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Results for “BOUNDARY LAYER SEPARATION”

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NASA NTRS1991

The effect of bleed configuration on shock/boundary layer interactions

The flow field characteristics are simulated numerically in an oblique shock wave/laminar boundary layer interaction for three different bleed configurations. The numerical solution for the flow field is obtained for the strong conservation-law form of the two-dimensional compressible Navier-Stokes equations using an implicit scheme. The computations model the flow in the interaction region and inside the bleed slot for an impinging oblique shock of sufficient strength to cause boundary layer separation on a flat plate boundary layer. The computed results are presented for a normal bleed slot at three different locations, upstream, across and downstream of the shock impingement point. The detailed flow characteristics in the interaction zone and inside the bleed slot are compared for the three cases. The resulting surface pressure and shear stress distributions are also compared for the three bleed cases and for the shock wave boundary layer interaction without bleed.

Hamed, A.↗

NASA NTRS1987

Flowfield measurements in a separated and reattached flat plate turbulent boundary layer

The separation and reattachment of a large-scale, two-dimensional turbulent boundary layer at low subsonic speed on a flat plate has been studied experimentally. The separation bubble was 55 cm long and had a maximum bubble thickness, measured to the height of the mean dividing streamline, of 17 cm, which was twice the thickness of the inlet boundary layer. A combination of laser velocimetry, hot-wire anemometry, pneumatic probing techniques, and flow visualization were used as diagnostics. Principal findings were that an outer inviscid rotational flow was defined which essentially convected over the blockage associated with the inner, viscously dominated bubble recirculation region. A strong backflow region in which the flow moved upstream 100 percent of the time was measured near the test surface over the central 35 percent of the bubble. A laminar backflow boundary layer having pseudo-turbulent characteristics including a log-linear velocity profile was generated under the highly turbulent backflow. Velocity profile shapes in the reversed flow region matched a previously developed universal backflow profile at the upstream edge of the separation region but not in the steady backflow region downstream. A smoke flow visualization movie and hot-film measurements revealed low frequency nonperiodic flapping at reattachment. However, forward flow fraction data at reattachment and mean velocity profiles in the redeveloping boundary layer downstream of reattachment correlated with backward-facing step data when the axial dimension was scaled by the distance from the maximum bubble thickness to reattachment.

Patrick, William P.↗

NASA NTRS2010

Boundary Layers, Transitions and Separation

Effects of roughness in boundary layers have to be addressed. Until adverse pressure gradient effects are understood, roughness will not significantly drive design. Mechanisms responsible for separation not understood. Effects on Zero Pressure Gradient boundary layers (shear stress). Effects on separation in pressure gradient (prediction of separation). Effect on scalar transport (heat transfer) not understood. Model for skin friction needed in simulations - first grid point likely to be in buffer layer. Definition of roughness important for useful experiments. A lot of validation experiments will be needed. How to get to ks for roughness of engineering interest? - depends on wavelength height, etc. for engineering interest? Re-discovering the wheel should be avoided: existing knowledge (theoretical and experimental) should find its way into the engineering models. It is a task of the industry to filter out the existing information in the literature for results relevant to its application, being external or internal.

Source record↗

NASA NTRS2016

Passive Rocket Diffuser Testing: Reacting Flow Performance of Four Second-Throat Geometries

Second-throat diffusers serve to isolate rocket engines from the effects of ambient back pressure. As one of the nation's largest rocket testing facilities, the performance and design limitations of diffusers are of great interest to NASA's Stennis Space Center. This paper describes a series of tests conducted on four diffuser configurations to better understand the effects of inlet geometry and throat area on starting behavior and boundary layer separation. The diffusers were tested for a duration of five seconds with a 1455-pound thrust, LO2/GH2 thruster to ensure they each reached aerodynamic steady state. The effects of a water spray ring at the diffuser exits and a water-cooled deflector plate were also evaluated. Static pressure and temperature measurements were taken at multiple axial locations along the diffusers, and Computational Fluid Dynamics (CFD) simulations were used as a tool to aid in the interpretation of data. The hot combustion products were confirmed to enable the diffuser start condition with tighter second throats than predicted by historical cold-flow data or the theoretical normal shock method. Both aerodynamic performance and heat transfer were found to increase with smaller diffuser throats. Spray ring and deflector cooling water had negligible impacts on diffuser boundary layer separation. CFD was found to accurately capture diffuser shock structures and full-flowing diffuser wall pressures, and the qualitative behavior of heat transfer. However, the ability to predict boundary layer separated flows was not consistent.

