Flow Separation Bubble Research Articles

Investigation of Chaotic Flutter Induced by Shock-Boundary-Layer Interactions

In this report, we investigate the onset of chaotic flutter in laminar and turbulent flows over a two-dimensional semi-infinite panel using time-accurate fluid–structure interaction (FSI) simulations of shock–boundary-layer interactions (SBLIs). Results indicate that the critical dynamic pressure and the nature of the panel dynamics strongly depend on the static pressure differential across the shock, the local loading, the viscous and turbulent damping in the flow, and the formation of dynamic flow separation bubbles due to the SBLIs. The structural system undergoes local instability (Hopf bifurcation) when the local loading due to panel deformation overcomes the static pressure differential across the shock. In the absence of external fluid instabilities, the structural oscillations will induce unsteadiness in the flowfield, resulting in low-frequency limit-cycle oscillations. Viscous and turbulent damping in the boundary layer also delays the bifurcation and reduces the amplitude of the oscillations. Sufficiently strong shocks can produce localized flow separations, which drive additional boundary-layer instabilities, resulting in an early bifurcation. In the presence of external fluid instabilities, the behavior of the FSI strongly depends on the nonlinear coupling of their instabilities. Chaotic oscillations are observed when the fluid instabilities are dominant enough to induce structural oscillations with broadband frequencies.

Spontaneous Laminar Separation Bubble Bursting on an Airfoil

An experimental investigation was conducted on a NACA 0018 airfoil at near-stall conditions, with stalled, stable laminar-separation bubble, and intermediate flow conditions were considered. Planar particle image velocimetry and surface pressure measurements were used to quantify the flow development at the midspan of the model. For the intermediate conditions, the airfoil is stalled in the time-average sense. However, conditionally averaged velocity field measurements reveal that reattachment of the separated flow and laminar separation bubble formation occur intermittently on the suction surface. The characteristics of the intermittently forming separation bubble and fully separated flow are analyzed based on conditional statistics, revealing variations in the transition process that may underpin the observed low-frequency flow oscillations. The growth rates of the most amplified wavenumbers in the separated shear layer vary with the angle between the separation streamline and the suction surface, leading to earlier transition when the separation angle increases. This change in transition location may play a role in spontaneous changes between fully separated and reattaching flow states.

Flow structure analysis of nanofluid impingement on modified target surface under different design parameters

The flow structures of jet impingement dominate heat and mass transfer process, even the whole thermal performance. In this study, we have inspected the flow structures and mechanism of nanofluid jet impingement onto a dimpled target surface with different design parameters. Investigations are performed for the relative depth of dimple ([Formula: see text]), the jet-to-plate spacing ([Formula: see text]), nanoparticle volume concentration ([Formula: see text]), and Reynolds number (Re) ranging to explore the mechanism of flow structure variations. Results indicate that these parameters have a significant effect on the flow structure of nanofluid jet impingement near the dimpled target surface. The flow begins to separate after passing the edge of the dimple along with the curvature of a dimple. [Formula: see text] will affect the form and location of flow separation and reattachment, and [Formula: see text] will affect the intensity of separation flow. The length of the flow separation bubble varies in different [Formula: see text] cases. When [Formula: see text] increases, the impinging energy and the velocity near the dimple edge decreases. The different Re has little effect on the length of the flow separation bubble and the tendency of the pressure coefficient (Cp). These results can provide further mechanism inspiration for the design of the flow structure of nanofluid jet impingement.

Combustion characteristics of line fire under cross wind: Effect of flow separation in the boundary layer

