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This paper proposes a new strategy for extending centrifugal compressor stall margin through shroud passive flow control placed at the impeller radial bend. Relative to the usual leading-edge location, the radial bend presents fewer geometrical constraints for flow control devices and allows them to have potentially more impact on low-momentum flow structures that lead to stall. Two centrifugal compressors with fishtail pipe diffusers are chosen for study: a high-speed transonic design exhibiting diffuser stall and a low-subsonic design with exducer stall. Four passive flow control techniques placed on the impeller shroud at the radial bend are evaluated for each compressor, in terms of stall margin improvement and efficiency penalty at design mass flow, using unsteady RANS CFD simulations. These are circumferential groove, slots, impeller recirculation pipe and diffuser-impeller recirculation pipe. The results indicate that passive flow control at the impeller shroud radial bend can be significantly more effective than at the impeller leading edge. For each compressor, the flow mechanisms associated with stall initiation and its delay by the most effective studied flow control technique as well as the associated efficiency penalty are elucidated. The findings indicate that techniques having flow injection with high spanwise penetration and low relative streamwise momentum are most effective for delaying diffuser stall, while the opposite applies to exducer stall. The design point efficiency penalty is due to mixing/shear losses from the jet-in-cross flow phenomenon in the impeller and shear losses from flow redistribution in the diffuser. [ABSTRACT FROM AUTHOR]

As a typical non-axisymmetric inflow, circumferential pressure inlet distortion was experimentally investigated in an axial flow compressor. Based on the obvious decreasing trend of stall margin with the enhancement of distorted intensity, the stall mechanism under the condition of maximum distorted intensity, namely, 24.57 %, was studied in detailed. The casing static pressure over the rotor blade tip along the chord and circumferential direction were measured using a collection of dynamic transducers. Results show that under the circumferential inlet distortion condition, the pressure disturbance is always generated downstream from the distorted region along the rotational direction, and circumferentially propagates until it disappears in the upstream of the distorted region. With the continuous throttling, if the pressure disturbance that has not completely disappeared in the undistorted region enters the distorted region again, its energy will be enhanced, finally induces the stall. Using the above understanding as a foundation to apply partial tip air injection to suppress the disturbance caused by the circumferential inlet distortion, and resist its adverse effect on stall margin to ensure the safe operation of the compressor. It demonstrated that the tip air injection applied in the downstream of the distorted region can obtained the largest stall margin on the premise of the same injected momentum. For the strongest distortion case (DC (60) equals to 24.57 %), the injectors fixed downstream the distorted region can reduce the number and scale of the spike disturbances, suppress the local tip leakage vortex, and decrease the amplitude of the pressure disturbance energy, thus a maximum stall margin improvement of 3.18 % is obtained. However, tip air injection in the undistorted region cannot contribute the stall margin improvement that it leads to stronger tip leakage vortex and earlier stall inception. This study provides a guideline to design the non-axisymmetric tip air injection for circumferential inlet distortion with a help of injected mass flow and energy reduction. [ABSTRACT FROM AUTHOR]

The impacts of the increased hub clearance size (HCS) in the cantilevered tandem stator on the overall performance and flow behaviors have been investigated based on numerical simulations in a small-scale axial flow compressor stage. The results indicate that the HCS variation in the front blade of the tandem stator has a more significant effect on the compressor peak efficiency and stall margin. However, there is no obvious change of the compressor performance with the increase of the HCS in the rear blade. Detailed flow analysis shows that the increase of the HCS in the front blade has remarkable impacts on the flow fields in the front blade, while it has an only slight influence on the hub flow fields in the rear blade. It is found that the compressor efficiency penalty is due to the increase of the loss generation near the hub caused by the larger extent of hub leakage flow, while the stall margin improvement is due to the redistribution of the mass flow rate along the spanwise direction. As a result, the inlet flow condition within the upper blade span is improved and the flow blockage is reduced. The resulting influence of the increased HCS in the rear blade is mainly located near the blade leading edge of the rear blade, and there is no remarkable effect on the flow field in the front blade passage. [ABSTRACT FROM AUTHOR]

