Showing papers on "Flow separation published in 2012"

Turbulent pipe flow at extreme Reynolds numbers.

Abstract: Both the inherent intractability and complex beauty of turbulence reside in its large range of physical and temporal scales. This range of scales is captured by the Reynolds number, which in nature and in many engineering applications can be as large as 10(5)-10(6). Here, we report turbulence measurements over an unprecedented range of Reynolds numbers using a unique combination of a high-pressure air facility and a new nanoscale anemometry probe. The results reveal previously unknown universal scaling behavior for the turbulent velocity fluctuations, which is remarkably similar to the well-known scaling behavior of the mean velocity distribution.

Low-frequency unsteadiness in shock wave–turbulent boundary layer interaction

Abstract: The low-frequency unsteadiness is characterized in the direct numerical simulation of a shock wave–turbulent boundary layer interaction generated by a compression ramp in Mach 2.9 flow. Consistent with experimental observations, the shock wave in the simulation undergoes a broadband streamwise oscillation at frequencies approximately two orders of magnitude lower than the characteristic frequency of the energetic turbulent scales in the incoming boundary layer. The statistical relation between the low-frequency shock motion and the upstream and downstream flow is investigated. The shock motion is found to be related to a breathing of the separation bubble and an associated flapping of the separated shear layer. A much weaker statistical relation is found with the incoming boundary layer. In order to further characterize the low-frequency mode in the downstream separated flow, the temporal evolution of the low-pass filtered flow field is investigated. The nature of the velocity and vorticity profiles in the initial part of the interaction is found to change considerably depending on the phase of the low-frequency motion. It is conjectured that these changes are due to an inherent instability in the downstream separated flow, and that this instability is the physical origin of the low-frequency unsteadiness. The low-frequency mode observed here is, in certain aspects, reminiscent of an unstable global mode obtained by linear stability analysis of the mean flow in a reflected shock interaction (Touber & Sandham, Theor. Comput. Fluid Dyn., vol. 23, 2009, pp. 79–107).

The onset of dynamic stall revisited

Abstract: Dynamic stall on a helicopter rotor blade comprises a series of complex aerodynamic phenomena in response to the unsteady change of the blade’s angle of attack. It is accompanied by a lift overshoot and delayed massive flow separation with respect to static stall. The classical hallmark of the dynamic stall phenomenon is the dynamic stall vortex. The flow over an oscillating OA209 airfoil under dynamic stall conditions was investigated by means of unsteady surface pressure measurements and time-resolved particle image velocimetry. The characteristic features of the unsteady flow field were identified and analysed utilising different coherent structure identification methods. An Eulerian and a Lagrangian procedure were adopted to locate the axes of vortices and the edges of Lagrangian coherent structures, respectively; a proper orthogonal decomposition of the velocity field revealed the energetically dominant coherent flow patterns and their temporal evolution. Based on the complementary information obtained by these methods the dynamics and interaction of vortical structures were analysed within a single dynamic stall life cycle leading to a classification of the unsteady flow development into five successive stages: the attached flow stage; the stall development stage; stall onset; the stalled stage; and flow reattachment. The onset of dynamic stall was specified here based on a characteristic mode of the proper orthogonal decomposition of the velocity field. Variations in the flow field topology that accompany the stall onset were verified by the Lagrangian coherent structure analysis. The instantaneous effective unsteadiness was defined as a single representative parameter to describe the influence of the motion parameters. Dynamic stall onset was found to be promoted by increasing unsteadiness. The mechanism that results in the detachment of the dynamic stall vortex from the airfoil was identified as vortex-induced separation caused by strong viscous interactions. Finally, a revised criterion to discern between light and deep dynamic stall was formulated.

Origins and structure of spike-type rotating stall

Abstract: In this paper we describe the structures that produce a spike-type route to rotating stall and explain the physical mechanism for their formation. The descriptions and explanations are based on numerical simulations, complemented and corroborated by experiments. It is found that spikes are caused by a loss of pressure rise capability in the rotor tip region, due to flow separation resulting from high incidence. The separation gives rise to shedding of vorticity from the leading edge and the consequent formation of vortices that span between the suction surface and the casing. As seen in the rotor frame of reference, near the casing the vortex convects toward the pressure surface of the adjacent blade. The approach of the vortex to the adjacent blade triggers a separation on that blade so the structure propagates. The above sequence of events constitutes a spike. The simulations show shed vortices over a range of tip clearances including zero. The implication is that they are not part of the tip clearance vortex, in accord with recent experimental findings. Evidence is presented for the existence of these vortex structures immediately prior to spike-type stall and, more strongly, for their causal connection with spike-type stall inception. Data from several compressors are shown to reproduce the pressure and velocity signature of the spike-type stall inception seen in the simulations.Copyright © 2012 by ASME

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

Abstract: 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.