The line fire under cross wind is a fundamental combustion phenomenon which could occur in wildland and urban fires. This paper systematically investigates the effect on the linear flame behavior of the flow separation in the boundary layer, by experimental and numerical simulations. The effects of the distance between the burner rim and table's edge, the wind speed and the heat release rate (HRR) are considered on the linear flame. A reverse flame is experimentally and numerically observed along the upstream direction, due to the interaction between the flow separation and the burning flame. The flow velocity and pressure fields, given by the numerical simulation, clarify that the positive pressure difference derives the reverse flame from the burner surface to the upstream in the bottom of flow separation bubble. In addition, the reverse flame would heat the air in the whole boundary layer and accelerate the flow speed to enlarge the size of the separation bubble. Two different regimes that relate to the laminar and turbulent separated flows, respectively, are clarified for the reverse flame at a critical state. Dimensionless analysis is conducted to develop the critical criterion that distinguishes the appearance and disappearance of the reverse flame. In a fully turbulent separated flow, the pulsation frequency of the reverse flame at a critical state seems to equal the frequency of vortex shedding from the separation point. It is also demonstrated that the flow separation would affect the flame geometry by enhancing the air-fuel mixing in the bottom of flame.

Peak suction on a wall-mounted square cylinder by conditional POD: Effects of incident angles

The wind incident angle (α) is one of the crucial factors that dominates the peak suction on wall-mounted square cylinders (e.g., high-rise buildings). With a focus on the incident angle effects, this study applies the conditional space–time proper orthogonal decomposition (conPOD) to investigate the temporal evolution of peak suction in a statistical sense. Local regions of peak suction are typically classified as the lower windward corner, upper windward corner, trailing-edge region, and lower central side wall, which occur at different ranges of incident angle. The results show that the occurrence of peak pressures at the upper and lower windward corners on the wall of flow separation bubble is not synchronized on average when the incident angle is smaller than the critical angle (α≈15°). However, at a greater angle, a strong correlation is observed between the two regions of peak pressures at α=15–35°. The peak suction on the lower central side wall observed when α=5–15° is verified to have the same flow mechanism (i.e., inverted conical vortex) as the lower windward corner. The location, time of occurrence, and the flow mechanism of the trailing-edge peak suction are systematically clarified at different incident angles by the conPOD method.

Fluid–structure interaction simulation of a flapping flag in a laminar jet

The flapping flag is a classical and important Fluid–Structure Interaction (FSI) problem with many applications in nature, industry, and biological systems. However, the interaction between the flow field and the flag response is not fully understood due to the complex and nonlinear nature of the problem. Although theoretical models qualitatively predict the flapping frequency and the onset velocity, there are still quantitative discrepancies between the theoretical predictions and the experimental data. In this study, numerical FSI simulations of a standard flag in a viscous uniform laminar flow are performed to characterize the flow field around the flapping flag and correlate the flow field with the flag response for six Reynolds numbers (Re= 4.5 ×104, Re= 5.3 ×104, Re= 5.7 ×104, Re= 5.9 ×104, Re= 6.6 ×104, and Re= 7.5 ×104). Two-way coupled FSI simulations are conducted using commercial software ANSYS®. Rigorous solution verification is conducted first on three Re to ensure that solutions are independent of mesh and time-step refinement. Then validation is achieved by comparing the simulation predictions with experimental data. Overall, results showed a good agreement with an error range between 0.7%–4.7% for oscillation amplitude (Am), 1.3%–6.41% for drag coefficient (Cd), and 1.3%–5.3% for frequency (f). Moreover, a comparison between inviscid and viscous models is performed to evaluate the accuracy of using inviscid theoretical models for the same flow. It is observed that the inviscid model accurately predicted f but underpredicted Am and Cd, and overpredicted lift coefficient Cl. In the case of the viscous model, it is observed that a flow separation bubble and a thin flow circulation region exist on the suction side of the flag during the limit cycle oscillation (LCO). Finally, an instantaneous negative drag is observed for a short time during the LCO, suggesting that flag reconfiguration and flow separation significantly affects the drag. These observations suggest that viscous models may be needed to improve the theoretical predictions.

Stability of hypersonic flow over a curved compression ramp

In this work, the stability of hypersonic flow over a curved compression ramp is studied using several stability analysis tools and direct numerical simulations (DNS). The free-stream Mach number and the unit Reynolds number are 7.7 and $4.2 \times 10^6$ m $^{-1}$ , respectively. Corner rounding is considered to alter the separation bubble flow so as to suppress the intrinsic instability of the compression-ramp flow. The variation of intrinsic instability is confirmed by global stability analysis. Subsequently, resolvent analysis is employed to examine the response of intrinsically stable flows to external disturbances. It is shown that the considered flows strongly amplify low-frequency streamwise streaks with a preferential spanwise wavelength. This result is verified using DNS by introducing a random forcing upstream of the separation point. Furthermore, both resolvent analysis and DNS demonstrate that the separation bubble contributes little to the selection of the spanwise wavelength of streamwise streaks. The combined effects of convective and intrinsic instabilities are also explored using DNS. A better agreement with experimental data is achieved after introducing upstream disturbances in an inherently unstable flow.