Jones, Daniel R.↗

NASA NTRS1972

Analytical investigation of conical diffusers

A study was made to determine if existing data for straight-channel single-plane divergence diffusers together with boundary-layer methods could be practically employed to predict conical diffuser performance in the high-speed flow regime. A semiempirical correlation technique, which uses the experimental performance of a straight-channel single-plane divergence diffuser, and a momentum integral boundary-layer method were employed to predict conical diffuser performance. Neither of these methods is applicable when the boundary layer separates. The momentum integral method was used to study the relative importance of various boundary layer parameters in correlating unstalled diffuser performance. Other investigators have found that the best diffuser performance occurs when the diffuser is operating in the transitory stall regime. Since the analytical methods presented in this report are not applicable after the point of boundary-layer separation, a criterion was established from examination of experimental tests which allows the designer to ascertain approximate diffuser performance in the transitory stall regime.

Means, J. L.↗

NASA NTRS1978

Upgraded viscous flow analysis of multielement airfoils

A description of an improved version of the NASA/Lockheed multielement airfoil analysis computer program is presented. The improvements include several major modifications of the aerodynamic model as well as substantial changes of the computer code. The modifications of the aerodynamic model comprise the representation of the boundary layer and wake displacement effects with an equivalent source distribution, the prediction of wake parameters with Green's lag-entrainment method, the calculation of turbulent boundary layer separation with the method of Nash and Hicks, the estimation of the onset of confluent boundary layer separation with a modified form of Goradia's method, and the prediction of profile drag with the formula of Squire and Young. The paper further describes the modifications of the computer program for which the structured approach to computer software development was employed. Important aspects of the structured program development such as the functional decomposition of the aerodynamic theory and its numerical implementation, the analysis of the data flow within the code, and the application of a pseudo code are discussed.

Brune, G. W.↗

NASA NTRS1979

Upgraded viscous flow analysis of multi-element airfoils

A description of an improved version of the NASA/Lockheed multi-element airfoil analysis computer program is presented. The improvements include several major modifications of the aerodynamic model as well as substantial changes of the computer code. The modifications of the aerodynamic model comprise the representation of the boundary layer and wake displacement effects with an equivalent source distribution, the prediction of wake parameters with Green's lag-entrainment method, the calculation of turbulent boundary layer separation with the method of Nash and Hicks, the estimation of the onset of confluent boundary layer separation with a modified form of Goradia's method, and the prediction of profile drag with the formula of Squire and Young. The modifications of the computer program for which the structured approach to computer software development was employed are also described. Important aspects of the structured program development such as the functional decomposition of the aerodynamic theory and its numerical implementation, the analysis of the data flow within the code, and the application of a pseudo code are discussed.

Brune, G. W.↗

NASA NTRS1985

The role of freestream turbulence scale in subsonic flow separation

The clarification of the role of freestream turbulence scale in determining the location of boundary layer separation is discussed. Modifications to the test facility were completed. Wind tunnel flow characteristics, including turbulence parameters, were determined with two turbulence generating grids, as well as no grid. These results are summarized. Initial results on the role of scale on turbulent boundary layer separation on the upper surface of an airfoil model are also discussed.

Potter, J. L.↗

NASA NTRS1983

A time-dependent approach for calculating steady inverse boundary-layer flows with separation

An unsteady inverse boundary-layer method is developed which can be used to calculate steady flows with separation. Two versions of Keller's box method with the Mechul function formulation developed by Cebeci (1976) are employed, depending on the complexity of the flow. The regular box is employed in regions of positive streamwise velocity component u, whereas the zigzag box is employed in regions where u becomes negative (t greater than 0). The regular box with the FLARE approximation is employed when t = 0 and u becomes negative in some region across the layer. Results of calculations show that the use of a time-dependent inverse boundary-layer method in which time is used as an iteration parameter provides a good approach in improving the accuracy of the solutions obtained from the FLARE approximation.

