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Abstract This paper presents results of aerodynamic enhancement of axial compressors utilizing optimum natural aspiration through a circumferential slot made within the casing wall upstream the rotor blades row. The method of investigation is based on numerical simulation of flow field. Geometries of the slot walls were optimized to find the maximum stall margin and pressure ratio of the compressor. The optimum case is accompanied by boosting more momentum within the blades tip region through the slot. Consequently, tip leakage vortex flow weakens and compressor stability margin increases. Final results showed that in comparison to the untreated casing, the optimum geometry of the slot causes the total pressure ratio and stall margin to increase by 4.2% and 3%, respectively. This state-of-the-art technique is simple and economic which can be easily implemented in practical cases while the compressor is exposed to commencement of flow instabilities in the forms of the rotating stall or surge phenomena.

This paper presents results of aerodynamic enhancement of axial compressors utilizing optimum natural aspiration through a circumferential slot made within the casing wall upstream the rotor blades row. The method of investigation is based on numerical simulation of flow field. Geometries of the slot walls were optimized to find the maximum stall margin and pressure ratio of the compressor. The optimum case is accompanied by boosting more momentum within the blades tip region through the slot. Consequently, tip leakage vortex flow weakens and compressor stability margin increases. Final results showed that in comparison to the untreated casing, the optimum geometry of the slot causes the total pressure ratio and stall margin to increase by 4.2% and 3%, respectively. This state-of-the-art technique is simple and economic which can be easily implemented in practical cases while the compressor is exposed to commencement of flow instabilities in the forms of the rotating stall or surge phenomena. [ABSTRACT FROM AUTHOR]

This study presents a design optimization of a circumferential feedback channel on the shroud casing between rotor trailing and leading edges to maximize the aerodynamic performance of a single-stage transonic axial compressor, NASA Stage 37, using three-dimensional Reynolds-averaged Navier–Stokes analysis (RANS) with the k-ε turbulence model. The location of the injection and bleeding ports, angles of injection, and bleeding were the most influential factors on the aerodynamic performance and were selected as the design variables. A hybrid genetic algorithm and particle swarm optimization algorithm coupled with three different surrogate models were utilized to maximize the peak adiabatic efficiency and stall margin of the transonic axial compressor. The optimization results for the circumferential feedback channel indicated that an increase in stall margin was achievable with a slightly reduced peak efficiency in comparison with that of the smooth casing. [ABSTRACT FROM AUTHOR]

The presence of bleed in an aero engine's compressor can significantly impact its flow characteristics and contribute to rotating instability. This study focuses on the impact of a typical cavity bleed structure on the internal flow characteristics of a compressor, specifically its circumferential non-uniformity and aerodynamic stability. A numerical simulation involving multiple flow passages was conducted on the transonic high-pressure rotor of the E3 compressor, considering its typical bleed structure. The study delves deep into the non-uniform flow characteristics and the mechanisms behind their generation in the compressor flow field. Furthermore, the influence of bleed on the rotating instability of the compressor is explored by comparing changes in compressor instability and adiabatic efficiency under uniform and non-uniform flow field conditions. The findings indicate that the axial position of the cavity bleed structure plays a crucial role in influencing key parameters such as rotor stall margin, peak efficiency, and total pressure ratio under near-stall conditions. The circumferential non-uniformity, resulting from the presence of the cavity bleed, intensifies with higher bleed air flow rates. For the upstream bleed configuration applied to the rotor, with a total bleed rate of 5%, the maximum variation in absolute flow angle at the inlet of different rotor channels can reach up to 1°. Additionally, the maximum difference in inlet flow coefficient can reach 0.0392. These findings demonstrate that the non-uniformity caused by the typical bleed structure leads to a loss in stall margin for the rotor when compared to a uniform flow field scheme. [ABSTRACT FROM AUTHOR]

This paper presents the effect of an inclined casing groove on the aerodynamic performance of a single-stage transonic axial compressor, NASA Stage 38, using three-dimensional Reynold-averaged-Navier-Stokes equations with the k-ε turbulence model. The research was carried out to examine the effects of four casing groove parameters: angle, width, depth, and location. Validation of a numerical model for a single-stage transonic axial compressor was conducted to evaluate the computational fluid dynamics method. Most of the simulations showed positive results with an increase in stall margin, adiabatic efficiency, and total pressure ratio, in which the maximum stall margin, adiabatic efficiency, and total pressure ratio can be raised by 87.09%, 0.13%, and 1.57%, respectively, as compared to the smooth casing case. [ABSTRACT FROM AUTHOR]

