Compressor Stability Research Articles

Effect of bent inlet pipe on the flow instability behavior of centrifugal compressors

Bent inlet pipes are often used in centrifugal compressors due to limited installation space, and an understanding of the effect on compressor stability is essential for safety and durability. This paper firstly investigates flow instability behaviors in two compressors, one with a straight inlet pipe and the other with an S-shaped bent pipe. In detail, it analyzes the resulting flow fields, instability evolution paths and surge boundaries. The results show that the S-shaped pipe obviously affects the flow field at high mass flow rates, while reverse flow mainly influences the flow field at low mass flow rates. Reverse flow first occurs at certain flow passages with a high pressure difference that is predominantly decided by the volute rather than the S-shaped bent pipe. In addition, centrifugal compressors can tolerate reverse flow to some extent so that surge would not occur immediately if reverse flow occurs unless the reverse flow region extends circumferentially and radially to a sufficiently large size. Since the S-shaped pipe is not dominant in the creation and extension of reverse flow, it does not exacerbate the stability of the central compressor to a great extent. Last but not least, the S-shaped pipe is noted to delay the occurrence of surge at 90% rotating speed, which suggests the possibility of improving compressor stability with bent inlet pipes. This result differs from the conventional understanding that inlet distortion usually deteriorates compressor stability and emphasizes the particularity of centrifugal compressors.

Surge and Stall Detection Using Acoustic Analysis for Gas Turbine Hybrid Cycles

Abstract Compressor dynamics were studied in a gas turbine—fuel cell hybrid power system having a larger compressor volume than traditionally found in gas turbine systems. This larger compressor volume adversely affects the surge margin of the gas turbine. Industrial acoustic sensors were placed near the compressor to identify when the equipment was getting close to the surge line. Fast Fourier transform (FFT) mathematical analysis was used to obtain spectra representing the probability density across the frequency range (0–5000 Hz). Comparison between FFT spectra for nominal and transient operations revealed that higher amplitude spikes were observed during incipient stall at three different frequencies, 900, 1020, and 1800 Hz. These frequencies were compared to the natural frequencies of the equipment and the frequency for the rotating turbomachinery to identify more general nature of the acoustic signal typical of the onset of compressor surge. The primary goal of this acoustic analysis was to establish an online methodology to monitor compressor stability that can be anticipated and avoided.

Wall-Modeled Large-Eddy Simulations of a Multistage High-Pressure Compressor

This study aims at evaluating the feasibility and the accuracy of a Wall-Modeled Large Eddy Simulation of an actual aeronautical multistage axial high-pressure compressor. The computational domain is composed of 37 blades and a geometrically complex recirculating cavity. The numerical method TurboAVBP is able to handle such a technical challenge thanks to its unstructured and massively parallel approach as well as dedicated rotor-stator interface treatments. The influence of grid resolution, from less than 100 millions to more than 1 billion of cells, is particularly evaluated. The intermediate grid correctly predicts the global aerodynamic performances up to the lowest mass flow rate regime. In comparing with time-resolved measurements, the finest grid is shown to accurately predict flow within rotors, especially in their tip regions that are critical for performances and stability of the whole compressor.

Evolution of the flow instabilities in an axial compressor rotor with large tip clearance: An experimental and URANS study

With modern engine designs trending toward smaller cores to increase propulsive efficiency, the gap-to-span ratio in compressors is expected to increase, especially for the rear stages. Detailed knowledge of stronger tip clearance flow and its impact on compressor instability under large clearance condition become increasingly important. The present study is aiming for illuminating the instability evolution and the underlying flow physics in a large-tip-gap compressor, with the use of casing-mounted pressure transducers to acquire the unsteady nature of tip flow, and full-annulus URANS to obtain the three-dimensional flow details related to instability generation. Results show that flow instability evolution of the compressor along speed-line experiences three stages: stable state, rotating instability, and rotating stall. Two critical behaviors of tip leakage vortex (TLV) are found to relate to the transition of instability pattern. As TLV moves to the adjacent blade trailing edge, it starts to oscillate under the interaction with the adjacent blade, leading to a short-length rotating disturbance, i.e. rotating instability. At near-stall condition, as the interface of TLV and main-flow exceeds the passage inlet plane, forward spillage occurs along with radial vortexes periodically shed from blade leading edge. The periodically generation, movement and decay of leading edge vortexes (LEVs) constitute an orderly propagating disturbance rather than evolve into a stall cell. However, as the compressor is further throttled, the orderly propagation of LEVs collapses due to their scattering, resulting in the generation of stall cells with local stronger blockage and higher entropy.

