Showing papers on "Flow separation published in 1992"

Boundary Layer Analysis

Abstract: Notation. 1. Introduction to Viscous Flows. 2. Integral Equations and Solutions for Laminar Flow. 3. Differential Equations of Motion for Laminar Flow. 4. Exact and Numerical Solutions for Laminar Constant-Property Incompressible Flows. 5. Compressible Laminar Boundary Layers. 6. Transition to Turbulent Flow. 7. Wall-Bounded, Incompressible Turbulent Flows. 8. Internal Flows. 9. Free Shear Flows. 10. Wall-Bounded Turbulent Flows with Variable Density and Heat and Mass Transfer. 11. Three-Dimensional External Boundary Layer Flows. Appendix A: Laminar Thermophysical Properties for Selected Fluids. Appendix B: Computer Codes for Students. References. Index.

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

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

The flow over a backward-facing step under controlled perturbation: laminar separation

Abstract: The flow over a backward-facing step with laminar separation was investigated experimentally under controlled perturbation for a Reynolds number of 11000, based on a step height h and a free-stream velocity UO. The reattaching shear layer was found to have two distinct modes of instability: the ‘shear layer mode’ of instability at Stθ ≈ 0.012 (Stθ ≡ fθ/UO, θ being the momentum thickness at separation and f the natural roll-up frequency of the shear layer); and the ‘step mode’ of instability at Sth ≈ 0.185 (Sth ≡ fh/U0). The shear layer instability frequency reduced to the step mode one via one or more stages of a vortex merging process. The perturbation increased the shear layer growth rate and the turbulence intensity and decreased the reattachment length compared to the unperturbed flow. Cross-stream measurements of the amplitudes of the perturbed frequency and its harmonics suggested the splitting of the shear layer. Flow visualization confirmed the shear layer splitting and showed the existence of a low-frequency flapping of the shear layer.

Turbulent flow between concentric rotating cylinders at large Reynolds number.

Abstract: Turbulent Taylor vortex flow is studied in experiments for Reynolds numbers ${10}^{3}$R${10}^{6}$. Simple scaling of the torque with Reynolds number is m/Inot observed for any range of R, although the characteristic time scales and the transport of passive scalars are found to scale with the global torque measurements. Above a nonhysteretic transition observed at R=1.3\ifmmode\times\else\texttimes\fi{}${10}^{4}$, the torque has a Reynolds number dependence similar to the drag observed in wall-bounded shear flows such as pipe flow and flow over a flat plate.

On the evolution of a turbulent boundary layer induced by a three-dimensional roughness element

Abstract: An experimental investigation is described which has as its objectives the extension of the technical data base pertaining to roughness-induced transition and the advancement of the understanding of the physical processes by which three-dimensional roughness elements induce transition from laminar to turbulent flow in boundary layers The investigation was carried out primarily with single hemispherical roughness elements surface mounted in a well-characterized zero-pressure-gradient laminar boundary layer on a flat plate The critical roughness Reynolds number at which turbulence is regarded as originating at the roughness was determined for the roughness elements herein considered and evaluated in the context of data existing in the literature The effect of a steady and oscillatory free-stream velocity on eddy shedding was also investigated The Strouhal behaviour of the ‘hairpin’ eddies shed by the roughness and role they play in the evolution of a fully developed turbulent boundary layer, as well as whether their generation is governed by an inflexional instability, are examined Distributions of mean velocity and intensity of the u-fluctuation demonstrating the evolution toward such distributions for a fully developed turbulent boundary layer were measured on the centreline at Reynolds numbers below and above the critical Reynolds number of transition A two-region model is postulated for the evolutionary change toward a fully developed turbulent boundary layer: an inner region where the turbulence is generated by the complex interaction of the hairpin eddies with the pre-existing stationary vortices that lie near the surface and are inherent to a flow about a three-dimensional obstacle in a laminar boundary layer; and an outer region where the hairpin eddies deform and generate turbulent vortex rings The structure of the resulting fully developed turbulent boundary layer is discussed in the light of the proposed model for the evolutionary process

