Suction Surface Boundary Layer Research Articles

Study of the influence of multiple factors on the boundary layer of a high-lift LPT with the RBF-GA method

In a high-altitude cruising state, boundary layer separation exists in high-lift low-pressure turbines, and inflow conditions corresponding to different blade designs can directly affect the working efficiency of low-pressure turbines. In particular, the reduced frequency of wake and free-stream turbulence intensity in an inlet flow can greatly influence boundary layer separation and transition development. In this paper, the influence of different inflow turbulence intensities and reduced wake frequencies on the development of suction surface boundary layers in high-lift low-pressure turbines under the influence of upstream wakes is studied by numerical simulations and experiments. Due to the combination of inflow free-stream turbulence intensity and reduced wake frequency, many inflow conditions can be chosen in the design process, and the unsteady influence of upstream wakes complicates the boundary layer flow. In this paper, an RBF (radial basis function)-GA (genetic algorithm) machine learning method is used to explore the optimal inlet conditions corresponding to the minimum profile loss of the Pak-B profile. The search region of the free-stream turbulence intensity is 2%–4%, and the reduced frequency of the wake is changed by changing the flow coefficient, whose variation range is 0.7–1.3. It is found that the RBF-GA machine learning method can attain an inflow condition with a lower profile loss while using the same amount of computation and effort.

Study on the Design Method of a New Type of Compressor Slotted Blade

the compressor has been developing towards the direction of high pressure ratio, high efficiency and wide stability margin. The flow in the air compressor blade channel presents a strong inverse pressure gradient. In order to reduce losses and improve the stability of the pressure aircraft, this paper adopts a slot structure from the leading edge of the blade to the suction surface. Air is injected from the front edge of the blade and discharged from the suction surface. The automatic optimization method is used to design the slot. The results show that the blade profile with inlet Mach number 0.4 and the diffusion factor 0.6, the total pressure loss of slotted blade profile is lower than that of unslotted blade profile, and the airflow Angle is greater than that of unslotted blade profile in the whole range of low loss incident Angle. In the incidence of 0 degree, due to the control effect of slotted jet on the boundary layer of the blade suction surface, the boundary layer separation zones on the blade suction surface becomes smaller and the wake also decreases, so the total pressure loss of the slotted blade profile decreases by 37%. When the incidence angle is 4°, the separation area of the suction surface boundary layer is larger, and the slotted jet has a better control effect on the boundary layer of the blade suction surface. However, When the incidence angle is -6°, boundary layer separation zones on the suction surface becomes smaller, and the boundary layer separation zones appears on the pressure surface, so the slotting effect is poor.

Numerical Study of Benchmark Experiments of Supersonic Compressor Cascades

Benchmark cases of turbulent supersonic flow through compressor cascades were computed using a modern upwind scheme (Roe with MUSCL pre-processing), with the k-ω turbulence model, and the classical MacCormack method. The test cases derive from experiments through ARL SL 19 blade cascades at Detroit Diesel Allison (DDA) at the design Mach number, and at an off-design condition from ONERA. A verification of the codes was obtained by computing an inviscid flow through a wedge cascade at similar conditions. Predictions are in very good agreement for the DDA test case for the blade surface pressure distributions. Cleaner solutions were obtained using the Roe scheme. The MacCormack solution exhibits oscillations corresponding to oscillations of the shock-induced separating boundary layer on the pressure surface near the trailing edge. Wake flow Mach number distributions were very well predicted, but not flow angle distributions. Predictions are less accurate for the ONERA test case which was at a slightly higher Mach number and which moves shock locations sufficiently to cause a more sensitive response from the suction surface boundary layer.

