling methods is of major importance for the design of new gas turbines blades. The conception of optimal cooling schemes requires a detailed knowledge of the heat transfer processes on the blade's surfaces. The thermal load of turbine blades is predominantly determined by convective heat transfer which is described by the local heat transfer coefficient. Heat transfer is closely related to the boundary layer development along the blade surface and hence depends on various flow conditions and geometrical parameters. Particularly Reynolds number, pressures gradient and turbulence level have great impact on the boundary layer development and the according heat transfer. Therefore, in the present study, the influence of Reynolds number, turbulence intensity, and periodic unsteady inflow on the local heat transfer of a typical low pressure turbine airfoil is experimentally examined in a plane cascade.

Mars Science Laboratory Heatshield Aerothermodynamics: Design and Reconstruction

NASA Technical Reports Server (NTRS)

Edquist, Karl T.; Hollis, Brian R.; Johnston, Christopher O.; Bose, Deepak; White, Todd R.; Mahzari, Milad

2013-01-01

The Mars Science Laboratory heatshield was designed to withstand a fully turbulent heat pulse based on test results and computational analysis on a pre-flight design trajectory. Instrumentation on the flight heatshield measured in-depth temperatures in the thermal protection system. The data indicate that boundary layer transition occurred at 5 of 7 thermocouple locations prior to peak heating. Data oscillations at 3 pressure measurement locations may also indicate transition. This paper presents the heatshield temperature and pressure data, possible explanations for the timing of boundary layer transition, and a qualitative comparison of reconstructed and computational heating on the as-flown trajectory. Boundary layer Reynolds numbers that are typically used to predict transition are compared to observed transition at various heatshield locations. A uniform smooth-wall transition Reynolds number does not explain the timing of boundary layer transition observed during flight. A roughness-based Reynolds number supports the possibility of transition due to discrete or distributed roughness elements on the heatshield. However, the distributed roughness height would have needed to be larger than the pre-flight assumption. The instrumentation confirmed the predicted location of maximum turbulent heat flux near the leeside shoulder. The reconstructed heat flux at that location is bounded by smooth-wall turbulent calculations on the reconstructed trajectory, indicating that augmentation due to surface roughness probably did not occur. Turbulent heating on the downstream side of the heatshield nose exceeded smooth-wall computations, indicating that roughness may have augmented heating. The stagnation region also experienced heating that exceeded computational levels, but shock layer radiation does not fully explain the differences.

Investigation of Low-Pressure Turbine Endwall Flows: Simulations and Experiments (Postprint)

DTIC Science & Technology

2015-01-01

direction) minor semiaxis of 0.0417Cx (0.25in). Measured in the axial direction, the flat plate leading edge was located at x=-3.958Cx (23.75in) where...for public release; distribution unlimited. plate boundary layer is δ∗ = 1.721s/Re0.5s . For s = 4.833Cx, Reδ∗ = 1.721 √ s/CxRe = 1, 200. At this...boundary which was located at x=-1.4Cx. The following approximations hold for a turbulent flat plate boundary layer: δ99 = 0.37s Re0.2s , δ∗ = 0.046s

Turbulent boundary layer on a convex, curved surface

NASA Technical Reports Server (NTRS)

Gillis, J. C.; Johnston, J. P.; Kays, W. M.; Moffat, R. J.

1980-01-01

The effects of strong convex curvature on boundary layer turbulence were investigated. The data gathered on the behavior of Reynolds stress suggested the formulation of a simple turbulence model. Three sets of data were taken on two separate facilities. Both rigs had flow from a flat surface, over a convex surface with 90 deg of turning, and then onto a flat recovery surface. The geometry was adjusted so that, for both rigs, the pressure gradient along the test surface was zero - thus avoiding any effects of streamwise acceleration on the wall layers. Results show that after a sudden introduction of curvature, the shear stress in the outer part of the boundary layer is sharply diminished and is even slightly negative near the edge. The wall shear also drops off quickly downstream. In contrast, when the surface suddenly becomes flat again, the wall shear and shear stress profiles recover very slowly towards flat wall conditions.

On investigating wall shear stress in two-dimensional plane turbulent wall jets

NASA Astrophysics Data System (ADS)

Mehdi, Faraz; Johansson, Gunnar; White, Christopher; Naughton, Jonathan

2012-11-01

Mehdi & White [Exp Fluids 50:43-51(2011)] presented a full momentum integral based method for determining wall shear stress in zero pressure gradient turbulent boundary layers. They utilized the boundary conditions at the wall and at the outer edge of the boundary layer. A more generalized expression is presented here that uses just one boundary condition at the wall. The method is mathematically exact and has an advantage of having no explicit streamwise gradient terms. It is successfully applied to two different experimental plane turbulent wall jet datasets for which independent estimates of wall shear stress were known. Complications owing to experimental inaccuracies in determining wall shear stress from the proposed method are also discussed.

Theory and operation of the Gould 32/27 programs ABLE-2A and EBLE for the tropospheric air motion measurement system

NASA Technical Reports Server (NTRS)

Butler, C.

1986-01-01

Software development for the Trospheric Air Motion Measurement Systems (TAMMS) is documented. In July/August the TAMMS was flown on the NASA/Goddard Flight Center Electra aircraft for 19 mission for the ABLE-2A (Amazon Boundary Layer Experiment) in Brazil. In December 1985, several flights were performed to assess the contamination and boundary layer of the Electra. Position data, flow angles, pressure transducer measurements were recorded. The programs written for the ABLE-2A were modified due to timing considerations for this particular program. The 3-step programs written for EBLE (Electra Boundary Layer Experiment) are described. Power up and log-on procedures are discussed. A few editing techniques are described for modification of the programs.