Control of low-speed turbulent separated flow using jet vortex generators

Abstract: A parametric study has been performed with jet vortex generators to determine their effectiveness in controlling flow separation associated with low-speed turbulent flow over a two-dimensional rearward-facing ramp. Results indicate that flow-separation control can be accomplished, with the level of control achieved being a function of jet speed, jet orientation (with respect to the free-stream direction), and jet location (distance from the separation region in the free-stream direction). Compared to slot blowing, jet vortex generators can provide an equivalent level of flow control over a larger spanwise region (for constant jet flow area and speed).

Forcing level effects of internal acoustic excitation on the improvement of airfoil performance

Abstract: The effects of internal acoustic excitation on the leading-edge, separated boundary layers and the aerodynamic performance of NACA 633-018 cross section airfoil are examined as a function of forcing level and forcing frequency of the introduced acoustics. Tests are separately conducted in two suction, open-typed wind tunnels at the Reynolds number of 3.0 x 10 s for the measurements and 1.0 x 10 4 for the visualization. Results indicate that the flow separation is delayed at the angles of attack higher than the stalled angle of small level excitation with the forcing frequency fe near the shear layer instability frequency ft. As the forcing level is increased to some extent, the velocity fluctuations around the slot exit are demonstrated to be the primary governing parameter for modifying the separated boundary layers. Data also show that the effective forcing frequency (and the Strouhal number, 50 extends over wider range as compared to the lower level excitation. Meanwhile, the pressure distributions on the airfoil surface exhibit recovery behaviors with different forcing frequencies. The corresponding boundary layers are visualized to be reattached to the surface to form a recirculation region when the airfoil is around at the stalled angles.

Transverse flow over a wavy cylinder

Abstract: Transverse flow over a wavy cylinder was investigated experimentally; surface‐pressure distributions and flow visualizations were obtained for a set of wavy cylinders with different axial wavelengths. Significant spanwise pressure gradients were present, resulting in three‐dimensional separation lines and the formation of streamwise trailing vortex structures near the geometric nodes. Despite the symmetry of the geometries, the separated flow structures near the geometric nodes were distinctly asymmetric a large fraction of time. Integration of the pressure data revealed greater sectional drag coefficients at the geometric nodes than at the geometric saddles.

An experimental investigation of the flow in a diffusing S-duct

Abstract: Compressible, subsonic flow through a diffusing S-duct has been experimentally investigated. Benchmark aerodynamic data are presented for flow through a representative S-duct configuration. The collected data would be beneficial to aircraft inlet designers and is suitable for the validation of computational codes. Measurements of the 3D velocity field and total and static pressures were obtained at five cross-sectional planes. Surface static pressures and flow visualization also helped to reveal flow field characteristics. All reported tests were conducted with an inlet centerline Mach number of 0.6 and a Reynolds number, based on the inlet centerline velocity and duct inlet diameter, of 2.6 x 10(exp 6). The results show that a larger region of streamwise flow separation occurred within the duct. Details about the separated flow region, including mechanisms which drive this complicated flow phenomenon, are discussed. Transverse velocity components indicate that the duct curvature induces strong pressure driven secondary flows, which evolve into a large pair of counter-rotating vortices. These vortices convect the low momentum fluid of the boundary layer towards the center of the duct, degrading both the uniformity and magnitude of the total pressure profile.

