Large-scale Flow Separation Research Articles

Experimental Research on Pressure Pulsation and Flow Structures of the Low Specific Speed Centrifugal Pump

The low specific speed centrifugal pump plays a crucial role in industrial applications, and ensuring its efficient and stable operation is extremely important for the safety of the whole system. The pump must operate with an extremely high head, an extremely low flow rate, and a very fast speed. The internal flow structure is complex and there is a strong interaction between dynamic and static components; consequently, the hydraulic excitation force produced becomes a significant factor that triggers abnormal vibrations in the pump. Therefore, this study focuses on a low specific speed centrifugal pump and uses a single-stage model pump to conduct PIV and pressure pulsation tests. The findings reveal that the PIV tests successfully captured the typical jet-wake structure at the outlet of the impeller, as well as the flow separation structure at the leading edge of the guide vanes and the suction surface. On the left side of the discharge pipe, large-scale flow separation and reverse flow happen as a result of the flow-through effect, producing a strong vortex zone. The flow field on the left side of the pressure chamber is relatively uniform, and the low-speed region on the suction surface of the guide vanes is reduced due to the reverse flow. The results of the pressure pulsation test showed that the energy of pressure pulsation in the flow passage of the guide vane occurs at the fBPF and its harmonics, and the interaction between the rotor and stator is significant. Under the same operating condition, the RMS value distribution and amplitude at fBPF of each measurement point are asymmetric in the circumferential direction. The amplitude of fBPF near the discharge pipe is lower, while the RMS value is higher. A complex flow structure is shown by the larger amplitude and RMS value of the fBPF on the left side of the pressure chamber. With the flow rate increasing, the energy at fBPF of each measurement point increases first and then decreases, while the RMS value decreases, indicating a more uniform flow field inside the pump.

Investigation into the Vortex Evolution characteristics of the Francis Turbine During the Runaway Process

Abstract The runaway process is a dynamic transition process that gradually deviates from the optimal operating condition. The flow passage components dissipate all of the energy corresponding to the hydraulic turbine’s head, causing the internal flow of the hydraulic turbine to gradually deteriorate and produce a large-scale vortex structure during the runaway process. A model Francis turbine is used as the research object in this paper to understand the evolution characteristics of an unsteady vortex during a runaway. The runaway process is investigated using high-precision simulation technology of gas-liquid two-phase flow, and the runaway speed is obtained, consistent with the experimental results. The results show that the cavitation volume increases as the rotating speed increases. The large incidence angle between the inlet flow and the blade causes large-scale flow separation in the runner, which leads to forming of a sheet vortex band near the hub of the runner hub and a columnar vortex in the blade passage of the runner blade suction surface. As the runaway process continues, the flow separation in the runner intensifies and the induced vortex size increases, obstructing the runner’s passage. In addition, pressure signal analysis reveals that the pressure fluctuation caused by the sheet vortex band is low-frequency. On the other hand, the pressure fluctuation caused by the columnar vortex is frequent. These two types of vortex structures with high-energy pressure fields degrade the flow in the runner and reduce the hydraulic performance of the turbine.

Experimental study on multi-angle pulsed jet actuators for controlling corner separation flow in a compressor cascade

Three-dimensional corner separation caused by the aerodynamic design of high-load compressor cascades is the main factor that limits their performance enhancement. Because of the unsteady characteristics of large-scale flow separation, injecting unsteady jets into the flow field has become a new flow control method. In this study, the essential flow characteristics of a compressor cascade were obtained through numerical simulations. Subsequently, the effect of pulsed jet angle on the aerodynamic performance of the cascade was experimentally investigated under different operating conditions using pulsed jet actuators. The results showed that at the design incidence angle of 0° and near-stall incidence angle of +4°, the starting point of the corner separation is located at 21 % and 9 % of the axial blade chord, respectively. The pulsation frequency downstream of the cascade trailing edge is 1454.5 Hz. Under the design conditions, the 20° pulsed jet exhibits the best control effect, reducing the total pressure loss coefficient by 13.4 % compared with that of the original cascade. Moreover, pulsed jets demonstrate strong applicability under variable operating conditions. At near-stall incidence angles, when the jet angle is 30°, the loss-reduction ability of the pulsed jet is maximized, decreasing the loss by 25.2 %. Pulsed jets reduce the cascade loss primarily by suppressing the secondary flow at the endwall and discretizing the vortex structures in the channel.

