Separation Bubble Research Articles

Control of Reverse Flow over Cantilevered Swept Blades Using Passive Camber Morphing

High-speed rotorcraft experience a region of reversed flow over the retreating blade at high advance ratios. This reverse flow region results in multiple unfavorable effects, including negative lift, increased drag, and a large pitching moment impulse. The effect of the sweep angle on a rotor blade in reverse flow was explored, and it was demonstrated that the introduction of a trailing edge reflex camber could mitigate the adverse effects seen in this regime. A cantilevered, finite-span NACA 63-218 blade was examined at a Reynolds number Rec=2.38×105 both with and without a 10° trailing edge reflex camber. Three blades were tested: one with 20° forward sweep, one with 20° backward sweep, and one with no sweep. The experiments consisted of load measurements across a range of angles of attack, as well as volumetric flowfield measurements using stereoscopic particle image velocimetry on the two swept models at 10° angle of attack in reverse flow. The flowfields exhibited a highly three-dimensional separation bubble that existed on the blades in both backward and forward sweep conditions. Cambering of the trailing edge led to almost complete flow attachment on the backward swept model and a significant reduction in separation over the forward swept model. Cambering also led to a large reduction in drag in the entire reverse flow regime in both swept and unswept conditions. A large reduction in pitching moment and a smaller reduction in negative lift were also observed at moderate angles of attack.

Water tunnel testing of downwind yacht sails

Downwind yacht sails, such as spinnakers, are low-aspect-ratio highly cambered wings with a sharp leading edge. They are characterised by substantial three-dimensional flow separation and are thus modelled with difficulty with numerical simulations. Furthermore, accurate full-scale validation data are not available. The first quantitative flow measurements have only recently been achieved in water tunnels. In this study, we aim to provide guidelines on this emerging sail testing methodology. We consider six model-scale rigid models at average-chord-based Reynolds numbers ranging from 5 870 to 61 870. A critical Reynolds number is identified, below which relaminarisation of the reattached boundary layer downstream of the leading edge separation bubble occurs. Both lift and drag increase monotonically at subcritical Reynolds numbers while remaining about constant at transcritical Reynolds numbers. The critical Reynolds number decreases with increasing incidence and is insensitive to the blockage ratio. Spinnakers are normally sailed in front of the mainsail, whose circulation is found to generate an approximately 3∘\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$3^{\\circ }$$\\end{document} upwash on the spinnaker and higher flow velocity on both sides of it. These findings provide guidelines for the experimental testing of spinnaker-like wings in water tunnels and provide new insights into the flow and experimental testing of highly cambered wings with massive flow separation at low Reynolds numbers.

Effects of herringbone riblets on shock-wave/turbulent boundary-layer interactions

In the experiment reported in this paper, a spanwise array of herringbone riblets strips is applied upstream of the interaction region between an impinging shockwave and a turbulent boundary layer developing along a flat plate boundary layer at a freestream Mach number of 1.85. High-speed schlieren photography, surface oil flow visualization and PSP are used to examine the effect of herringbone riblets on the characteristics of the shockwave-boundary layer interactions. It shows that despite that the height of the riblets is less than 2 % of the local thickness of the undisturbed boundary layer, a profound modification to the shockwave-boundary layer interaction zone is observed, and the overall size of the separation bubble is decreased. The observed control effects are attributed to the large-scale secondary flow structures produced by the directional riblets. To the best knowledge of the author, this is the first study evaluating the control effect of herringbone riblets on shock wave and turbulent boundary layer interactions in supersonic flow using the experimental method.

Averaging theory for heat transfer in circular hydraulic jumps with a separation bubble