Nonaxisymmetric Study of Tip Leakage Flow in a Centrifugal Compressor With a Volute During Stall Process

Abstract Tip leakage flow (TLF) patterns, which affect compressor performance, are closely related to compressor stability. To date, minimal attention has been given to circumferential nonuniformity of the TLF in a centrifugal compressor with a nonaxisymmetric volute structure. In this study, the circumferential difference of the TLF in a centrifugal compressor with a volute during the stall process is analyzed. The circumferential nonuniformity of tip leakage vortex (TLV) trajectories, loading distribution near the tip, and distance between the TLV core and the leading edge (LE) of splitter blades were also investigated. It is shown that in the circumferential direction, there are two peaks associated with the angle (α) between the TLV trajectory of the seven main blades and the axial direction. As the stall process progresses, the blade whose LE is affected by the high static pressure band (PP) induced by the volute tongue (VT) loses its work capacity first and the α difference between this blade and the other blades increases. In addition, the tip loading and TLF velocity of the blade whose LE is affected by the high static pressure band induced by the VT are at a minimum, and the flow loss in the tip clearance is higher. There is a phenomenon of the TLV breakdown. When the blade trailing edge (TE) is located in the low static pressure region, TLV streamlines appear as a significant turn at the breakdown point. However, the TLV streamlines at other circumferential positions do not exhibit this phenomenon.

The effect of high temperature steam ingestion on aerodynamic performance and stability of a low-speed axial compressor

To investigate the effect of high temperature steam ingestion on the aerodynamic stability of a multistage axial compressor, a two-stage low-speed axial compressor was studied, and full-annulus steady-state and unsteady-state numerical simulations were carried out. The effect of the high temperature steam mass fraction and the distribution of steam at the inlet boundary on the aerodynamic stability of a two-stage low-speed axial compressor was investigated. From the simulation results, we found that high temperature steam ingestion has an adverse effect on the low-speed axial compressor. The larger the steam mass fraction is, the greater the impact of the steam ingestion on the stability boundary and stall margin will be. When the steam mass fraction is equal to 0.35 and 0.7%, the stability margin decreases from 36.07 to 29.72% and 28.05%, respectively. The distribution of steam at the inlet boundary will change the performance and stability. When the steam ingestion range is less than 90°, the steam ingestion area increases and the stability margin will decrease gradually. After 90°, the stability margin is almost unchanged. The difference between the calculated and experimental values of the stability margin reduction caused by steam ingestion is 0.87%. In addition, with the ingestion of high temperature steam, the blockage in the corresponding passages is intensified and the loss is increased, which leads to the occurrence of the stall in advance. It is evident that steam ingestion has a significant impact on compressor stability, ensuring that the steam mass fraction and steam ingestion range are close to the actual value.

Improving the operating range using a centrifugal compressor with a tandem impeller

The centrifugal compressor is being utilized in both commercial and military applications; however, the compressor performance is limited by flow instability affecting the stable operating range. In this paper, a tandem-bladed impeller method is employed to achieve passive flow control to improve the stable operating range of a compressor without control costs or performance penalties. The results inspiringly indicate that the tandem impeller method can achieve a maximum operating range gain of 16.7%. The baseline compressor instability is induced by tip region flow deterioration, which results from tip leakage flow and low-momentum fluid radial transport. For a tandem impeller, the inducer suction side always has a higher trailing edge static pressure compared to the baseline impeller due to the exducer potential influence and the restriction of the Kutta-Joukowski condition, which weakens the radial transport of low-momentum fluid and constrains the leading-edge separation. As a result, the tandem-design compressor operates far from the impeller instability point. Further investigations show that the inducer pressure field varies with the clocking fraction λs, which results in the change of the radial transport at the inducer suction side, finally affecting the flow stability.