Low-Speed Space Launch System Computational Fluid Dynamics: A Comprehensive Overview

In this review paper, low-speed computational work from NASA Langley Research Center in support of the Space Launch System (SLS) is discussed. This information includes both historical and present efforts with the Kestrel computational fluid dynamics solver. The low-speed aerodynamics of SLS is highly complex, and the analysis of the unsteady flowfield requires significant computational efforts. The SLS mission profile varies from the vehicle static on the launch pad through high-speed ascent. This study focuses on the prelaunch as well as liftoff and transition portions of the flight both in proximity to the launch tower and in isolation. High-alpha conditions, as large as 90 deg, result in a flowfield dominated by massive, large-scale flow separation and asymmetric vortices. High-fidelity solutions require an unsteady computational formulation to accurately capture the aerodynamics of the vehicle. A detailed discussion of the computational approach is presented, followed by key efforts to support the program, both historic and present, including information that has been either published previously or that has never before been published external to NASA.

Aeroacoustic simulation of bluff bodies with protrusions at moderate Reynolds number

This paper presents an evaluation of passive control methods that employ surface protrusions to mitigate the aerodynamic sound generated from a cylinder wake flow. Building on previous designs optimized for low Reynolds numbers (Re = 150) through adjoint-based aeroacoustic shape optimization, this study investigated the performance under a moderate Reynolds number (Re = 67 000) condition typical of mechanical engineering applications using aeroacoustic simulations based on the lattice Boltzmann method. Three configurations of surface protrusions were tested, all of which were found to significantly reduce the mean drag by at least 45% compared with that of an unmodified circular cylinder. Designs featuring rear protrusions outperformed the conventional splitter plate in terms of the sound reduction performance, with symmetrical protrusions on both the front and rear surfaces achieving a tonal sound reduction of 13 dB. However, a specific protrusion design increased the low-frequency sound owing to the intensified large-scale flow separation. These findings highlight the effectiveness of rear protrusions in suppressing wake oscillations and dipole sound generation in the subcritical Reynolds number range. Moreover, the study revealed the need to tailor the front protrusion shape to the Reynolds number for performance optimization.

Numerical study on the periodic control of supersonic compression corner flow using a nanosecond pulsed plasma actuator

This study investigates the effectiveness of a pulsed nanosecond dielectric barrier discharge (NSDBD) plasma actuator for flow control over a supersonic compression corner through numerical simulations. The effects of varying applied voltages, repetitive frequencies, and activated locations of the plasma actuator are examined under large-scale laminar flow separation conditions around a compression corner. The unit Reynolds number and Mach number are 7.8 × 106 m−1 and 4, respectively. The results indicate that the discharge induces a pressure rise and leads to misalignment between the pressure gradient and density gradient in the residual heat region. Because of the interaction between the supersonic freestream and the actuation-induced shock/compression flow, convection, compressibility of the fluid element, and baroclinicity of the residual heat region collectively lead to the formation of an induced spanwise vortex, which in turn enables momentum migration. The induced vortex disrupts the initial flow structures and entrains high-energy fluid from the main flow into the boundary layer, promoting momentum mixing between the main flow and the separated flow, which increases the energy of the boundary layer to resist the adverse pressure gradient. The time-averaged flow structures imply that it is possible to totally eliminate the flow separation near the supersonic compression corner. For aerodynamics on the surface, the normal force produces a pitching moment that can be potentially utilized to control the body's orientation and trajectory. Additionally, total drag on the surface can be reduced by 5 %. This suggests that choosing the most appropriate position based on its local fluid characteristics can strongly increase the control effectiveness.