Viscous/potential flow about multi-element two-dimensional and infinite-span swept wings: Theory and experiment

NASA Technical Reports Server (NTRS)

Olson, L. E.; Dvorak, F. A.

1975-01-01

The viscous subsonic flow past two-dimensional and infinite-span swept multi-component airfoils is studied theoretically and experimentally. The computerized analysis is based on iteratively coupled boundary layer and potential flow analysis. The method, which is restricted to flows with only slight separation, gives surface pressure distribution, chordwise and spanwise boundary layer characteristics, lift, drag, and pitching moment for airfoil configurations with up to four elements. Merging confluent boundary layers are treated. Theoretical predictions are compared with an exact theoretical potential flow solution and with experimental measures made in the Ames 40- by 80-Foot Wind Tunnel for both two-dimensional and infinite-span swept wing configurations. Section lift characteristics are accurately predicted for zero and moderate sweep angles where flow separation effects are negligible.

Vortex Formation During Unsteady Boundary-Layer Separation

NASA Astrophysics Data System (ADS)

Das, Debopam; Arakeri, Jaywant H.

1998-11-01

Unsteady laminar boundary-layer separation is invariably accompanied by the formation of vortices. The aim of the present work is to study the vortex formation mechanism(s). An adverse pressure gradient causing a separation can be decomposed into a spatial component ( spatial variation of the velocity external to the boundary layer ) and a temporal component ( temporal variation of the external velocity ). Experiments were conducted in a piston driven 2-D water channel, where the spatial component could be be contolled by geometry and the temporal component by the piston motion. We present results for three divergent channel geometries. The piston motion consists of three phases: constant acceleration from start, contant velocity, and constant deceleration to stop. Depending on the geometry and piston motion we observe different types of unsteady separation and vortex formation.

Using Model Helicopters for Meteorological Observations in Support of Tornado Forecasting

NASA Astrophysics Data System (ADS)

Harrison, William; Roscoe, Bryan; Schafer, David; Bluestein, Howard; Lary, David

2012-10-01

In order to gain a better understanding of the physical factors involved in tornadogenesis, a complete 3-D profile of winds, temperature, and humidity in the forward-flank and rear-flank gust front regions in supercells is required. Conventional methods of making comparative measurements in and around storms are very limited. Measurements that comprehensively profile the boundary layer winds and thermodynamics are valuable but rare. A better understanding of the physical properties in these boundary layers will improve forecasts and increase warning times in affected areas. Remote-controlled model helicopters are a uniquely qualified platform for this application, allowing us to fully profile these boundary layers. Our system will consist of a swarm of autonomous acrobatic helicopters, each outfitted with temperature, pressure, humidity, and wind speed sensors.

DOE Office of Scientific and Technical Information (OSTI.GOV)

Bühler, Stefan; Obrist, Dominik; Kleiser, Leonhard

We investigate numerically the effects of nozzle-exit flow conditions on the jet-flow development and the near-field sound at a diameter-based Reynolds number of Re{sub D} = 18 100 and Mach number Ma = 0.9. Our computational setup features the inclusion of a cylindrical nozzle which allows to establish a physical nozzle-exit flow and therefore well-defined initial jet-flow conditions. Within the nozzle, the flow is modeled by a potential flow core and a laminar, transitional, or developing turbulent boundary layer. The goal is to document and to compare the effects of the different jet inflows on the jet flow development and themore » sound radiation. For laminar and transitional boundary layers, transition to turbulence in the jet shear layer is governed by the development of Kelvin-Helmholtz instabilities. With the turbulent nozzle boundary layer, the jet flow development is characterized by a rapid changeover to a turbulent free shear layer within about one nozzle diameter. Sound pressure levels are strongly enhanced for laminar and transitional exit conditions compared to the turbulent case. However, a frequency and frequency-wavenumber analysis of the near-field pressure indicates that the dominant sound radiation characteristics remain largely unaffected. By applying a recently developed scaling procedure, we obtain a close match of the scaled near-field sound spectra for all nozzle-exit turbulence levels and also a reasonable agreement with experimental far-field data.« less

Pressure measurements in a rapidly sheared turbulent wall layer

NASA Astrophysics Data System (ADS)

Diwan, Sourabh; Morrison, Jonathan

2014-11-01

The aim of the present work is to improve understanding of the role of pressure fluctuations in the generation of coherent structures in wall-bounded turbulent flows, with particular regard to the rapid and slow source terms. The work is in part motivated by the recent numerical simulations of Sharma et al. (Phy. Fluids, 23, 2011), which showed the importance of pressure fluctuations (and their spatial gradients) in the dynamics of large-scale turbulent motions. Our experimental design consists of first generating a shearless boundary layer in a wind tunnel by passing a grid-generated turbulent flow over a moving floor whose speed is matched to the freestream velocity, and then shearing it rapidly by passing it over a stationary floor further downstream. Close to the leading edge of the stationary floor, the resulting flow is expected to satisfy the approximations of the Rapid Distortion Theory and therefore would be an ideal candidate for studying linear processes in wall turbulence. We carry out pressure measurements on the wall as well as within the flow - the former using surface mounted pressure transducers and the latter using a static pressure probe similar in design to that used by Tsuji et al. (J. Fluid. Mech. 585, 2007). We also present a comparison between the rapidly sheared flow and a more conventional boundary layer subjected to a turbulent free stream. We acknowledge the financial support from EPSRC (Grant No. EP/I037938).