The impacts of casing aspiration with circumferential slot over the rear rotor on the overall performance and unsteady flow behaviors have been studied with numerical simulations in a counter-rotating axial flow compressor. The results indicate that casing aspiration is an effective active flow control method to expand the safe operating range of the compressor. The peak efficiency is always decreased to a different extent in the aspiration schemes and the best aspiration scheme is located at near 10% tip axial chord without significant peak efficiency loss. After the casing aspiration, both the relative inlet flow angle and the leading edge blockage are reduced obviously in the tip region. The angle between the exit direction of tip leakage flow (TLF) and axial direction is also reduced over the whole blade chord range, especially below the aspiration slots due to the local blade unloading effect. Additionally, the interface between the TLF and incoming main flow is pushed more downstream in the aspiration schemes and it is more effective to improve the flow stability by controlling the TLF released around mid-chord through casing aspiration. The intensity of both TLF and double leakage flow is attenuated and the overall oscillations inside are alleviated to a different extent in the aspiration schemes. Frequency analysis shows that the fluctuation of TLF with lower frequency component is suppressed by casing aspiration which is thought to be beneficial for the enhancement of stall margin. [ABSTRACT FROM AUTHOR]

In order to improve the operating stability of a centrifugal compressor, discrete cavities are introduced just upstream of the impeller and optimized for further improve the operating stability. Aerodynamic analysis is performed using a commercial code, ANSYS-CFX 15.0, to solve three-dimensional Reynolds-averaged Navier–Stokes equations with a shear stress transport turbulence model. The stall margin of the centrifugal compressor is used as the objective function, and three geometric parameters associated with the shape of the discrete cavities (the width and angle of the cavity port, and the axial distance between cavities) were selected as design variables. The Latin hypercube sampling method is used to compose the design space with 27 design points, and a radial basis neural network model is constructed as a surrogate model of the objective function. The optimization results show that the stall margin of the centrifugal compressor with the optimum discrete cavities is improved by 15.4% compared to the compressor without cavity. [ABSTRACT FROM AUTHOR]

Development of casing treatment geometries which will enhance the stable operating envelope of the compressor stage without sacrificing the efficiency is a great challenge for the designers. Bend skewed casing treatment with relatively lower porosities in the presence of plenum chamber is proven to be capable of improving the stall margin and gain in the compressor stage efficiency compared to other conventional casing treatments. Experimental investigations presented in this paper show very encouraging results with substantial improvement in the stable margin of the compressor stage with increase in the stage efficiency. The three bend skewed casing treatments designed for the experimental evaluations have porosities of 21%, 33% and 45%. The steady and transient performances of these three casing treatment geometries are evaluated at two axial extensions of 20% and 40%. Flow survey at the rotor exit shows distinct behaviors of the aerodynamic parameters in the tip region at near stall flow condition. The unsteady velocity measurements are performed with hot wire probe in the tip region ahead of the rotor leading edge to capture the fluctuations in the inlet axial velocity and the results are presented at the stall flow condition. The presence of the casing treatment changes the stalling behavior of the compressor stage. Highest porosity casing treatment results in 40.62% improvement in the stall margin with 0.88% improvement in the compressor stage efficiency at 20% axial extension. [ABSTRACT FROM AUTHOR]

The circumferentially non-uniform clearance has significant effect on the overall performance and operation safety of high loading compressor rotors. However, both the variation rules and mechanism of flow characteristic in the circumferential, radial and streamwise directions caused by the asymmetric clearance are not fully clarified, which restricts the performance degradation assessment and rubust design of high loading compressors. Therefore, different types of asymmetric tip clearance were designed in this paper to reveal the circumferential fluctuation laws and induced mechanism of key characteristic parameters in full-channel domains of compressor rotor with circumferentially non-uniform clearance. It is shown that the circumferentially non-uniform clearance has less effect on the total pressure ratio and efficiency of compressor rotor, while it causes an obvious deterioration of the stall margin. Furthermore, the circumferentially non-uniform clearance causes the circumferential fluctuations of performance parameters, and the fluctuation amplitude has been affected by the eccentricity ratio and operating condition of rotor directly. In addition, the leakage flow originated from a single channel may pass through about three adjacent channels in rotor with uniform clearance, while it covers almost seven adjacent channels in the rotor with circumferentially non-uniform clearance. In consequence, it is precisely the difference of strength variation of leakage flow in tip region caused by the circumferential distributions of tip clearance that produces an alternated variation of low-velocity zone in rotor prototype and a concentrated distribution of low-velocity zone in the range of 270° to 360° of rotor with non-uniform clearance. Due to the serious blockage appearing in this range, the rotating stall of transonic rotor with circumferentially non-uniform clearance has occurred in advance, even with a maximum drop of 35.94% in stall margin. [ABSTRACT FROM AUTHOR]