Experimental investigation on the turbulent wake flow in fully established transonic buffet conditions

The transonic flight regime is often dominated by transonic buffet, a highly unsteady and complex shock-wave/boundary-layer interaction involving major parts of the flow field. The phenomenon is associated with a large-amplitude periodic motion of the compression shock coupled with large-scale flow separation and intermittent re-attachment. Due to the resulting large-scale variation of the global flow topology, also the turbulent wake of the airfoil or wing is severely affected, and so are any aerodynamic devices downstream on which the wake impinges. To analyze and understand the turbulent structures and dynamics of the wake, we performed a comprehensive experimental study of the near wake of the supercritical OAT15A airfoil in transonic buffet conditions at a chord Reynolds number of 2×106\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$2\ imes 10^{6}$$\\end{document}. Velocity field measurements reveal severe global influences of the buffet mode on both the surface-bound flow field on the suction side of the airfoil and the wake. The flow is intermittently strongly separated, with a significant momentum deficit that extends far into the wake. The buffet motion induces severe disturbances and variations of the turbulent flow, as shown on the basis of phase-averaged turbulent quantities in terms of Reynolds shear stress and RMS-values. The spectral nature of downstream-convecting fluctuations and turbulent structures are analyzed using high-speed focusing schlieren sequences. Analyses of the power spectral density pertaining to the vortex shedding in the direct vicinity of the trailing edge indicate dominant frequencies one order of magnitude higher than those associated with shock buffet (Stc=O(1)\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$St_c=\\mathcal {O}({1})$$\\end{document}) vs. Stc=O(0.1)\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$St_c=\\mathcal {O}({0.1})$$\\end{document}). It is shown that the flapping motion of the shear layer is accompanied by the formation of a von Kármán-type vortex street of fluctuating strength. These wake structures and dynamics will impact any downstream aerodynamic devices affected by the wake. Our study, therefore, allows conclusions regarding the incoming flow of devices such as the tail plane.

Flow Separation Control of Nacelle Inlets in Crosswinds by Dielectric Barrier Discharge Plasma Actuation

Crosswinds will lead to large-scale flow separation in the nacelle inlets, which seriously affects the flight safety of the aircraft; there is an urgent need to develop flow control measures. As a plasma flow control method, the application of surface dielectric barrier discharge in the field of nacelle inlet separation control is of great significance for improving the intake quality. Based on the characteristic law of the baseline flow field, the flow control effect of the nacelle inlet separation flow field experiments with NS-DBD, and the influence of the actuation frequency on the flow control is discussed. A comparative experimental study of NS-DBD and AC-DBD is carried out. Finally, the flow control mechanisms for both are discussed. The results show that under the condition that the flow velocity of the wind tunnel is 35 m/s and the crosswind angle is 10°, the average total pressure loss coefficient and distortion index decrease by 29.62% and 44.14% by NS-DBD actuation. At the same time, exists an inherent optimal coupling frequency in NS-DBD, and the control effect of NS-DBD is better than that of AC-DBD. NS-DBD mainly through shock waves and induced vortices, while AC-DBD mainly through the induced generation of near-wall jets to reduce the inverse pressure gradient and improve nacelle flow separation.

Approach for Aerodynamic Gust Load Alleviation by Means of Spanwise-Segmented Flaps

Active gust load alleviation techniques exhibit a high potential in significantly reducing the transient gust loads on aircraft. In this work the aerodynamic potential of trailing-edge flaps and leading-edge flaps is numerically studied with the purpose to significantly reduce the structural gust loads. The utilized spanwise-segmented flaps represent slight modifications of existing devices for high-lift and maneuvering. The investigations based on unsteady Reynolds-averaged Navier–Stokes computations are conducted by employing a generic wing–fuselage aircraft configuration at transonic flow conditions. Idealized discrete “”-type vertical gusts that are relevant for the certification process are used as representative atmospheric disturbances. The focus of this paper is to introduce a practicable prediction method for required trailing- and leading-edge flap deflections for a significant mitigation of gust-induced wing loads. The three-dimensional flap deflections are determined by parametric two-dimensional simulations at representative wing sections. Different extensions of the estimation approach are investigated to assess the influence of the wing planform, the finite wing span, the aerodynamic phase lags, and the flap scheduling. It is shown that the trailing- and leading-edge flaps are promising in terms of alleviation of gust-induced wing bending and wing torsional moments, respectively. However, at high leading-edge flap deflections that are necessary for a full compensation of the wing torsional moment large-scale flow separation is identified. The introduced gust load alleviation approach indicates a good transferability between two-dimensional airfoil and three-dimensional wing aerodynamics for unsteady flap deflections.