Unsteady Effects of Wake on Downstream Rotor at Low Reynolds Numbers

In a compressor, the periodic wake is an inherently unsteady phenomenon that affects the downstream flow conditions and loading distribution. Thus, understanding the physical mechanisms of these unsteady effects is important for eliminating flow losses and improving compressor performance, particularly at low Reynolds numbers. To understand the influence of the upstream wake on the downstream flow field structure, this paper describes numerical simulations of a one-stage high-pressure compressor at altitudes of 10–20 km. The influence of the wake on rotor flow blockage at different Reynolds numbers is analyzed, and the unsteady interaction between the upstream wake and boundary layer or tip leakage flow is discussed. The results indicate that the wake has a beneficial effect on the efficiency of the rotor at high Reynolds numbers, but this weakens and becomes negative as the Reynolds number decreases. The wake can reduce the flow blockage in the mainflow region. Due to the wake, the length of the laminar separation bubble at high Reynolds numbers decreases and that at low Reynolds numbers increases. In addition, the unsteadiness of the wake causes separation bubbles to appear periodically at high Reynolds numbers and induces an open separation bubble at low Reynolds numbers. The Kelvin–Helmholtz instability can dominate the transition process of the boundary layer, which is also affected by the disturbance vortex induced by the wake. Regarding the tip leakage flow, the wake can reduce the flow blockage at high Reynolds numbers but increase the flow blockage at low Reynolds numbers. The interaction at low Reynolds numbers causes a double-leakage flow, which finally leads to the large-scale separation of the suction surface boundary layer. The large-scale separation causes flow blockage in the tip region and prevents the rotor wake from propagating downstream. On the contrary, the unsteady wake can pass through the tip clearance vortex and inhibit the separation of the suction boundary layer at high Reynolds numbers, which is reflected in a larger amplitude of one blade passage frequency. Therefore, the flow loss in the downstream flow field at high Reynolds numbers is significantly reduced at high Reynolds numbers.

Critical endwall blockage attenuation-based automatic optimization of casing treatment design for transonic axial flow compressor

Tip leakage flow (TLF) and endwall blockage formation are closely associated with the inception of rotating stall in tip-critical transonic axial flow compressor (AFC) rotors. The present study proposes a critical endwall blockage attenuation-based optimization method of casing treatment design for transonic AFC. First, a well-defined endwall blockage quantification method is developed to reveal the development of the endwall blockage at near-stall, and two blockages of the maximum endwall blockage and the near trailing edge endwall blockage are identified as the critical values for stall inception. Next, both adiabatic efficiency at design point and critical endwall blockages at near-stall are used as optimization objectives in an adaptive surrogate-based optimization procedure to deal with the design of nonuniform multiple grooves casing treatment (NMGCT) for the NASA Stage 35, thus avoiding the extremely time-consuming simulations of compressor performance curves. The analysis of the representative optimal designs demonstrates that the NMGCT can act on the tip flow field in a variety of ways, e.g. lowering the maximum endwall blockage can suppress the intensity of the TLF-shock wave interaction, or lowering the near trailing edge endwall blockage is able to control both the trajectory of the TLF closer to the suction surface and the thickness of the suction surface boundary layer. These achievements can help improving the stall margin of the tip-critical transonic AFC. However, in the case of AFC similar to Stage 35, reducing the maximum endwall blockage can achieve greater stall margin improvement and smaller peak efficiency penalty.

Numerical investigation of centrifugal compressor stability enhancement using vaned diffuser endwall contouring technology

To improve the stable operating range of centrifugal compressor stage, endwall contouring technology is introduced on the hub-side wall of the vaned diffuser for NASA CC3 centrifugal compressor as a promising technique to redistribute the flow field near the endwall in this paper. The main contents of this paper are listed as follows. First, a specific parameterized endwall contouring guideline is developed based on sinusoidal function and Bézier curve, and three factors are considered, namely, peak height, peak radial position, and frequency coefficient. Then, centrifugal compressor stages with baseline diffuser and contoured endwall diffusers are numerically investigated to reveal the influence of endwall contouring on the stable operating range. Results show that endwall contouring is an effective method to achieve stability enhancement with no evident reduction in stage performance on the design point. CEW13-12 has the most successful combination of endwall contouring parameters, and stall margin is increased by up to 40% for CEW13-12. Finally, the influence mechanism of endwall contouring on the stable operating range is discussed. At the near stall condition, the contouring endwall forces the fluid within the semi vaneless space to deflect toward the suction side to increase the momentum of low-energy fluid in the suction surface boundary layer and reduce the adverse pressure gradient on the suction side, which delay the development of flow separation on the vane leading edge near the shroud side to suppress the onset of rotating stall. At the near choke condition, the contouring endwall has more potential to increase the throat area, resulting in a decrease of throat blockage to improve choke margin. The results show that the endwall contouring technology is a reliable method to achieve stability enhancement of centrifugal compressor stage.

