Laminar Boundary Layer Separation Research Articles

Impact of Low Freestream Turbulence and Surface Roughness on the Flat-Plate Boundary Layer Transition Under a High-Lift Airfoil Pressure Gradient

Abstract Effects of low incoming freestream turbulence level (Tu < 1%) and surface roughness on the transition of flat-plate boundary layers under a high-lift airfoil pressure gradient have been investigated. Time-resolved streamwise and wall-normal velocity fields with surface roughness values of Ra/C = 0.065 × 10−5, 4.417 × 10−5, and 7.428 × 10−5 have been measured at a fixed Reynolds number of 5.2 × 105 and freestream turbulence intensity of 0.2%. For the reference smooth surface of Ra/C = 0.065 × 10−5, a laminar separation bubble forms from about 64% to 83% of the chord length. Increasing surface roughness has little impact on the laminar boundary layer separation onset but reduces the height and length of the separation bubble and induces earlier transition. For Ra/C = 4.417 × 10−5, displacement thickness during transition is slightly thinner and the overall momentum deficit is slightly lower than those for Ra/C = 0.065 × 10−5. For Ra/C = 7.428 × 10−5, the separation bubble becomes hardly visible as the transition mode approaches the attached mode, and turbulent mixing by the wall-bounded turbulence becomes dominant. In addition, the portion of turbulent wetted area increases due to earlier transition, and momentum deficit increases more rapidly in the turbulent wetted area. Thus, the overall momentum deficit for Ra/C = 7.428 × 10−5 is larger than that for Ra/C = 0.065 × 10−5.

Propeller-noise reduction by microfiber coating on a blade surface

The popularity of small aerial vehicles has dramatically increased in recent years and propeller noise from such vehicles is a public health concern. Further advancement and utilization of small aerial vehicles require a substantial focus on noise reduction. Microstructures using surface and coating technology are applied in a variety of ways to address this engineering challenge. This study investigates a microfiber coating as a passive means for reducing propeller noise. The microfiber coating is comprised of a fibrous structure and has been previously shown to be a passive mean for reducing drag on a circular cylinder. To begin testing the efficacy of the microfiber coating for propeller noise reduction, microfiber-coated strips are placed at different spanwise locations on propeller blades. The sound pressure level produced by the rotating propeller is measured using a sound-level meter. The microfiber-coated propeller exhibited a lower sound pressure level than that of the uncoated propeller. At a Reynolds number of 7.4 × 104 based on the chord at the 75% spanwise station of the propeller blade, the microfiber-coated propeller achieved a noise reduction of up to 1.6 dBA compared to that of the uncoated propeller. The microfiber coating is effective in reducing broadband noise associated with the interaction of the vortex shedding associated with laminar boundary layer separation. It is found that the noise-reduction performance is a function of the spanwise location of the microfiber-coated strips.

Piezoelectric Based Adaptive Wing for Flow Control

An attempt is made to develop a piezoelectric based adaptive wing concept that can be used in Micro Air Vehicle (MAV) and Unmanned Air Vehicle (UAV) applications. Laminar boundary layer separation is considered to be a significant problem in low Reynolds number vehicles like MAV. Recently, periodic excitation has gained much interest in the control of flow separation through energizing the boundary layer. In the present study, a piezoelectric bimorph actuator has been used to develop a mechanism that will deflect or vibrate a segment of the wing to produce the required thickness variation and achieve flow control. A Finite Element model of the adaptive wing is developed in ANSYS and numerical results are compared with experiments. A five percent increase in the camber has resulted in 4.39 percent additional lift at higher angles of attack.

A new concept of diminishing cavitation inception number for hydrofoils

A new concept of diminishing cavitation inception number is applied to sheet cavitation over contours traditional for sections of marine propeller blades. An inverse hydrodynamic problem is solved for a segment located downstream from the point of separation of laminar boundary layer at highest angles of attack of the contour. The pressure along the forward part of this segment will be constant and it will have a prescribed increase along its rear part up to the point of merging the corrected streamline with the original contour. Description of employed numerical method is supplemented by examples of enhancement of traditional hydrofoils. The substantial improvement of their baskets of minimal pressure is obtained.