Intermittent Turbulence in the Stable Boundary Layer over Land. Part III: A Classification for Observations during CASES-99.

NASA Astrophysics Data System (ADS)

van de Wiel, B. J. H.; Moene, A. F.; Hartogensis, O. K.; de Bruin, H. A. R.; Holtslag, A. A. M.

2003-10-01

In this paper a classification of stable boundary layer regimes is presented based on observations of near-surface turbulence during the Cooperative Atmosphere-Surface Exchange Study-1999 (CASES-99). It is found that the different nights can be divided into three subclasses: a turbulent regime, an intermittent regime, and a radiative regime, which confirms the findings of two companion papers that use a simplified theoretical model (it is noted that its simpliflied structure limits the model generality to near-surface flows). The papers predict the occurrence of stable boundary layer regimes in terms of external forcing parameters such as the (effective) pressure gradient and radiative forcing. The classification in the present work supports these predictions and shows that the predictions are robust in a qualitative sense. As such, it is, for example, shown that intermittent turbulence is most likely to occur in clear-sky conditions with a moderately weak effective pressure gradient. The quantitative features of the theoretical classification are, however, rather sensitive to (often uncertain) local parameter estimations, such as the bulk heat conductance of the vegetation layer. This sensitivity limits the current applicability of the theoretical classification in a strict quantitative sense, apart from its conceptual value.

Observation of a brine layer on an ice surface with an environmental scanning electron microscope at higher pressures and temperatures.

PubMed

Krausko, Ján; Runštuk, Jiří; Neděla, Vilém; Klán, Petr; Heger, Dominik

2014-05-20

Observation of a uranyl-salt brine layer on an ice surface using backscattered electron detection and ice surface morphology using secondary-electron detection under equilibrium conditions was facilitated using an environmental scanning electron microscope (ESEM) at temperatures above 250 K and pressures of hundreds of Pa. The micrographs of a brine layer over ice grains prepared by either slow or shock freezing provided a complementary picture of the contaminated ice grain boundaries. Fluorescence spectroscopy of the uranyl ions in the brine layer confirmed that the species exists predominately in the solvated state under experimental conditions of ESEM.

Internal hypersonic flow. [in thin shock layer

NASA Technical Reports Server (NTRS)

Lin, T. C.; Rubin, S. G.

1974-01-01

An approach for studying hypersonic internal flow with the aid of a thin-shock-layer approximation is discussed, giving attention to a comparison of thin-shock-layer results with the data obtained on the basis of the imposition theory or a finite-difference integration of the Euler equations. Relations in the case of strong interaction are considered together with questions of pressure distribution and aspects of the boundary-layer solution.

Martian Dust Devils: Laboratory Simulations of Particle Threshold

NASA Technical Reports Server (NTRS)

Greeley, Ronald; Balme, Matthew R.; Iverson, James D.; Metzger, Stephen; Mickelson, Robert; Phoreman, Jim; White, Bruce

2003-01-01

An apparatus has been fabricated to simulate terrestrial and Martian dust devils. Comparisons of surface pressure profiles through the vortex core generated in the apparatus with both those in natural dust devils on Earth and those inferred for Mars are similar and are consistent with theoretical Rankine vortex models. Experiments to determine particle threshold under Earth ambient atmospheric pressures show that sand (particles > 60 micron in diameter) threshold is analogous to normal boundary-layer shear, in which the rotating winds of the vortex generate surface shear and hence lift. Lower-pressure experiments down to approx. 65 mbar follow this trend for sand-sized particles. However, smaller particles (i.e., dust) and all particles at very low pressures (w 10-60 mbar) appear to be subjected to an additional lift function interpreted to result from the strong decrease in atmospheric pressure centered beneath the vortex core. Initial results suggest that the wind speeds required for the entrainment of grains approx. 2 microns in diameter (i.e., Martian dust sizes) are about half those required for entrainment by boundary layer winds on both Earth and Mars.

Flow separation of currents in shallow water

USGS Publications Warehouse

Signell, Richard P.

1989-01-01

Flow separation of currents in shallow coastal areas is investigated using a boundary layer model for two-dimensional (depth-averaged) tidal flow past an elliptic headland. If the shoaling region near the coast is narrow compared to the scale of the headland, bottom friction causes the flow to separate just downstream of the point where the pressure gradient switches from favoring to adverse. As long as the shoaling region at the coast is well resolved, the inclusion of eddy viscosity and a no-slip boundary condition have no effect on this result. An approximate analytic solution for the pressure gradient along the boundary is obtained by assuming the flow away from the immediate vicinity of the boundary is irrotational. On the basis of the pressure gradient obtained from the irrotational flow solution, flow separation is a strong function of the headland aspect ratio, an equivalent Reynolds number, and a Keulegan-Carpenter number.

Natural laminar flow flight experiments on a swept wing business jet-boundary layer stability analyses

NASA Technical Reports Server (NTRS)

Rozendaal, R. A.

1986-01-01

The linear boundary layer stability analyses and their correlation with data of 18 cases from a natural laminar flow (NLF) flight test program using a Cessna Citation 3 business jet are described. The transition point varied from 5% to 35% chord for these conditions, and both upper and lower wing surfaces were included. Altitude varied from 10,000 to 43,000 ft and Mach number from 0.3 to 0.8. Four cases were at nonzero sideslip. Although there was much scatter in the results, the analyses of boundary layer stability at the 18 conditions led to the conclusion that crossflow instability was the primary cause of transition. However, the sideslip cases did show some interaction of crossflow and Tollmien-Schlichting disturbances. The lower surface showed much lower Tollmien-Schlichting amplification at transition than the upper surface, but similar crossflow amplifications. No relationship between Mach number and disturbance amplification at transition could be found. The quality of these results is open to question from questionable wing surface quality, inadequate density of transition sensors on the wing upper surface, and an unresolved pressure shift in the wing pressure data. The results of this study show the need for careful preparation for transition experiments. Preparation should include flow analyses of the test surface, boundary layer disturbance amplification analyses, and assurance of adequate surface quality in the test area. The placement of necessary instruments and usefulness of the resulting data could largely be determined during the pretest phase.

