Experimental study of the boundary layer over an airfoil in plunging motion

This is an experimental study on the boundary layer over an airfoil under steady and unsteady conditions.It specifically deals with the effect of plunging oscillation on the laminar/turbulent characteristics of the boundary layer.The wind tunnel measurements involved surfacemounted hot-film sensors and boundary-layer rake.The experiments were conducted at Reynolds numbers of 0.42×10 6 to 0.84 × 10 6 and the reduced frequency was varied from 0.01 to 0.11.The results of the quasi-wall-shear stress as well as the boundary layer velocity profiles provided important information about the state of the boundary layer over the suction surface of the airfoil in both static and dynamic cases.For the static tests,boundary layer transition occurred through a laminar separation bubble.By increasing the angle of attack,disturbances and the transition location moved toward the leading edge.For the dynamic tests,earlier transition occurred with increasing rather than decreasing effective angle of attack.The mean angle of attack and the oscillating parameters significantly affected the state of the boundary layer.By increasing the reduced frequency,the boundary layer transition was promoted to the upstroke portion of the equivalent angle of attack,but the quasi skin friction coefficient was decreased.     

Measurement of unsteady boundary layer developed on an oscillating airfoil using multiple hot-film sensors

 The spatial-temporal progressions of the leading-edge stagnation, separation and reattachment points, and the state of the unsteady boundary layer developed on the upper surface of a 6 in. chord NACA 0012 airfoil model, oscillated sinusoidally within and beyond the static-stall angle, were measured using 140 closely-spaced, multiple hot-film sensors (MHFS). The MHFS measurements show that (i) the laminar separation point and transition were delayed with increasing α and the reattachment and relaminarization were promoted with decreasing α, relative to the static case, (ii) the pitchup motion helped to keep the boundary layer attached to higher angles of attack over that could be obtained statically, (iii) the dynamic stall process was initiated by the turbulent flow separation in the leading-edge region as well as by the onset of flow reversal in the trailing-edge region, and (iv) the dynamic stall process was found not to originate with the bursting of a laminar separation bubble, but with a breakdown of the turbulent boundary layer. The MHFS measurements also show that the flow unsteadiness caused by airfoil motion as well as by the flow disturbances can be detected simultaneously and nonintrusively. The MHFS characterizations of the unsteady boundary layers are useful in the study of unsteady separated flowfields generated by rapidly maneuvering aircraft, helicopter rotor blades, and wing energy machines. Received: 17 June 1997 / Accepted: 10 December 1997     

Parametric study of separation and transition characteristics over an airfoil at low Reynolds numbers

Time-resolved surface pressure measurements are used to experimentally investigate characteristics of separation and transition over a NACA 0018 airfoil for the relatively wide range of chord Reynolds numbers from 50,000 to 250,000 and angles of attack from 0° to 21°. The results provide a comprehensive data set of characteristic parameters for separated shear layer development and reveal important dependencies of these quantities on flow conditions. Mean surface pressure measurements are used to explore the variation in separation bubble position, edge velocity in the separated shear layer, and lift coefficients with angle of attack and Reynolds number. Consistent with previous studies, the separation bubble is found to move upstream and decrease in length as the Reynolds number and angle of attack increase. Above a certain angle of attack, the proximity of the separation bubble to the location of the suction peak results in a reduced lift slope compared to that observed at lower angles. Simultaneous measurements of the time-varying component of surface pressure at various spatial locations on the model are used to estimate the frequency of shear layer instability, maximum root-mean-square (RMS) surface pressure, spatial amplification rates of RMS surface pressure, and convection speeds of the pressure fluctuations in the separation bubble. A power-law correlation between the shear layer instability frequency and Reynolds number is shown to provide an order of magnitude estimate of the central frequency of disturbance amplification for various airfoil geometries at low Reynolds numbers. Maximum RMS surface pressures are found to agree with values measured in separation bubbles over geometries other than airfoils, when normalized by the dynamic pressure based on edge velocity. Spatial amplification rates in the separation bubble increase with both Reynolds number and angle of attack, causing the accompanying decrease in separation bubble length. Values of the convection speed of pressure fluctuations in the separated shear layer are measured to be between 35 and 50% of the edge velocity, consistent with predictions of linear stability theory for separated shear layers.     