The standing hydraulic jump: theory, computations and comparisons with experiments

Abstract: In this theoretical and computational study of the flow of a liquid layer, under the influence of surface tension and gravity most notably, the nonlinear equations governing an interaction between viscous effects and the effects of surface tension, gravity and streamline curvature for the limit of large Reynolds numbers are derived. The aim is to make a comparison between the predictions of this theory and the experiments of Craik et al. on the axisymmetric hydraulic jump. Such a jump is commonly encountered in the everyday context of the initial filling of a kitchen sink, for example, and it is found in the present work that initially all the effects listed above can play a primary role in practice in the local jump phenomenon. As a first step here, the flow of the layer over a small obstacle is considered. It is seen that as surface tension becomes increasingly significant the upstream influence becomes more wave-like. Second, calculations and analysis of the nonlinear free interaction are presented and show wave-like behaviour upstream, followed downstream by a depth profile not unlike that in the typical hydraulic jump. The effects of gravity dominate those of surface tension downstream. Finally, comparisons are made with the experiments and show fair quantitative agreement, supporting the present proposition that these hydraulic jumps are caused by boundary-layer separation due to a viscous–inviscid interaction forced by downstream boundary conditions on, in this case, a fully developed, high-Froude-number liquid layer.

Boundary element method for the analysis of the unsteady flow aroundextreme propeller geometries

Abstract: The unsteady flow around a marine propeller subject to a spatially nonuniform inflow is analysed by utilizing a time-marching potential-based low-order boundary element method. Constant strength dipole or source distributions are used on each of the quadrilateral panels representing the propeller blades and their trailing wakes. Linear dipole distributions are used at the first wake panels adjacent to the blade trailing edge in order to render the method insensitive to the time step size. An efficient algorithm is implemented in order to ensure an explicit Kutta condition (i.e., pressure equality) at the blade trailing edge at each time step. The numerical method is shown to be consistent with known analytic solutions for two-dimensional unsteady flows. The robustness of the method is tested in the case of a highly skewed propeller in a given wake inflow and the results are shown to converge quickly with number of panels for a broad range of reduced frequencies.

Effect of acoustic excitation on stalled flows over an airfoil

Abstract: The effect of acoustic excitation on poststalled flows over an airfoil, ie, flows that are fully separated from near the leading edge, is investigated The excitation results in a tendency toward reattachment, which is accompanied by an increased lift and reduced drag, although the flow may still remain fully separated It is found that with increasing excitation amplitude, the effect becomes more pronounced but shifts to a Strouhal number which is much lower than that expected from linear, inviscid instability of the separated shear layer

Distortion of a flat-plate boundary layer by free-stream vorticity normal to the plate

Abstract: A nominally uniform flow over a semiinfinite flat plate is considered. The analysis shows how a small streamwise disturbance in the otherwise uniform flow ahead of the plate is amplified by leading-edge bluntness effects and eventually leads to a small-amplitude but nonlinear spanwise motion far downstream from the leading edge of the plate. This spanwise motion is then imposed on the viscous boundary-layer flow at the surface of the plate - causing an order-one change in its profile shape. This ultimately reduces the wall shear stress to zero, causing the boundary layer to undergo a localized separation, which may be characterized as a kind of bursting phenomenon that could be related to the turbulent bursts observed in some flat-plate boundary-layer experiments.

Measurements in a leading-edge separation bubble due to a simulated airfoil ice accretion

Abstract: The separation bubble formed on an airfoil at low Reynolds number behind a simulated leading-edge glaze ice accretion is studied experimentally. Surface pressure and split hot-film measurements as well as flow visualization studies of the bubble reattachment point are reported. The simulated ice generates an adverse pressure gradient that causes a laminar separation bubble of the long bubble type to form. The boundary layer separates at a location on the ice accretion that is independent of angle of attack and reattaches at a downstream location 5-40 percent chord behind the leading edge, depending on the angle of attack. Velocity profiles show a large region of reverse flow that extends up from the airfoil surface as much as 2.5 percent chord. After reattachment, a thick distorted turbulent boundary layer exists. The separation bubble growth and reattachment are clearly seen in the plots of boundary-layer momentum thickness vs surface distance. Local minima and maxima in the boundary-layer momentum thickness development compare well with the shear layer transition point as indicated by the surface pressures and the reattachment point as measured from surface oil flow, respectively.