Transonic Relief in Fans and Compressors

Abstract Every supersonic fan or compressor blade row has a streamtube, the “sonic streamtube,” which operates with a blade relative inlet Mach number of one. A key parameter in the design of the “sonic streamtube” is the area ratio between the blade throat area and the upstream passage area, Athroat/Ainlet. In this article, it is shown that one unique value exists for this area ratio. If the area ratio differs, even slightly, from this unique value, then the blade either chokes or has its suction surface boundary layer separated due to a strong shock. Therefore, it is surprising that in practice designers have relatively little problem designing blade sections with an inlet relative Mach number close to unity. This article shows that this occurs due to a physical mechanism known as “transonic relief.” If a designer makes a mistake and designs a blade with a “sonic streamtube,” which has the wrong area ratio, then “transonic relief” will self-adjust the spanwise streamtube height automatically moving it toward the unique optimal area ratio, correcting for the designer’s error. Furthermore, as the blade incidence changes, the spanwise streamtube height self-adjusts, moving the area ratio toward its unique optimal value, effectively controlling the blade’s incidence range. Without “transonic relief,” supersonic and transonic fan and compressor design would be impossible. This article develops a simple model that allows “transonic relief” to be decoupled from other mechanisms and to be systematically studied. The physical mechanism on which it is based is thus determined and a universal relationship for core compressor preliminary blade design is presented. Finally, its implications in relaxing manufacturing tolerances and in the design of future distortion tolerant blades are discussed.

Unsteady fluidic oscillators for active controlling boundary layer separation in an ultra-high-lift low-pressure turbine

This study used large eddy simulation (LES) to investigate the control mechanism of a fluidic oscillator in the suppression of non-reattachment separation of a boundary layer from an ultra-high-lift, aft-loaded, low-pressure turbine under low Reynolds number. The internal flow characteristics of the fluidic oscillator were examined, with a focus on the influence mechanism of the pulse jet from the fluidic oscillator on boundary layer separation and transition. The effects of different jet angles (θ) on the suppression of separation bubbles were also compared in detail. Finally, the mechanism of interaction between the fluidic oscillator jet and the blade suction surface boundary layer was assessed. It was found that the pulse jet of the fluidic oscillator generated a streamwise vortex pair above the surface of the blade suction, which promoted an exchange of momentum between the boundary layer for the blade suction surface and the main flow, and clearly suppressed boundary layer separation. The streamwise vortex pair at θ=30° showed the characteristics of higher strength, smaller size, slower decay, and longer penetration distance than the results at θ=90°, such that the momentum exchange between the low-energy fluid inside the boundary layer and the main flow is sustainable, which promoted the laminar boundary layer to produce bypass transition and significantly improved the aerodynamic performance of the turbine.

Investigation of shock wave control by suction in a supersonic cascade

The numerical simulation of suction control on a supersonic cascade is presented in this paper. The purpose of this study is to enhance the resistance backpressure characteristics of a flow field by suction control and demonstrate the evolution of shock structures and parameter distributions in the flow field controlled by suction. The study is based on the SAV21 supersonic cascade designed by the German Aerospace Centre and operated under an incoming Mach number of 2.4 and a flow turning angle of 45°. The three-dimensional Reynolds-averaged Navier-Stokes (RANS) equations in a Cartesian coordinate system are successfully applied to the cascade flow. The two-equation shear-stress transport (SST) k-ω turbulence model of Menter is employed to model the turbulent velocity profile. The results show that when the backpressure is greater than the critical backpressure, the shock train leading edge crosses the throat, causing a decrease in the mass-capturing coefficient and the occurrence of stall. To control supersonic cascade by suction, the most effective location is at the interaction of the shock train leading edge and the suction surface boundary layer. By applying suction control at the key position, the shock train is stabilized at the downstream boundary of the suction slot with a small loss of suction mass flow. At mild backpressure ratio range, the increased backpressure is compensated by the local enhanced barrier wave. By increasing the pressure gradient downstream of the barrier wave, the maximum backpressure is improved by suction control. By using the improved combined suction scheme with one spanwise slot on the suction surface and one streamwise slot on the endwall, the maximum backpressure can be improved by 20% under a suction mass flow of 8% of the capture mass.