Wall-Modeled Large Eddy Simulation and Detached Eddy Simulation of Wall-Mounted Separated Flow via OpenFOAM

Considering grid requirements of high Reynolds flow, wall-modeled large eddy simulation (WMLES) and detached eddy simulation (DES) have become the main methods to deal with near-wall turbulence. However, the flow separation phenomenon is a challenge. Three typical separated flows, including flow over a cylinder at ReD = 3900 based on the cylinder diameter, flow over a wall-mounted hump at Rec = 9.36 × 105 based on the hump length, and transonic flow over an axisymmetric bump with shock-induced separation at Rec = 2.763 × 106 based on the bump length, are used to verify WMLES, shear stress transport k-ω DES (SST-DES), and Spalart–Allmaras DES (SA-DES) methods in OpenFOAM. The three flows are increasingly challenging, namely laminar boundary layer separation, turbulent boundary layer separation, and turbulent boundary layer separation under shock interference. The results show that WMLES, SST-DES, and SA-DES methods in OpenFOAM can easily predict the separation position and wake characteristics in the flow around the cylinder, but they rely on the grid scale and turbulent inflow to accurately simulate the latter two flows. The grid requirements of Larsson et al. (δ/Δx,δ/Δy,δ/Δz≈(12,50,20)) are the basis for simulating turbulent boundary layers upstream of flow separation. A finer mesh (δ/Δx,δ/Δy,δ/Δz≈(40,75,40)) is required to accurately predict the separation and reattachment. The WMLES method is more sensitive to grid scales than the SA-DES method and fails to obtain flow separation under a coarser grid, while SST-DES method can only describe the vortices generated by the separating shear layer, but not within the turbulent boundary layer, and overestimates the separation-reattachment zone based on the grid system in this paper.

Hypersonic shock impingement studies on a flat plate: flow separation of laminar boundary layers

Interactions between shock waves and boundary layers produce flow separations and augmented pressure/thermal loads in hypersonic flight. This study provides details of Mach 7 impinging-shock-flat-plate experiments conducted in the T4 Stalker Tube. Measurements were taken at flow conditions of Mach 7.0 (2.44 MJ kg $^{-1}$ ) and Mach 7.7 (2.88 MJ kg $^{-1}$ ) flight enthalpies with a range of freestream unit Reynolds numbers from $1.43 \times 10^{6}$ m $^{-1}$ to $5.01 \times 10^{6}$ m $^{-1}$ . A shock generator at $12^{\circ }$ or $16^{\circ }$ to the freestream created an oblique shock which impinged on a boundary layer over a flat plate to induce flow separation. The flow field was examined using simultaneous measurements of wall static pressure, heat transfer and schlieren visualisation. Measured heat transfer along the flat plate without the shock impingement indicated that the boundary layer remained laminar for all flow conditions. The shock impingement flow field was successfully established within the facility test duration. The onset of separation was observed by a rise in wall pressure and a decrease in heat transfer at the location corresponding to the stem of the separation shock. Downstream of this initial rise, an increased pressure and higher heating loads were observed. The heat-transfer levels also indicated an immediate boundary layer transition due to the shock impingement. The separation data of the present work showed good agreement with our previous work on shock impingement on heated walls (Chang et al., J. Fluid Mech., vol. 908, 2021, pp. 1–13). A comparison with the previous scaling indicated that the separation also relates to the pressure ratio and the wall temperature parameter.

Experimental study of laminar-to-turbulent transition in an adverse pressure gradient flow

Experimental results on laminar-to-turbulent transition in a separated laminar boundary layer formed over a flat plate in an adverse pressure gradient are provided. Experiments have been performed for a range of Reynolds numbers (1.4 − 2.4 · 105) and for two freestream turbulence levels (Tu =3.6 and 5.3%). The measurements of mean and fluctuating velocity profiles have been carried out at inlet to the test section using Constant Temperature Anemometry (CTA). The data might be useful for validation and development of transition models. In order to characterize the dynamics leading to the transition process in the separated boundary layer on the plate surface, the Particle Image Velocimetry (PIV) measurements have been performed. The results show that the transition process in the separated boundary layer is influenced by breakdown of the Klebanoff streaks.