Atmospheric tides on Venus. III - The planetary boundary layer

NASA Technical Reports Server (NTRS)

Dobrovolskis, A. R.

1983-01-01

Diurnal solar heating of Venus' surface produces variable temperatures, winds, and pressure gradients within a shallow layer at the bottom of the atmosphere. The corresponding asymmetric mass distribution experiences a tidal torque tending to maintain Venus' slow retrograde rotation. It is shown that including viscosity in the boundary layer does not materially affect the balance of torques. On the other hand, friction between the air and ground can reduce the predicted wind speeds from about 5 to about 1 m/sec in the lower atmosphere, more consistent with the observations from Venus landers and descent probes. Implications for aeolian activity on Venus' surface and for future missions are discussed.

On the growth of turbulent regions in laminar boundary layers

NASA Technical Reports Server (NTRS)

Gad-El-hak, M.; Riley, J. J.; Blackwelder, R. F.

1981-01-01

Turbulent spots evolving in a laminar boundary layer on a nominally zero pressure gradient flat plate are investigated. The plate is towed through an 18 m water channel, using a carriage that rides on a continuously replenished oil film giving a vibrationless tow. Turbulent spots are initiated using a solenoid valve that ejects a small amount of fluid through a minute hole on the working surface. A novel visualization technique that utilizes fluorescent dye excited by a sheet of laser light is employed. Some new aspects of the growth and entrainment of turbulent spots, especially with regard to lateral growth, are inferred from the present experiments. To supplement the information on lateral spreading, a turbulent wedge created by placing a roughness element in the laminar boundary layer is also studied both visually and with probe measurements. The present results show that, in addition to entrainment, another mechanism is needed to explain the lateral growth characteristics of a turbulent region in a laminar boundary layer. This mechanism, termed growth by destabilization, appears to be a result of the turbulence destabilizing the unstable laminar boundary layer in its vicinity. To further understand the growth mechanisms, the turbulence in the spot is modulated using drag-reducing additives and salinity stratification.

β-distribution for Reynolds stress and turbulent heat flux in relaxation turbulent boundary layer of compression ramp

NASA Astrophysics Data System (ADS)

Hu, YanChao; Bi, WeiTao; Li, ShiYao; She, ZhenSu

2017-12-01

A challenge in the study of turbulent boundary layers (TBLs) is to understand the non-equilibrium relaxation process after sep-aration and reattachment due to shock-wave/boundary-layer interaction. The classical boundary layer theory cannot deal with the strong adverse pressure gradient, and hence, the computational modeling of this process remains inaccurate. Here, we report the direct numerical simulation results of the relaxation TBL behind a compression ramp, which reveal the presence of intense large-scale eddies, with significantly enhanced Reynolds stress and turbulent heat flux. A crucial finding is that the wall-normal profiles of the excess Reynolds stress and turbulent heat flux obey a β-distribution, which is a product of two power laws with respect to the wall-normal distances from the wall and from the boundary layer edge. In addition, the streamwise decays of the excess Reynolds stress and turbulent heat flux also exhibit power laws with respect to the streamwise distance from the corner of the compression ramp. These results suggest that the relaxation TBL obeys the dilation symmetry, which is a specific form of self-organization in this complex non-equilibrium flow. The β-distribution yields important hints for the development of a turbulence model.

Boundary-layer and wake measurements on a swept, circulation-control wing

NASA Technical Reports Server (NTRS)

Spaid, Frank W.; Keener, Earl R.

1987-01-01

Wind-tunnel measurements of boundary-layer and wake velocity profiles and surface static pressure distributions are presented for a swept, circulation-control wing. The model is an aspect-ratio-four semispan wing mounted on the tunnel side wall at a sweep angle of 45 deg. A full-span, tangential, rearward blowing, circulation-control slot is located ahead of the trailing edge on the upper surface. Flow surveys were obtained at mid-semispan at freestream Mach numbers of 0.425 and 0.70. Boundary-layer profiles measured on the forward portions of the wing are approximately streamwise and two dimensional. The flow in the vicinity of the jet exit and in the near wake is highly three dimensional. The jet flow near the slot on the Coanda surface is directed normal to the slot. Near-wake surveys show large outboard flows at the center of the wake. At Mach 0.425 and a 5-deg angle of attack, a range of jet-blowing rates was found for which an abrupt transition from incipient separation to attached flow occurs in the boundary layer upstream of the slot. The variation in the lower-surface separation location with blowing rate was determined from boundary-layer measurements at Mach 0.425.