Experimental Study on the Behaviour of Local Laminar Separation Bubble at a Rectangular Wing Section

This paper identifies laminar separation bubbles at the root or span-wise midsection of a rectangular wing using direct surface pressure measurements in the wind tunnel and analyses their behavior at different Reynolds numbers and angles of attack. The separation, transition, and reattachment locations are determined as functions of the angles of attack and the Reynolds number. The transition structure and turbulence characteristics in the separated shear layer are studied using laser Doppler velocimetry. Surface pressure data and simultaneously acquired velocity signals are correlated to show the pattern of growing disturbances in the shear layer. Surface oil flow visualizations clarified the wingtip and separation bubble’s interactions near the leading edge of the wing at the higher angles of attack. Turbulence statistics are also calculated from the streamwise velocity distributions, and an apparent deviation is observed for the skewness and flatness values from the normal distributions in the near-wall region. The separation bubble effect on aerodynamic coefficients of a 3D rectangular wing root section is studied and reported.

     

FLOW CHARACTERISTICS AROUND AN INCLINED ELLIPTIC CYLINDER IN A TURBULENT BOUNDARY LAYER

The flow characteristics around an inclined elliptic cylinder located near a flat plate were investigated experimentally. The axis ratio of the elliptic cylinder was AR=2. The pressure distributions along the surface of the cylinder and the flat plate were measured by varying the angle of attack of the elliptic cylinder. The velocity profiles behind the cylinder were measured using hot-wire anemometry. When the angle of attack varies, the peak pressure location on the windward cylinder surface moves towards the rear edge of the cylinder, while that on the leeward surface moves towards the front edge of the cylinder. The vortex-shedding frequency also gradually decreases, defining a critical angle of attack for each gap ratio. The location of the minimum pressure on the flat plate surface moves downstream for positive angles of attack, while it moves upstream for negative angles of attack. Negative angles of attack cause a greater disturbance in the boundary layer near the wall compared to positive angles of attack. This shows that the separated wall shear layer from the boundary layer and the lower shear layer of the cylinder wake are strongly merged compared to other cases.     

Multiple laminar-turbulent transition cycles around a swept leading edge

Certain interesting flow features involving multiple transition/relaminarization cycles on the leading edge of a swept wing at low speeds are reported here. The wing geometry tested had a circular nose and a leading edge sweep of 60°. Tests were made at a chord Reynolds number of 1.3 × 10⁶ with model incidence α varied in the range of 3°?18° in discrete steps. Measurements made included wing chord-wise surface pressure distributions and wall shear stress fluctuations (using hot-film gages) within about 10 % of the chord in the leading edge zone. Results at α = 16° and 18° showed that several (often incomplete) transition cycles between laminar-like and turbulent-like flows occurred. These rather surprising results are attributable chiefly to the fact that the Launder acceleration parameter K (appropriately modified for swept wings) can exceed a critical range more than once along the contour of the airfoil in the leading edge region. Each such crossing results in a relaminarization followed by direct retransition to turbulence as K drops to sufficiently low values. It is further shown that the extent of each observed transition zone (of either type) is consistent with earlier data acquired in more detailed studies of direct transition and relaminarization. Swept leading edge boundary layers therefore pose strong challenges to numerical modelling.     