Unsteady flow mechanism of the integrated aggressive inter-turbine duct in low Reynolds number condition

Aggressive inter-turbine duct, which has ultra-high bypass ratio and ultra-short axial length, is widely applied in the modern turbofan engine because it can reduce engine weight and improve low-pressure rotor dynamic characteristics. However, the aggressive inter-turbine duct that has swirling flow, wake, shock, and tip clearance leakage flow of upstream high-pressure turbine, and even has structs in its flow channel, is liable to separate, especially in high-altitude low Reynolds number (Re) condition. In addition, its downstream low-pressure turbine is on the edge of separation too. In this paper, an integrated aggressive inter-turbine duct embedded with wide-chord low-pressure turbine nozzle is adopted to eliminate the aggressive inter-turbine duct's end-wall separation. Since there are many studies on suppressing the blade suction surface's separation by upstream wake, in this study inherent wake is utilized to suppress the boundary layer separation on low-pressure turbine nozzle's suction surface in the integrated aggressive inter-turbine duct. The paper studies the unsteady flow mechanisms of the integrated aggressive inter-turbine duct (especially the separation and transition mechanisms of low-pressure turbine nozzle's suction surface boundary layer) by the computatioinal simulation method.

Effect of Camber on Non-axisymmetric End-wall Contouring Performance in Compressor Linear Cascades

In order to explore non-axisymmetric end-wall’s mechanism of secondary flow control in compressor cascade, based on three-dimensional numerical calculation, this paper analysed effects of camber on non-axisymmetric end-wall contouring performance in compressor linear cascades. Results showed that non-axisymmetric end-wall reduces the corner separation loss mainly by suppressing the interaction between the crossflow and the suction surface boundary layer and decreasing the adverse pressure gradient on suction surface. Meanwhile, non-axisymmetric end-wall partly enhances the cross flow, increasing the secondary flow loss. At the incidence of 0°, the increase of the secondary flow loss caused by the non-axisymmetric end-wall plays the leading role when the camber angle is 20°; with the increase of camber angle, the reduction of corner separation loss becomes the leading role. At the incidence of 7°, the hub-corner stall is removed by the non-axisymmetric end-wall when the camber angle increases to 40°, decreasing the loss remarkably.

Impact of turbulence intensity on the unsteady flow characteristics of the integrated aggressive inter-turbine duct

Aggressive inter-turbine duct (AITD) which has ultra-short axial length and ultra-large area ratio is widely applied in the modern turbofan engine to meet the demand of higher efficiency and better low pressure rotor dynamic characteristics. However, the adverse stream-wise pressure gradient in this duct is enhanced, and there are complicate upstream and downstream flow conditions. The flow field of the AITD and its downstream low pressure turbine is seriously deteriorated and prone to separate, especially in high-altitude cruise state. The integrated AITD which integrates the AITD with strut had been proved to can restrain the traditional AITD's three-dimensional flow separation. Besides, periodic wake can suppress the downstream blade suction surface's flow separation by “calmed region” effect according to former studies. The AITD of this research taken from a real engine utilizes integrated design and inherent upstream wake to suppress the flow separation of the AITD's end wall and low pressure turbine nozzle's suction surface correspondingly. In this article, the computational simulation method is used to quantify the impact of turbulence intensity ( T u) on the integrated AITD's unsteady flow characteristics, especially the separation and transition mechanism of low pressure turbine nozzle's suction surface boundary layer. Research shows that the impact of T u on the integrated AITD's flow property depends on the balance of T u's dual effect.

The Effect of Nonequilibrium Boundary Layers on Compressor Performance

The paper investigates the effect of nonequilibrium behavior of boundary layers on the profile loss of a compressor. The investigation is undertaken using both high fidelity simulations of a midheight section of a compressor blade and a reduced order model, MISES. The solutions are validated using experimental measurements made in the embedded stage of a multistage low speed compressor. The paper shows that up to 35% of the suction surface boundary layer of the compressor blade exhibits nonequilibrium behavior. The size of this region reduces as the Reynolds number is increased. The nonequilibrium behavior was found to reduce profile loss in cases of attached transition and raise loss where transition occurs through separation.