Shed vortex structure and phase-averaged velocity statistics in symmetric/asymmetric turbulent flat plate wakes

NASA Astrophysics Data System (ADS)

Rai, Man Mohan

2018-05-01

The near wake of a flat plate is investigated via direct numerical simulations. Many earlier experimental investigations have used thin plates with sharp trailing edges and turbulent boundary layers to create the wake. This results in large θ/DTE values (θ is the boundary layer momentum thickness toward the end of the plate and DTE is the trailing edge thickness). In the present study, the emphasis is on relatively thick plates with circular trailing edges (CTEs) resulting in θ/D values less than one (D is the plate thickness and the diameter of the CTE) and vigorous vortex shedding. The Reynolds numbers based on the plate length and D are 1.255 × 106 and 10 000, respectively. Two cases are computed: one with turbulent boundary layers on both the upper and lower surfaces of the plate (statistically the same, symmetric wake, Case TT) and the other with turbulent and laminar boundary layers on the upper and lower surfaces, respectively (asymmetric case, Case TL). The data and understanding obtained are of considerable engineering interest, particularly in turbomachinery where the pressure side of an airfoil can remain laminar or transitional because of a favorable pressure gradient and the suction side is turbulent. Shed-vortex structure and phase-averaged velocity statistics obtained in the two cases are compared here. The upper negative shed vortices in Case TL (turbulent separating boundary layer) are weaker than the lower positive ones (laminar separating boundary layer) at inception (a factor of 1.27 weaker in terms of peak phase-averaged spanwise vorticity at the first appearance of a peak). The upper vortices weaken rapidly as they travel downstream. A second feature of interest in Case TL is a considerable increase in the peak phase-averaged, streamwise normal intensity (random component) with increasing streamwise distance (x/D) that occurs near the positive vortex cores. This behavior is observed for a few diameters in the near wake. This is counter to Case TT where the peak value essentially decreases with increasing x/D. Both these effects are examined in detail, and the important contributors are identified.

Unsteady turbulent boundary layers in swimming rainbow trout.

PubMed

Yanase, Kazutaka; Saarenrinne, Pentti

2015-05-01

The boundary layers of rainbow trout, Oncorhynchus mykiss, swimming at 1.02±0.09 L s(-1) (mean±s.d., N=4), were measured by the particle image velocimetry (PIV) technique at a Reynolds number of 4×10(5). The boundary layer profile showed unsteadiness, oscillating above and beneath the classical logarithmic law of the wall with body motion. Across the entire surface regions that were measured, local Reynolds numbers based on momentum thickness, which is the distance that is perpendicular to the fish surface through which the boundary layer momentum flows at free-stream velocity, were greater than the critical value of 320 for the laminar-to-turbulent transition. The skin friction was dampened on the convex surface while the surface was moving towards a free-stream flow and increased on the concave surface while retreating. These observations contradict the result of a previous study using different species swimming by different methods. Boundary layer compression accompanied by an increase in local skin friction was not observed. Thus, the overall results may not support absolutely the Bone-Lighthill boundary layer thinning hypothesis that the undulatory motions of swimming fish cause a large increase in their friction drag because of the compression of the boundary layer. In some cases, marginal flow separation occurred on the convex surface in the relatively anterior surface region, but the separated flow reattached to the fish surface immediately downstream. Therefore, we believe that a severe impact due to induced drag components (i.e. pressure drag) on the swimming performance, an inevitable consequence of flow separation, was avoided. © 2015. Published by The Company of Biologists Ltd.

Aeroheating Measurement of Apollo Shaped Capsule with Boundary Layer Trip in the Free-piston Shock Tunnel HIEST

NASA Technical Reports Server (NTRS)

Hideyuki, TANNO; Tomoyuki, KOMURO; Kazuo, SATO; Katsuhiro, ITOH; Lillard, Randolph P.; Olejniczak, Joseph

2013-01-01

An aeroheating measurement test campaign of an Apollo capsule model with laminar and turbulent boundary layer was performed in the free-piston shock tunnel HIEST at JAXA Kakuda Space Center. A 250mm-diameter 6.4%-scaled Apollo CM capsule model made of SUS-304 stainless steel was applied in this study. To measure heat flux distribution, the model was equipped with 88 miniature co-axial Chromel-Constantan thermocouples on the heat shield surface of the model. In order to promote boundary layer transition, a boundary layer trip insert with 13 "pizza-box" isolated roughness elements, which have 1.27mm square, were placed at 17mm below of the model geometric center. Three boundary layer trip inserts with roughness height of k=0.3mm, 0.6mm and 0.8mm were used to identify the appropriate height to induce transition. Heat flux records with or without roughness elements were obtained for model angles of attack 28º under stagnation enthalpy between H(sub 0)=3.5MJ/kg to 21MJ/kg and stagnation pressure between P(sub 0)=14MPa to 60MPa. Under the condition above, Reynolds number based on the model diameter was varied from 0.2 to 1.3 million. With roughness elements, boundary layer became fully turbulent less than H(sub 0)=9MJ/kg condition. However, boundary layer was still laminar over H(sub 0)=13MJ/kg condition even with the highest roughness elements. An additional experiment was also performed to correct unexpected heat flux augmentation observed over H(sub 0)=9MJ/kg condition.

Development of a Perfectly Matched Layer Technique for a Discontinuous-Galerkin Spectral-Element Method

NASA Technical Reports Server (NTRS)

Garai, Anirban; Murman, Scott M.; Madavan, Nateri K.