In this study, the stator geometry of a single-stage transonic fan was optimized to improve its aerodynamic performance. Steady and unsteady simulations were performed on the transonic fan stage to maximize the efficiency and stall margin. In a transonic fan stage, the rotor blade generates different shock waves at various operating points. When the operating condition is changed from the design point to the stall point, the passage shock wave moves upstream, leading to a change in the location of flow separation on the suction surface of the rotor blade toward the upstream. This results in a larger rotor wake near the stall and complicates the stator inlet flow after interaction between the rotor and stator blade. In this study, an optimal geometry of the stator blade is proposed to improve the aerodynamic performance and stall margin of the fan stage. The central composite face-centered method as the design of experiment was used to sample the test cases, and the response surface method was applied to find the optimal case. The transient blade row-time transformation model was applied as an unsteady simulation approach. The optimized fan was found to increase the efficiency at the near-stall point and the stall margin by 0.24 % and 0.36 %, respectively. [ABSTRACT FROM AUTHOR]

In this work, a circumferential groove on the shroud was applied to a centrifugal compressor and optimized to improve the stability of the compressor. To reduce the number of design variables before optimization, the position of the casing groove that maximizes the stall margin of the compressor was firstly selected. The width and height of the casing groove were optimized using the RSM (Response Surface Method) and the full factorial design of experiments to increase the stall margin further. It was observed that the optimized case with the casing groove can increase the stall margin by 7.0% but the adiabatic efficiency at the design condition decreases by 0.99%, in comparison to the reference case without the groove. Finally, the flow field of both cases with and without the optimized casing groove was compared to each other to analyze effects of the casing groove on the stall margin and design efficiency. The casing groove decreases the leakage flow across the impeller tip and weakens the tip leakage vortex, consequently decreasing the blockage near the shroud and improving the stall margin. However, it was found that the spillage of the flow from the casing groove increases the mixing loss near the shroud. [ABSTRACT FROM AUTHOR]

The aim of the paper is to explore the influence of the reversed blade-angle slot casing treatment (RBSCT) and its axial deflected angle (ADA) on the compressor performance and stability, and to reveal the mechanism that the change in ADA of the RBSCT influences the effect to broaden the compressor stable working range. The NASA Rotor 35 is used as the object of the investigation, and four RBSCTs with ADA of −15°, −30°, −45° and −60° are designed and investigated by unsteady numerical simulation. The results show that as the absolute value of the axial deflected angle increases, the capacity to improve the compressor stability of the RBSCT increases and then decreases. The unsteadiness of the injection and suction flows formed by the reversed blade angle slot plays an important role in the removal of the low-velocity zone. When ADA is −30°, the unsteadiness amplitude of the injection and suction flows is significantly higher than those of the other three. Consequently, the RBSCT with −30° ADA obtains the maximum stall margin improvement of 17.41% and the maximum design point efficiency improvement of 1.06% among the four RBSCTs. [ABSTRACT FROM AUTHOR]

The effects of root fillet on the flow behavior of high loading compressor rotor tends to be much more crucial in practice, and it's necessary to explore the internal relations between the geometric effects of root fillet and the flow behaviors of rotor blade. Therefore, eight types of root fillet with different radius were designed and installed around the blade root of NASA Rotor67. With the aids of fillet, the corner separation near suction side of blade root has been suppressed significantly in that the root fillet reconstructs the circumferential bending distributon of the suction-side curve from leading edge to trailing edge, and reduces the genmetric turning angle in the latter part of root section near trailing edge. However, apart from the improvement of corner flow characteristic caused by root fillet, both the tip flow deterioration and the decrease of stall margin occur in the new rotors, which indicates an indirect correlation between tip flow characteristic and root fillet exists indeed in the three-dimensional flowfields of transonic rotor. Actually, by means of the new radial pressure equilibrium affected by root fillet, a larger radius of root fillet contributes to much larger blade loading and stronger leakage flow in tip region of compressor rotor. As a result, a monotonic decrease of stall margin was present in the transonic rotor with increase of the root fillet radius. Subsequently, the positive bending of blade tip was introduced to deal with the negative effect caused by the root fillet indirectly. Combined with the effects of root fillet and positive tip-bending on the radial pressure equilibrium existing in channels, both the radial and streamwise loading distributions tend to be much more reasonable in new rotors, and the static pressure difference in former 1/3 chord of blade tip has decreased clearly which benefits to reduce the strength of leakage flow in tip region. Therefore, the flow deterioration in tip region of transonic rotor induced by root fillet has been well suppressed, with an obvious improvement of overall performance occurring in new rotors. [ABSTRACT FROM AUTHOR]