Analysis of laminar boundary-layer separation in retarded flow over bodies of revolution

The laminar boundary-layer separation phenomenon is an interesting and important aspect of boundary-layer flows. It occurs in various physical situations because of a decrease in wall shear stress caused by retarded flow velocity, among other things. Flow separation can be prevented or delayed by utilizing bodies of revolution because the surface transverse curvature produces a favorable pressure gradient, which in turn increases the wall shear stress that keeps the flow attached to the surface. Bodies of revolution, whose body contours follow a power-law form, also play a vital role in delaying flow separation. Bodies of revolution of varying cross sections and involving surface transverse curvature are utilized to examine their effects on flow separation. In particular, a convex transverse curvature has been considered owing to its effects on the nature of the favorable pressure gradient, which delays the flow separation. A Görtler-type retarded flow velocity was considered in this study to investigate the flow separation process. A detailed analysis is provided to understand the flow separation by calculating the separation points under various assumptions. It has been observed that the body contour exponent n and the convex transverse curvature parameter k play an assistive role in the delay of boundary-layer separation even under the influence of strong retardation. The results are presented through various tables and graphs to highlight the role of the power-law exponent of external velocity m, convex transverse curvature parameter k, and body contour exponent n on the separation points.

The influence of micro vortex generator on inception cavitation

The objective of the present paper is to investigate the influence of the micro vortex generator (mVG) on the inception cavitation number and mode around a National Advisory Committee for Aeronautics 66 hydrofoil. Two different sets of mVG with varying position are employed in this paper, i.e., the mVG-1 (located upstream of the laminar separation point of the baseline hydrofoil) and the mVG-2 (located in the laminar separation zone of the baseline hydrofoil). A high-speed camera is applied to visualize the inception cavitating structures, and numerical simulation is assisted to the effect of mVG. The results indicate that compared to the baseline hydrofoil, the mVG-1 can promote the earlier inception cavitation while the mVG-2 delays the inception, especially for the cases with smaller angle of attack (α = 4°–8°). For the mVG-1 hydrofoil, there are two reasons to be responsible for this phenomenon. One is that the fingerlike vortex at the rear of mVG-1 induces the fingerlike vortex cavitation earlier. The other is that the mVG-1 increases the length of the laminar separation bubble (LSB) by comparison with the baseline hydrofoil, thus causing a cavitation due to the laminar boundary layer separation. For the mVG-2 hydrofoil, it is located at the high-pressure zone of leading edge and reduces the length of the LSB. More precisely, the fingerlike vortex in the high-pressure zone is not enough to induce a fingerlike vortex cavitation, and the smaller length of the LSB than that of the baseline hydrofoil suppressing the cavitation at some angles of attack.

Numerical simulation of separation induced laminar to turbulent transition over an airfoil

The article deals with the numerical simulation of flows with laminar to turbulent transition due to the separation of boundary layer. Mathematical model consists of the Reynolds averaged Navier–Stokes equations which are completed by the explicit algebraic Reynolds stress model (EARSM) of turbulence. The EARSM model is enhanced with algebraic model of bypass transition which is further modified by the additional source term for prediction of laminar boundary layer separation induced transition. Numerical solution is obtained by the finite volume method based on the second order HLLC scheme and explicit Runge–Kutta method. The proposed method is then tested on several cases of low-Reynolds subsonic transitional flows including flows over the SD 7003 airfoil with various angles of attack and flow past the NACA 0012 airfoil.

Adaptive separation control of a laminar boundary layer using online dynamic mode decomposition

Abstract

Numerical Investigation of the Transient Nature of a Laminar Separation Bubble in Hypersonic Flow

Shock-wave/boundary-layer interaction is a prime research topic in the design of hypersonic vehicles. For the proper designing of hypersonic vehicles, especially, their thermal protection systems, it is necessary to understand the time-dependent behavior of the wall properties and the corresponding pressure and thermal loads. Hence investigation is carried out to understand the unsteady nature of shock-induced laminar boundary layer separation for a simple canonical configuration of a two-dimensional ramp. A density-based non-reactive Navier-Stokes solver named rhoCentralFoam in openFOAM is employed in the present investigation. The detailed physics of laminar boundary layer separation in hypersonic flow is investigated through timewise behavior of the streamline patterns and the surface properties, such as the pressure, heat flux, and friction coefficients. It is found that at the onset of fluid flow it is its inviscid characteristics that are predominant and there is occurrence of very small separation bubble. The separation bubble grows, as the involvement of viscous characteristics increases gradually, and ultimately it attains a steady state. Because of separated boundary layer the heat flux and skin friction coefficients are found to follow the diffused V and deformed W shapes, respectively. The peaks of pressure and thermal loads are found to exist in the vicinity of the reattachment region. These peaks are higher in the initial stage of the process and attain steady state eventually. The high pressure and thermal loads may cause structural damage to the vehicle and hence their correct prediction is necessary. Thus, the current investigation is helpful in the design of thermal protection systems of hypersonic vehicles.