2016-01-01

The numerical simulation of many aerodynamic non-periodic flows of practical interest involves discretized computational domains that often must be artificially truncated. Appropriate boundary conditions are required at these truncated domain boundaries, and ideally, these boundary conditions should be perfectly "absorbing" or "nonreflecting" so that they do not contaminate the flow field in the interior of the domain. The proper specification of these boundaries is critical to the stability, accuracy, convergence, and quality of the numerical solution, and has been the topic of considerable research. The need for accurate boundary specification has been underscored in recent years with efforts to apply higher-fidelity methods (DNS, LES) in conjunction with high-order low-dissipation numerical schemes to realistic flow configurations. One of the most popular choices for specifying these boundaries is the characteristics-based boundary condition where the linearized flow field at the boundaries are decomposed into characteristic waves using either one-dimensional Riemann or other multi-dimensional Riemann approximations. The values of incoming characteristics are then suitably modified. The incoming characteristics are specified at the in flow boundaries, and at the out flow boundaries the variation of the incoming characteristic is zeroed out to ensure no reflection. This, however, makes the problem ill-posed requiring the use of an ad-hoc parameter to allow small reflections that make the solution stable. Generally speaking, such boundary conditions work reasonably well when the characteristic flow direction is normal to the boundary, but reflects spurious energy otherwise. An alternative to the characteristic-based boundary condition is to add additional "buffer" regions to the main computational domain near the artificial boundaries, and solve a different set of equations in the buffer region in order to minimize acoustic reflections. One approach that has been used involves modeling the pressure fluctuations as acoustic waves propagating in the far-field relative to a single noise-source inside the buffer region. This approach treats vorticity-induced pressure fluctuations the same as acoustic waves. Another popular approach, often referred to as the "sponge layer," attempts to dampen the flow perturbations by introducing artificial dissipation in the buffer region. Although the artificial dissipation removes all perturbations inside the sponge layer, incoming waves are still reflected from the interface boundary between the computational domain and the sponge layer. The effect of these refkections can be somewhat mitigated by appropriately selecting the artificial dissipation strength and the extent of the sponge layer. One of the most promising variants on the buffer region approach is the Perfectly Matched Layer (PML) technique. The PML technique mitigates spurious reflections from boundaries and interfaces by dampening the perturbation modes inside the buffer region such that their eigenfunctions remain unchanged. The technique was first developed by Berenger for application to problems involving electromagnetic wave propagation. It was later extended to the linearized Euler, Euler and Navier-Stokes equations by Hu and his coauthors. The PML technique ensures the no-reflection property for all waves, irrespective of incidence angle, wavelength, and propagation direction. Although the technique requires the solution of a set of auxiliary equations, the computational overhead is easily justified since it allows smaller domain sizes and can provide better accuracy, stability, and convergence of the numerical solution. In this paper, the PML technique is developed in the context of a high-order spectral-element Discontinuous Galerkin (DG) method. The technique is compared to other approaches to treating the in flow and out flow boundary, such as those based on using characteristic boundary conditions and sponge layers. The superiority of the current PML technique over other approaches is demonstrated for a range of test cases, viz., acoustic pulse propagation, convective vortex, shear layer flow, and low-pressure turbine cascade flow. The paper is structured as follows. We first derive the PML equations from the non{linear Euler equations. A short description of the higher-order DG method used is then described. Preliminary results for the four test cases considered are then presented and discussed. Details regarding current work that will be included in the final paper are also provided.

Mitigation of Adverse Effects Caused by Shock Wave Boundary Layer Interactions Through Optimal Wall Shaping

NASA Technical Reports Server (NTRS)

Liou, May-Fun; Lee, Byung Joon

2013-01-01

It is known that the adverse effects of shock wave boundary layer interactions in high speed inlets include reduced total pressure recovery and highly distorted flow at the aerodynamic interface plane (AIP). This paper presents a design method for flow control which creates perturbations in geometry. These perturbations are tailored to change the flow structures in order to minimize shock wave boundary layer interactions (SWBLI) inside supersonic inlets. Optimizing the shape of two dimensional micro-size bumps is shown to be a very effective flow control method for two-dimensional SWBLI. In investigating the three dimensional SWBLI, a square duct is employed as a baseline. To investigate the mechanism whereby the geometric elements of the baseline, i.e. the bottom wall, the sidewall and the corner, exert influence on the flow's aerodynamic characteristics, each element is studied and optimized separately. It is found that arrays of micro-size bumps on the bottom wall of the duct have little effect in improving total pressure recovery though they are useful in suppressing the incipient separation in three-dimensional problems. Shaping sidewall geometry is effective in re-distributing flow on the side wall and results in a less distorted flow at the exit. Subsequently, a near 50% reduction in distortion is achieved. A simple change in corner geometry resulted in a 2.4% improvement in total pressure recovery.

Investigation of the effects of pressure gradient, temperature and wall temperature ratio on the stagnation point heat transfer for circular cylinders and gas turbine vanes

NASA Technical Reports Server (NTRS)

Nagamatsu, H. T.; Duffy, R. E.

1984-01-01

Low and high pressure shock tubes were designed and constructed for the purpose of obtaining heat transfer data over a temperature range of 390 to 2500 K, pressures of 0.3 to 42 atm, and Mach numbers of 0.15 to 1.5 with and without pressure gradient. A square test section with adjustable top and bottom walls was constructed to produce the favorable and adverse pressure gradient over the flat plate with heat gages. A water cooled gas turbine nozzle cascade which is attached to the high pressure shock tube was obtained to measuse the heat flux over pressure and suction surfaces. Thin-film platinum heat gages with a response time of a few microseconds were developed and used to measure the heat flux for laminar, transition, and turbulent boundary layers. The laminar boundary heat flux on the shock tube wall agreed with Mirel's flat plate theory. Stagnation point heat transfer for circular cylinders at low temperature compared with the theoretical prediction, but for a gas temperature of 922 K the heat fluxes were higher than the predicted values. Preliminary flat plate heat transfer data were measured for laminar, transition, and turbulent boundary layers with and without pressure gradients for free-stream temperatures of 350 to 2575 K and flow Mach numbers of 0.11 to 1.9. The experimental heat flux data were correlated with the laminar and turbulent theories and the agreement was good at low temperatures which was not the case for higher temperatures.