Performance of a NASA 37 transonic compressor stage with a symmetric Airfoil Vortex Generator (AVG) positioned in the upstream of rotor is investigated through a numerical simulation. Steady state flow simulations were performed with k-ω SST turbulence model in ANSYS CFX flow solver. Grid independence study for the baseline compressor was performed besides CFD predicted performance characteristics were validated against available experimental data. The results of parametric study shows that AVG is able to improve compressor stall margin with a penalty on the stage efficiency. Incorporation of an AVG on the casing surface upstream of rotor reduces its specific work capacity, incurs flow losses and decreases stage peak efficiency by 3.34%. However few AVG configurations have negligible reduction, nearly 0.27-0.4%, in stage efficiency compared to baseline case. At near stall operating point, an AVG affects both rotor and stator flow field. An AVG under loads rotor tip region and decreases tip leakage mass flow rate, leading to the reduction of flow blockage. The flow swallowing capacity of the stator passage increases by diffusing the flow considerably to a low velocity. Consequently, AVG is able to increase the compressor stage stall margin by 8.06%. [ABSTRACT FROM AUTHOR]

Counter-rotating axial-flow compressor (CRAC) technology is an important technical way to improve the thrust-toweight ratio of aero-engines. To explore the stall margin improvement potential of axial slot casing treatment (ASCT) in the CRAC, a two-stage CRAC is selected to investigate the influence of axial position and radial deflection angle of slots on the stabilization capability of ASCT by numerical method. The paper also revealed the stability expansion mechanism of ASCT in the CRAC and the effect of casing treatment on the rotor/rotor interference effect between the front and rear rotors. Results show that the axial position of slots has little effect on stall margin improvement capacity of ASCT but has a great influence on peak efficiency. Increasing the radial deflection angle of slots to further improve the stabilization ability of ASCT is not remarkable, but it is beneficial to decline the efficiency penalty. The suppression of tip leakage flow by slots and the transport of low-energy fluid near the blade tip can delay the occurrence of the stall, and the maximum stall margin is improved by about 10.75%. Besides, the ASCT decreases the unsteady interference effect between the rotors and inhibits the potential flow effect of the downstream blade passage and improves the inlet flow field quality of the rear rotors. [ABSTRACT FROM AUTHOR]

Counter-rotating axial-flow compressor (CRAC) is a promising technology to increase the thrust-to-weight ratio of aero-engines. Self-recirculating casing treatment (SRCT) is selected to explore its stability enhancement capacity in the CRAC. In the present work, an effective SRCT is designed by the design of experiment method, and the stability improvement potential of the SRCT and its mechanism of stall margin enhancement are studied. Results show that the optimal SRCT scheme increases the stall margin by about 7.73% and without remarkable peak efficiency loss, and the overall performance of the compressor is enhanced at the near-stall condition. The SRCT improves the quality of the tip flow field by sucking out low-energy fluid and restrains the spillage of the tip leakage flow by the jet effect from the injection port. In addition, the SRCT also reduces the unsteady interference between the adjacent rotors. The effect of speed ratio on the effectiveness of SRCT is further investigated, and results indicate that the SRCT increases the stall margin of the CRAC by about 4.13–5.80% at some off-design speed ratio conditions. The speed ratio affects the first stall stage of the CRAC, thus affecting the effectiveness of the SRCT.

