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A numerical study of casing treatments on a centrifugal compressor to improve stability and stall margin is presented. High efficiency, high pressure ratio, and a wide operating range are required for a high-performance centrifugal compressor. A ring groove casing treatment is effective for flow range enhancement in centrifugal compressors. Compressor performance was analyzed according to the ring groove location, and the results were compared with the case without a ring groove. The effect of guide vanes in the ring groove was also investigated. Four variants of grooves were modeled and simulated using computational fluid dynamics to optimize the groove location. Numerical analysis was performed using a commercial code ANSYS-CFX program. The simulation results showed that the ring groove increased the operating range of the compressor. The ring groove with guide vanes improved both performance of the compressor at low flow rates and the stall margin of the compressor. [ABSTRACT FROM AUTHOR]

This article presents an investigation on the performance characteristics of a transonic axial compressor with the application of the circumferential casing grooves. Three-dimensional Reynolds-averaged Navier–Stokes equations with the shear stress turbulence model are discretized by finite volume approximations and solved on hexahedral grids for the flow analysis. The numerical results for the pressure ratio and the adiabatic efficiency show good agreement with the experimental data. The numerical stall inception point is identified from the last converged point by the convergence criterion, and the stall margin (SM) is numerically predicted. The performances of an axial compressor with rectangular shape of the circumferential casing grooves are analysed in comparison with the case of smooth casing. Additionally, the SMs and the peak adiabatic efficiencies with various shapes of the circumferential casing grooves are compared and evaluated in order to explore the influence of groove shape. The application of the circumferential casing grooves leads to significant improvement in the operating stability of compressors with slight reduction in the peak efficiency. [ABSTRACT FROM AUTHOR]

Reliability-based design optimization (RBDO) of the NASA stage 37 axial compressor is performed using an uncertainty model for stall margin in order to guarantee stable operation of the compressor. The main characteristics of RBDO for the axial compressor are summarized as follows: First, the values of mass flow rate and pressure ratio in stall margin calculation are defined as statistical models with normal distribution for consideration of the uncertainty in stall margin. Second, Monte Carlo Simulation is used in the RBDO process to calculate failure probability of stall margin accurately. Third, an approximation model that is constructed by an artificial neural network is adopted to reduce the time cost of RBDO. The present method is applied to the NASA stage 37 compressor to improve the reliability of stall margin with both maximized efficiency and minimized weight. The RBDO result is compared with the deterministic optimization (DO) result which does not include an uncertainty model. In the DO case, stall margin is slightly higher than the reference value of the required constraint, but the probability of stall is 43%. This is unacceptable risk for an aircraft engine, which requires absolutely stable operation in flight. However, stall margin obtained in RBDO is 2.7% higher than the reference value, and the probability of success increases to 95% with the improved efficiency and weight. Therefore, RBDO of the axial compressor for aircraft engine can be a reliable design optimization method through consideration of unexpected disturbance of the flow conditions. [ABSTRACT FROM AUTHOR]

The stall margin of compressor could be improved effectively by rotor tip injection, and the periodic injection is commonly used in the research. The purpose of this work is to investigate the influence of injection frequency on the rotor stall margin. An unsteady CFD code was employed to simulate the flow field of the rotor with injections of different frequencies. Comparing the stall margin of the rotor with injections of different frequencies, it is shown that there is an optimal injection frequency, around which the rotor stability enhancement is the largest. When the injection frequency is away form the optimal frequency, the improvement in stable flow range decreases correspondingly. For the rotor in this paper, the optimal frequency was 1.5 times the frequency of tip leakage vortex (for short, TLV) fluctuation. Time-averaged loading distribution at 98.5% span indicates that the loading of the rotor near the leading edge is decreased through injection with the optimal frequency, and therefore, the stall could be delayed. [ABSTRACT FROM AUTHOR]

Tilting the blade sections to the flow direction (blade sweep) would increase the operating range of an axial compressor due to modifications in the pressure and velocity fields on the suction surface. On the other hand, blade tip gap, though finite, has great influence on the performance of a turbomachine. The present paper investigates the combined effect of these two factors on various flow characteristics in a low speed axial flow compressor. For this present study, nine computational domains were modeled; three rotor sweep configurations (0°, 20° and 30°) and for three different clearance levels for each rotor. Commercial CFD solver ANSYS CFX 11.0 is used for the simulations. Results indicated that tip chordline sweep is found to improve the stall margin of the compressor by modifying the suction surface boundary layer migration phenomenon. Diffusion Factor (DF) contours showed the severity of stalling with unswept rotor. For the swept rotors, the zones of high probable stall are less severe and they become less in size with increasing sweep. Increment in the tip gap is found to gradually affect the performance of unswept rotor, while the effect is very high for the two swept rotors for the earlier increments. As a minimum clearance is unavoidable, swept rotors suffer relatively higher deviation from the idealistic behavior than the unswept rotor due to tip clearance. [ABSTRACT FROM AUTHOR]