Rotor tip clearance (TC) shows significant effects on the stability and performance of aero-engine compressors. TC is also related to the producing of noise and tip leakage flow which is an indicator of critical operating conditions for rotating stall or surge. In this paper, the effect of TC on stall evolution process under different TC values is studied on a single-stage compressor with low speed and high load, and a frequency-domain interferometer diastimeter for high-resolution measurements is applied to accurately measure the TC value. In the present study, a soft-wall casing made of grindable material is applied to achieve the "zero clearance" condition, and casings with different diameters are adopted to reach different clearance values. Experimental results show that pressure rise and efficiency are improved with a small TC. With the TC varying from zero to a certain large scale, experimental results have proved that there is an optimal TC with the presence of maximum flow margin. In addition, eighty measurement points are evenly distributed around the casing circumferentially at the leading edge of the blade to analyze the circumferential flow characteristics of the tip flow field. And twenty wall-flush mounted microphones are evenly distributed circumferentially to monitor the induct sound field of the compressor, and the experimental results show that the sound pressure level (SPL) of the compressor is lowest at a relatively small TC rather than zero. The influence of different TC values on compressor is discussed based on the analysis of the pressure fields, stall process, and the acoustic signal. Furthermore, a detailed discussion of rotating instability (RI) brought about by large clearances is provided. [ABSTRACT FROM AUTHOR]

A new view of internal blade loading match is employed in this paper to reveal the essential relationship between sweep and complicated flow phenomenon in compressors. A forward swept rotor named as SF and an aft swept rotor named as SA, are designed originating from NASA Rotor 37. The numerical results indicate the isentropic efficiency and total pressure ratio of the new rotor with forward sweep are a little larger than rotor Baseline at peak efficiency point, and its stall margin has an over 10% increase as well. By contrast, the opposite change appears in the aft swept rotor SA. In addition, the choking massflow of swept rotors changes clearly as a result of the spatial change of aerodynamic throat and the forward sweep produces a larger throat area and choking massflow. A noted phenomenon is shown that both forward and aft sweep benefit a certain range of blade span for flow characteristic improvement, and the sweep also affects the inflow condition, 3-D shock structure and internal loading distributions of swept rotors. Both the new radial equilibrium in flowfield and transportation of low-energy fluids in boundary layer induced by sweep appear conductive to reduce shock wave and flow blockage in tip region of forward swept rotor SF. Meanwhile, the strength of leakage flow is also suppressed in the forward swept rotor SF because of the decreased blade loading near leading edge in tip region. In fact, all the change in swept rotors tends to be a new characteristic match of blade airfoil at different span, and the re-matching of flow characteristic for blade airfoil determines the stable operation range and the overall performance of swept rotors at design point directly. [ABSTRACT FROM AUTHOR]

In this paper, a recessed blade tip was applied to a transonic axial compressor in expectation of increasing the stable operating range. The internal flow has been calculated by a commercial flow solver for different recess cavities on the blade tip and the predicted performances for each case have been compared with one another. It was found that the recessed blade tip with a proper depth and length is effective to increase the stall margin. The tip leakage flow became weak due to a strong vortex generated in the recess cavity so that it was pushed backward by the main flow. This is why the recess cavity can increase the stall margin without any reduction of the adiabatic efficiency at the design point. Additionally, based on the comparison of the circumferential velocity between the numerical and the theoretical results, it was found that the rim of the recessed blade tip can work as the labyrinth seal and consequently reduce the mass flow rate through the tip clearance. [ABSTRACT FROM AUTHOR]