This paper presents a new investigation of the recirculation channel combined with bleeding channels, where the recirculation airflow recirculates air from downstream high-pressure region to an upstream through a recirculation channel located on shroud surface in rotor domain and the bleeding airflow through 36 bleeding channels sited on the recirculation channels ejects airflow to the by-pass channel in a single-stage transonic axial compressor used in turbofan engine. The object of this investigation is to reduce the vortex flow in the tip leakage region by finding the effects of recirculation and bleeding channels design on aeromechanic performance of a single-stage axial compressor, NASA Stage 37. A parametric study of the recirculation and bleeding channels were performed using various geometric parameters related to the locations and shapes of the recirculation and bleeding channels to find the effects of recirculation and bleeding channels design on aeromechanic performance, such as pressure ratio, efficiency, stall margin, blade tip deformation and Von-Mises stress. The numerical results showed an increase in aerodynamic performance and decrease in blade tip deformation with recirculation-bleeding channels, compared to the smooth casing. [ABSTRACT FROM AUTHOR]

Counter-rotating axial-flow compressor (CRAC) is a promising technology to increase the thrust-to-weight ratio of aero-engines. Self-recirculating casing treatment (SRCT) is selected to explore its stability enhancement capacity in the CRAC. In the present work, an effective SRCT is designed by the design of experiment method, and the stability improvement potential of the SRCT and its mechanism of stall margin enhancement are studied. Results show that the optimal SRCT scheme increases the stall margin by about 7.73% and without remarkable peak efficiency loss, and the overall performance of the compressor is enhanced at the near-stall condition. The SRCT improves the quality of the tip flow field by sucking out low-energy fluid and restrains the spillage of the tip leakage flow by the jet effect from the injection port. In addition, the SRCT also reduces the unsteady interference between the adjacent rotors. The effect of speed ratio on the effectiveness of SRCT is further investigated, and results indicate that the SRCT increases the stall margin of the CRAC by about 4.13–5.80% at some off-design speed ratio conditions. The speed ratio affects the first stall stage of the CRAC, thus affecting the effectiveness of the SRCT. [ABSTRACT FROM AUTHOR]

Foam metal is a foam-like substance with a high porosity; it has been used in flow control, vibration abatement, and acoustic absorption, mainly based on its physical properties. The aim of the current paper is to investigate the effect of foam metal casing treatments on the stability and acoustic level of a low-speed axial flow compressor. The experimental results show that the casing treatment improves the stall margin by 14.9%, without any efficiency loss. In terms of noise, the SPL of the tonal noise at the third order of BPF is reduced by 3.2 dB, while the SPL of the broadband noise is reduced up to 2.4 dB. The comparison in evolutions of the tip structure in a smooth casing condition and with a casing treatment indicates that the casing treatment affects the origination and the development of the tip leakage vortex. The working mechanism is also discussed. [ABSTRACT FROM AUTHOR]

Single-channel unsteady numerical simulation was carried out on NASA Rotor 35 to study the influence of radial skewed angle of reversed blade-angle slot casing treatment (RBSCT) on the aerodynamic performance and stall margin of a transonic compressor. Moreover, the influence mechanism is explained by detailed flow field analysis. The radial skewed angles were set to +0°, +30°, +60°, and +75° in the research. The calculated result shows that stall margin improvement (SMI) generated by the slots is increased when the radial skewed angle is gradually increased. The SMI of 17.47% for the slots with +75° radial skewed angle is the biggest among the four RBSCTs. As for the design efficiency improvement (DEI), it is increased first and then decreased with the radial skewed angle increasing. RBSCT with +60°radial skewed angle achieves the greatest DEI of 1.11%. The flow field analysis shows that the radial momentum of the injected and sucked flows is improved with the increasing of radial skewed angle. The improvement can reduce the relative airflow angle of mainstream near the stall condition. Furthermore, the excessive radial skewed angle will increase the flow losses in the slots. Consequently, the design efficiency is reduced at +75°. [ABSTRACT FROM AUTHOR]