Fully resolved numerical simulation of particle-laden turbulent flow in a horizontal channel at a low Reynolds number

Abstract: A fictitious domain method is used to perform fully resolved numerical simulations of particle-laden turbulent flow in a horizontal channel. The effects of large particles of diameter 0.05 and 0.1 times the channel height on the turbulence statistics and structures are investigated for different settling coefficients and volume fractions (0.79 %–7.08 %) for the channel Reynolds number being 5000. The results indicate the following. (a) When the particle sedimentation effect is negligible (i.e. neutrally buoyant), the presence of particles decreases the maximum r.m.s. of streamwise velocity fluctuation near the wall by weakening the intensity of the large-scale streamwise vortices, while increasing the r.m.s. of the streamwise fluctuating velocity in the region very close to the wall and in the centre region. On the other hand, the particles increase the r.m.s. of transverse and spanwise fluctuating velocities in the near-wall region by inducing the small-scale vortices. (b) When the particle settling effect is so substantial that most particles settle onto the bottom wall and form a particle sediment layer (SL), the SL plays the role of a rough wall and parts of the vortex structures shedding from the SL ascend into the core region and substantially increase the turbulence intensity there. (c) When the particle settling effect is moderate, the effects of particles on the turbulence are a combination of the former two situations, and the Shields number is a good parameter for measuring the particle settling effects (i.e. the particle concentration distribution in the transverse direction). The average velocities of the particle are smaller in the lower half-channel and larger in the upper half-channel compared to the local fluid velocities in the presence of gravity effects. The effects of the smaller particles on the turbulence are found to be stronger at the same particle volume fractions.

Direct numerical simulation of a 30R long turbulent pipe flow at R + = 685: Large-and very large-scale motions

Abstract: Fully developed incompressible turbulent pipe flow at Reynolds number (based on bulk velocity) and Karman number is simulated in a periodic domain with a length of pipe radii . While single-point statistics match closely with experimental measurements, questions have been raised of whether streamwise energy spectra calculated from spatial data agree with the well-known bimodal spectrum shape in premultiplied spectra produced by experiments using Taylor’s hypothesis. The simulation supports the importance of large- and very large-scale motions (VLSMs, with streamwise wavelengths exceeding ). Wavenumber spectral analysis shows evidence of a weak peak or flat region associated with VLSMs, independent of Taylor’s hypothesis, and comparisons with experimental spectra are consistent with recent findings (del Alamo & Jimenez, J. Fluid Mech., vol. 640, 2009, pp. 5–26) that the long-wavelength streamwise velocity energy peak is overestimated when Taylor’s hypothesis is used. Yet, the spectrum behaviour retains otherwise similar properties to those documented based on experiment. The spectra also reveal the importance of motions of long streamwise length to the energy and Reynolds stress and support the general conclusions regarding these quantities formed using experimental measurements. Space–time correlations demonstrate that low-level correlations involving very large scales persist over in time and indicate that these motions convect at approximately the bulk velocity, including within the region approaching the wall. These very large streamwise motions are also observed to accelerate the flow near the wall based on force spectra, whereas smaller scales tend to decelerate the mean streamwise flow profile, in accordance with the behaviour observed in net force spectra of prior experiments. Net force spectra are resolved for the first time in the buffer layer and reveal an unexpectedly complex structure.