Rotating Stall Induced Non-Synchronous Blade Vibration Analysis for an Unshrouded Industrial Centrifugal Compressor.

Rotating stall limits the operating range and stability of the centrifugal compressor and has a significant impact on the lifetime of the impeller blade. This paper investigates the relationship between stall pressure wave and its induced non-synchronous blade vibration, which will be meaningful for stall resonance avoidance at the early design phase. A rotating disc under a time-space varying load condition is first modeled to understand the physics behind stall-induced vibration. Then, experimental work is conducted to verify the model and reveal the mechanism of stall cells evolution process within flow passage and how blade vibrates when suffering such aerodynamic load. The casing mounted pressure sensors are used to capture the low-frequency pressure wave. Strain gauges and tip timing sensors are utilized to monitor the blade vibration. Based on circumferentially distributed pressure sensors and stall parameters identification method, a five stall cells mode is found in this compressor test rig and successfully correlates with the blade non-synchronous vibration. Furthermore, with the help of tip timing measurement, all blades vibration is also evaluated under different operating mass flow rate. Analysis results verify that the proposed model can show the blade forced vibration under stall flow condition. The overall approach presented in this paper is also important for stall vibration and resonance free design with effective experimental verification.

Flow instability suppression and deep surge delay by non-axisymmetric vaned diffuser in a centrifugal compressor

The performance of centrifugal compressors is significantly limited by the deep surge with strong flow oscillations, and the occurrence of deep surge is found to be related to flow instabilities. A novel in-house designed non-axisymmetric vaned diffuser (NAVD) is applied in a centrifugal compressor, and this paper experimentally analyzes influences of the NAVD on suppressing the flow instability and delaying the deep surge. Results show that the flow instability at the diffuser inlet region are distinctly suppressed by the NAVD, meanwhile the deep surge is delayed at 90% speed which extends the stable flow range (SFR) by 25.0%. Nevertheless, the NAVD neither delays the deep surge at other rotating speeds, nor does it suppress the flow instability at the impeller inlet region. According to a previous research, only at 90% speed, the deep surge onset initiates from the diffuser inlet flow instability, and this explains why the NAVD fails to delay the deep surge occurrence at other rotating speeds except for 90% speed. This paper experimentally proves that the NAVD has a positive influence on the flow instability suppression and deep surge delay in some situations, and it provides a new method to improve the compressor stability by non-axisymmetric vaned diffuser.

Coupling method of stability enhancement based on casing treatments in an axial compressor

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.

Experimental study on performance of centrifugal compressor exposed to pulsating backpressure

Pressure fluctuations in intake manifold of Internal Combustion Engine introduce a pulsating boundary condition to the centrifugal compressor. In this study, the performance of centrifugal compressor exposed to pulsating backpressure is experimentally investigated. Unsteady performance and surge characteristic of the compressor at different pulsating conditions are analyzed. Results show that instantaneous performance of the compressor forms a hysteresis loop encapsulating steady performance. A correlation of compressor unsteadiness with pulse frequency, magnitude, and local slope of the characteristic curve is obtained based on measurements. The results show that cycle-average performance of the compressor is notably lower than that at constant conditions. Specifically, the averaged peak efficiency at pulsating condition drops by 3%–7%. Moreover, Fast Fourier Transformation method is applied to study surge characteristics at pulsating conditions. Results manifest that compressor surge is postponed remarkably by pulsating backpressure at the three speeds. As the compressor approaches to small mass flow rate, flow fluctuations decay quickly and then increase dramatically when the surge happens. This phenomenon is resulted from the influence of the slope of characteristic curve on the filling-empty effect and compressor stability. Larger slope produces weaker filling-empty effect and hence smaller size of the loop, but reduces the stability of compressing system and triggers compressor surge.