Large Eddy Simulation of a Controlled Diffusion Compressor Cascade

In this research a Controlled Diffusion (CD) compressor cascade stator blade is simulated at a Reynolds number of ??700,000, based on inflow velocity and chord length, using Large Eddy Simulation (LES). A wide range of flow inlet angles are computed, including conditions near the design angle, and at high negative and positive incidence. At all inlet angles the surface pressure distributions are well-predicted by the LES. Near the design angle the computed suction side boundary layer thickness agrees well with experimental data, whilst the pressure side boundary layer is poorly predicted due to the inability of LES to capture natural boundary layer transition on the present grid. A good estimation of the loss is computed near the design angle, whilst at both high positive and negative incidences the loss is less well predicted owing to discrepancies between the computed and experimental boundary layer thickness. At incidences above the design angle a laminar separation bubble forms near the leading edge of the suction surface, which undergoes a transition to turbulence. Similar behaviour is noted on the pressure surface at negative incidence. At high negative incidence contra-rotating vortex pairs are found to form around the leading edge in response to an unsteady stagnation line across the span of the blade. Such structures are not apparent in time-averaged statistical data due to their highly-transient nature.     

Transient Characterization of the Reattachment of a Massively Separated Turbulent Boundary Layer Under Flow Control

The transient dynamics of a high Reynolds number separated flow over a two-dimensional ramp submitted to pulsed fluidic control is investigated. A spanwise array of 22 round jets, located upstream of the flap leading edge, is used as actuator to generate co-rotating vortical structures. Simultaneous measurements of wall friction using hot-film anemometry and phase-averaged velocity using 2D2C PIV are conducted. The PIV plane encompasses the incoming boundary layer upstream the flap leading edge, the separation bubble and the natural reattachment region. The dynamics of the separated flow is studied for successive sequences of pulsed actuation. Pockets of turbulence are periodically generated by the separation process and pushed downstream. After the transition period, the controlled flow shows large amplitude oscillations around a steady mean, particularly for the separation area. The transient dynamics of the flow at the actuation activation is also studied. The separated flow is strongly modified by the actuation from the first pulse. Characteristic times of the transient dynamics can be determined by fitting a first-order model with delay on the data. For the reattachment, the dimensionless characteristic rising times defined as \(\tau _{r}^{+} = \tau _{r} ~ U_{0} ~/~ L_{sep}\) of 11.7 for the friction gain, 4.8 for the separation length and 4.1 for the first mode of a Conditional Proper Orthogonal Decomposition analysis of the phase-averaged velocity fields were found. These values are in good agreement with previous studies and are of particular interest for modeling the transients and for further closed-loop control applications.     

Measurements with a flow direction boundary-layer probe in a two-dimensional laminar separation bubble

 Measurements with a directional sensitive hot-wire probe have been carried out in a two-dimensional laminar separation bubble caused by an adverse pressure gradient. The probe has three parallel, in plane wires and can be traversed in the boundary layer in all spatial directions. The central wire, operated as a conventional hot-wire in CTA mode, and two surrounding resistance wires measure the instantaneous magnitude and direction of the flow, respectively. The probe is calibrated and operated in a similar way as a single hot-wire probe for boundary layer measurements. The frequency response is high enough for measurements of naturally occurring instability waves in the bubble. The flow direction intermittency was measured inside the bubble and regions with reversed flow were mapped out. Prior to reattachment periodical oscillations of the flow direction are found associated with shedding of vortical structures from the bubble. Received: 13 March 1998/Accepted: 22 April 1998     

Stability of preseparation boundary layer on the leading edge of a thin airfoil

Numerous experiments on subsonic flow of gas past thin wing profiles (see the reviews [1, 2]) have shown that the flow near the leading edge of an airfoil is separationless only at angles of attack less than a certain critical value, which depends on the shape of the airfoil. If the angle of attack reaches the critical value, a closed region of recirculation flow of small extension is formed on the upper surface of the airfoil. Under ordinary flow conditions, the boundary layer on the leading edge of the airfoil remains laminar in the entire preseparation range of angles of attack. However, the appearance of the closed separation region is, as a rule, accompanied by transition from a laminar to a turbulent flow regime. Moreover, generation of turbulence is observed precisely in the flow separation region. The present paper is devoted to a study of the stability of the boundary layer on the leading edge of a thin airfoil in a flow of incompressible fluid. The case when the angle of attack of the airfoil relative to the oncoming flow differs little from the critical value is considered.Translated from Izvestiya Akademii Nauk SSSR, Mekhanika Zhidkosti i Gaza, No. 6, pp. 55–63, November–December, 1982.     