Separated shear layer transition over an airfoil at a low Reynolds number

Abstract: Shear layer development over a NACA 0018 airfoil at a chord Reynolds number of 100 000 was investigated using a combination of flow visualization, velocity field mapping, surface pressure fluctuation measurements, and stability analysis. The results provide a detailed description of shear layer transition on an airfoil at low Reynolds numbers. An extensive comparison of measured surface pressure and velocity fluctuations demonstrated that time-resolved surface pressure sensor arrays can be used to identify the presence of flow separation, estimate the extent of the separated flow region, and measure disturbance growth rate spectra in significantly less time than is required by conventional techniques. Surface pressure sensor measurements of disturbance growth rate, wave number, and convection speed are found to compare well with predictions of linear stability theory, supporting the claim that convection speeds measured in separation bubbles over low Reynolds number airfoils are associated with wave packe...

Pressure fluctuation in high-Reynolds-number turbulent boundary layer: results from experiments and DNS

Abstract: We have developed a small pressure probe and measured both static pressure and wall pressure simultaneously in turbulent boundary layers up to Reynolds numbers based on the momentum thickness . The measurements were performed at large experimental facilities in Sweden, Australia, and Japan. We find that the measured pressure data are contaminated by the artificial background noise induced by test section and are also affected by the flow boundary conditions. By analyzing data from different wind tunnels acquired at the same Reynolds number, we evaluate the effect of background noises and boundary conditions on the pressure statistics. We also compare the experimental results with results of direct numerical simulations and discuss differences in boundary conditions between real and simulated wind tunnels.

Formation of Turbulent Vortex Breakdown: Intermittency, Criticality, and Global Instability

Abstract: This study provides quantitative insight into the formation of vortex breakdown and the onset of global instability in a turbulent swirling jet. A water jet is guided through a rotating honeycomb that imparts the rotational motion, passed through a contraction, and discharged into a large water tank. The flow states evolving at increasing swirl are mapped out via time-resolved particle image velocimetry. The experimental results scale properly with the swirl number based on the axial momentum flux when the commonly used boundary-layer approximations are omitted. The instantaneous velocity field reveals that vortex breakdown occurs intermittently at a wide range of swirl numbers before it appears in the mean flow. At this intermittent state, the evolving breakdown bubble oscillates heavily between two streamwise locations where the vortex core is subcritical. Upon further increasing the swirl, the breakdown oscillations decay and a region of reversed flow appears in the mean flowfield. The formation of this socalled axisymmetric breakdown state is accompanied by a supercritical-to-subcritical transition of the inflowing vortex core. The reversed flow region is found to grow linearly with increasing swirl until the flow undergoes a supercritical Hopf bifurcation to a global single-helical mode, and vortex breakdown adopts a spiral shape. The global mode shape is extracted from the particle image velocimetry snapshots by means of proper orthogonal decomposition and Fourier analysis. The present experiment reveals that, at gradually increasing swirl, the jet first transitions to an axisymmetric breakdown state that remains globally stable until a critical swirl number is exceeded. This sequence of flow states agrees well with the transient formation of vortex breakdown observed in laminar flows.

Mechanism of flow separation control using plasma actuation

Abstract: The mechanism of flow separation control was investigated experimentally and computationally using pulse-modulated dielectric barrier discharge (DBD) plasma actuation on a stalled flat plate airfoil at a Reynolds number of 3000. Load measurements were complimented with two-dimensional phase-averaged particle image velocimetry performed in the flowfield above the airfoil. A parametric study was carried out where the pulse-modulation frequency, duty cycle, and peak plasma body-force were varied. The two-dimensional Navier-Stokes equations, with no turbulence modeling, were solved directly using a commercial flow solver and a simple but satisfactory heuristic DBD plasma body-force model was incorporated. The overall experimental trends were well predicted by the computations, where the frequencies that produced the largest increases in lift coefficient excited bluff-body shedding at a frequency corresponding approximately to its unforced sub-harmonic. At non-dimensional frequencies most effective for increas...

The effect of an external transmitter on the drag coefficient of a bird’s body, and hence on migration range, and energy reserves after migration