Leakage and Cavity Pressures in an Interlocking Labyrinth Gas Seal: Measurements Versus Predictions

Gas labyrinth seals (LS) restrict secondary flows (leakage) in turbomachinery and their impact on the efficiency and rotordynamic stability of high-pressure compressors and steam turbines can hardly be overstated. Among seal types, the interlocking labyrinth seal (ILS), having teeth on both the rotor and the stator, is able to reduce leakage up to 30% compared to other LSs with either all teeth on the rotor (TOR) or all teeth on the stator. This paper introduces a revamped facility to test gas seals for their rotordynamic performance and presents measurements of the leakage and cavity pressures in a five teeth ILS. The seal with overall length/diameter L/D = 0.3 and small tip clearance Cr/D = 0.00133 is supplied with air at T = 298 K and increasing inlet pressure Pin = 0.3–1.3 MPa, while the exit pressure/inlet pressure ratio PR = Pout/Pin is set to range from 0.3 to 0.8. The rotor speed varies from null to 10 krpm (79 m/s max. surface speed). During the tests, instrumentation records the seal mass flow (m˙) and static pressure in each cavity. In parallel, a bulk-flow model (BFM) and a computational fluid dynamics (CFD) analysis predict the flow field and deliver the same performance characteristics, namely leakage and cavity pressures. Both measurements and predictions agree closely (within 5%) and demonstrate that the seal mass flow rate is independent of rotor speed. A modified flow factor Φ¯=m˙T/(PinD1−PR2) characterizes best the seal mass flow with a unique magnitude for all pressure conditions, Pin and PR.

Inlet bent torsional pipe effect on the performance and stability of a centrifugal compressor with volute

Inlet bent torsional (BT) pipes are applied to complex turbocharger systems and industrial centrifugal compressors owing to space and weight constraints. To determine the interaction mechanism between the inlet BT pipe and outlet volute, and its effect on the centrifugal compressor performance and stability, three compressor models (model with clean inlet (M0 model) and models with distorted inlets induced by BT pipes (M1 and M2 models)) were chosen to conduct a performance experiment and static pressure measurement of the casing wall. The results show that the M1 and M2 cause a 6.4% increment and 22% reduction in the stable operating range, respectively. At the near-choke and peak efficiency points, both M1 and M2 can change the casing static pressure distribution evidently and thus influence the compressor pressure ratio and efficiency significantly. At the near-stall point, the high static pressure at the inlet region induced by the BT pipe of M1 can weaken the pressure peak strip induced by the volute. Therefore, both the considerable increase in circumferential pressure uniformity and the positive pre-whirl distortion of M1 are responsible for the delay of compressor stall. Finally, the interaction relationship of the inlet/outlet distortions in the entire operating range was sketched.

Roles of vanes in diffuser on stability of centrifugal compressor

The stability considerations of centrifugal compressors become increasingly severe with the high pressure ratios, especially in aero-engines. Diffuser is the major subcomponent of centrifugal compressor, and its performance greatly influences the stability of compressor. This paper experimentally investigates the roles of vanes in diffuser on component instability and compression system instability. High pressure ratio centrifugal compressors with and without vanes in diffuser are tested and analyzed. Rig tests are carried out to obtain the compressor performance map. Dynamic pressure measurements and relevant Fourier analysis are performed to identify complex instability phenomena in the time domain and frequency domain, including rotating instability, stall, and surge. For component instability, vanes in diffuser are capable of suppressing the emergence of rotating stall in the diffuser at full speeds, but barely affect the characteristics of rotating instability in the impeller at low and middle speeds. For compression system instability, it is shown that the use of vanes in diffuser can effectively postpone the occurrence of compression system surge at full speeds. According to the experimental results and the one-dimensional flow theory, vanes in diffuser turn the diffuser pressure rise slope more negative and thus improve the stability of compressor stage, which means lower surge mass flow rate.