Experimental study of compressible boundary layer on a cone at angles of attack

Experimental study was conducted for boundarylayers on a sharp 5° half-angle cone of 400mm length at angles of attack. The model was tested in the T-326 hypersonic wind tunnel （ITAM） at freestream Mach number M = 5.95. Mean and fluctuation wall characteristics of the boundary layer are measured at 0°, 2°, 3° and 4° angles of attack for different stagnation pressures. Pulsation measurements are carried out by means of ALTP sensor. Pressure and temperature distributions along the model are obtained, and transition beginning and end locations have been found. Boundary layer stabilization with the increase of angle of attack and the decrease of stagnation pressure is observed. High frequency pulsations inherent to hypersonic boundary layer （second mode） have been detected.     

Analysis of DNS and LES of Flow in a Low Pressure Turbine Cascade with Incoming Wakes and Comparison with Experiments

The flow around a low-pressure turbine rotor blade with incoming periodic wakes is computed by means of DNS and LES. The latter adopts a dynamic sub-grid-scale model. The computed results are compared with time-averaged and instantaneous measured quantities. The simulation sreveal the presence of elongated flow structures, stemming from the incoming wake vorticity, which interact with the pressure side boundary layer. As the wake approaches the upstream half of the suction side, its vortical structures are stretched and align with the main flow, resulting in an impingement at virtually zero angle of attack. Periodically, in the absence of impinging wakes, the laminar suction side boundary layer separates in the adverse pressure gradient region. Flow in the laminar separation bubble is found to undergo transition via a Kelvin–Helmholtz instability. Subsequent impingement of the wake inhibits separation and thus promotes boundary layer reattachment. LES provides a fair reproduction of the DNS results both in terms of instantaneous, phase-averaged, and time-averaged flow fields with a considerable reduction in computational effort. This revised version was published online in July 2006 with corrections to the Cover Date.     

Investigation of the vortex induced unsteadiness of a separation bubble via time-resolved and scanning PIV measurements

A transitional separation bubble on the suction side of an SD7003 airfoil is considered. The transition process that forces the separated shear layer to reattach seems to be governed by Kelvin–Helmholtz instabilities. Large scale vortices are formed due to this mechanism at the downstream end of the bubble. These vortices possess a three-dimensional structure and detach from the recirculation region, while other vortices are formed within the bubble. This separation of the vortex is a highly unsteady process, which leads to a bubble flapping. The structure of these vortices and the flapping of the separation bubble due to these vortices are temporally and spatially analyzed at angles of attack from 4° to 8° and chord-length based Reynolds numbers Re _c = 20,000–60,000 using time-resolved PIV measurements in a 2D and a 3D set-up, i.e., stereo-scanning PIV measurements are done in the latter case. These measurements complete former studies at a Reynolds number of Re _c = 20,000. The results of the time-resolved PIV measurements in a single light-sheet show the influence of the angle of attack and the Reynolds number. The characteristic parameters of the separation bubble are analyzed focusing on the unsteadiness of the separation bubble, e.g., the varying size of the main recirculation region, which characterizes the bubble flapping, and the corresponding Strouhal number are investigated. Furthermore, the impact of the freestream turbulence is investigated by juxtaposing the current and former results. The stereo-scanning PIV measurements at Reynolds numbers up to 60,000 elucidate the three-dimensional character of the vortical structures, which evolve at the downstream end of the separation bubble. It is shown that the same typical structures are formed, e.g., the c-shape vortex and the screwdriver vortex at each Reynolds number and angle of attack investigated and the occurrence of these patterns in relation to Λ-structures is discussed. To evidence the impact of the freestream turbulence, these results are compared with findings of former measurements.     