Cebeci, T.↗

NASA NTRS1986

A new, improved method for separating turbulent boundary layer for aerodynamic performance prediction of trailing edge stall airfoils

A method based on integral techniques is presented to determine the turbulent separation location and separation pressure for trailing edge stall airfoils exhibiting extensive separated flow regions during prestall and poststall conditions. The method and experimental laser velocimetry data in the turbulence amplification region are used to establish a relationship between fluctuating velocity components and mean velocity in the boundary layer. The method is demonstrated for symmetrical, cambered, conventional and supercritical airfoils in the 5-to-21-percent thickness ratio range, and over a Reynolds number range of 2 to 22 million. Good agreement with experimental data is found, and the method is used in conjunction with the NASA Multi-Component Airfoil Computer Program to predict maximum L/D ratios and maximum section/lift coefficients.

Goradia, S.↗

NASA NTRS1995

An experimental study of a turbulent boundary layer in the trailing edge region of a circulation-control airfoil

This report describes the development of a new near-wall LDV technique and its use to measure the full turbulent stress tensor in a highly three-dimensional separated boundary layer flow. Laser Doppler velocimetry, with small diameter laser beams allowing small beam-crossing angles (i.e., larger fringe spacings), was used, resulting in measurements close to solid surfaces. Dennis Johnson of NASA/Ames combined the advantages offered by FFT signal processors with a beam-steering probe inserted in the flow to devise a new self-aligning, near-wall LDV approach capable of making direct measurements of the transverse velocity component. This new technique was applied to a 2D, zero-pressure-gradient, flat-plate boundary layer in the High Reynolds Number Pilot Channel at Ames. Accurate and repeatable measurements of U and uu to within 30 mm of the wall were obtained. Next, the near-wall LDV technique was employed in a 3D separated boundary layer flow, consisting of a 2-dimensional wall boundary layer approaching a right-angle wedge. The resulting lateral and adverse streamwise pressure gradients generated significant spanwise velocities and turbulent stresses, and a large reverse-flow region. The near-wall LDV was modified further to facilitate measurements in the 3D Wedge Flow.

Brown, Jeffrey↗

NASA NTRS1982

Modeling parameter influences on MHD swirl combustion nozzle design

Attention is given to a research project which has the goal to develop a two-stage slagging gasifier-combustor in the form of a high-intensity combustor, taking into account a suitable aerodynamic design of the second stage nozzle which will prevent the separation of the boundary layer as the flow turns from axial to radial direction. The specific objectives of the present investigation are to test the effect of various second-stage nozzle geometries, flow rates, swirl number, and distribution in the first and second stages upon the corresponding flowfield in the second stage. Special emphasis is given to the avoidance of boundary layer separation as the flow turns from axial to radial direction into the MHD disk generator.

Lilley, D. G.↗

NASA NTRS1987

Transition from laminar to turbulent flow in separated shear layers

The location of boundary layer separation may be predicted by either semiempirical criteria or by various computer codes written for such calculations; attention is presently given to the streamwide length from separation to transition, denoted by Re(sL). It is demonstrated that the unit Reynolds number, a feature of the transition phenomenon implicit in the data of Schmidt et al. (1985) and overlooked in previous investigations, must be considered in future work.

Potter, J. Leith↗

NASA NTRS1995

Reduction of fluctuating pressure loads in shock wave turbulent boundary layer interactions

Fluctuating surface pressure measurements have been made to investigate the effectiveness of boundary layer separators (BLS's) in reducing the fluctuating pressure loads produced by separated shock wave turbulent boundary layer interactions. Measurements have been made under unswept and swept compression corner interactions in a Mach 5 flow. BLS's fix the separation location and eliminate the large-amplitude, low-frequency fluctuating pressure loads upstream of the compression corners. The loads on the unswept compression corner face are reduced by as much as 59%. The BLS's also shift the mean pressure distribution on the unswept corner face in the streamwise direction. Results show that the loads on the corner face vary with the BLS height and the distance between the BLS and the compression corner. Suggestions for the optimum placement and the use of the BLS's are also made.