Operating Conditions Effect on Aerodynamic Instabilities in Axial Compressors

This work aims at analyzing the flow behaviour through an axial compressor and predicting the instabilities appearance with the help of Computational Fluid Dynamics using the FLUENT software. The aerodynamic instabilities produced in axial flow compressors are one of the most important issues for compressor designers causing energy losses within the compressor. In this paper we focus on the effect of the operating conditions on the instabilities inception. In order to well demarcate the conditions favouring the flow instabilities the flow around compressor blades is investigated in different operating conditions characterized by the inlet Mach number the flow attack angle and the compressor rotation velocity. Three cases are considered a 2D, 3D isolated blade, and a 2D stage cascade. An attempt has been made to study the effect of the aerodynamic instabilities on various flow parameters. Particularly, four principle parameters, viz. the static pressure, the velocity, the lift and drag coefficients, are considered in this analysis. The obtained results show that the compressor instability margin is encountered between two range of inlet flow velocity, where operating at low inlet Mach number, especially with higher flow angles can involve the stall of the blades and operating at higher inlet Mach number, produce shock waves, in both situations the flow endure a separation and creation of vortices which leads to an aerodynamic instability reducing compressor efficiency characterized here by the lift and drag aerodynamic coefficients. The results also show that the compressor blade complex geometry is a factor controlling instabilities, where those instabilities differ from the root to the tip of the blade. It emerges that the aerodynamic simulation over a compressor stage can make appear within the flow new phenomena such as the rotating stall which does not appear while analysing the flow over an isolated blade.

МЕТОД РАСЧЁТА ТЕРМОГАЗОДИНАМИЧЕСКИХ ПАРАМЕТРОВ ТУРБОВАЛЬНОГО ГАЗОТУРБИННОГО ДВИГАТЕЛЯ НА ОСНОВЕ ПОВЕНЦОВОГО ОПИСАНИЯ ЛОПАТОЧНЫХ МАШИН. ЧАСТЬ II. ОПРЕДЕЛЕНИЕ ПАРАМЕТРОВ СТУПЕНЕЙ И МНОГОСТУПЕНЧАТЫХ КОМПРЕССОРОВ

Gas Turbine Engine (GTE) operating characteristics such as thrust (or power), specific fuel consumption and other cycle parameters on different regimes, can be determined by engine modeling and applying correspondent calculation method. Its accuracy is the function of the engine’s element maps definition precision. So these maps representations influence for engines investigation results significantly. Main points and equation system for engine performances calculation method were represented in Part I of this article. The method gives an opportunity for the flow path thermodynamical parameters and engine integral values analyzing by using multistage axial blade machines blade-to-blade descriptions. The compressor and gas turbine and parameters are getting by special program modules, adding to the engine operating characteristics investigation program complex. These modules use the flow path and cascade middle radius geometrical parameters as the data for calculation. The goal of this article is the representation of the method for axial stages and multistage compressors performances definition. The calculation technique is based on one-dimensional (1D) multistage axial compressor flow description. Proposed 1D flow analysis method allows to get the multistage axial compressor maps taking into account the blade-to-blade gaps flow bleeding and by-pass. The method including is founded on the thermal and gas dynamic equations and turbomachinery theoretical dependences and empirical functions for losses and deviation angles determination. Besides, the representing method allows to calculate gas dynamic parameters, velocity triangles, angles of attack, evaluate their deviations from optimal values, hydraulic losses. Also, it can show accordance of stages working on different regimes, find the stage, which is a reason for compressor instability, and stall margin. This method can be used in GTE mathematic simulation, founded on blade-to-blade description multistage blade machines or also in multistage compressor designing. The proposed method gives the opportunity to control the stator variable vanes stagger angles control and to analyze its influence for stage and multistage compressor gas dynamic parameters and maps.

Mechanism investigation of enhancing the stability of an axial flow rotor by blade angle slots