Blade tip ejection was introduced to a casing groove combined with tip injection in a transonic axial compressor with NASA Rotor 37. A parametric study of compressor performances was performed using three-dimensional (3-D) Reynolds-averaged Navier–Stokes equations with the k – ϵ turbulence model. Effects of three geometric parameters and one operating parameter, i.e., the front and rear lengths and height of the casing groove, and the mass flow rate of groove ejection, on the performance parameters, i.e., the total pressure ratio, adiabatic efficiency, stall margin, and stable range extension, were investigated in the parametric study. The application of a casing groove combined with tip injection and ejection was found to be effective in the simultaneous improvements of adiabatic efficiency, stall margin, and stable range extension without loss of the total pressure ratio of the transonic axial compressor, compared to the case with a casing groove with injection only, and a smooth casing. [ABSTRACT FROM AUTHOR]

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. [ABSTRACT FROM AUTHOR]

In order to enhance the aerodynamic performance of a centrifugal compressor with a vaned diffuser while also improving its stall margin, squealer tip cavities were applied to the tips of the impeller blades. Computational analysis was carried out by solving the compressible three-dimensional Reynolds-averaged Navier-Stokes equations. The numerical approach was validated by comparing the results from the flat tip reference case with experimental measurements. Various squealer tips with different cavity depths were applied to both the main and splitter blades and were numerically tested to investigate the influence of cavity depth on the performance and stall margin of the compressor. The results show that squealer tips with a proper cavity depth were effective in enhancing compressor efficiency by 0.32% at the design point, in addition, they were able to improve the stall margin by 1.02%. It was found that the squealer tips reduce tip leakage flow and circumferential velocity near the impeller shroud compared to using flat tips due to the generation of vortical structure within the cavity, which blocks leakage flow in a similar way to labyrinth seals. However, it was observed that by increasing the cavity depth, the size of this vortex becomes larger and the entropy production is increased inside the cavity. Moreover, it was found that the proper cavity depth decreases the mixing tip loss by about 4% compared to flat tips, this plays an important role in the stall margin improvement. [ABSTRACT FROM AUTHOR]

Abstract: A transonic axial compressor with circumferential casing grooves is optimized to improve operating stability. Numerical analysis is conducted by solving three-dimensional Reynolds-averaged Navier–Stokes equations using the shear stress transport turbulence model. An optimization process based on a weighted-average surrogate model and steady flow analysis are performed with three design variables defining the tip clearance, blade tip angle, and depth of the grooves. The steady stall inception point is identified from the last converged point, and values of the stall margin as the objective function are predicted using steady flow analysis at the design points sampled by Latin hypercube sampling in the design space. The surrogate model is constructed based on these objective function values. Optimization of this model found the optimum design, which yields a considerable increase in the stall margin compared to the smooth casing. To investigate unsteady behavior of the flow in the optimized compressor with casing grooves, an unsteady flow analysis is performed, and the stall inception point is re-predicted using this analysis. [Copyright &y& Elsevier]

Inlet guide vanes (IGVs) affect compressor performance by characterizing the upstream flow of the rotor blade row. After operating a long period, the IGV may suffer from blade deformation and consequently influence the downstream flow fields. The deformation of IGV always occurs locally, and the deformed structure can be regarded as a partially-flapped variable-camber IGV. In the current study, six typical configurations, with different span ranges of variable-camber IGVs, were selected. The impact on the compressor performance was discussed through numerical simulations. The dynamic mode decomposition (DMD) approach was applied over the computational domain for investigating the dynamic features. Results show that the stall margin and the total pressure rise of the compressor rotor are sensitive to the span range of the variable-camber IGV. For the unsteady feature in the flow field, the variation of IGV has a limited effect on the IGV itself while significantly changes the dynamic characteristics over the rotor domain. Compared to the case with conservative non-variable IGV, the extension in the flapped span yields a stall margin ranging from −4.1% to 8.1% as well as a monotone drop in the total pressure ratio. The mechanism behind the effect of IGV on the stall margin and the total pressure rise is dependent on the particular range of the modified IGV span. [ABSTRACT FROM AUTHOR]