Characteristics of internal flow of nozzle integrated with aircraft under transonic flow

Abstract To reveal how aircraft affects the internal flow of the ejector nozzle, we have constructed three model types in this article. These include the model of SR-71 aircraft, the model that only contains ejector nozzle with third auxiliary valve, and the model that integrates the previous two. The results showed that in the transonic regime (M a = 1.2), the third auxiliary flow mainly stems from the boundary layer of the aircraft body. Indeed, a large-scale flow separation phenomenon near the third auxiliary door may require a more nuanced description. The mainstream flow is always in an overexpansion state and results in a Mach plate structure at the exit of the nozzle. However, after integration, the rates of the third auxiliary and the secondary flow are reduced by 18.15% and 5.26%, respectively. Meanwhile, the mainstream flow demonstrates higher overexpansion levels, the position of the Mach plate further downstream changes, and the thrust coefficient decreases by 1.75%. It is worthwhile noting that a strong pressure gradient occurs in the circumferential direction near the connecting structure, which induces lateral flow. This lateral flow breaks away from the wall under the reverse pressure gradient of the nozzle, thus forming three vortex pairs.

Experimental Investigation of the Vortex Dynamics in Circular Jet Impinging on Rotating Disk

A circular jet impinging perpendicularly onto a rotating disk is studied in order to understand the influence of centrifugal forces on the radial wall jet. Time-resolved Particle Image Velocimetry (TR-PIV) measurements are conducted in different jet regions in order to investigate the flow physics of the large-scale vortical structures and the boundary layer development on the impinging wall for both stationary and rotating impinging disks. The Reynolds number is ReD = 2480, the orifice-to-plate distance H = 4D (D is the jet-orifice diameter) and the rotation rate is 200 RPM. It is found that the rotation of the impinging wall results in strong centrifugal effects, which affect different regions of the jet. Both radial velocity profiles and turbulence intensity distributions show different behavior when comparing the stationary and rotating cases. Finite Time Lyapunov Exponent (FTLE) analysis is implemented to describe the time-resolved behavior of the large-scale vortical structures and flow separation.

Experimental investigation of expansion effect on shock wave boundary layer interaction near a compression ramp

The experiment is conducted to investigate the effect of expansion on the shock wave boundary layer interaction near a compression ramp. The small-angle expansion with an angle degree of 5° occurs at different positions in front of the compression ramp. The particle image velocimetry and flow visualization technology show the flow structures, velocity field, and velocity fluctuation near the compression ramp. The mean pressure distribution, pressure fluctuation, and power spectral density are measured by high-frequency response pressure transducers. The experimental results indicate that the expansion before the compression ramp position affects the shock wave boundary layer interaction to induce a large-scale separation. But the velocity fluctuation and pressure fluctuation are attenuated near the large-scale flow separation region. When the expansion occurs closer to the compression ramp, the expansion has a more significant impact on the flow. The fluctuation of velocity and pressure is significantly attenuated, and the wall pressure rise of the separation point is reduced obviously. And the characteristic low-frequency spectrum signal related to the unsteadiness of the shock wave boundary layer interaction is significantly suppressed. In addition, variation of the separation region scale at different compression angle degrees is distinctive with the effect of expansion.

Supersonic compressor cascade flow control using plasma actuation at low Reynolds number

To control the supersonic compressor cascade flow at low Reynolds numbers, this paper describes the use of nanosecond dielectric barrier discharge (NS-DBD) plasma actuation for flow control, and presents the results of large-eddy simulations conducted to investigate the corresponding flow control effects. NS-DBD plasma actuation on both the blade pressure surface and suction surface induces a distorted flow structure (DFS) within the blade passage. In the case of NS-DBD plasma actuation on the blade pressure surface, the influence of the DFS on the flow is suppressed by a shock wave. Even so, the DFS can still trigger the instability in the shear layer between the separated flow and the mainstream flow. Shock-wave-induced large-scale flow separation on the blade pressure surface is then suppressed, and the overall total pressure loss of the blade passage is reduced by 7.4%, despite the increased shock wave loss from the reduced flow blockage within the blade passage. In the case of NS-DBD plasma actuation on the blade suction surface, the DFS is less effective in suppressing the shock-wave-induced small-scale flow separation on the blade suction surface. However, the DFS on the blade suction surface enhances the shock wave oscillations within the blade passage, and this suppresses the flow separation on the blade pressure surface.