Analytical investigations of heat transfer during the vertical impingement of an unsubmerged axisymmetric liquid jet on a horizontal plate have been limited to the regions ahead of the jump. This limitation is due to the complex flow physics in the jump region arising from sudden changes in the flow field. This is addressed in here by extending the averaging theory (AT) introduced by Bohr et al. (Phys. Rev. Lett., vol. 79, issue 6, 1997, pp. 1038–1041) which was further developed by Watanabe et al. (J. Fluid Mech., vol. 480, 2003, pp. 233–265), to describe the heat transfer problem in circular hydraulic jumps including separation. The applicability of the resulting theory to determine the temperature field in the jump region is evaluated using the data available in the literature and also by means of fully resolved numerical solutions. Good agreement is observed for moderate Prandtl numbers. However, for sufficiently high Prandtl numbers, deviations become notable. The reasons for the deviations according to their relevance are (i) monotonically decreasing temperature profile inherent to the AT, whereas the fully resolved numerical solutions exhibit a local maximum in the temperature profile away from the plate; and (ii) inapplicability of the concept of dividing the flow field into a region affected and a region unaffected by heat transfer according to the thermal boundary layer thickness. This concept leads to the overestimation of the temperature close to the wall and to the existence of a threshold Prandtl number, for which the thermal boundary layer thickness does not meet the free surface anymore. Around this threshold Prandtl number, the temperature field shows a discontinuous behaviour.

An Experimental Database for the Analysis of Bursting of a Laminar Separation Bubble

The bursting phenomenon consists in the switch of a laminar separation bubble from a short to a long configuration. In the former case, reduced effects on profile pressure distribution are typically observed with respect to the attached condition. On the contrary, long bubbles provoke significant variations in the loading coefficient upstream of the separation position, with increased risk of stall of the lifting surfaces. The present work presents an experimental database describing separated boundary layers evolving under different Reynolds numbers, adverse pressure gradients and free-stream turbulence levels. Overall, more than 80 flow conditions were tested concerning short and long bubbles for the characterization of separated flows under turbine-like conditions. Measurements were performed on a flat plate geometry using a fast-response Particle Image Velocimetry (PIV) system. For each flow case, two sets of 6000 flow records were acquired with an acquisition frequency equal to 300 and 1000 Hz. Based on existing criteria for the identification of the bursting phenomenon, the flow cases were clustered in terms of short and long bubble states. Additionally, the kind of instability (i.e., convective or absolute) developing into the separated boundary layer was identified based on flow statistics. The present data captures the existing link between the bursting of a laminar separation bubble and the onset of the absolute instability of the separated shear layer, with stationary vortices forming in the dead air region.

Experimental investigation of instantaneous lift on NACA 0018 airfoil section due to a non-harmonic pitching motion

In this study, we explore the problem of an airfoil in a harmonic and arbitrary periodic pitching motion and seek to understand how theoretical estimation with first-order potential theory match experiments. We adopt Jones’ approximation for the Wagner function for lift computations in the time domain and Theodorsen’s model in the frequency domain. We experimentally investigate a two-dimensional symmetric NACA 0018 airfoil in pitching motion under attached flow with reduced frequencies up to 0.25 and Reynolds number varying from 1.5×105 to 3×105. To eliminate the influence of the laminar separation bubble, free-stream turbulence intensity was elevated above the baseline level by placing turbulence generating grid, thus, obtaining better compatibility between the flow conditions in the experiments and the theoretical assumptions. Time-resolved sectional lift is determined by simultaneous static pressure measurement with miniature pressure transducers. The transient response is captured with accuracy in the time domain by quantifying the non-circulatory contribution. Excellent agreement is achieved in the time and frequency domains due to the highly accurate measurement of the instantaneous magnitude and phase of the time-resolved surface pressures. We further explore the applicability of the theory for high-pitching amplitudes, which results in a temporary deviation of the measured lift from the theoretical predictions. As the frequency domain unsteady theory is more widely studied than its time domain counterpart, this study provides a unique opportunity to highlight the significance of time-domain analysis in estimating instantaneous lift in non-harmonic motion.

Research on the influence of spanwise cross-flow on the boundary layer transition of compressor cascade

The cross-flow perpendicular to the inviscid main flow in the boundary layer has potential instability, causing the transition from laminar flow to turbulent flow. In order to explore the mechanism of cross-flow in the blade boundary layer on transition, this paper studies the rectangular cascade of a certain compressor stator blade. Large eddy simulation calculations and flow display experiments for six attack angles with end wall cascades were carried out. It is found that the disturbance is dominated by the two-dimensional Kelvin–Helmholtz (K–H) instability. The transition begins at the position where the separation bubble begins to fall off into a two-dimensional K–H vortex and is completed where the K–H vortex breaks. The closer to the blade root, the later the transition occurs and the smaller the total pressure loss. The cross-flow velocity develops alternately between positive and negative, showing severe instability with more than 4 inflection points. The study on variable angles of attack shows that there is a superposition of two mechanisms, namely, separation bubble transition and cross-flow transition, at an angle of attack from −4° to 10°. In summary, although the separation bubble transition is dominated by K–H vortices, the occurrence of cross-flow instability is closely related to the transition position.