Physics of Vortical Flows

Abstract: Separation in three-dimensional flows leads to the formation of vortical structures resulting from rolling up of the viscous flow "sheet," initially contained in a thin boundary layer, which springs up from the surface into the outer perfect fluid flow. A clear physical understanding of this phenomenon must be based on a rational analysis of the flowfield structure using the critical-point theory. With the help of this theory, it is possible to interpret correctly the surface flow patterns that constitute the imprints of the outer flow and to give a rational and coherent description of the vortical system generated by separation. This kind of analysis is applied to separated flows forming on typical obstacles, the field of which has been thoroughly studied by means of visualizations and probings using multihole pressure probes and laser velocimetry. Thus, the skin friction line patterns of a transonic channel flow and of a multibody launcher are interpreted. Then, the vortical systems of a delta wing and an afterbody at an incidence are considered. The last two configurations are a missile fuselage-type body and an oblate ellipsoid. I. Introduction F LIGHT at high-incidenc e of combat aircraft or hypersonic vehicles during re-entry, as well as that of tactical missiles, raises practical interest on the study of three-dimensional separated flows. Applications also concern internal flows, in particular air intakes and turbomachines in which the often complex geometry of the channel and the existence of shock waves almost inevitably lead to boundary-laye r separation. In three-dimensional flows, separation entails the formation of vortical structures—frequently, but improperly, called vortices to simplify—form ed by rolling up of the viscous flow "sheet," previously confined in a thin layer attached to the wall, which suddenly springs into the outer nondissipative flow. Although it has been known for a long time, this phenomenon is still incompletely understood from a physical point of view and it is delicate to model due to the flowfield complexity, all the components of which are difficult to capture properly. Many predictive methods are based on perfect fluid models, the first of which use the vortex sheet concept. Such a sheet is defined as a surface of tangential discontinuity for the velocity field. The computational method can use different schemes: doublets, vortex filaments, vortex particles, and so forth. Publications in this domain are too numerous to be cited here. A greater accuracy in flow prediction can be obtained in the solution of the complete Euler equations, which allows, in theory, automatic capture of sheet-like disconti

Inlet distortion effects in aircraft propulsion system integration

Abstract: A tutorial survey of inlet flow distortion effects on engine performance and stability is presented. Inlet distortions in aero engines arise through a variety of causes. They can be essentially steady, due to non-axisymmetric intake duct geometry, or time varying, for example from flow separation off the lip of the inlet during maneuvers or shock-induced separation during supersonic flight. Whatever the cause, the result is generally a decrease in performance and, more importantly, a lessening of the stable flow range of the compressor. The distortions are generally three-dimensional. It is an extremely useful simplification to break them, at least conceptually, into radial and circumferential non-uniformities and approach each separately. Purely radial distortions can be treated by the methods that were developed for designing compressors in nominally axisymmetric inlet flow, and this type of distortion will be only briefly discussed. Circumferential non-uniformities, however, introduce additional fluid dynamic features into the analysis of compressor behavior and often have the larger impact on performance and stability. Thus we concentrate mainly on the effects of steady circumferential inlet flow distortion.

Sources of High Alpha Vortex Asymmetry at Zero Sideslip

Abstract: A review of existing experimental results for slender bodies and delta wings, tested at high angles of attack, reveals that no physical evidence exists that vortex asymmetry on slender pointed bodies or delta wings has ever occurred through the so-called hydrodynamic instability process. It will be shown that in the numerous tests performed, asymmetric flow separation and/or asymmetric flow reattachment, were the flow mechanisms triggering the vortex asymmetry. Slender wing rock is found to result from a basic lack of roll damping, existing for attached leading-edge vortices, and the vortex-asymmetry is generated at nonzero roll angle, i.e., for asymmetric flow conditions. Nomenclature b = wingspan c = reference length, d CQ = delta wing center chord d = maximum diameter of body of revolution € = rolling moment, coefficient C€ = €/(p00U£/2) Re = Reynolds number based on d and freestream conditions S — reference area, ird2/4 or projected wing area U = horizontal velocity Y = side force, coefficient CY = Y/(pJJl/2) a = angle of attack OA = aPex half-angle Oc = cone half-angle A = leading-edge sweep p = air density = body roll angle Subscripts A = apex c = cone oo = freestream conditions