Wavelet analysis of the coherent structures in airfoil leading-edge separation control by bionic coverts

We experimentally investigate leading-edge separation control effect by bionic coverts with various materials and sawteeth shapes in wind tunnel tests. The artificial flexible coverts, bio-inspired by bird covert feathers on upper wings, are hinged at the trailing-edge of a NACA 0018 airfoil at a constant high angle-of-attack of 15°. The chord-based Reynolds number is 1.0 × 105 in the generic range of bird flight in nature and low-speed fixed-wing unmanned aerial vehicles. The velocity profiles in the wake flow are measured by multi-channel hot-wire anemometer. By comparing the mean velocity profiles and root-mean-square velocities, we find the trailing-edge coverts reduced the thickness of the shear layers by 0.05 chord length. The turbulence intensity of the trailing- and leading-edge shear layers are reduced 34% and 5%, respectively. Further wavelet analysis reveals that the large sizes of vortices are considerably suppressed in the time-frequency spectrum. Based on the hot-wire datasets, we develop a novel multi-dimensional genetic algorithm to analyze the featured ordered structures in the shear layers and quantitatively characterize the amplitude modulation between the large- and small-scale flow structures. As a result, we find that the coverts-generated perturbations induce an increase in the high-frequency ( f = 91.2 Hz) coherence between the leading- and trailing-edge shear layers from 40% to 70%, leading to a reduction of the flow separation bubble on the upper wing. The present work reveals that the artificial bionic coverts have leading-edge flow separation control effectiveness and shows the engineering potential for aircrafts and unmanned aerial vehicles.

Effect of 2D ice accretion on turbulent boundary layer and trailing-edge noise

Trailing-edge noise is known to be sensitive to airfoil shapes, and ice accretion is one cause of an airfoil shape deformation. This paper investigates how trailing-edge noise is affected by the airfoil shape deformation due to ice accretion. The formation of ice-induced flow separation and the development of a turbulent boundary layer are analyzed to understand the correlation between the altered flow physics due to ice accretion inside the boundary layer and trailing-edge noise. The near-wall flow behind the leading-edge ice accretion is analyzed by using Reynolds-Averaged Navier Stokes CFD in OpenFOAM, and trailing-edge noise is investigated using an empirical wall pressure spectrum model in conjunction with Amiet’s trailing-edge noise theory. Validations of tools against measurement data are presented. Liquid water content, freestream velocity, and ambient temperature are varied to investigate the impact of flow conditions on the ice accretion shape and the resulting boundary layer flow characteristics at the trailing edge. It is found that a more significant leading edge deformation due to ice accretion generates larger ice-induced flow separation bubbles, which increases the trailing-edge boundary layer thickness. As a result, an increase in low- and mid-frequency noise is observed. The purpose of this paper is not only to understand the effect of ice accretion on trailing-edge noise but also to comprehensively analyze how flow physics inside the turbulent boundary layer is altered by the presence of various ice accretion shapes.

Method for utilizing PIV to investigate high curvature and acceleration boundary layer flows around the compressor blade leading edge