Abstract: Externally mounted transmitters or loggers may adversely affect migration performance for reasons other than the effects of added mass. The added frontal area of a payload box increases drag, and if the box triggers separation of the boundary layer over the posterior body, the drag coefficient could also be increased, possibly by a large amount. Any such effects would lead directly to a decreased migration range and reduced energy reserves on completion of migration. We measured the body drag coefficients of Rose-coloured Starlings in the Seewiesen wind tunnel by the wingbeat-frequency method. The speed at which the wingbeat frequency passed through a minimum was taken to be an estimate of the minimum-power speed (V mp), from which the body drag coefficient was calculated in turn. Dummy transmitter boxes were mounted on the bird’s back by attaching them with Velcro to a side-loop harness pad. The pad alone projected 6 mm above the bird’s back, and increased the drag coefficient by nearly 50%, as compared to the “clean” configuration with no harness. Adding boxes (square-ended or streamlined) produced no further significant increase in the drag coefficient, but the addition of a sloping antenna increased it to nearly twice the clean value. These increases are attributed to separation of the boundary layer over the posterior upper body, triggered by the payload. We then ran computer simulations of a particular Barnacle Goose, for which detailed information was available from an earlier satellite-tracking project, to see how its migration range and reserves on arrival would be affected if its transmitter installation also caused flow separation and affected the body drag coefficient in a similar way. By representing the range calculation in terms of energy height, we separated the effect of the transmitter’s mass, which reduces the fat fraction (and hence also energy height) at departure, from that of flow separation, which steepens the energy gradient. The effect of the mass is small, and increases only slightly with increasing distance, whereas a steeper energy gradient not only reduces the range but also reduces the reserves remaining on arrival, to an extent that increases with migration distance. Energy height is related to the fat fraction rather than the fat mass, and is therefore preferable to energy as such, for expressing reserves in birds of different sizes.

Dielectric-barrier-discharge vortex generators: characterisation and optimisation for flow separation control

Abstract: We investigated the use of dielectric-barrier-discharge plasma actuators as vortex generators for flow separation control applications. Plasma actuators were placed at a yaw angle to the oncoming flow, so that they produced a spanwise wall jet. Through interaction with the oncoming boundary layer, this created a streamwise longitudinal vortex. In this experimental investigation, the effect of yaw angle, actuator length and plasma-induced velocity ratio was studied. Particular attention was given to the vortex formation mechanism and its development downstream. The DBD plasma actuators were then applied in the form of co-rotating and counter-rotating vortex arrays to control flow separation over a trailing-edge ramp. It was found that the vortex generators were successful in reducing the separation region, even at plasma-to-free-stream velocity ratios of less than 10%.

Experimental investigation of the micro-ramp based shock wave and turbulent boundary layer interaction control

Abstract: Using a nanoparticle-based planar laser-scattering technique and supersonic particle image velocimetry, we investigated the effects of micro-ramp control on incident shockwave and boundary-layer interaction (SWBLI) in a low-noise supersonic wind-tunnel with Mach number 27 and Reynolds number Rθ = 5845 High spatiotemporal resolution wake structures downstream of the micro-ramps were detected, while a complex evolution process containing a streamwise counter-rotating vortex pair and large-scale hairpin-like vortices with Strouhal number Stδ of about 05–065 was revealed The large-scale structures could survive while passing through the SWBLI region Reflected shockwaves are clearly seen to be distorted accompanied by high-frequency fluctuations Micro-ramp applications have a distinct influence on flow patterns of the SWBLI field that vary depending on spanwise locations Both the shock foot and separation line exhibit undulations corresponding with modifications of the velocity distribution of the inco

Revival of classical vortex generators now for transition delay.

Abstract: Classical vortex generators, known for their efficiency in delaying or even inhibiting boundary layer separation, are here shown to be coveted devices for transition to turbulence delay. The present devices are miniature with respect to classical vortex generators but are tremendously powerful in modulating the laminar boundary layer in the direction orthogonal to the base flow and parallel to the surface. The modulation generates an additional term in the perturbation energy equation, which counteracts the wall-normal production term and, hence, stabilizes the flow. Our experimental results show that these devices are really effective in delaying transition, but we also reveal their Achilles' heel.

Numerical Simulation of Fluidic Actuators for Flow Control Applications

Abstract: Active flow control technology is finding increasing use in aerospace applications to control flow separation and improve aerodynamic performance. In this paper we examine the characteristics of a class of fluidic actuators that are being considered for active flow control applications for a variety of practical problems. Based on recent experimental work, such actuators have been found to be more efficient for controlling flow separation in terms of mass flow requirements compared to constant blowing and suction or even synthetic jet actuators. The fluidic actuators produce spanwise oscillating jets, and therefore are also known as sweeping jets. The frequency and spanwise sweeping extent depend on the geometric parameters and mass flow rate entering the actuators through the inlet section. The flow physics associated with these actuators is quite complex and not fully understood at this time. The unsteady flow generated by such actuators is simulated using the lattice Boltzmann based solver PowerFLOW R . Computed mean and standard deviation of velocity profiles generated by a family of fluidic actuators in quiescent air are compared with experimental data. Simulated results replicate the experimentally observed trends with parametric variation of geometry and inflow conditions.