A Power-Law Formulation of Laminar Flow in Short Pipes

Abstract: In the quantification of air flow through penetrations in buildings, it is necessary to be able to characterize the flow without detailed knowledge of the geometry of the paths. At the conditions typical of buildings, the flow regime is partially developed laminar flow. This report develops a theoreticfal description of the hydrodynamic relationship based an a power-law representation between the air flow and applied pressure for laminar flow in short pipes

Scaling of adverse-pressure-gradient turbulent boundary layers

Abstract: An asymptotic analysis is developed for turbulent boundary layers in strong adverse pressure gradients. It is found that the boundary layer divides into three distinguishable regions: these are the wall layer, the wake layer and a transition layer. This structure has two key differences from the zero-pressure-gradient boundary layer: the wall layer is not exponentially thinner than the wake; and the wake has a large velocity deficit, and cannot be linearized. The mean velocity profile has a y½ behaviour in the overlap layer between the wall and transition regions.The analysis is done in the context of eddy viscosity closure modelling. It is found that k-e-type models are suitable to the wall region, and have a power-law solution in the y½ layer. The outer-region scaling precludes the usual e-equation. The Clauser, constant-viscosity model is used in that region. An asymptotic expansion of the mean flow and matching between the three regions is carried out in order to determine the relation between skin friction and pressure gradient. Numerical calculations are done for self-similar flow. It is found that the surface shear stress is a double-valued function of the pressure gradient in a small range of pressure gradients.

On Turbulent Flows Dominated by Curvature Effects

Abstract: A technique for improving the numerical predictions of turbulent flows with the effect of streamline curvature is developed. Separated flows and the flow in a curved duct are examples of flowfields where streamline curvature plays a dominant role. New algebraic formulations for the eddy viscosity incorporating the k-epsilon turbulence model are proposed to account for various effects of streamline curvature. The loci of flow reversal of the separated flows over various backward-facing steps are employed to test the capability of the proposed turbulence model in capturing the effect of local curvature.

Numerical investigation of vortex breakdown on a delta wing

Abstract: A numerical investigation of leading-edge vortex breakdown on a delta wing at high angles of attack is presented. The analysis has been restricted to low-speed flows on a flat-plate wing with sharp leading edges. Both Euler and Navier-Stokes (assuming fully laminar and turbulent flows) equations have been used in this study and the results are compared against experimental data. Predictions of vortex breakdown progression with angle of attack with both Euler and Navier-Stokes equations are shown to be consistent with the experimental data. However, the Navier-Stokes predictions show significant improvements in breakdown location at angles of attack where the vortex breakdown approaches the wing apex. The location of the primary vortex and the level of vorticity in the prebreakdown regions are affected very little by the viscous effects. In the postbreakdown regions, however, the levels of vorticity in the primary vortex have increased differences between the Euler and Navier-Stokes solutions. Navier-Stokes solutions indicate the presence of a secondary vortex even after the primary vortex is burst. The predicted trajectories of the primary vortex are in very good agreement with the test data with the laminar solutions providing the overall best comparison.