Study of boundary-layer transition using transonic cone Preston tube data

NASA Technical Reports Server (NTRS)

Reed, T. D.; Abu-Mostafa, A.

1982-01-01

Laminar layer Preston tube data on a sharp nose, ten degree cone obtained in the Ames 11 ft TWT and in flight tests are analyzed. During analyses of the laminar-boundary layer data, errors were discovered in both the wind tunnel and the flight data. A correction procedure for errors in the flight data is recommended which forces the flight data to exhibit some of the orderly characteristics of the wind tunnel data. From corrected wind tunnel data, a correlation is developed between Preston tube pressures and the corresponding values of theoretical laminar skin friction. Because of the uncertainty in correcting the flight data, a correlation for the unmodified data is developed, and, in addition, three other correlations are developed based on different correction procedures. Each of these correlations are used in conjunction with the wind tunnel correlation to define effective freestream unit Reynolds numbers for the 11 ft TWT over a Mach number range of 0.30 to 0.95. The maximum effective Reynolds numbers are approximately 6.5% higher than the normal values. These maximum values occur between freestream Mach numbers of 0.60 and 0.80. Smaller values are found outside this Mach number range. These results indicate wind tunnel noise affects the average laminar skin friction much less than it affects boundary layer transition. Data on the onset, extent, and end of boundary layer transition are summarized. Application of a procedure for studying the relative effects of varying nose radius on a ten degree cone at supercritical speeds indicates that increasing nose radius promotes boundary layer transition and separation of laminar boundary layers.

Sensitivity of High-Resolution Simulations of Hurricane Bob (1991) to Planetary Boundary Layer Parameterizations

NASA Technical Reports Server (NTRS)

Braun, Scott A.; Tao, Wei-Kuo

1999-01-01

The MM5 mesoscale model is used to simulate Hurricane Bob (1991) using grids nested to high resolution (4 km). Tests are conducted to determine the sensitivity of the simulation to the available planetary boundary layer parameterizations, including the bulk-aerodynamic, Blackadar, Medium-RanGe Forecast (MRF) model, and Burk-Thompson boundary-layer schemes. Significant sensitivity is seen, with minimum central pressures varying by up to 17 mb. The Burk-Thompson and bulk-aerodynamic boundary-layer schemes produced the strongest storms while the MRF scheme produced the weakest storm. Precipitation structure of the simulated hurricanes also varied substantially with the boundary layer parameterizations. Diagnostics of boundary-layer variables indicated that the intensity of the simulated hurricanes generally increased as the ratio of the surface exchange coefficients for heat and momentum, C(sub h)/C(sub M), although the manner in which the vertical mixing takes place was also important. Findings specific to the boundary-layer schemes include: 1) the MRF scheme produces mixing that is too deep and causes drying of the lower boundary layer in the inner-core region of the hurricane; 2) the bulk-aerodynamic scheme produces mixing that is probably too shallow, but results in a strong hurricane because of a large value of C(sub h)/C(sub M) (approximately 1.3); 3) the MRF and Blackadar schemes are weak partly because of smaller surface moisture fluxes that result in a reduced value of C(sub h)/C(sub M) (approximately 0.7); 4) the Burk-Thompson scheme produces a strong storm with C(sub h)/C(sub M) approximately 1; and 5) the formulation of the wind-speed dependence of the surface roughness parameter, z(sub 0), is important for getting appropriate values of the surface exchange coefficients in hurricanes based upon current estimates of these parameters.

Heat Transfer Due to Shock Wave Turbulent Boundary Layer Interactions on High Speed Weapon Systems

DTIC Science & Technology

1974-04-01

Thermocouple Location. ............ 20 8 Pressure Tap LoceL.lons ............ 23 9 Pitot Pressurp Rake Detail and Profile Measurement Stations... Pitot Pressure Profile Measurements - Pitot pressure measure- ments were made with a NASA Langley five-probe pitot pressure rake at the axial location...defined in Figure 6. Figure 9 is a schematic illustration providing the dimensions of the rake and the pitot pressure profile measure- ment stations. The

Numerical and experimental investigation of VG flow control for a low-boom inlet

NASA Astrophysics Data System (ADS)