An explanation for flow features of spike-type stall inception in an axial compressor rotor

Abstract: The unsteady behavior and three-dimensional flow structure of spike-type stall inception in an axial compressor rotor were investigated by experimental and numerical analyses. Previous studies revealed that the test compressor falls into a mild stall after emergence of a spike, in which multiple stall cells, each consisting of a tornado-like vortex, are rotating. However, the flow mechanism from the spike onset to the mild stall remains unexplained. The purpose of this study is to describe the flow mechanism of a spike stall inception in a compressor. In order to capture the transient phenomena of spike-type stall inception experimentally, an instantaneous casing pressure field measurement technique was developed, in which 30 pressure transducers measure an instantaneous casing pressure distribution inside the passage for one blade pitch at a rate of 25 samplings per blade passing period. This technique was applied to obtain the unsteady and transient pressure fields on the casing wall during the inception process of the spike stall. In addition, the details of the three-dimensional flow structure at the spike stall inception were analyzed by a numerical approach using the detached-eddy simulation (DES). The instantaneous casing pressure field measurement results at the stall inception show that a low-pressure region starts traveling near the leading edge in the circumferential direction just after the spiky wave was detected in the casing wall pressure trace measured near the rotor leading edge. The DES results reveal the vortical flow structure behind the low-pressure region on the casing wall at the stall inception, showing that the low-pressure region is caused by a tornado-like separation vortex resulting from a leading-edge separation near the rotor tip. A leading-edge separation occurs near the tip at the onset of the spike stall and grows to form the tornado-like vortex connecting the blade suction surface and the casing wall. The casing-side leg of the tornado-like vortex generating the low-pressure region circumferentially moves around the leading-edge line. When the vortex grows large enough to interact with the leading edge of the next blade, the leading-edge separation begins to propagate, and then the compressor falls into a stall with decreasing performance.

Aerofoil tones at moderate Reynolds number

Abstract: It is known experimentally that an aerofoil immersed in a uniform stream at a moderate Reynolds number emits tones. However, there have been major differences in the experimental observations in the past. Some experiments reported the observation of multiple tones, with strong evidence that these tones are most probably generated by a feedback loop. There is also an experiment reporting the observation of a single tone with no tonal jump or other features associated with feedback. In spite of the obvious differences in the experimental observations published in the literature, it is noted that all the dominant tone frequencies measured in all the investigations are in agreement with an empirically derived Paterson formula. The objective of the present study is to perform a direct numerical simulation (DNS) of the flow and acoustic phenomenon to investigate the tone generation mechanism. When comparing with experimental studies, numerical simulations appear to have two important advantages. The first is that there is no background wind tunnel noise in numerical simulation. This avoids the signal-to-noise ratio problem inherent in wind tunnel experiments. In other words, it is possible to study tones emitted by a truly isolated aerofoil computationally. The second advantage is that DNS produces a full set of space–time data, which can be very useful in determining the tone generation processes. The present effort concentrates on the tones emitted by three NACA0012 aerofoils with a slightly rounded trailing edge but with different trailing edge thickness at zero degree angle of attack. At zero degree angle of attack, in the Reynolds number range of to , the boundary layer flow is attached nearly all the way to the trailing edge of the aerofoil. Unlike an aerofoil at an angle of attack, there is no separation bubble, no open flow separation. All the flow separation features tend to increase the complexity of the tone generation processes. The present goal is limited to finding the basic tone generation mechanism in the simplest flow configuration. Our DNS results show that, for the flow configuration under study, the aerofoil emits only a single tone. This is true for all three aerofoils over the entire Reynolds number range of the present study. In the literature, it is known that Kelvin–Helmholtz instabilities of free shear layers generally have a much higher spatial growth rate than that of the Tollmien–Schlichting boundary layer instabilities. A near-wake non-parallel flow instability analysis is performed. It is found that the tone frequencies are the same as the most amplified Kelvin–Helmholtz instability at the location where the wake has a minimum half-width. This suggests that near-wake instability is the energy source of aerofoil tones. However, flow instabilities at low subsonic Mach numbers generally do not cause strong tones. An investigation of how near-wake instability generates tones is carried out using the space–time data provided by numerical simulations. Our observations indicate that the dominant tone generation process is the interaction of the oscillatory motion of the near wake, driven by flow instability, with the trailing edge of the aerofoil. Secondary mechanisms involving unsteady near-wake motion and the formation of discrete vortices in regions further downstream are also observed.