绕振荡水翼流动及其转捩特性的数值计算研究简

通过对比标准k-ω SST 湍流模型和基于标准k-ω SST 湍流模型修正的γ-Re_θ 转捩湍流模型对绕振荡NACA66 水翼流动的数值计算结果与实验结果,对水翼振荡过程的水动力特性和流场结构变化进行了分析研究. 结果表明：与标准k-ω SST 湍流模型的数值计算结果相比,基于标准k-ω SST 湍流模型修正的γ-Re_θ 转捩湍流模型能有效预测绕振荡翼型流场结构和水动力特性,捕捉流场边界层发生的流动分离和转捩现象;绕振荡水翼的流动过程可分为5 个特征阶段,当来流攻角较小时,在水翼前缘发生层流向湍流的转捩现象,水翼动力特征曲线出现变化拐点;随着来流攻角的增大,顺时针尾缘涡逐渐形成并向水翼前缘发展;当攻角较大时,前缘涡分离导致动力失速,水翼的动力特征曲线出现大幅波动;水翼处于顺时针向下旋转阶段,绕水翼的流动状态逐渐由湍流过渡为层流.     

绕振荡水翼流动及其转捩特性的数值计算研究

通过对比标准k-ω SST 湍流模型和基于标准k-ω SST 湍流模型修正的γ-Re_θ 转捩湍流模型对绕振荡NACA66 水翼流动的数值计算结果与实验结果,对水翼振荡过程的水动力特性和流场结构变化进行了分析研究. 结果表明:与标准k-ω SST 湍流模型的数值计算结果相比,基于标准k-ω SST 湍流模型修正的γ-Re_θ 转捩湍流模型能有效预测绕振荡翼型流场结构和水动力特性,捕捉流场边界层发生的流动分离和转捩现象;绕振荡水翼的流动过程可分为5 个特征阶段,当来流攻角较小时,在水翼前缘发生层流向湍流的转捩现象,水翼动力特征曲线出现变化拐点;随着来流攻角的增大,顺时针尾缘涡逐渐形成并向水翼前缘发展;当攻角较大时,前缘涡分离导致动力失速,水翼的动力特征曲线出现大幅波动;水翼处于顺时针向下旋转阶段,绕水翼的流动状态逐渐由湍流过渡为层流.      

Scanning PIV investigation of the laminar separation bubble on a SD7003 airfoil

A laminar separation bubble occurs on the suction side of the SD7003 airfoil at an angle of attack α =  4–8° and a low Reynolds number less than 100,000, which brings about a significant adverse aerodynamic effect. The spatial and temporal structure of the laminar separation bubble was studied using the scanning PIV method at α =  4° and Re = 60,000 and 20,000. Of particular interest are the dynamic vortex behavior in transition process and the subsequent vortex evolution in the turbulent boundary layer. The flow was continuously sampled in a stack of parallel illuminated planes from two orthogonal views with a frequency of hundreds Hz, and PIV cross-correlation was performed to obtain the 2D velocity field in each plane. Results of both the single-sliced and the volumetric presentations of the laminar separation bubble reveal vortex shedding in transition near the reattachment region at Re = 60,000. In a relatively long distance vortices characterized by paired wall-normal vorticity packets retain their identities in the reattached turbulent boundary layer, though vortices interact through tearing, stretching and tilting. Compared with the restricted LSB at Re = 60,000, the flow at Re = 20,000 presents an earlier separation and a significantly increased reversed flow region followed by “huge” vortical structures.     

Effect of Local DBD Plasma Actuation on Transition in a Laminar Separation Bubble

This work examines the effect of local active flow control on stability and transition in a laminar separation bubble. Experiments are performed in a wind tunnel facility on a NACA 0012 airfoil at a chord Reynolds number of 130 000 and an angle of attack of 2 degrees. Controlled disturbances are introduced upstream of a laminar separation bubble forming on the suction side of the airfoil using a surface-mounted Dielectric Barrier Discharge plasma actuator. Time-resolved two-component Particle Image Velocimetry is used to characterise the flow field. The effect of frequency and amplitude of plasma excitation on flow development is examined. The introduction of artificial harmonic disturbances leads to significant changes in separation bubble topology and the characteristics of coherent structures formed in the aft portion of the bubble. The development of the bubble demonstrates strong dependence on the actuation frequency and amplitude, revealing the dominant role of incoming disturbances in the transition scenario. Statistical, topological and linear stability theory analysis demonstrate that significant mean flow deformation produced by controlled disturbances leads to notable changes in stability characteristics compared to those in the unforced baseline case. The findings provide a new outlook on the role of controlled disturbances in separated shear layer transition and instruct the development of effective flow control strategies.     