Barter, John W.↗

NASA NTRS2017

Shed Vortex Structure and Phase-Averaged Velocity Statistics in Symmetric/Asymmetric Turbulent Flat Plate Wakes

The near wake of a flat plate is investigated via direct numerical simulations (DNS). Many earlier experimental investigations have used thin plates with sharp trailing edges and turbulent boundary layers to create the wake. This results in large theta divided by D (sub TE) values (theta is the boundary layer momentum thickness towards the end of the plate and D (sub TE) is the trailing edge thickness). In the present study the emphasis is on relatively thick plates with circular trailing edges (CTE) resulting in theta divided by 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 x 10 (sup 6) 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, a second with turbulent and laminar boundary layers on the upper and lower surfaces, respectively (asymmetric case, Case TL). The data and understanding obtained is 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 1.27 weaker in terms of peak phase-averaged spanwise vorticity at 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 divided by D) that occurs nears 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 divided by D. Both these effects are examined in detail and the important contributors are identified.

direct numerical simulations↗

NASA NTRS1993

Numerical study of shock-wave/boundary-layer interactions with bleed

A numerical study was conducted to investigate how bleed through a two-dimensional slot affects shock-wave induced, boundary-layer separation on a flat plate. This study is based on the ensemble-averaged, compressible, Navier-Stokes equations closed by the Baldwin-Lomax, algebraic turbulence model. The algorithm used to obtain solutions was the implicit, partially split, two-factored scheme of Steger. This study examined the effects of the following parameters in controlling shock-wave induced flow separation: location of slot in relation to where the incident shock wave impinged on the boundary layer, size of slot in relation to the boundary-layer thickness, number of slots, spacings between slots, and strength of the incident shock wave. This study also showed the nature of the very complex flowfield about the slot or slots and how the plenum affects the bleed process. The results of this study are relevant to problems where bleed is used to control shock-wave induced, boundary-layer separation (e.g., inside jet engine inlets and wind tunnels).

Hahn, T. O.↗

NASA NTRS2011

Direct Numerical Simulations of Transitional/Turbulent Wakes

The interest in transitional/turbulent wakes spans the spectrum from an intellectual pursuit to understand the complex underlying physics to a critical need in aeronautical engineering and other disciplines to predict component/system performance and reliability. Cylinder wakes have been studied extensively over several decades to gain a better understanding of the basic flow phenomena that are encountered in such flows. Experimental, computational and theoretical means have been employed in this effort. While much has been accomplished there are many important issues that need to be resolved. The physics of the very near wake of the cylinder (less than three diameters downstream) is perhaps the most challenging of them all. This region comprises the two detached shear layers, the recirculation region and wake flow. The interaction amongst these three components is to some extent still a matter of conjecture. Experimental techniques have generated a large percentage of the data that have provided us with the current state of understanding of the subject. More recently computational techniques have been used to simulate cylinder wakes, and the data from such simulations are being used to both refine our understanding of such flows as well as provide new insights. A few large eddy and direct numerical simulations (LES and DNS) of cylinder wakes have appeared in the literature in the recent past. These investigations focus on the low Reynolds number range where the cylinder boundary layer is laminar (sub-critical range). However, from an engineering point of view, there is considerable interest in the situation where the upper and/or lower boundary layer of an airfoil is turbulent, and these turbulent boundary layers separate from the airfoil to contribute to the formation of the wake downstream. In the case of cylinders, this only occurs at relatively large unit Reynolds numbers. However, in the case of airfoils, the boundary layer has the opportunity to transition to turbulence on the airfoil surface at a relatively lower unit Reynolds number because the characteristic length of the airfoil is typically one to two orders of magnitude larger than the trailing edge diameter. This transition to turbulence would occur unless there is a strong favorable pressure gradient that results in the boundary layer remaining laminar or transitional over the surface of the airfoil. This presentation will focus on two direct numerical simulations that have been performed at NASA ARC. The first is of a cylinder wake with laminar separating boundary layers. The second is the wake of a flat plate with a circular trailing edge. The upper and lower plate surface boundary layers are both turbulent and statistically identical. Thus the computed wake is symmetric in a statistical sense. This flow is more representative of airfoil wakes than cylinder wakes. Results from the two simulations including flow visualization and turbulence statistics in the near wake will be presented at the seminar.

wakes↗

NASA NTRS1992

Separation control on high Reynolds number multi-element airfoils

Small surface-mounted vortex generators were investigated as means for the control of a boundary-layer separation on a 2D single-flap three-element high-lift system at near-flight Reynolds numbers and in landing configurations. Wind-tunnel results obtained for small vane-type vortex generators mounted on a multielement airfoil showed that vortex generators as small as 0.18 percent of total chord can effectively reduce or eliminate boundary-layer separation on the flap at approach conditions. It was found that both the outerrotating and the corotating streamwise vortices were effective in reducing flow separation.

Lin, John C.↗