The compressor is a critical component that determines the aerodynamic stability of an aero-engine. Total pressure inlet distortion decreases the thrust and shrinks the stability margin, thus inducing severe performance degradation or even flameout. Generally, tip air injection is used to reduce the adverse influence of total pressure inlet distortion on the aerodynamic stability. In the present work, an experimental investigation on the effects of tip air injection on the stability of a two-stage low-speed axial compressor with total pressure inlet distortion was carried out. A flat baffle generated the total pressure distortion at the inlet of the compressor. The stall margin of the compressor was reduced significantly by the total pressure distortion. When the dimensionless insertion depth of the flat baffle was 0.45, the stall margin decreased to 11.4%. Under the total pressure inlet distortion, tip air injection effectively improved the distortion resistance capability of the compressor. The circumferential layout of the nozzle played a critical role in the stability expansion effect of tip air injection under the inlet flow condition of the total pressure distortion. The modal wave disturbance was likely to occur in the distortion-affected region (the low-pressure region and the mixing region). Tip air injection did not inhibit the generation of the modal wave but restrained the development of the modal wave into the stall cell. It improved the low-speed compressor's tolerance to the modal wave and allowed a higher amplitude modal wave to occur. [ABSTRACT FROM AUTHOR]

To deeply understand the hub leakage flow and its influence on the aerodynamic performance and flow behaviors of a small-scale transonic axial compressor, variations of the performance and the flow field of the compressor with different hub clearance sizes and clearance shapes were numerically analyzed. The results indicate that the hub clearance size has remarkable impacts on the overall performance of the compressor. With the increase of the hub clearance, the intensity of the hub leakage flow increases, resulting in more intense flow blockage near the stator hub, which reduces the compressor efficiency. However, the flow field near the blade mid-span is modified due to the more convergent flow as the reduced effective flow area caused by the passage blockage, and the flow separation range is narrowed, thus the flow stability of the compressor is enhanced. On this basis, two kinds of non-uniform clearance cases of expanding clearance and shrinking clearance with the same circumferential leakage area as the design clearance were investigated. The occurrence position of the double leakage flow which is closely connected with the flow loss and blockage is shifted backward by the expanding clearance, the flow capacity near the stator hub is enhanced, and the unsteady fluctuation intensity of the flow field is attenuated but fluctuation frequency remains. Similarly, the modification of the stator blade root flow field may result in the reduction of stall margin. The effect of the shrinking clearance on compressor performance is opposite to that of the expanding clearance, which reduces the peak efficiency and delays the stall inception. [ABSTRACT FROM AUTHOR]

• Proper injection opening position of the recirculation can broaden the stall margin. • Discussing the impact of the recirculation injection opening position by parameterization. • Analyzing the stabilizing mechanism of the recirculation applied to the diffuser. • Quantitative analysis of loss distribution is discussed in detail. Pressure ratio, efficiency, and the stability operating range are the major criteria used to assess a centrifugal compressor. In this paper, a high pressure ratio single-stage centrifugal compressor is considered as a research object. A parametric investigation of recirculating flow control technology is used to treat the casing of the vaned diffuser in order to improve the compressor stall margin. The maximum achievable improvement in the comprehensive stall margin is 3.90%. The paper discusses how different injection configurations (circumferential and radial distributions) of the recirculation structure in the diffuser impact the flow field. The mechanism of influence of the injection opening position on the stall margin is explored. By implementing this technology, a reduction in the inlet Mach number at a high spanwise position of the diffuser, results in an increased critical incidence angle. Additionally, the technology effectively reduces flow separation at the throat of the diffuser and increased the flow area within this region. A comprehensive analysis is conducted to investigate the mechanisms underlying the influence of injection opening position on different energy loss categories in diffuser. A detailed discussion on the impact of casing treatment on high pressure ratio centrifugal compressors is provided, focusing on two aspects: expanding stall margin and efficiency loss. [ABSTRACT FROM AUTHOR]

Compressor design is always a prevalent research issue as nowadays the compressor demands higher performance. Designing afresh a new compressor of higher aerodynamic performance will cost a tremendous amount of time and expense. Consequently, the wiser method of acquiring a higher performance compressor is to optimize a current compressor, discussed in this paper. First, a detailed numerical simulation of al compressor is done. A parametric rotor model is created and the variations of lean and swept parameters are limited. Then, an optimization calculation is carried out based on the combination of an artificial neural network and a genetic algorithm. The optimization iteration is started at three specified working points – a choke point, a design point and a near stall point – to acquire a comprehensive improved solution. A novel penalty function considering both pressure ratio and efficiency is constructed to find the optimal model. Finally, a detailed review is made of the original and the optimized compressor. The computational results show that the isentropic efficiency is successfully increased and the stall margin is remarkably improved. Ultimately, a detailed vortex structure analysis with increasing outlet pressure is discussed. This helps the performance increment mechanism to be understood further.