Particle Image Velocimetry (PIV) is a well-developed and contactless technique in experimental fluid mechanics, but the strong velocity gradient and streamline curvature near the wall substantially limits its accuracy improvement. This paper presents a data processing procedure combining conventional PIV and newly developed Mirror Interchange (MI) based Interface-PIV for the measurement of the boundary layer parameter development in the blade leading edge region. The synthetic particle images are used to analyze the measurement errors in the entire procedure. Overall, three types of errors, namely the errors caused by the Window Deformation Iterative Multigrid (WIDIM) algorithm, the discrete data interpolation and integration, and the wall offset uncertainty, comprise the main measurement error. Specifically, the errors due to the discrete data interpolation and integration and the WIDIM algorithm comprise the mean bias, which can be corrected through the error analysis method proposed in the present work. Meanwhile, the errors due to the WIDIM algorithm and the wall offset uncertainty contribute to the measurement uncertainty. Computational fluid dynamics-based synthetic particle flows were generated to verify the newly developed PIV data processing procedure and the corresponding error analysis method. Results showed that the data processing method could improve the accuracy of PIV measurements for boundary layer flows with high curvature and acceleration and even with significant flow separation bubbles. Finally, the data processing method is also applied in a PIV experiment to investigate the boundary layer flows around a compressor blade leading edge, and several credible boundary flow parameters were obtained.

Transition-based constrained large-eddy simulation method with application to an ultrahigh-lift low-pressure turbine cascade flow

A transition-predictive Reynolds-averaged Navier–Stokes (RANS) model is introduced as the Reynolds controlling condition to constrained large-eddy simulation (TrCLES) to improve the ability of this method to predict wall-bounded laminar–turbulent transition flows. In the TrCLES method, the constraint conditions for total Reynolds stress and heat flux are only imposed to the near-wall region in transitional and turbulent boundary layer. The newly proposed method recovers direct numerical simulation in the laminar boundary region, and retrieves the traditional large-eddy simulation method in the far-wall regions. The TrCLES method is validated in simulations of external flow around the Eppler 387 (E387) airfoil and internal flow past the ultrahigh-lift low-pressure turbine T106C cascade. The improved delayed detached-eddy simulation (IDDES) method, CLES method based on full-turbulence shear stress transport model and RANS method with a three-equation transition model are also evaluated in comparison with the available experimental and numerical data. As expected, the TrCLES method can predict laminar separation bubble and separation-induced transition process in both the E387 and T106C flows pretty well. In contrast, neither IDDES nor the original CLES can provide reasonable prediction for the laminar separation-induced transition phenomenon. The validity and fidelity of the TrCLES method are further verified by simulations of the T106C cascade flows in a wider range of exit Reynolds and Mach numbers. It is shown that the TrCLES method can not only predict the time-averaged aerodynamic quantities such as isentropic Mach number and exit kinetic energy loss very well, but also capture the laminar separation bubble and unsteady flow structures with satisfactory accuracy.

Computational Analysis of Active and Passive Flow Control for Backward Facing Step

The internal steady and unsteady flows with a frequency and amplitude are examined through a backward facing step (expansion ratio 2), for low Reynolds numbers (Re=400, Re=800), using the immersed boundary method. A lower part of the backward facing step is oscillating with the same frequency as the unsteady flow. The effect of the frequency, the amplitude, and the length of this oscillation is investigated. By suitable active control regulation, the recirculation lengths are reduced, and, for a percentage of the time period, no upper wall, negative velocity, region occurs. Moreover, substituting the prescriptively moving surface by a pressure responsive homogeneous membrane, the fluid–structure interaction is examined. We show that, by selecting proper values for the membrane parameters, such as membrane tension and applied external pressure, the upper wall flow separation bubble vanishes, while the lower one diminishes significantly in both the steady and the unsteady cases. Furthermore, for the time varying case, the length fluctuation of the lower wall reversed flow region is fairly contracted. The findings of the study have applications at the control of confined and external flows where separation occurs.

The influence of the flow separation bubble and transition location on the profile drag of three 4-digit NACA aerofoil profiles

Modeling of the flow over aerofoil profiles at low Reynolds numbers is difficult due to the complex physics associated with the laminar flow separation mechanism. Two major problems arise in the estimation of profile drag: (1) the drag force at low Reynolds numbers is extremely small to be measured in a wind tunnel by force balance techniques, (2) the profile drag is usually calculated by pressure integration, hence the skin friction component of drag is excluded. In the present work, three different 4-digit NACA aerofoils are investigated. Measurements are conducted in an open-ended subsonic wind tunnel, while numerical work is performed by time Reynolds-averaged Navier Stokes (RANS) coupled with the laminar-kinetic-energy ( K-kl-w) turbulence model. The influence of the flow separation bubbles and transition locations on the profile drag is discussed and addressed. This paper gives important insights into importance of measurements at low Reynolds numbers for better aerodynamic loads predictions.