Rybalko, Michael

The application of vortex generators (VGs) for shock/boundary layer interaction flow control in a novel external compression, axisymmetric, low-boom concept inlet was studied using numerical and experimental methods. The low-boom inlet design features a zero-angle cowl and relaxed isentropic compression centerbody spike, resulting in defocused oblique shocks and a weak terminating normal shock. This allows reduced external gas dynamic waves at high mass flow rates but suffers from flow separation near the throat and a large hub-side boundary layer at the Aerodynamic Interface Plane (AIP), which marks the inflow to the jet engine turbo-machinery. Supersonic VGs were investigated to reduce the shock-induced flow separation near the throat while subsonic VGs were investigated to reduce boundary layer radial distortion at the AIP. To guide large-scale inlet experiments, Reynolds-Averaged Navier-Stokes (RANS) simulations using three-dimensional, structured, chimera (overset) grids and the WIND-US code were conducted. Flow control cases included conventional and novel types of vortex generators at positions both upstream of the terminating normal shock (supersonic VGs) and downstream (subsonic VGs). The performance parameters included incompressible axisymmetric shape factor, post-shock separation area, inlet pressure recovery, and mass flow ratio. The design of experiments (DOE) methodology was used to select device size and location, analyze the resulting data, and determine the optimal choice of device geometry. Based on the above studies, a test matrix of supersonic and subsonic VGs was adapted for a large-scale inlet test to be conducted at the 8'x6' supersonic wind tunnel at NASA Glenn Research Center (GRC). Comparisons of RANS simulations with data from the Fall 2010 8'x6' inlet test showed that predicted VG performance trends and case rankings for both supersonic and subsonic devices were consistent with experimental results. For example, experimental surface oil flow visualization revealed a significant post-shock separation bubble with flow recirculation for the baseline (no VG) case that was substantially broken up in the micro-ramp VG case, consistent with simulations. Furthermore, the predicted subsonic VG performance with respect to a reduction in radial distortion (quantified in terms of axisymmetric incompressible shape factor) was found to be consistent with boundary layer rake measurements. To investigate the unsteady turbulent flow features associated with the shock-induced flow separation and the hub-side boundary layer, a detached eddy simulation (DES) approach using the WIND-US code was employed to model the baseline inlet flow field. This approach yielded improved agreement with experimental data for time-averaged diffuser stagnation pressure profiles and allowed insight into the pressure fluctuations and turbulent kinetic energy distributions which may be present at the AIP. In addition, streamwise shock position statistics were obtained and compared with experimental Schlieren results. The predicted shock oscillations were much weaker than those seen experimentally (by a factor of four), which indicates that the mechanism for the experimental shock oscillations was not captured. In addition, the novel supersonic vortex generator geometries were investigated experimentally (prior to the large-scale inlet 8'x6' wind tunnel tests) in an inlet-relevant flow field containing a Mach 1.4 normal shock wave followed by a subsonic diffuser. A parametric study of device height and distance upstream of the normal shock was undertaken for split-ramp and ramped-vane geometries. Flow field diagnostics included high-speed Schlieren, oil flow visualization, and Pitot-static pressure measurements. Parameters including flow separation, pressure recovery, centerline incompressible boundary layer shape factor, and shock stability were analyzed and compared to the baseline uncontrolled case. While all vortex generators tested eliminated centerline flow separation, the presence of VGs also increased the significant three-dimensionality of the flow via increased side-wall interaction. The stronger streamwise vorticity generated by ramped-vanes also yielded improved pressure recovery and fuller boundary layer velocity profiles within the subsonic diffuser. (Abstract shortened by UMI.)

Measurements in the near-wall region of a relaxing three-dimensional low speed turbulent air boundary layer

NASA Technical Reports Server (NTRS)

Hebbar, K. S.; Melnik, W. L.

1976-01-01

An experimental investigation was conducted at selected locations of the near-wall region of a three dimensional turbulent air boundary layer relaxing in a nominally zero external pressure gradient behind a transverse hump (in the form of a 30 deg swept, 5-foot chord wing-type model) faired into the side wall of a low speed wind tunnel. Wall shear stresses measured with a flush-mounted hot-film gage and a sublayer fence were in very good agreement with experimental data obtained with two Preston probes. With the upstream unit Reynolds number held constant at 325,000/ft. approximately one-fourth of the boundary layer thickness adjacent to the wall was surveyed with a single rotated hot-wire probe mounted on a specially designed minimum interference traverse mechanism. The boundary layer (approximately 3.5 in thick near the first survey station where the length Reynolds number was 5.5 million) had a maximum crossflow velocity ratio of 0.145 and a maximum crossflow angle of 21.875 deg close to the wall.

Characterization of a highly efficient chevron-shaped anti-contamination device

NASA Astrophysics Data System (ADS)

Fiore, M.; Vermeersch, O.; Forte, M.; Casalis, G.; François, C.

2016-04-01

This paper is devoted to the characterization of an optimized chevron-shaped anti-contamination device (ACD). This device can prevent efficiently the propagation of turbulence from the fuselage along the attachment line (hypothetical streamline that spreads the flow going to suction side and the one going to pressure side) of swept wings and enables the development of a new laminar boundary layer downstream. More specifically, the aim is to prevent boundary-layer transition along the attachment line by a contamination process. This process is characterized by the typical Reynolds number overline{R} and the associated Poll's criterion. Thus, ACD efficiency will be expressed in terms of overline{R} values. Some experiments performed on a new numerically optimized ACD have shown its ability to prevent leading-edge contamination up to overline{R} values close to the natural transition process of the laminar boundary layer along the attachment line. The corresponding stability analysis of the laminar boundary layer is made using the Görtler-Hämmerlin stability approach. The study is completed with the different transition processes that can occur downstream the attachment line, around the airfoil, especially with crossflow analysis.

Performance tests for the NASA Ames Research Center 20 cm x 40 cm oscillating flow wind tunnel

NASA Technical Reports Server (NTRS)

Cook, W. J.; Giddings, T. A.

1984-01-01

An evaluation is presented of initial tests conducted to assess the performance of the NASA Ames 20 cm x 40 cm oscillating flow wind tunnel. The features of the tunnel are described and two aspects of tunnel operation are discussed. The first is an assessment of the steady mainstream and boundary layer flows and the second deals with oscillating mainstream and boundary layer flows. Experimental results indicate that in steady flow the test section mainstream velocity is uniform in the flow direction and in cross section. The freestream turbulence intensity is about 0.2 percent. With minor exceptions the steady turbulent boundary layer generated on the top wall of the test section exhibits the characteristics of a zero pressure gradient turbulent boundary layer generated on a flat plate. The tunnel was designed to generate sinusoidal oscillating mainstream flows. Experiments confirm that the tunnel produces sinusoidal mainstream velocity variations for the range of frequencies (up to 15 Hz). The results of this study demonstrate that the tunnel essentially produces the flows that it was designed to produce.