Effect of swirl on the performance and stability of transonic axial compressor

To investigate the effect of twin swirl and bulk swirl on the performance and stability of a transonic axial compressor, a blade swirl generator was designed and simulated with a transonic single rotor using steady and unsteady numerical calculation methods. The bulk swirl intensity was adjusted by replacing the blades with different camber angles. The twin swirl intensity was decreased by reducing the blade number. The counter-rotating bulk swirl generated a significant drop in both the efficiency and stall margin, and resulted in an increase in the choked mass flow, and total pressure ratio. The co-rotating bulk swirl generated a decrease in the mass flow, total pressure ratio and stable operating range. The counter-rotating bulk swirl resulted in suction surface boundary layer separation beyond 50% of the span-wise height as well as more serious tip leakage blockage. The twin swirl resulted in a decrease in the total pressure ratio, maximum efficiency and stable operating range. The steady and unsteady numerical calculation results were consistent, though some differences were observed in the values. For the steady calculation, the maximum efficiency and choked mass flow decreased by 0.88% and 1.74%, respectively, and the mass flow at the stable boundary increased by 3.92% as compared to the uniform inlet flow at twin swirl intensity of 24°. During the unsteady calculation, the mass flow exhibited an increase of only 2.2% at the stable boundary. Under twin swirl and co-rotating bulk swirl and uniform inlet flow, the leading edge spillage of the tip leakage flow resulted in compressor instability. The counter-rotating bulk swirl changed the mechanism of instability. The characterisation of the swirl distortion presented a difference between the steady and unsteady calculations near the stable boundary. The unsteady calculation exhibited a lower mass flow at the stable boundary point and a higher total pressure ratio.

Numerical Investigation of Three-Dimensional Separation in an Axial Flow Compressor: The Influence of Freestream Turbulence Intensity and Endwall Boundary Layer State

Regions of three-dimensional separations are an inherent flow feature of the suction surface-endwall corner in axial compressors. These corner separations can cause a significant total pressure loss and reduce the compressor's efficiency. This paper uses wall-resolved LES to investigate the loss sources in a corner separation, and examines the influence of the inflow turbulence on these sources. Different subgrid scale (SGS) models are tested and the choice of model is found to be important. The σ SGS model, which performed well, is then used to perform LES of a compressor endwall flow. The time-averaged data are in good agreement with measurements. The viscous and turbulent dissipation are used to highlight the sources of loss, with the latter being dominant. The key loss sources are seen to be the 2D laminar separation bubble and trailing edge wake, and the 3D flow region near the endwall. Increasing the freestream turbulence (FST) intensity changes the suction surface boundary layer transition mode from separation induced to bypass. However, it does not significantly alter the transition location and therefore the corner separation size. Additionally, the FST does not noticeably interact with the corner separation itself, meaning that in this case the corner separation is relatively insensitive to the FST. The endwall boundary layer state is found to be significant. A laminar endwall boundary layer separates much earlier leading to a larger passage vortex. This significantly alters the endwall flow and loss. Hence, the need for accurate boundary measurements is clear.

Effects of periodic wakes on boundary layer development on an ultra-high-lift low pressure turbine airfoil

The laminar-turbulent transition process in the boundary layer is of significant practical interest because the behavior of this boundary layer largely determines the overall efficiency of a low pressure turbine. This article presents complementary experimental and computational studies of the boundary layer development on an ultra-high-lift low pressure turbine airfoil under periodically unsteady incoming flow conditions. Particular emphasis is placed on the influence of the periodic wake on the laminar-turbulent transition process on the blade suction surface. The measurements were distinctive in that a closely spaced array of hot-film sensors allowed a very detailed examination of the suction surface boundary layer behavior. Measurements were made in a low-speed linear cascade facility at a freestream turbulence intensity level of 1.5%, a reduced frequency of 1.28, a flow coefficient of 0.70, and Reynolds numbers of 50,000 and 100,000, based on the cascade inlet velocity and the airfoil axial chord length. Experimental data were supplemented with numerical predictions from a commercially available Computational Fluid Dynamics code. The wake had a significant influence on the boundary layer of the ultra-high-lift low pressure turbine blade. Both the wake’s high turbulence and the negative jet behavior of the wake dominated the interaction between the unsteady wake and the separated boundary layer on the suction surface of the ultra-high-lift low pressure turbine airfoil. The upstream unsteady wake segments convecting through the blade passage behaved as a negative jet, with the highest turbulence occurring above the suction surface around the wake center. Transition of the unsteady boundary layer on the blade suction surface was initiated by the wake turbulence. The incoming wakes promoted transition onset upstream, which led to a periodic suppression of the separation bubble. The loss reduction was a compromise between the positive effect of the separation reduction and the negative effect of the larger turbulent-wetted area after reattachment due to the earlier boundary layer transition caused by the unsteady wakes. It appeared that the successful application of ultra-high-lift low pressure turbine blades required additional loss reduction mechanisms other than “simple” wake-blade interaction.