This paper investigates a novel casing treatment for an axial compressor that combines a circumferential casing groove with flow recirculation channels. This casing treatment is located in the rotor's tip region and comprises a single shallow circumferential groove with 36 recirculation channels that are distributed equally around the annulus. The effects of the casing treatment on the stability and aerodynamic performance were evaluated numerically on a NASA Stage 37 single-stage transonic axial compressor. The results show that the addition of the recirculation channels improved the gain in stall margin while mitigating the penalty in the adiabatic efficiency and pressure rise compared to a casing groove without recirculation channels. Six geometric parameters were examined in a parametric study. The maximum increase in stall margin was 42.5% with small reductions in the efficiency and pressure ratio. To maximize the stall margin and minimize efficiency loss, two design variables were selected to perform a multi-objective optimization using a multi-objective genetic algorithm coupled with surrogate models using a radial basis neural network. The optimization successfully improved both the objective functions for all representative designs compared to the reference design. [ABSTRACT FROM AUTHOR]

Inclined discrete cavities are proposed to improve the operating stability of a centrifugal compressor while minimizing the flow loss, and optimized to simultaneously maximize the adiabatic efficiency and stall margin. Aerodynamic analysis was performed using three-dimensional Reynolds-averaged Navier–Stokes equations. The numerical results obtained for the total pressure ratio and adiabatic efficiency were validated with experimental data for the centrifugal compressor with a smooth casing. The angle of the cavity port, axial distance between cavities, and inclined angle of the cavity were selected as design variables based on sensitivity analysis. The stall margin and adiabatic efficiency at the design point were used as the objective functions for the design. Latin hypercube sampling was used to select 24 design points in the design space. Two neural network models, that is, the radial basis and deep neural networks, were tested with respect to surrogate models and their performances were compared using statistical analysis. A hybrid PSO–GA algorithm was used to identify the optimal solutions for the surrogate models. The deep neural network model has a better overall performance than the radial basis neural network model. The optimization results show that the stall margin increases, with increments of up to 6.41% and 11.32% compared with the reference design and compressor with a smooth casing, respectively. [ABSTRACT FROM AUTHOR]

Casing treatment is a powerful method for enhancing stability of certain aircraft compressors in which tip leakage vortex (TLV) triggers stall. An optimized slot-type casing treatment was installed on the first rotor of a highly loaded two-stage compressor. The experimental result showed that the casing treatment could not increase the compressor stability at three tested speeds. This study aims to understand the failure mechanism of the casing treatment for developing an effective method to depress stall. A "three-step method" is proposed to determine the stall-limiting location and understand the underlying stall mechanism. At 90% speed, there is a spike in the aerodynamic flow parameters in the tip region of the first rotor under the near stall condition compared with that under the design condition. The spike is induced by an interaction among the TLV, boundary layer at the blade suction surface, and passage shockwave. Both the TLV and the shockwave–boundary layer interaction contribute to the accumulation of tip blockage which triggers compressor stall. The TLV does not break down under the near stall condition, but unsteady forward spillage of the TLV at the blade tip leading edge can be detected, resulting in unsteady tip flow. When the casing treatment is installed on the first rotor, a typical flow circulation is established in the slots. The intensity of the TLV and the flow unsteadiness are greatly depressed, and the discharge capability in the tip region is remarkably improved. However, the flow circulation cannot manipulate the fierce boundary layer separation induced by the shockwave, which causes increased tip blockage and reduced compressor stability. A coupled casing treatment, which can bleed the shockwave induced blockage, is proposed and found to be very effective in improving compressor stability. Therefore, the technology for damping the compressor stall should be capable of manipulating both the TLV and the shockwave induced boundary layer separation. At 100% speed, a shedding vortex is initiated at the hub of the second stator. This vortex expands sharply in the stalling process and triggers compressor stall. Variation of the location at which the compressor stalls is the reason for the failure of the casing treatment at 100% speed. [ABSTRACT FROM AUTHOR]