This paper seeks to reveal the mechanisms of enhancing the stability of a subsonic axial flow rotor by applying blade angle slots casing treatment (BSCT). When blade angle slots are applied, there is about 9% stall margin improvement for the experiment and about 8% stall margin improvement for the calculation, but the decrease in the rotor maximum efficiency is about 11% for the experiment and the calculation. The compared results between smooth wall and blade angle slots indicate that the backflow in the rotor top passage is weakened by the injected and sucked flows formed inside the slots of BSCT. Moreover, the injected flows inside the slots interfere with the flows in the rotor passage upstream, and this interference leads to large flow losses. Therefore, the rotor efficiency for blade angle slots is much lower than that for smooth wall. To confirm that the structural optimization of blade angle slots can effectively improve the compressor stability with small efficiency losses, optimized blade angle slots casing treatment (BSCT1) was designed according to the past experience of slot casing treatment. The calculated result shows that the optimized blade angle slots generate about 59% stall margin improvement, and the compressor maximum efficiency with the optimized blade angle slots is about 0.05% more than that for smooth wall. The flow field analyses show that the strong sucked flows formed inside the slots for BSCT1 can prevent the backflow, which exists in the rotor top passage for BSCT, from appearing. In addition, the level of interference of the flows in the rotor passage upstream for BSCT1 is much lower than that for BSCT, and the corresponding losses with BSCT1 become lower. Therefore, the rotor with BSCT1 has a larger stable operating range and better efficiencies than that with BSCT.

On the Effect of Engine Pulsations on the Performance of a Turbocharger Centrifugal Compressor

In an internal combustion engine, the centrifugal compressor is placed upstream of the inlet manifold and therefore, it is exposed an unsteady flow regime caused by the inlet valves of the cylinder arrangement. This valve motion sets a pulsating state at the compressor exit, having greater influence when the operation is near the surge margin of the compressor. This paper presents the experimental results of the evaluation of the surge dynamics on a compressor with induced downstream pulsating flow. Different pulsation levels are achieved by the variation of three different parameters on the induced pulse: pulse frequency, amplitude, and system storage volume (plenum). Each pulse parameter was evaluated independently in order to assess its effect on the compressor stability limit. The main effect on the surge margin of the compressor was found to be due to the presence of a storage volume in the system for all cases (steady/pulsating condition) and at all frequencies. It was found that the magnitude of the pulse frequency determines the hysteresis behavior of the system that leads to a phase difference between the convected terms and the acoustic dominated terms, and therefore this affects the onset of flow instability, surge, in the compression system under study.

The Effect of Trailing Edge Clearance in Suppressing Hub-Corner Stall

Hub corner is a high-loss area located within in the blade passages of turbomachinery. It is well known that flow separation and vortex development in this area directly affects, not only energy loss and efficiency, but also the stability of axial compressors. In this paper, a high diffusion factor cascade is used in order to analyze the vortex structure which causes the hub-corner stall in cascade. The trailing edge clearance method is proposed for controlling the hub-boundary layer and this is compared with full uniform clearance. Our research highlights that the trailing edge clearance suppress the radial development of the suction separation vortex (SSV) in the hub corner and restrains the concentrated separation vortex (CSV). The influence of trailing edge clearance parameters and its effect on control of hub-corner stall is discussed in detail. It is revealed that clearance height avoids flow separation and suppresses radial vortex development, while the clearance length resists the overturning that is caused by transverse pressure gradient and controls the development of CSV. As the radial vortex plays a dominant role in hub-corner stall, clearance height has a better effect in controlling hub-corner stall in comparison to length at the constant flow area.

Improvements of performance and stability of a single-stage transonic axial compressor using a combined flow control approach

The performance and the stability of axial compressors are seriously concerned as it's widely used in aero-engines and power machines. A new combined flow control approach was developed previously using blade slot and vortex generator, which was used to reduce flow loss and improve flow stability. The approach was experimentally proved to be very effective in linear compressor cascades. To further evaluate its availability in actual compressors, the new approach is implemented in a single-stage transonic compressor, which introduces a part-span slot into the rotor while a full-span slot and a vortex generator into the stator. CFD simulations are performed to evaluate the gains of compressor performance, and flow structures are analyzed to explain the corresponding mechanisms. Results show great improvements of performance and stability of the compressor. The averaged total pressure ratio and isentropic efficiency improve by 1.82% and 0.88% respectively due to the great reduction of separations in both the rotor and the stator. Because the compressor instability is initiated by the stator corner stall, the delay of the corner separations leads to a great increase of the stable mass-flow range by 26.1%. The compressor flow structures are also well improved because the detrimental backflows and large-scale vortices are mostly removed by the combined approach.