The dynamics of flow-induced flutter of a thin flexible sheet

The dynamics of the flow-induced flutter of a thin flexible sheet attached to a streamlined support was experimentally studied in a low-speed wind tunnel. In this study, both the structural dynamics and the fluid dynamics aspects of flutter were considered. The kinematics of the oscillating sheet was investigated using high-speed imaging and the flowfield was examined using hotwire anemometry and particle image velocimetry (PIV). The small-scale perturbation in the flow over the sheet was found to induce a low-amplitude vibration, which changed to a large-amplitude flutter as the wind speed was increased to a critical value. The initiation of flutter occurs with the second mode limit cycle oscillation (LCO), bypassing the first mode, and changes to third mode LCO at a higher wind speed. Based on the behavior of the sheet, five different regimes are identified and discussed in this paper. The natural frequencies of the sheet were found to have a significant role in the initiation of the LCO and its transition to the higher modes. The PIV results show a highly accelerated flow over the curved surface of the oscillating sheet, which induces a lift force that acts as a driving force. The accelerated flow over the sheet separates at its tail and forms a large-scale undulating wake. In the LCO regimes, any large-scale flow separation over the sheet could not be observed and the flow appears to be attached even at high deflection of the sheet.

Sound generation by entropy perturbations passing through a sudden flow expansion

Abstract

High-Fidelity Finite Element Modeling and Analysis of Adaptive Gas Turbine Stator-Rotor Flow Interaction at Off-Design Conditions

ABSTRACTThe objective of this work is to computationally investigate the impact of an incidence-tolerant rotor blade concept on gas turbine engine performance under off-design conditions. When a gas turbine operates at an off-design condition such as hover flight or takeoff, large-scale flow separation can occur around turbine blades, which causes performance degradation, excessive noise, and critical loss of operability. To alleviate this shortcoming, a novel concept which articulates the rotating turbine blades simultaneous with the stator vanes is explored. We use a finite-element-based moving-domain computational fluid dynamics (CFD) framework to model a single high-pressure turbine stage. The rotor speeds investigated range from 100% down to 50% of the designed condition of 44,700 rpm. This study explores the limits of rotor blade articulation angles and determines the maximal performance benefits in terms of turbine output power and adiabatic efficiency. The results show articulating rotor blades can achieve an efficiency gain of 10% at off-design conditions thereby providing critical leap-ahead design capabilities for the U.S. Army Future Vertical Lift (FVL) program.

Diagnosis of internal energy characteristics of mixed-flow pump within stall region based on entropy production analysis model

In order to investigate the influence of tip clearance size on the characteristics of mixed-flow pump under stall conditions, the entropy production analysis method is utilized to visualize the high energy loss area of main flow domains in internal flow fields. The results shown that the tip clearance size will not only affect the internal flow fields of mixed-flow pump under designed point but also play a remarkable role in the hydraulic loss with stall region. As the tip clearance increases from 0.2 mm to 1.1 mm, the head difference between critical stall condition and deep stall condition is about 0.79 m, 0.58 m, 0.52 m and 0.35 m, indicating that the intensity of tip leakage flow (TLF) is not linear with tip clearance size. The inflow instability at pump inlet will increase as the growth of tip clearance size, and it comes from the inlet swirling flow induced by the return of TLF. The high loss in impeller is caused by the TLF and stall vortex under stall conditions while the large-scale flow separation vortex and back flow from the annular volute are responsible for the high loss area in guide vane domain. The results could be referred to optimize the hydraulic structure of mixed-flow pump.