The present work performed a comprehensive investigation to find the effects of a dual-bleeding port recirculation channel on the aerodynamic performance of a single-stage transonic axial compressor, NASA Stage 37, and optimized the channel's configuration to enhance the operating stability of the compressor. The compressor's performance was examined using three parameters: The stall margin, adiabatic efficiency, and pressure ratio. Steady-state three-dimensional Reynolds-averaged Navier–Stokes analyses were performed to find the flow field and aerodynamic performance. The results showed that the addition of a bleeding channel increased the recirculation channel's stabilizing effect compared to the single-bleeding channel. Three design variables were selected for optimization through a parametric study, which was carried out to examine the influences of six geometric parameters on the channel's effectiveness. Surrogate-based design optimization was performed using the particle swarm optimization algorithm coupled with a surrogate model based on the radial basis neural network. The optimal design was found to increase the stall margin by 51.36% compared to the case without the recirculation channel with only 0.55% and 0.28% reductions in the peak adiabatic efficiency and maximum pressure ratio, respectively. [ABSTRACT FROM AUTHOR]

In the recent past, experimental studies have shown some advantages of blade lean and sweep in axial compressors. As most of the experimental results are combined with other features, it is difficult to determine the effect of individual parameters on the performance of the compressor. The present numerical studies are aimed at understanding the performance and three-dimensional flow pattern within and at the exit of swept and unswept rotors. Three rotors, namely, unswept, 20° forward swept, and 20° backward swept rotors, are analysed with a specific intention of understanding the three-dimensional flow pattern within the rotors and also the pattern of the blade boundary layer flow. The analysis was done using a fully three-dimensional viscous CFD code CFX-5. Results indicated a reduction in pressure rise with sweep. Backward sweep adversely affects the stall margin. Forward sweep changes the streamline pattern in such a way that the suction surface streamlines are deflected towards the hub and the pressure surface streamlines are deflected towards the casing. An opposite behaviour is observed in the backward swept rotors. High axial velocities reduce the secondary losses near the hub, resulting in a high pressure rise in forward swept rotor. [ABSTRACT FROM AUTHOR]

In the recent past, experimental studies have shown some advantages of blade lean and sweep in axial compressors. As most of the experimental results are combined with other features, it is difficult to determine the effect of individual parameters on the performance of the compressor. The present numerical studies are aimed at understanding the performance and three-dimensional flow pattern at the exit of swept and unswept rotors. Three rotors, namely; unswept, 200 forward swept and 200 backward swept rotors are analysed with a specific intention of understanding the pattern of the blade boundary layer flow. The analysis was done using a fully three-dimensional viscous CFD code CFX-5. Results indicated reduction in pressure rise with sweep. Backward sweep is detrimental as far as the performance near endwalls is considered. On the other hand total pressure loss in the wake in mid span region is less with backward sweep, which favours its application here. However, backward sweep adversely affects the stall margin. The ability of the forward sweep to deflect the streamlines towards hub gets diminished at low flow rates. Forward sweep changes the streamline pattern in such a way that the suction surface streamlines are deflected towards the hub and the pressure surface streamlines are deflected towards the casing. An opposite behaviour is observed in backward swept rotors. [ABSTRACT FROM AUTHOR]

On the basis of the test results of discrete axial and blade angle slot casing treatment, a new type of casing treatment was designed for a subsonic axial flow compressor rotor by optimizing various geometry parameters. To obtain a wide operating range and to minimize penalties in terms of isentropic efficiency, seven compressor configurations incorporating casing treatments of 0, 16.6, 33.3, 50, 66.6, 83.3, and 100 per cent rotor exposures were experimentally investigated. The results showed that significant improvements in stall margin are possible in all exposures and insignificant isentropic efficiency sacrifices are recorded in some exposures. Nearly 21.43 per cent stall margin improvement in terms of the corrected mass flow-rate was achieved with 33.3 per cent rotor blade tip axial chord exposure. The compressor built with 16.6 per cent rotor exposure was the best configuration in terms of maximum isentropic efficiency gain.The second issue of the paper was to offer a contribution to the understanding of the physical mechanism by which bend-skewed slot-casing treatment improves stall margin under subsonic conditions. By applying a concept similar to ‘Domain scaling’ approach (as often used in multistage turbomachinery flow-fields) to the interface between the rotor blade passage and end-wall treatments, a time-dependent three-dimensional numerical simulation was performed for the subsonic axial-flow compressor rotor with bend-skewed slot-casing treatment. The numerical results agreed well with the available experimental results. Detailed analyses of the coupled flow through bend-skewed slot-casing treatment and rotor blade passage under subsonic conditions led to some preliminary conclusions as to the flow physics involved in the stall margin improvements afforded by the use of bend-skewed slot-casing treatment. [ABSTRACT FROM AUTHOR]