Low-reaction (LR) aspirated compressor is a promising design methodology to enhance the work capacity substantially. However, for the high-load transonic compressor, to achieve LR design, the effect of different combinations of inlet pre-swirl and meridian contraction on performance has not been investigated. In addition, the quantitative advantage of LR design over convention-reaction (CR) design is also not available. This paper presents a series of aerodynamic designs of the high-load transonic compressor stage, including a CR stage and three LR stages with a comprehensive application of the LR design method and the aspiration technology. 3D calculation reveals that under the same tip speed and rotational speed and a total bleed flow fraction of 2.84% of the stage inlet mass flow, the preliminary LR design achieves a total pressure ratio of 2.34 and a throughflow adiabatic efficiency of 88.11%, which is respectively 30% and 0.19% higher over CR design. The corresponding loading coefficient rises from 0.45 to 0.66. Through the modified design of introducing non-uniform counter-swirl along the span and increasing the meridian contraction, the loading coefficient is further increased to 0.69. Due to the insufficient stall margin induced by ultra-high-load LR design, a further casing contraction of the optimized design relative to the modified one is conducted to decrease the loading and thus the stall margin is increased by 1%. [ABSTRACT FROM AUTHOR]

To prevent the deterioration of the inlet tip and hub radial distortion on stall margin, tip air injection was experimentally studied in a three-stage axial flow compressor. By generating a total pressure distortion intensity of 16.8% and 8.6%, the inlet tip and hub radial distortion can obtain stall margin variation of -5.1% and 10.5%, respectively. With the help of time-resolved pressure transducers installed on the casing wall and five-hole probes installed downstream of each rotor blade row, the unsteady measurement results demonstrate that the first stage stalls first under different inflow conditions and only the unsteady tip leakage flow at the first stage becomes stronger near the stall limit. Therefore, tip air injection is applied at the first stage under the distorted inflow to ensure the stable operation. Under inlet tip radial distortion, which strengths the unsteady tip leakage flow and blade tip separation reflected by the outlet axial velocity and flow angle, tip air injection can weaken the unsteady tip leakage flow, and remove the tip separation at the first stage, thereby making the second stage stall first such that only stall margin improvement of 19.2% can be obtained compared with uniform inflow. Although hub radial distortion can delay the occurrence of stall by weakening unsteady tip leakage flow and increasing the tip flow capability, overlarge injection can still improve the stall margin by 20.5% through further weakening the unsteady tip leakage flow and reducing the blade tip separation. [ABSTRACT FROM AUTHOR]

Casing treatment utilized in aircraft compressors is characterized by its powerful capacity of enhancing stability but causes remarkable efficiency penalty. To deal with this problem, we tested several types of casing treatments, including slots, grooves, recirculating casing treatments, and a composite structure to understand their characteristics. The test results show that different casing treatments can be combined to obtain a higher compressor stability, and the efficiency penalty can be decreased by implementing a novel flow management technology. Based on the results, a coupled casing treatment (CCT) that is constructed with slots, injectors, and a plenum chamber is proposed and optimized in the present study. The optimized CCT improves compressor stability by 16.7% with no penalty on the compressor efficiency. Two flow circulations, namely, an inner circulation in the slots and an outer circulation from the slots to the injectors, which dampen the development of tip leakage vortex, account for the excellent stability enhancement. The outer circulation, defined as the coupling effect, has generally a positive effect on compressor stability and a negative effect on the compressor efficiency. The effects of the length and numbers of the slots on the coupling effect are more significant compared to the slot location. The hysteresis effect that the recovery of tip blockage lags behind that of the tip leakage vortex cannot be established due to inadequate injection energy actuated in the CCT. As a result, a circumferentially discrete distribution of recirculating loops cannot improve compressor stability satisfactorily. As the CCT covers the full annulus, a much higher stability is obtained, and the boundary layer at the rear part of blade passages is severely separated when approaching the stall limit, which ultimately triggers compressor stall rather than the tip leakage vortex. The stall inception behaves as a full-annulus-dimension collapse with the flow phenomena of backflow at the trailing edge and forward spillage at the leading edge of the blade. [ABSTRACT FROM AUTHOR]