Control of leading-edge vortices on a delta wing

Abstract: The unsteady flow structure of leading-edge vortices on a delta wing has been investigated using new types of experimental techniques, in order to provide insight into the consequences of various forms of active control. These investigations involve global control of the entire wing and local control applied at crucial locations on or adjacent to the wing. Transient control having long and short time-scales, relative to the convective time-scale C/U(sub infinity), allows substantial modification of the unsteady and time-mean flow structure. Global control at long time-scale involves pitching the wing at rates an order of magnitude lower than the convective time-scale C/U(sub infinity), but at large amplitudes. The functional form of the pitching maneuver exerts a predominant influence on the trajectory of the feeding sheet, the instantaneous vorticity distribution, and the instantaneous location of vortex breakdown. Global control at short time-scales of the order of the inherent frequency of the shear layer separating from the leading-edge and the natural frequency of vortex breakdown shows that 'resonant' response of the excited shear layer-vortex breakdown system is attainable. The spectral content of the induced disturbance is preserved not only across the entire core of the vortex, but also along the axis of the vortex into the region of vortex breakdown. This unsteady modification results in time-mean alteration of the axial and swirl velocity fields and the location of vortex breakdown. Localized control at long and short time-scales involves application of various transient forms of suction and blowing using small probes upstream and downstream of the location of vortex breakdown, as well as distributed suction and blowing along the leading-edge of the wing applied in a direction tangential to the feeding sheet. These local control techniques can result in substantial alteration of the location of vortex breakdown; in some cases, it is possible to accomplish this without net mass addition to the flow field.

Simplified Linear Stability Transition Prediction Method for Separated Boundary Layers

Abstract: An existing transition prediction method for attached, two-dimensional, incompressible boundary layers based on linear stability analysis is extended to separated, two-dimensional, incompressible boundary layers such as those found in laminar (transitional) separation bubbles. It is shown why the present method, which tracks the growth of disturbances at many different frequencies, is more accurate than the so-called envelope methods for nonsimilar boundary-layer developments. Reliance on a database of precalculated stability characteristics of known velocity profiles makes this method much faster than traditional stability calculations of similar accuracy. The Falkner-Skan self-similar profiles are used for attached flow, and a new, very general family of profiles is used for separated flow. Comparisons with measured transition locations inside the bubble show good agreement over the range of chord Reynolds numbers and airfoil angles of attack of interest.

A Viscous Flow Study of Shock-Boundary Layer Interaction, Radial Transport, and Wake Development in a Transonic Compressor

Abstract: A numerical study based on the 3D Reynolds-averaged Navier-Stokes equation has been conducted to investigate the detailed flow physics inside a transonic compressor. 3D shock structure, shock-boundary layer interaction, flow separation, radial mixing, and wake development are all investigated at design and off-design conditions. Experimental data based on laser anemometer measurements are used to assess the overall quality of the numerical solution. An additional experimental study to investigate end-wall flow with a hot-film was conducted, and these results are compared with the numerical results. Detailed comparison with experimental data indicates that the overall features of the 3D shock structure, the shock-boundary layer interaction, and the wake development are all calculated very well in the numerical solution. The numerical results are further analyzed to examine the radial mixing phenomena in the transonic compressor. A thin sheet of particles is injected in the numerical solution upstream of the compressor. The movement of particles is traced with a 3D plotting package. This numerical survey of tracer concentration reveals the fundamental mechanisms of radial transport in this transonic compressor.

Turbulence model effects on separated flow about a prolate spheroid

Abstract: The three-dimensional separated flow about a prolate spheroid at high incidence is numerically investigated using the F3D thin-layer Navier-Stokes code. The effect of different turbulence models on the flowfield solution and the characteristics of the predicted flow are analyzed. The models used in this study are the Baldvvin-Lomax algebraic model, the Baldvvin-Lomax model as modified for crossflow separation by Degani and Schiff, and a modified version of the Johnson-King model with and without the Degani-Schiff crossflow modifications applied. The Johnson-King model is applied to assess the importance of modeling nonequilibrium effects in predicting flow about a slender body at high incidence. The computations are made for steady-state, fully turbulent flow. The results are compared with experimental pressure data and with computational results obtained by Panaras and Steger, using the identical code and grid, but with their own modifications to the Baldwin-Lomax model. In addition, the computed solutions are analyzed using surface flow patterns, helicity density contours, and turbulent eddy-viscosity profiles. The results of this analysis provide insight into the effects of the turbulence models on flow characteristics and demonstrate the effect of the models on the accurate prediction of highly separated and vortical flows about a slender body.