Prediction of mean flow and over-tip shock distribution in pressure-driven tip leakage flows

The flow across the blade tip clearance in turbomachinery is simplified as pressure-driven tip leakage flow (TLF) by isolating it from the mainstream. Based on schlieren visualization and numerical simulations, several common features of TLF are achieved. Consequently, a diffusion model is proposed to evaluate the mean flow and shock motions within the clearance. It takes into consideration the effects of relative wall motion by superposing a fully developed Couette flow. In addition, the over-tip shock waves are treated as repeated sawtooth wave to model the propagation. This approach enables quick and accurate evaluations of the meanflow and shock motions under configurations of stationery and moving casing wall. Given the flow variables at boundaries of the shock region, the meanflow and the evolution of the over-tip shock waves can be achieved instantly with an error less than 2%. Another advantage of this model is it can be non-intrusive. Hence, the challenges, arising from spatial constraints in direct measuring of TLF within the clearance, are surmounted. This is beneficial for locating the tip flow loss and the shock-induced heat load. Two flow mechanisms are unveiled from the predictions: (1) The strongest shock–boundary interaction accompanied by strong momentum exchange occurs above the separation bubble. (2) The oscillation of over-tip shock waves is self-sustained by a feedback loop formed by the pressure-side vortex shedding, shock generation, and shock–boundary interactions.

Direct numerical simulation of turbulence amplification in a strong shock-wave/turbulent boundary layer interaction

In this study, the turbulence amplification mechanism within the strong shock-wave/turbulent boundary layer interaction is investigated using direct numerical simulation (DNS) over a 24° compression ramp with Mach 2.9 flow. A new in-house solver based on the compact finite difference scheme is introduced, and its accuracy is validated by comparing the flow statistics with existing DNS and experimental data. Within the DNS findings, two distinct turbulence kinetic energy (TKE) hotspots are identified. In contrast to previous studies, this study sheds light on shocklets, characterized by mid-frequency features, as a key factor contributing to the second TKE amplification, which occurs near the reattachment point. Streamline coordinate analysis reveals that shear effects dominate TKE production over the flow deceleration effect in the shock-wave/turbulent boundary layer interaction. The shear effect induced by the rolling up of the boundary layer initiates the first TKE amplification near the wall region in proximity to the separation point, followed by flow deceleration due to the main shock wave contributing to TKE generation. The initial detachment of the shear layer enhances the shear contribution. While TKE decreases above the separation bubble due to the positive mean velocity gradient, TKE amplifies again due to the flow deceleration caused by the secondary shock wave. In addition, the intermittently spawning shocklets above the bulge structures enhance the shear effect on the TKE production. Moreover, the generated TKE subsequently transfers to the local pressure minimum line, created by the bulges effect, thereby establishing a spatially converged maximum TKE line.

Assimilating experimental data of a mean three-dimensional separated flow using physics-informed neural networks

In this article, we address the capabilities of physics-informed neural networks (PINNs) in assimilating the experimentally acquired mean flow of a turbulent separation bubble occurring in a diffuser test section. The training database contains discrete mean pressure and wall shear-stress fields measured on the diffuser surface as well as three-component velocity vectors obtained with particle image velocimetry throughout the volumetric flow domain. Imperfections arise from the measurement uncertainty and the inability to acquire velocity data in the near-wall region. We show that the PINN methodology is suited to handle both of these issues thanks to the incorporation of the underlying physics that, in the present study, are taken into account by minimizing residuals of the three-dimensional incompressible Reynolds-averaged Navier–Stokes equations. As a result, measurement errors are rectified and near-wall velocity profiles are predicted reliably. The latter benefits from the incorporation of wall shear-stress data into the PINN training, which has not been attempted so far to the best of our knowledge. In addition to demonstrating the influence of this novel loss term, we provide a three-dimensional, highly resolved, and differentiable model of a separating and reattaching flow that can be readily used in future studies.