Design of Supercritical Low-Reynolds-Number Airfoils for Fixed-Wing Flight on Mars

Aerodynamic shape optimization for the high-subsonic low-Reynolds number flow regime represents an area of on-going research. The interaction between supercritical compressible flow and laminar boundary layer separation is not well understood due to the high challenges associated with setting up relevant experimental work. However, in the design of future fixed wing aircraft for flight in extra-terrestrial atmospheres, such flow conditions might commonly occur. The present study presents a family of single-point and multi-point optimized airfoils designed for high-subsonic flight at a high-lift condition in the Martian atmosphere. A gradient based optimizer is used, with a second-order finite-volume flow solver and a second-order continuous adjoint solver for determining surface sensitivities with respect to the objective function of minimizing drag. Both fully turbulent and transitional flow are considered, to evaluate the impact on the resulting design and to stress the importance of continuing research to develop robust shape optimization including laminar boundary layer and transition prediction. Both on-design and off-design conditions are evaluated, the airfoils obtained when considering transition effects demonstrating good overall performance.

A review on hydrodynamics of horseshoe vortex at a vertical cylinder mounted on a flat bed and its implication to scour at a cylinder

A comprehensive review of the local scour due to vortical flow around a cylindrical bridge pier under steady current is presented in this paper. The mechanism of the formation of vortices, the size, velocity and strength of horseshoe vortex (HSV), formation of the HSV by the separation of laminar and turbulent boundary layer and the scour around a cylindrical pier due to vortices have been presented. The complexity involved in the scour-related calculations, and the scope for future research are discussed in the last section.

Low Reynolds number effects on aerodynamic loads of a small scale wind turbine

Small-scale or scaled-down wind turbines for model experiments mostly operate in low-Reynolds-number flow. The nonlinear variations of aerodynamic coefficients with respect to the angle of attack caused by viscous effects and laminar boundary layer separation affect the wind turbine performance under these conditions. Although the vortex lattice method (VLM) is an efficient way to predict rotor performance, it tends to suffer from numerical error because nonlinear aerodynamic characteristics cannot be considered. In this study, the nonlinear vortex lattice method (NVLM) is adopted to compute the aerodynamic loads of two small-scale wind turbines. This method involves a sectional airfoil look-up table and vortex strength correction and can be applied to a wide range of operating conditions. The simulations of TU Delft and NTNU wind turbines are conducted to validate the prediction capability of numerical models by comparing predictions with the measurements. It was found that the overall results from the NVLM simulation are more accurate than the VLM results, which implies that the nonlinear aerodynamic characteristics associated with low-Reynolds-number flow should be considered to accurately assess the aerodynamic performance of small-sized wind-turbines, particularly at the low tip speed ratio at which the rotor blade may experience flow separation and dynamic stall.

Effect of active oscillation of local surface on the performance of low Reynolds number airfoil

In low Reynolds number conditions, laminar boundary layer separation and the consequent stall are usual, which are harmful to the aerodynamic characteristics of an airfoil. Energy addition to the flow by active oscillation of local surface can effectively restrain the flow separation and improve the performance. In the present paper, flow control based on active oscillation of local surface on an E387 airfoil is studied with numerical methods. The Reynolds number is Re = 30000 and the location of the oscillation region is a part of the upper surface near the leading edge. The respective relationships between the flow control effect and excitation parameters such as oscillation amplitude and excitation frequency are obtained. The flow mechanisms such as the behavior of the separation vortex and the variation of the equivalent shape are analyzed in both time-averaged and unsteady aspects. Furthermore, flow control effect at different angles of attack is also investigated. The results show that laminar separation can be effectively restrained, and the lift enhancement and drag reduction are evident.