Low-Reynolds-Number Effects in Passive Stall Control Using Sinusoidal Leading Edges

Abstract: The objective of the presentwork is to investigate the application of a sinusoidal leading edge to the design ofmicro air vehicles. Wind-tunnel tests of wings with low aspect ratios of 1 and 1.5, rectangular planforms, and five distinct leading edges [four sinusoidal leading edges and one baseline (straight) leading edge for each aspect ratio] have been conducted. The Reynolds numbers of 70,000 and 140,000 have been analyzed. For the higher Reynolds number, a proper combination of amplitude andwavelength can lead to a substantial increase in lift for angles of attack greater than the baseline stall angle. Maximum lift coefficient gains of the order of 45% were achieved by combining both large amplitude and large wavelength. At the lower Reynolds number, the benefits can be extended to low angles of attack, leading to a dramatic increase in the range of operation. The results depend strongly on the aspect ratio.

Vortex Generators for Wind Turbine Blades: A Combined Wind Tunnel and Wind Turbine Parametric Study

Abstract: Vortex generators (VGs) are passive flow control devices commonly employed to prevent flow separation on wind turbine blades. They mitigate the damaging fatigue loads resulting from stall while increasing lift and consequently lead to rotor torque increase. This work summarizes a research project aimed at optimizing the sectional as well as the full rotor-blade aerodynamics using VGs.The effects of chordwise position, spanwise spacing and VG size were studied with force balance measurements of a 2D wing section. Reducing the distance between adjacent VGs produced large increases in the static stall angle and maximum lift, but also resulted in a significant increase in drag as well as sharp lift excursions at angles exceeding the static stall angle. The optimal chordwise position of the vortex generators was found to be in the range of x/c = 15%–20%, where a comparatively low parasitic drag and a smooth post-stall lift curve were achieved. Particle Image Velocimetry measurements were conducted at various chordwise positions to provide insight into the interaction between adjacent streamwise vortices.The experimental aerodynamic performance curves of the optimal VG configuration were used to project their effect on wind turbine blade aerodynamics. Three different rotorblades were designed and several stall and pitch regulated wind turbine models were simulated by means of a Blade Element Momentum (BEM) code (QBlade) developed by Smart Blade GmbH. The performance of the rotorblades with and without VGs was simulated in order to assess their effect on the aerodynamic performance and loads. Finally, previously measured steady state performance curves under high-roughness conditions were used to simulate the detrimental effect of surface roughness on the performance of the aforementioned rotorblades. This allows for an estimate of the potential of the VGs to be employed as retrofit elements for performance recovery of blades with a contaminated surface.© 2012 ASME

Changes in turbulent dissipation in a channel flow with oscillating walls

Abstract: Harmonic oscillations of the walls of a turbulent plane channel flow are studied by direct numerical simulations to improve our understanding of the physical mechanism for skin-friction drag reduction. The simulations are carried out at constant pressure gradient in order to define an unambiguous inner scaling: in this case, drag reduction manifests itself as an increase of mass flow rate. Energy and enstrophy balances, carried out to emphasize the role of the oscillating spanwise shear layer, show that the viscous dissipations of the mean flow and of the turbulent fluctuations increase with the mass flow rate, and the relative importance of the latter decreases. We then focus on the turbulent enstrophy: through an analysis of the temporal evolution from the beginning of the wall motion, the dominant, oscillation-related term in the turbulent enstrophy is shown to cause the turbulent dissipation to be enhanced in absolute terms, before the slow drift towards the new quasi-equilibrium condition. This mechanism is found to be responsible for the increase in mass flow rate. We finally show that the time-average volume integral of the dominant term relates linearly to the drag reduction.