Effect of the Inlet Boundary Conditions on the Flow over Complex Terrain Using Large Eddy Simulation

For large eddy simulation, it is critical to choose the suitable turbulent inlet boundary condition as it significantly affects the calculated flow field. In this paper, the effect of different inlet boundary conditions, including random method (RAND), Lund method, and divergence-free synthetic eddies method (DFSEM), on the flow in a channel with a hump are investigated through large-eddy simulation. The simulation results are further compared with experimental data. It has been found that turbulence is nearly fully developed in the case based on the Lund method, not fully developed in the case based on DFSEM, and not developed in the case based on the RAND method. In the flow region before the hump, mean velocity profiles in the case applying the Lund method gradually fit the law of the wall as the main flow moves towards the hump, but the simulation results based on the RAND and DFSEM methods cannot fit the wall function. In the flow region after the hump, cases applying Lund and DFSEM methods could relative precisely predict the size of turbulent bubble and turbulent statistics profiles. Meanwhile, the case based on the RAND method cannot capture the positions of flow separation and re-attachment point and overestimates the turbulent bubble size. From this research, it could be found that different turbulent inflow generation methods have a manifested impact on the flow separation and re-attachment after the hump. If the coherent turbulence is maintained in the approach flow, even though turbulent intensity is not large enough, the simulation can still predict the flow separation and turbulent bubble size relative precisely.

High-resolution direct numerical simulations of flow structure and aerodynamic performance of wind turbine airfoil at wide range of Reynolds numbers

The objective of this study is to develop direct numerical simulations (DNS) to investigate the aerodynamic performance, transition to turbulence, and to capture the laminar separation bubble occurring on a wind turbine blade. Simulations are conducted with spectral/hp element method to investigate the details of flow separation bubble over wind turbine blades with NACA-4412 airfoil at wide range of design parameters. This airfoil is chosen because recent studies have shown that it is challenging to capture the details of the flow instabilities and pressure fluctuations in the separated shear layer of wind turbines by experimental methods. Furthermore, owing to more accurate development of DNS, the separated bubbles at high Reynolds numbers are captured. The results show that the vortex structures shed from the trailing edge of the airfoil by raising the angle of attack (α). Consequently, the fully turbulent flow develops downstream of the trailing edge (Karman vortex). Moreover, the pressure fluctuation significantly increased by raising α. However, some rolling up of the flow structures, similar to Kelvin–Helmholtz rolls, on the pressure surface near the trailing edge, are observed at α>12°. The separation point was delayed from Xsep/C = 0.19 to 0.58 by decreasing α from 16 to 0 at Re = 5 × 104.

Heat transfer and flow around cylinder: Effect of corner radius and Reynolds number

This numerical study aims to investigate the flow around, heat transfer from, and aerodynamic forces on a square cylinder with the corner radius ratio r/R = 0.0 – 1.0 and Reynolds number Re = 40 180, where r is the cylinder corner radius and R is the half side width of the cylinder. The critical Reynolds number (Recr) for the onset of vortex shedding is found to vary from 49.5 to 46.75 when r/R is increased from 0.0 to 1.0. The dependence of the flow topology on the Re – r/R domain leads to four major flow patterns: steady trailing-edge separated flow, unsteady trailing-edge separated flow, unsteady separation-bubble flow, and unsteady leading-edge separated flow which are distinguished by flow separation, reattachment, boundary layer, and separation bubble. Accordingly, forces, Strouhal number, Nusselt number, and wake structure parameters all are dependent on r/R and Re. The dependence of each parameter is presented in the r/R – Re domain. The Strouhal number and Nusselt number are equally sensitive to both r/R and Re, increasing with increasing r/R and Re. Forces, on the other hand, are more sensitive to Re than r/R. The maximum velocity in the shear layer increases with r/R and Re while the boundary-layer displacement thickness behaves oppositely, shrinking with increasing r/R and Re.