In multi-spool engines, the downstream compressor experiences a rotating inlet distortion if rotating stall occurs in upstream compressor, which may induce the instability of the whole compressor. In this paper, the compressor dynamic behavior is the major focus. The experiment was carried out on a single-stage low-speed axial-flow compressor. A rotating distortion generator equipped with different distortion sector(s) was designed to produce different rotating inlet distortion clockwise or counterclockwise with up to 100% of the compressor design speed. The distortion sector can be installed single or in some combination such as four sectors together. Three types of distortion sector/combination are used in the research work, which are single 30 degree sector, four 30 degree sectors and single 120 degree sector. It is found that the total pressure loss caused by rotating sector(s) increases when the distortion speed rises. For co-rotating distortions, all the three types of inlet distortion exhibited a peak in stall margin degradation when the distortion speed corresponded to roughly 50% of rotor speed. The two-dimensional numerical simulations of the compressor flow field clearly show the propagation of the disturbances trigged by the distortion sector(s). [ABSTRACT FROM AUTHOR]

The design of an axial-flow compressor stage for subcritical Mach numbers has been formulated as a non-linear multi-objective mathematical programming problem with the objective of minimizing the aerodynamic losses and the weight of the stage, while maximizing the compressor''s stall margin. Aerodynamic as well as mechanical constraints are considered in the optimization solution. The prediction model for estimating the performance characteristics, such as efficiency, weight and stall margin, of the compressor stage is presented. The present design optimization procedure can be applied to a multi-stage compressor. [Copyright &y& Elsevier]

Axial slot casing treatment is a common method to extend the stall margin of a compressor. Based on the mechanism of unsteady flow control, this paper redesigns axial slots with large circumferential interval and small opening area. To test the effect of this axial slot structure, unsteady numerical simulations were carried out with different slot areas and circumferential intervals. The results show that this novel axial slot casing treatment can significantly improve compressor stall margin. Meanwhile, compared with the traditional axial slot, the efficiency loss is greatly reduced. The flow field analysis shows that the new axial slot structure proposed in this paper can suppress the development of tip leakage vortex and unsteadiness in the tip region at the near stall condition through decreasing the tip loading periodically. Moreover, we find that the slot area is proportional to the improvement of stability margin. Under the same slot area, an excessive number of slots is not conducive to the improvement of the stability margin. [ABSTRACT FROM AUTHOR]

Numerical simulations have been performed to study the effect of the circumferential single-grooved casing treatment (CT) at multiple locations on the tip-flow stability and the corresponding control mechanism at three tip-clearance-size (TCS) schemes in a transonic axial flow compressor rotor. The results show that the CT is more efficient when its groove is located from 10% to 40% tip axial chord, and G2 (located at near 20% tip axial chord) is the best CT scheme in terms of stall-margin improvement for the three TCS schemes. For effective CTs, the tip-leakage-flow (TLF) intensity, entropy generation and tip-flow blockage are reduced, which makes the interface between TLF and mainstream move downstream. A quantitative analysis of the relative inlet flow angle indicates that the reduction of flow incidence angle is not necessary to improve the flow stability for this transonic rotor. The control mechanism may be different for different TCS schemes due to the distinction of the stall inception process. For a better application of CT, the blade tip profile should be further modified by using an optimization method to adjust the shock position and strength during the design of a more efficient CT. [ABSTRACT FROM AUTHOR]

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]