Flow mechanism of a highly loaded low pressure turbine cascade with integrated-optimized end wall contouring and root lean

The end wall loss of modern highly loaded low pressure turbine (LPT) has been greatly increased, due to the enhanced secondary flow loss and boundary layer separation loss. Thus, it is of great significance to develop effective flow control strategies to improve the end wall flow condition and aerodynamic performance of modern LPT. This research carried out a numerical investigation on the coupled flow control strategy, which combined non-axisymmetric end wall contouring (NEC) and root tangential lean (RTL), based on a highly loaded LPT cascade (Zweifel = 1.59). Meanwhile, the optimization process was used to get the optimal design parameters of the coupled method NEC&RTL. The results indicate that the optimal coupled configuration can reduce the total pressure loss coefficient by 12.68% and the non-dimensional secondary kinetic energy by 23.91%. Compared with the reference cascade without modification, the coupled method is found to improve the end wall flow conditions: the passage vortex is weakened both in size and strength, mainly attributed to the smaller cross-passage pressure gradient resulting from NEC; the closed separation bubble near end wall and the three-dimensional separation flow before trailing edge are eliminated, due to the great downward pressure gradient near end wall resulting from RTL; and the counter vortex is eliminated and the slender back flow is weakened under the additional coupling flow control effect of NEC&RTL. Therefore, the coupled flow control method can not only highlight the advantages of the independent methods, but also induce external flow control superiorities, demonstrating the application prospect of the coupled flow control strategy on the highly loaded LPT.

Generalized prediction for self-starting performance of two-dimensional hypersonic inlets

The acceleration self-starting performance of a hypersonic inlet is pivotal for ensuring the stable operation of a scramjet. While the geometric configuration of the internal contraction section (ICS) considerably influences the inlet's self-starting performance, the current prediction model solely considers the internal contraction ratio. To encompass the entire geometric configuration's influence on ICS, a generalized prediction model for the acceleration self-starting Mach number of critical hard unstart two-dimensional hypersonic inlets is proposed. This model calculates the theoretical reattachment pressure rise of the main separation bubble within the actual unstarted flow structure of the hypersonic inlet. Additionally, it computes the theoretical pressure rise assuming the main separation bubble is in a critical state. By comparing these pressures, the model evaluates whether the main separation bubble can be sustained under given incoming flow conditions, predicting the self-starting Mach number. This modeling approach offers broad adaptability to various ICS configurations and incoming flow Reynolds numbers. Each step of the prediction model and the final computational results underwent rigorous evaluation through unsteady numerical simulations. Remarkably, the prediction results demonstrated exceptional alignment with simulation outcomes, surpassing the accuracy of previous prediction methods.

A review and reassessment of the aerodynamics of cricket ball swing

This paper reviews 70 years of cricket ball swing literature and reassesses the results in light of new measurements. A comparison of ball tracking data with published experimental results shows that current understanding does not explain the behaviour observed in professional cricket matches. Descriptions of cricket ball boundary layer aerodynamics are updated with new results which show that the seam acts more like a series of vortex generators than a trip and that there is a laminar separation bubble on the seam side of the ball. Previous results for reverse swing are consolidated and compared with new data, showing that different magnitudes of swing occur depending on the condition of both sides of the ball. Variation in pressure and temperature should be included when the Reynolds number, a non-dimensional ball speed, is calculated, while humidity can be neglected. Studies on the effect of wind speed and direction are summarised and results considering the effect of free-stream turbulence are compared with new measurements. Throughout the paper, recommendations for future work are suggested. These include quantitative study of the effect of surface defects and roughness, an assessment of whether atmospheric turbulence can affect swing and investigation into the effect of backspin on swing.

NN-augmented k-ω shear stress transport turbulence model for high-speed flows with shock-wave/boundary layer interaction

Menter k-ω shear stress transport (SST) turbulence model demonstrates excellent performance for incompressible, subsonic and transonic flows with mild separation but shows overprediction of the separation bubble of supersonic shock-wave/boundary layer interaction (SWBLI). Some efforts focus on the effect of the structure parameter in stress limiter in an ad-hoc way. Few studies attempt to construct the relation between the structure parameter and flow field variables. The motivation of this work is to construct such a relation to augment the prediction performance of the SST model by introducing a correction factor. Machine learning methods are used since the physical mechanism of SWBLI is complex and unclear. The simulation results show that the constructed relation enhances the structure parameter near the shock wave in the boundary layer when applied to the SST model. Compared with direct numerical simulation and experimental data, the pressure and skin friction coefficients along the wall and the velocity field are significantly improved. In addition, the introduced correction factor can automatically degrade for the subsonic benchmark case of NACA4412 airfoil and maintain the prediction accuracy of the original SST model, but delay the shock location of the transonic case.