The capability of large eddy simulation to predict relaminarization

The boundary layer which represents the narrow zone between a solid body and the free stream can have a laminar or a turbulent state. This state influences on the one hand the properties of the near-wall flow like skin friction or heat transfer and on the other hand also the free-stream flow itself, e.g. the downstream flow angle of a turbomachinery blade. Thus it is important for designers of fluid machinery to understand and predict the state of the boundary layer as well as the transition processes between the two states.In this work the so-called relaminarization is investigated which represents a reverse transition from a turbulent to a laminar boundary layer. At the Institute for Thermal Turbomachinery and Machine Dynamics at Graz University of Technology a test bench has been designed in order to produce a highly accelerated flow, thus triggering relaminarization. In the present work, the flow in this test bench is numerically investigated with Reynolds-averaged Navier-Stokes (RANS) flow simulation as well as with a large eddy simulation (LES).An outcome of this paper is, that the LES shows a very good agreement to the measurement results and is capable of predicting relaminarization.     

Low Reynolds number aerodynamics of free-to-roll low aspect ratio wings

The aerodynamics of thin, flat-plate wings of various planforms (rectangular, elliptical and Zimmerman) have been studied in free-to-roll experiments in a wind tunnel. Non-zero trim angles at low angles of attack, self-induced roll oscillations with increasing angle of attack and even autorotation in some cases were observed. The rectangular wings with round leading-edge had non-zero trim angles at low incidences due to the asymmetric development of the three-dimensional separation bubble at these low Reynolds numbers. With increasing angle of attack, the bubble increases in length and once reattachment is lost, large amplitude roll oscillations develop. The Strouhal number of the roll oscillations is of the order of 10⁻², which is in the same range as those expected for small aircraft experiencing atmospheric gusts. Velocity measurements revealed that variations in the strength of the vortices drove the rolling motion. At the mean roll angle, because of the time lag in the strength of the vortices, an asymmetric flow is generated, which results in a net rolling moment in the direction of the rolling motion.     

基于涡量矩理论的绕振荡水翼涡动力学分析

文章采用标准k-ω SST湍流模型和动网格技术, 实现了绕俯仰振荡NACA66水翼非定常流动结构与水动力特性的数值模拟, 并基于有限域涡量矩理论定量表征了局部旋涡结构对水翼动力特性的影响. 研究结果表明: 在水翼升程阶段, 当攻角较小时, 层流向湍流的转捩点由水翼尾缘向前缘移动; 在较大攻角时, 顺时针尾缘涡?TEV在水翼吸力面上生成并向前缘发展, 同时与吸力面上的顺时针前缘涡?LEV融合发展为附着在整个吸力面上的新前缘涡?LEV, 新的?LEV与逆时针尾缘涡+TEV相互作用直至完全脱落, 直接导致了水翼的动力失速, 在回程阶段, 绕振荡水翼的流场结构逐渐由湍流转变为层流. 基于有限域涡量矩理论的定量分析发现, 有限域内附着的?LEV和?TEV提供正升力, 当?LEV发展覆盖整个吸力面时对升力的贡献最大, 占总升力近50%, 而+TEV提供负升力. 同时发现, 有限域内各旋涡内部的不同区域提供的升力有正有负; 而逸出有限域的旋涡内部不同区域提供的升力方向均保持一致, 其中顺时针涡提供正升力, 而逆时针涡提供负升力. 在失速阶段, 域外旋涡整体对升力贡献较小且存在小幅波动, 体现了流动的非定常性.