Unsteady effects in a hypersonic compression ramp flow with laminar separation

Abstract

Aerodynamic Characteristics and Laminar Bubble Separation Study on a Generic Light Aircraft Model

This paper presents the aerodynamic characteristics, flow separation and laminar bubble analysis of a generic UTM-LST light aircraft model at low Reynolds number. The complex interaction between flow separation and laminar bubble is unclear to date. The model has overall length of 1.3m and wingspan of 1.5m and has been designed for wind tunnel experiments in Universiti Teknologi Malaysia Low Speed Wind Tunnel, Aerolab. The aircraft model is equipped with several control surfaces such as ailerons, rudder, elevators and flaps. The experiments were conducted at the speed of 35 m/s corresponding to Reynolds number of 0.515 x 106 and at angles of attack ranging from 0° to 16°. The experiments were performed at several pitching and yawing configurations. In order to investigate the effects of control surfaces, several control surfaces were changed during the experiments; for this paper, however, only elevator changes will be highlighted. Three measurement techniques were employed during the experiments; the first one was the Steady balance, the second was the surface pressure while the last one was the tuft flow experiment. The main observation from steady balance data was that the aircraft possesses longitudinal static stability for all test cases. The main observation from the surface pressure measurement and tuft experiments is that the laminar bubble separation occurred at lower angles of attack of the wing. This separation is seen to be travelling towards the leading edge as the angle of attack is increased and eventually results in flow separation.

Shock–Boundary Layer-Interaction Control Through Recirculation in a Hypersonic Inlet

Abstract This technical brief presents a flow separation mitigation device, called cavity-recirculator that can be used to control flow separation during shock wave–boundary layer interaction (SBLI) in high-speed intake flows. The cavity-recirculator isolates the flow separation bubble generated at the SBLI spot, thereby thinning the boundary layer and reducing the blockage of the inviscid stream in the duct. The device has a potential application in scramjet engine intakes/isolators. The cavity-recirculator was tested on a single-ramp-compression intake model in a hypersonic shock tunnel, in a freestream of Mach 8 (±0.1). The device operation and effectiveness were assessed by flow visualization and pressure measurements in the test model. The measurements and visualization displayed a mitigation of flow separation through an improved flow field with a single shock train and the absence of flow separation shock, in the inlet.

Numerical studies of the flow structure and aerodynamic forces on two tandem square cylinders with different chamfered-corner ratios

Three-dimensional large eddy simulations were carried out to investigate the flow over two tandem square cylinders for a fixed spacing ratio of L/D = 4 and a Reynolds number of 5.3 × 103, where L is the cylinder center-to-center distance between the cylinders and D is the cylinder width. The corners of each cylinder were chamfered with a ratio of ξ = B/D = 0%, 5%, 10%, and 15%, where B is the chamfered corner dimension. The focus is given on how ξ influences the flow structure, wake recirculation bubble, flow separation bubble, Strouhal number (St), aerodynamic force, and phase lag (ϕ) between vortex sheddings from the cylinders. With increasing ξ, the recirculation bubble length and minimum velocity in the wake of the upstream cylinder remain more or less constant and are close to those of a single cylinder, while the minimum velocity in the wake of the downstream cylinder dramatically drops between ξ = 0% and 5%. While the flow over the downstream remains attached on the side surfaces, that over the upstream cylinder forms primary and secondary side-surface bubbles on the side surface, and both shrink with increasing ξ. In addition, the leading chamfered corners are accompanied by corner bubbles that play an important role in the flow structure modification. Therefore, the time-mean drag and fluctuating lift of the upstream cylinder remarkably decrease for 0% ≤ ξ ≤ 5% and remain almost unchanged for 5% < ξ ≤ 15%, whereas those on the downstream cylinder diminish in the whole range of 0% ≤ ξ ≤ 15%. The Strouhal number (St) being identical for the two cylinders grows from 0.104 to 0.132 when ξ is increased from 0% to 15%. Overall, the influence of ξ on the wake structure and aerodynamics is remarkable for 0% ≤ ξ ≤ 5% because of the corner bubbles and is less for 5% < ξ ≤ 15%.