Couette-Poiseuille flow over a backward-facing step: Investigating hydrothermal performance and irreversibility analysis

The present study investigates the hydrothermal performance of Couette-Poiseuille flow over a backward-facing step. With a focus on the irreversibility analyses, the study sheds light on the flow behavior and explores the impact of key parameters, including Reynolds number, and wall terminal velocities. The study investigates a range of laminar Reynolds numbers between 100 and 800, as well as wall terminal velocities, Ut between −3 and 3. Notably, the results reveal a unique flow structure appears at negative Ut, where multiple separation zones (bounces) form downstream of the step. The number of bounces depends on both Re and Ut. An enhancement in the Nusselt number can be observed as the value of -Ut increases, which can be directly correlated to the increase in the number of bounces and the decrease in the reattachment length. On the contrary, the increase in the +Ut results in a decrease in heat transfer, driven by the increase in the reattachment length. The thermal performance factor, TPF shows an improvement for Re of 200 and 400 at Ut of −0.5, while TPF worsens as the value of -Ut increases due to the pressure drop induced by separation bubbles. Thermal entropy generation was identified as the main source of irreversibility processes for all cases.

An Unsteady Reynolds–Averaged Navier–Stokes–Large Eddy Simulation Study of Propeller–Airframe Interaction in Distributed Electric Propulsion

A conceptual framework is presented to determine the improvement in the aerodynamic performance of a canard aircraft fitted with distributed propellers along its main wing. A preliminary study is described with four airframe–propeller configurations predominantly studied in academic and commercial designs. The leading edge–based tractors and trailing edge–based pushers are identified as configurations of interest for the main study. Subsequently, a Navier–Stokes solver is used to simulate the flow using two numerical approaches–a modified steady-state actuator disk and an unsteady rotating propeller profile. Moving meshes with rotating sub-domains are used with a hybrid RANS-LES-based turbulence model while the actuator disks are modified to include viscous swirl effects. The preliminary study shows a local minimum in the change in CL and CD at 10∘ for the pusher and tractor configurations. The main study then demonstrates the outperformance of the pushers over tractors quantified using CL and CL/CD. There is a clear preference for the pushers as they increase the lifting capacity of the aircraft without disproportionately increasing the drag due to the flow smoothening by the suction of the pusher propellers over the main wing. The pushers also delay the separation of the boundary layer whereas the tractors are unable to prevent the formation of the separation bubble despite injecting momentum through their slipstreams into the flow. The results from the two numerical approaches are then compared for accuracy in designing DEP configurations for an airframe.

Mean Reattachment Length of Roof Separation Bubbles Using Proper Orthogonal Decomposition

Investigating flow separation regions on the surfaces of three-dimensional bluff bodies in turbulent flows is important because these regions can induce significant aerodynamic loads. Separation bubbles can generate extreme pressures, making the roof components of low-rise buildings vulnerable. In this study, proper orthogonal decomposition (POD) was applied to wind-induced roof pressures to elucidate the physical significance of the dominant modes. Based on the interpretation of the first mode from the POD, the mean reattachment length of the roof separation bubbles on a low-rise building model in turbulent flow was determined. The mean reattachment length derived from the POD was then compared with the length obtained from an aerodynamic database. For the centerline of the roof, the mean reattachment length based on the POD aligned well with that from the aerodynamic database, showing a difference of less than 5%. This study highlights the efficacy of POD as a powerful tool for estimating the reattachment length of separation bubbles on bluff bodies.