The change of the blade tip clearance size has an important impact on the performance of the compressor. Considering that the performance curve of the compressor is often limited by surge and stall boundaries, this paper used the numerical simulation method to investigate the influence mechanism of the blade tip clearance size change on the stall margin of transonic axial flow compressor rotor. By mathematically decomposing the calculation formula of the stall margin of rotor, the approximate calculation formula of the change of rotor's stall margin was obtained. Then, the detailed quantitative analysis of the factors that affect the rotor's stall margin was carried out, the influence weights of various factors on the rotor's stall margin was also obtained. Finally, the physical mechanism of the change of the rotor's performance parameters was obtained by the analysis of rotor tip flow field after the blade tip clearance size change. [ABSTRACT FROM AUTHOR]

Engine size and weight optimization have always been high-priority design objectives for designers. Compressors occupy a relatively large part of the gas turbine engine. Owing to the adverse pressure gradient in the compressor, achieving the required pressure ratio within fewer stages has been a challenging task for compressor designers. Tandem blading is one of the novel concepts, which could be used to increase the pressure ratio by means of higher flow turning through the blade passages. This paper presents the performance characteristics of a tandem stage based on results from experiments and numerical analyses. The investigation is further extended to analyze the effect of a radial hub and tip distortion on the performance of the tandem stage. The experimental results are very well supported with some interesting numerical results, particularly near the hub and tip region. It is observed that the tandem stage demonstrates higher pressure rise and stall margin under clean inflow. The tandem stage is also observed to be more sensitive to radial distortion leading to a significant loss in the total pressure and the stall margin. [ABSTRACT FROM AUTHOR]

Tip leakage flow has become one of the major triggers for rotating stall in tip region of high loading transonic compressor rotors. Comparing with active flow control method, it's wise to use blade tip modification to enlarge the stable operating range of rotor. Therefore, three pressure-side winglets with the maximum width of 2.0, 2.5 and 3.0 times of the baseline rotor, are designed and surrounded the blade tip of NASA rotor 37, and the three new rotors are named as RPW1, RPW2, and RPW3 respectively. The numerical results show that the width of pressure-side winglet has significant influence on the stall margin and the minimum throttling massflow of rotor, while it produces less effect on the choking massflow and the peak efficiency of new rotors. As the width of the pressure-side winglet increases from new rotor RPW1 to RPW3, the strength of leakage massflow has been attenuated dramatically and a reduction of 20% in leakage massflow rate has appeared in the new rotor RPW3. By contrast, the extended blade tip caused by winglet has not introduced much more aerodynamic losses in tip region of rotor, and the new rotors with different width of pressure-side winglet have the similar peak efficiency to the baseline. The new shape of the leakage channel over blade tip which replaces of the static pressure difference near blade tip has dominated the behavior of the leakage flow in tip gap. As both the new aerodynamic boundary and throat in tip gap have reshaped by the low-velocity flow near the solid wall of extended blade tip, the discharging velocity and massflow rate of leakage flow have been suppressed obviously in new rotors. In addition, the increasing inlet axial velocity at the entrance of new rotor has increased slightly as well, which is attributed to the less blockage in the tip region of new rotor. In consideration of the increased inlet axial velocity and the weakened leakage flow, the new rotor presents an appropriately linear increase of the stall margin when the width of pressure-side winglet increases, and has a nearly 15% increase in new rotor RPW3. [ABSTRACT FROM AUTHOR]

Splitter blades located between stator blades in a single-stage axial compressor were proposed and investigated in this work to find their effects on aerodynamic performance and operating stability. Aerodynamic performance of the compressor was evaluated using three-dimensional Reynolds-averaged Navier-Stokes equations using the k-turbulence model with a scalable wall function. The numerical results for the typical performance parameters without stator splitter blades were validated in comparison with experimental data. The numerical results of a parametric study using four geometric parameters (chord length, coverage angle, height and position) of the stator splitter blades showed that the operational stability of the single-stage axial compressor enhances remarkably using the stator splitter blades. The splitters were effective in suppressing flow separation in the stator domain of the compressor at near-stall condition which affects considerably the aerodynamic performance of the compressor.