Effect of Partial Span Aspiration on the Performance of a Transonic Axial Compressor Rotor: A Numerical Study

Aspiration in an axial compressor is normally regarded as sucking out the low momentum boundary layer from blade suction surface, thus lowering the chances of flow separation and consequently that of stall under off-design operation. However, the suction mass flow does not take part in useful work and leads to loss of engine power output. This paper deals with a new concept of natural aspiration to energize blade suction surface boundary layer by injecting some fluid from pressure to suction side through a part span slot on the blade. The energized boundary layer has lesser tendency to separate, thus enhancing stall margin. Numerical simulations were carried out to study the effect of aspiration slot location and geometry on the performance and stall margin of a transonic axial compressor rotor. The computational results without aspiration slot were in fair agreement with the published experimental data. The modified rotor, with part span aspiration, showed ~3.2% improvement in stall margin at design rotational speed. The pressure ratio and efficiency of the aspirated rotor dropped by ~1.42% and ~2.0%, respectively, whereas the structural analysis did not indicate any adverse effect on the blade stress distribution in the presence of aspiration slot.

Control of separations in a highly loaded diffusion cascade by tailored boundary layer suction

In order to explore the control mechanism of boundary layer suction on the separated flows of highly loaded diffusion cascades, a linear compressor cascade, which has separated flows on the whole span and three-dimensional separations over the suction surface/endwall corner, was investigated by tailored boundary layer suction. Three suction surface-slotted schemes and two combined suction surface/endwall-slotted schemes were designed. The original cascade and the cascade with part blade span suction were experimentally investigated on a high-subsonic cascade wind tunnel. In addition, numerical simulation was employed to study the flow fields of different suction schemes in detail. The results shows that while tailored boundary layer suction at part blade span can effectively remove the separations at the suction span, the flow fields of other spans deteriorated. The reasons are the ‘C’ shape or reverse ‘C’ shape spanwise distribution of static pressure after part blade span boundary layer suction. Suction surface boundary layer suction over the whole span can obviously eliminate the separation at the suction surface. However, because of the endwall boundary layer, suction surface boundary layer suction cannot effectively remove the corner three-dimensional separation. The separation over the whole span and the three-dimensional separation at the corner are completely eliminated by combined suction surface/endwall boundary layer suction. After combined boundary layer suction, the static pressure distribution over the blade span just like the shape of ‘C’ is good for the transport of the low-energy fluid near the endwall to the midspan.

Experimental study of the impact of hole-type suction on the flow characteristics in a high-load compressor cascade with a clearance

In a low-speed linear cascade wind tunnel, effects of boundary layer suction on the flow characteristic in a highly-loaded compressor cascade with a stator clearance are studied experimentally, and two types of clearance including “long clearance” and “short clearance” are considered separately. In particular, hole-type suctions with different locations and distributions on the endwall are applied to control clearance flow and loss. The nature of flow separation is described by analyzing the properties of the limiting streamlines, based on the ink-trace flow visualizations on the cascade surface and lower endwall. The configuration of suction holes was associated with the kinetic energy of a clearance flow and the interaction of tip clearance flow with the endwall boundary layer and the suction surface boundary layer. The separation line and reattachment line form the base of a separation structure for corners with clearance. The results show that the clearance flow can be greatly improved and the losses reduced to some extent when boundary layer suction is adopted appropriately. The maximum amplitude of the loss reduction in different schemes with clearance is 32%.

Endwall Loss and Mixing Analysis of a High Lift Low Pressure Turbine Cascade

A high lift low pressure turbine (LPT) profile designated L2A is used as a test bed for studying the origin of endwall mixing loss and the role of vortical structures in loss development. It is shown analytically and experimentally that the mixing forces within the endwall wake can be decoupled into either mean flow or turbulent forces and can be further classified as either reversible or irreversible. Among the irreversible forces, mean flow shear is negligible compared to turbulent shear, suggesting that turbulence dissipation is the dominant cause of loss generation. As a result, the mean flow components of the vortical structures do not generate significant mixing losses. Rather than mixing effects, the mean flow of the vortices causes the suction surface boundary layer to separate inside the passage, thereby generating the large low energy regions typical of endwall flows. Losses are generated as the low energy regions mix out. This vortex separation effect is demonstrated with an experiment using a profile fence and pressure surface modification near the endwall. The findings in this paper suggest that profile modifications near the endwall that suppress flow separation may provide loss reductions additive to modifications aimed at weakening vortical structures, such as endwall contouring.