Nonlinear shock-induced flutter of a compliant panel using a fully coupled fluid-thermal-structure interaction model

A comprehensive study has been conducted on the fully coupled aero-thermo-elastic interaction between a shock wave and a laminar boundary layer over a flexible surface with different thermal conditions. In this investigation, the flow is characterized by an oblique impinging shock on a laminar boundary layer over a compliant panel with a cavity pressure and temperature. To model the fluid–structure interaction (FSI), we have combined a finite element model of the structure with a high-order sharp interface immersed boundary method implemented in a finite-difference flow solver. We have analyzed the behavior of the FSI coupling over a wide range of parameters and various boundary conditions while the panel undergoes relaxation to its unexcited state or exhibits limit cycle oscillations. Notably, the panel oscillation and the fluctuations in the shear layer exhibit strong coupling and demonstrate a lock-in resonance frequency response. The colder wall conditions result in a reduced separation bubble and a raised aerodynamic pressure difference between the maximum and minimum instantaneous dynamic pressures, consequently amplifying the amplitude of panel oscillation. For the cases with thermal coupling, the behavior of panel oscillation is notably influenced by specific heat and thermal expansion coefficients. In these scenarios, the thermal state of the solid material emerges as the pivotal factor governing the sustained oscillation of the panel. The non-linear impact of the thermal dependency of the structural properties plays a crucial role in determining the stability of the coupled FSI system. The study revealed that the stable decaying sinusoidal oscillation undergoes a transformation into a limit-cycle oscillation when panel stiffness is affected by reduced temperature. During the sustained oscillation, a mode switch is observed between the first and second modes of deformation of the panel.

Transient Flow Evolution of a Hypersonic Inlet/Isolator with Incoming Windshear

In this paper, a novel flow perturbation model meant to investigate the effects of incoming wind shear on a hypersonic inlet/isolator is presented. This research focuses on the transient shock/boundary layer interaction and shock train flow evolution in a hypersonic inlet/isolator with an on-design Mach number of 6.0 under incoming wind shear at high altitudes, precisely at an altitude of 30 km with a magnitude speed of 80 m/s. Despite the low intensity of wind shear at high altitudes, the results reveal that wind shear significantly disrupts the inlet/isolator flowfield, affecting the shock wave/boundary layer interaction in the unthrottled state, which drives the separation bubble at the throat to move downstream and then upstream. Moreover, the flowfield behaves as a hysteresis phenomenon under the effect of wind shear, and the total pressure recovery coefficients at the throat and exit of the inlet/isolator increase by approximately 10% to 12%. Furthermore, this research focuses on investigating the impact of wind shear on the behavior of the shock train. Once the inlet/isolator is in a throttled state, wind shear severely impacts the motion of the shock train. When the downstream backpressure is 135 times the incoming pressure (p0), the shock train first moves upstream and gradually couples with a cowl shock wave/boundary layer interaction, resulting in a more significant separation at the throat, and then moves downstream and decouples from the separation bubble at the throat. However, if the downstream backpressure increases to 140 p0, the shock train enlarges the separation bubble, forcing the inlet/isolator to fall into the unstart state, and it cannot be restarted. These findings emphasize the need to consider wind shear effects in the design and operation of hypersonic inlet/isolator.

Control of shock–boundary layer interaction in a transonic compressor cascade via heat transfer at a low Reynolds number

The shock–boundary layer interaction mechanism of a transonic compressor profile at different Reynolds numbers (Re) was investigated by a three-dimensional Reynolds averaged Navier–Stokes simulation. The reason behind the deterioration of the transonic blade performance at low Re was revealed. Based on this, the mechanism of wall heat transfer regulating the interaction between the shock and the boundary layer on the blade surface was investigated to clarify the impact of wall heat transfer on transonic blade performance. The results show that as the Re decreases, the blade performance deteriorates, and a critical Re exists. When the Re exceeds this critical value, the separation structure on the suction side becomes the dominant factor contributing to losses. At this point, wall heat transfer reduces the gas kinematic viscosity near the wall, weakening the stabilizing effect of turbulence dissipation on the separated shear layer. This accelerates the transition process, consequently reducing the size of the separation bubble and decreasing blade losses. Below the critical Re, the passage normal shock approaches the exit of the blade passage and increases in strength, causing severe separation on the pressure side and resulting in a sharp decline in blade performance. At this stage, wall heat transfer can restore the blade’s boosting capability and weaken the passage normal shock. Although this may deteriorate the performance on the suction side to some extent, the degree of separation on the pressure side is significantly reduced, causing the wake region to narrow and improving blade performance. Comparatively, the control effect is more significant at a low Re (Re = 1.5 × 105), with losses reduced by 40.6 % compared to adiabatic wall conditions.