Investigation of oblique shock/boundary-layer bleed interaction

Abstract: The detailed flowfield characteristics in an oblique shock-wave/laminar-boundary-layer interaction with bleed were investigated. The numerical solution for the flowfield was obtained for the strong conservation-law form of the two-dimensional compressible Navier-Stokes equations using an implicit scheme. The computations mod- eled the flow in the interaction region and inside the bleed slot for an impinging oblique shock on a flat-plate boundary layer. The computed results for the streamlines and the pressure and Mach number contours inside the bleed slot indicate that the flow is choked in the slot, with a recirculation zone near the upstream slot corner. The bleed results in the interaction zone demonstrate that flow separation is controlled. The interaction length . is reduced and the downstream velocity profiles are more favorable than the separated flow results at the same shock strength without bleed. HE control of shock/boundary-layer interactions in inlets and nozzles and over vehicle surfaces is accomplished through bleed and/or blowing in the interaction zone. In the case of mixed compression supersonic inlets, the bleed system design is critical to the efficient and stable operation of the system. Hamed and Shang1 reviewed the existing experimen- tal data for shock-wave/boundary-layer interactions in super- sonic inlets and other related configurations. According to this survey, most of the experimental measurements in mixed compression supersonic inlets consisted of total pressure re- covery surveys at the engine face and static pressure distri- butions over the inner surfaces. In the few cases involving velocity profile measurements,2-3 the latter were obtained up- stream and downstream of the interactions. Comparisons of internal flow computational results3'5 with the experimental measurements in supersonic inlets2-3 revealed reasonable agreement between the computed and measured surface pres- sures upstream of the ramp bleed. However, discrepancies in the predicted shock locations and velocity profiles were ob- served downstream of shock/boundary-layer interactions with bleed. There is enough experimental evidence6'11 to indicate that local bleed can control flow separation in shock-wave/bound- ary-layer interactions. There are disagreements,1 however, among the different experimental studies regarding the effects of bleed hole size,7-8 and the location of the bleed holes in relation to the shock.6-9'11 The experimental data in these studies are not sufficient, however, to resolve these discrep- ancies. Strike and Rippy9 measured the surface pressure in the interaction zone of an oblique shock wave impinging a tur- bulent boundary layer over a flat plate, with suction. They determined that less suction is required to control separation, when applied upstream of the shock. Seebaugh and Childs11 investigated experimentally the axisymmetric flow in the in- teraction region of the boundary layer inside a duct. Contrary to the conclusions of Strike and Rippy,9 suction within the

Natural laminar flow and laminar flow control

Abstract: The present volume discusses the development history and basic concepts of laminar flow control, laminar flow flight experiments, subsonic laminar-flow airfoils, and a design philosophy for long-range laminar flow-control commercial transports with advanced supercritical airfoils. Also discussed are the relationship of wave-interaction theory to laminar flow control, supersonic laminar flow control, and the NASA-Langley 8-ft Transonic Pressure Tunnel.

Parallel spectral-element—Fourier simulation of turbulent flow over riblet-mounted surfaces

Abstract: The flow in a channel with its lower wall mounted with streamwise V-shaped riblets is simulated using a highly efficient spectral-element—Fourier method. The range of Reynolds numbers investigated is 500 to 4000, which corresponds to laminar, transitional, and turbulent flow states. Our results suggest that in the laminar regime there is no drag reduction, while in the transitional and turbulent regimes drag reduction up to 10% exists for the riblet-mounted wall in comparison with the smooth wall of the channel. For the first time, we present detailed turbulent statistics in a complex geometry. These results are in good agreement with available experimental data and provide a quantitative picture of the drag-reduction mechanism of the riblets.