Laminar boundary layer separation over a fluttering panel induced by an oblique shock wave

Panel flutter in the presence of oblique shock waves and the associated shock wave/boundary-layer interaction have been identified as one of typical phenomena in the design and optimization of air-breathing, high-speed flight vehicles. The current study investigates this phenomenon using a previously considered two-dimensional model where an elastic panel with both ends pinned is impinged at the mid-point by an oblique shock wave with specified strength. An in-house code was used to solve the Euler or the full viscous compressible Navier–Stokes equation and nonlinear structural dynamics of the panel, where the conventional serial staggered algorithm was adopted for the fluid–structure interaction. As compared with previous studies of this topic, we focus on the effect of surface velocity feedback (i.e., using boundary blowing and suction for flow control), as well as the effect of the upstream boundary layer thickness, on the panel’s dynamic behavior, surface pressure distribution, and boundary-layer separation. The results show that for inviscid flow, boundary control with feedback gain above one can suppress the panel vibration; however, this effect is not clear for viscous flow, where feedback gain below one has some effect on attenuation of vibration. Further study shows that the panel velocity based control introduces a phase shift for the pressure in the inviscid flow as a damping effect, but the effect is not as strong in the viscous flow. Finally, the boundary layer thickness has a non-monotonic effect on the panel flutter and flow separation. At intermediate thicknesses considered here, the panel flutter is reduced and separation becomes less oscillatory.

An experimental study of trailing edge noise from a pitching airfoil.

In this study, the far-field noise from a pitching NACA 0012 airfoil was measured at a Reynolds number of 6.6 × 104. The pitching motion was in sinusoidal functions with a mean incident angle of 0°. Cases with the pitching amplitude varying from 7.5° to 15° and frequency from 3 to 8 Hz were tested, corresponding to the reduced frequency from 0.094 to 0.25. A microphone was placed in the far-field and the particle image velocimetry technique was utilized to study the flow structures near the trailing edge. The short-time Fourier transformation results of the noise signals revealed that a high-level narrow-band noise hump occurred at a specific angle of attack in a pitching cycle. At the corresponding moment, a coherent vortex street convecting on the airfoil surface was observed, and the vortex shedding frequency was in good agreement with the central frequency of the noise hump. The occurrence of the noise humps was attributed to the laminar boundary layer separation. In one pitching period, the moment when the narrow-band noise hump occurs is independent from the pitching amplitude and it is delayed as the pitching frequency increases. Larger pitching frequency or amplitude results in lower peak level of the noise humps.

Numerical Investigation of Steady and Harmonic Vortex Generator Jets Flow Separation Control

Active flow control of canonical laminar separation bubbles by steady and harmonic vortex generator jets (VGJs) was investigated using direct numerical simulations. Both control strategies were found to be effective in controlling the laminar boundary-layer separation. However, the present results indicate that using the same blowing amplitude, harmonic VGJs were more effective and efficient at reducing the separated region than the steady VGJs considering the fact that the harmonic VGJs use less momentum than the steady case. For steady VGJs, longitudinal structures forming immediately downstream of the injection location led to the formation of hairpin-type vortices, causing an earlier transition to turbulence. Symmetric hairpin vortices were shown to develop downstream of the forcing location for the harmonic VGJs, as well. However, the increased control effectiveness for harmonic VGJs’ flow control strategy is attributed to the fact that the shear-layer instability mechanism was exploited. As a result, disturbances introduced by VGJs were strongly amplified, leading to the development of large-scale coherent structures, which are very effective at increasing the momentum exchange, thus limiting the separated region.

POD Analysis of Entropy Generation in a Laminar Separation Boundary Layer

Separation of laminar boundary layer is a great source of loss in energy and power machinery. This paper investigates the entropy generation of the boundary layer on the flat plate with pressure gradient. The velocity of the flow field is measured by a high resolution and time related particle image velocimetry (PIV) system. A method to estimate the entropy generation of each mode extracted by proper orthogonal decomposition (POD) is introduced. The entropy generation of each POD mode caused by mean viscous, Reynolds normal stress, Reynolds sheer stress, and energy flux is analyzed. The first order mode of the mean viscous term contributes almost 100% of the total entropy generation. The first three order modes of the Reynolds sheer stress term contribute less than 10% of the total entropy generation in the fore part of the separation bubble, while it reaches to more than 95% in the rear part of the separation bubble. It indicates that the more unsteady that the flow is, the higher contribution rate of the Reynolds sheer stress term makes. The energy flux term plays an important role in the turbulent kinetic energy balance in the transition region.