To explain the effect of tip leakage flow on the performance of an axial-flow transonic compressor, the compressors with different rotor tip clearances were studied numerically. The results show that as the rotor tip clearance increases, the leakage flow intensity is increased, the shock wave position is moved backward, and the interaction between the tip leakage vortex and shock wave is intensified, while that between the boundary layer and shock wave is weakened. Most of all, the stall mechanisms of the compressors with varying rotor tip clearances are different. The clearance leakage flow is the main cause of the rotating stall under large rotor tip clearance. However, the stall form for the compressor with half of the designed tip clearance is caused by the joint action of the rotor tip stall caused by the leakage flow spillage at the blade leading edge and the whole blade span stall caused by the separation of the boundary layer of the rotor and the stator passage. Within the investigated varied range, when the rotor tip clearance size is half of the design, the compressor performance is improved best, and the peak efficiency and stall margin are increased by 0.2% and 3.5%, respectively. [ABSTRACT FROM AUTHOR]

A casing treatment using inclined oblique slots (INOS) is proposed to improve the stability of the single-stage transonic axial compressor, NASA Stage 37, during operation. The slots are installed on the casing of the rotor blades. The aerodynamic performance was estimated using three-dimensional steady Reynolds-Averaged Navier-Stokes analysis. The results showed that the slots effectively increased the stall margin of the compressor with slight reductions in the pressure ratio and adiabatic efficiency. Three geometric parameters were tested in a parametric study. A single-objective optimization to maximize the stall margin was carried out using a Genetic Algorithm coupled with a surrogate model created by a radial basis neural network. The optimized design increased the stall margin by 37.1% compared to that of the smooth casing with little impacts on the efficiency and pressure ratio. [ABSTRACT FROM AUTHOR]

Downsizing in engine size is pushing the automotive industry to operate compressors at low mass flow rate. However, the operation of turbocharger centrifugal compressor at low mass flow rate leads to fluid flow instabilities such as stall. To reduce flow instability, surface roughness is employed as a passive flow control method. This paper evaluates the effect of surface roughness on a turbocharger centrifugal compressor performance. A realistic validation of SRV2-O compressor stage designed and developed by German Aerospace Center (DLR) is achieved from comparison with the experimental data. In the first part, numerical simulations have been performed from stall to choke to study the overall performance variation at design conditions: 2.55 kg/s mass flow rate and rotational speed of 50,000 rpm. In second part, surface roughness of magnitude range 0–200 μm has been applied on the diffuser shroud to control flow instability. It was found that completely rough regime showed effective quantitative results in controlling stall phenomena, which results in increases of operating range from 16% to 18% and stall margin from 5.62% to 7.98%. Surface roughness as a passive flow control method to reduce flow instability in the diffuser section is the novelty of this research. Keeping in view the effects of surface roughness, it will help the turbocharger manufacturers to reduce the flow instabilities in the compressor with ease and improve the overall performance. [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 provide a guideline for the optimal selection of forward sweep height (FSH) in terms of operating range and performance, numerical simulations are utilized to investigate the effects of different FSHs on a new-type low-reaction ultra-high-load compressor rotor. The results illustrate that it is an effective measure to amplify stall margin improvement (SMI) by adequately increasing FSH (no more than 50% span). Nonetheless, when FSH exceeds 50% span, there appears a drop in SMI. Both total pressure ratio (TPR) and peak efficiency (PE) demonstrate a continual downward trend with increasing FSH. It is noteworthy that compared with the unswept rotor, forward sweep enhances stall margin in all rotors but reduces PE and the corresponding TPR when FSH is more than 50% span. There exists an optimal FSH (50% span) that could maximize SMI (15.12%) and simultaneously achieve a negligible performance change at PE condition. A deep insight into the flow field reveals that as FSH increases, the shock gradually migrates downstream and the separation bubble on suction side shrinks at PE point and shortens in streamwise direction near stall. If FSH continually increases above 50% span, SMI brought by the change in shock structure and tip leakage flow will drop. [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]

This paper analyzes the health and performance of the 12-stage axial compressor of the TV3-117VM/VMA turboshaft operated in a desert environment. The results of the dimensional control of 4800 worn blades are analyzed to model the wear process. Operational experience and two-phase flow simulations are used to assess the effectiveness of an inlet particle separator. Numerical modal analysis is performed to generate the Campbell diagram of the worn blades and identify resonant blade vibrations which can lead to high cycle fatigue (HCF): mode 7 engine order 30 in the first stage and mode 8 engine order 60 in the fourth. It is also shown that the gradual loss of the stall margin over time determines the serviceability limits of compressor blades. In particular, the chord wear of sixth-stage blades as high as 6.19 mm results in a reduction of the stall margin by 15–17% and a permanent stall at 770–790 flight hours. In addition, recommendations setting out go/no-go criteria are made to maintenance and repair organizations. [ABSTRACT FROM AUTHOR]