Numerical Investigation of a Two-Element Wingsail for Ship Auxiliary Propulsion

The rigid wingsail is a new type of propulsion equipment which greatly improves the performance of the sailboat under the conditions of upwind and downwind. However, such sail-assisted devices are not common in large ships because the multi-element wingsail is sensitive to changes in upstream flow, making them difficult to operate. This problem shows the need for aerodynamic study of wingsails. A model of two-element wingsail is established and simulated by the steady and unsteady RANS approach with the k-ω SST turbulence model and compared with the known experimental data to ensure the accuracy of the numerical simulation. Then, some key design and structural parameters (camber, the rotating axis position of the flap, angle of attack, flap thickness) are used to characterize the aerodynamic characteristics of the wingsail. The results show that the position of the rotating shaft of the flap has little influence on the lift coefficient at low camber. When stall occurs, the lift coefficient first increases and then decreases as the flap axis moves backward, which also delays the stall angle at a low camber. At the high camber of AOA = 6°, the lift coefficient always increases with the increase of the rotating axis position of the flap; especially between 85% and 95%, the lift coefficient increases suddenly, which is caused by the disappearance of large-scale flow separation on the suction surface of the flap. It reflects the nonlinear coupling effect between camber of wingsail and the rotating axis position of the flap

An experimental study on dynamic ice accretion process over the surfaces of rotating aero-engine spinners

An experimental study was conducted to investigate the dynamic ice accretion process over the surfaces of rotating aero-engine spinners and to examine the detriment effects of the ice accretion on the airflow to be inhaled by aero-engines. Three scaled spinner-fan models with different spinner shapes (i.e., conical-shaped, coniptical-shaped, and elliptical-shaped spinners) were manufactured and exposed under typical rime and glaze icing conditions for a comparative study. During the experiments, while a high-speed imaging system was used to record the dynamic ice accretion process over the rotating spinner models, a high-resolution particle image velocimetry (PIV) system was utilized to examine the trajectories of super-cooled water droplets as they approach to the surfaces of the spinner models. It was found that, under typical rime conditions, while accreted ice layers were found to conform well with the original shapes of the spinner models in general, the total amount of the ice mass accreted over the spinner surfaces were found to be a strong dependent on the spinner shapes. While the conical-shaped spinner was found to have the largest amount of ice accretion (i.e., 60–80% more than those over the other two spinner models) over almost entire spinner surface, ice accretion was found to take place mainly in the front portion of the coniptical-shaped and elliptical-shaped spinners. Under the glaze icing condition, in addition to forming ice layers over the spinner surfaces, very complicated, needle-shaped icicles were also found to grow rapidly out from the spinner surfaces and extrude into the incoming airflow, due to the effects of the centrifugal forces associated with the rotation motion. The complex glaze ice structures accreted over the spinner surface were found to induce significant disturbances/distortions and even cause large-scale flow separations for the airflow near the iced spinner surfaces, which would significantly degrade the quality of the inlet airflow to be inhaled by aero-engines, thereby, adversely affecting the performance of aero-engines.

Numerical investigation on flow control effects of dynamic hump for turbine cascade at different Reynolds number and hump oscillating frequency

From takeoff to cruise, the operating Reynolds number (Re) of the low-pressure turbine (LPT) in an aero-engine decreases significantly, making the boundary layer near the aft portion of the blade surface susceptible to separation. This paper discusses the feasibility of using local dynamic surfaces as an active flow control method to suppress the laminar flow separation on the suction surface in PAKB LPT cascade at low Re. A local dynamic surface, shaped as a half-sinusoidal hump with an oscillation amplitude of 1 mm, was positioned just upstream of the peak velocity point. Unsteady Reynolds-Averaged Navier–Stokes (RANS) simulations were performed to investigate the aerodynamic performance of the cascades with and without dynamic hump. At first, the effect of Re was studied with the hump oscillating frequency at 200 Hz. Compared with uncontrolled cascade at various Re from 25,000 to 150,000, the profile losses of the cascade controlled by dynamic hump reduced significantly at low Re condition where the separation bubble was very large, and increased slightly at high Re condition where the separation bubble was small. Then the effect of the hump oscillation frequency was investigated under the Re of 25,000. It was found that the controlled cascade loss reached a minimum in a certain hump oscillating frequency range. At low Re of 25,000, the loss of controlled cascade was mainly composed of low-loss continuous laminar vortices that were attached to the suction surface. Continuous vortices with a certain spacing confined the low-energy fluid near the suction surface by energy exchange between the free flow and near-wall flow, which suppressed the large-scale flow separation.