Compressor design is always a prevalent research issue as nowadays the compressor demands higher performance. Designing afresh a new compressor of higher aerodynamic performance will cost a tremendous amount of time and expense. Consequently, the wiser method of acquiring a higher performance compressor is to optimize a current compressor, discussed in this paper. First, a detailed numerical simulation of al compressor is done. A parametric rotor model is created and the variations of lean and swept parameters are limited. Then, an optimization calculation is carried out based on the combination of an artificial neural network and a genetic algorithm. The optimization iteration is started at three specified working points – a choke point, a design point and a near stall point – to acquire a comprehensive improved solution. A novel penalty function considering both pressure ratio and efficiency is constructed to find the optimal model. Finally, a detailed review is made of the original and the optimized compressor. The computational results show that the isentropic efficiency is successfully increased and the stall margin is remarkably improved. Ultimately, a detailed vortex structure analysis with increasing outlet pressure is discussed. This helps the performance increment mechanism to be understood further. [ABSTRACT FROM AUTHOR]

In this paper, the influence of a shallow reversed slot-type casing treatment on the performance of a tip-critical transonic compressor has been numerically investigated. Firstly, the complex flow fields in the rotor tip region are studied in details. It shows the severe blockage induced by suction surface boundary separation triggers compressor stall at 100% design speed, while the blockage due to tip leakage vortex dominates at 80% design speed. Secondly, the mechanism of stability extension is presented at different rotating speeds. The casing treatment alleviates greatly the tip blockage by manipulating the tip leakage flow, accompanied by the redistribution of aerodynamic loading and mass flux. As a result, the casing treatment is more efficient for the blockage induced by tip leakage vortex (at 80% design speed). Further analysis of the pressure field and passage shock distribution demonstrates that the passage shock intensity and its location will affect the effectiveness of casing treatment. Finally, the instability characteristics of compressor with casing treatment are revealed. The numerical results reflect when the mass flow approaching the instability boundary, the stator passage blockage presumably is dominant for triggering the compressor stall. [ABSTRACT FROM AUTHOR]

Casing treatment is an efficient technique that is used to increase the compressor stall margin with a minor reduction in efficiency. The interaction between the blade passage and the groove is the main cause of the stall margin improvement by various researches. A numerical study is conducted on a single circumferential casing groove using NASA rotor 37 in the current study. The performance of the two circumferential groove models is studied by discretizing RANS 3D equations using a finite volume technique. The inception of the stall is predicted according to the criteria of convergence. Two models of the circumferential groove have been suggested and numerically tested. A single passage simulation is selected for the two models. The performance of the smooth casing and the two models are analyzed. Moreover, the stall margin, total pressure ratio, and peak adiabatic efficiency of the normal casing, and the two models are analyzed to determine the influence of the groove on the axial compressor performance and stability. Models 1 and 2 stall margins are enhanced by 6.62 % and 4.45% respectively. The adiabatic efficiency of model 1 and model 2 are decreased by 0.79 % and 1.08 % respectively. [ABSTRACT FROM AUTHOR]

Foam metal is a foam-like substance with a high porosity; it has been used in flow control, vibration abatement, and acoustic absorption, mainly based on its physical properties. The aim of the current paper is to investigate the effect of foam metal casing treatments on the stability and acoustic level of a low-speed axial flow compressor. The experimental results show that the casing treatment improves the stall margin by 14.9%, without any efficiency loss. In terms of noise, the SPL of the tonal noise at the third order of BPF is reduced by 3.2 dB, while the SPL of the broadband noise is reduced up to 2.4 dB. The comparison in evolutions of the tip structure in a smooth casing condition and with a casing treatment indicates that the casing treatment affects the origination and the development of the tip leakage vortex. The working mechanism is also discussed.

Numerical investigations are carried out in a single-stage subsonic axial flow compressor to unravel the influence of blade tip surface roughness on the tip leakage flow characteristics and hence the compressor performance. The studies were carried out at different tip clearance of 0.38τ, 0.77τ, 1.15τ, and 1.54τ and blade tip surface roughness of 0.31ε and 0.62ε. The tip clearance of 0.38τ with blade tip surface roughness of 0.62ε resulted in the highest stall margin and pressure rise of 20.3% and 4.3%, respectively. The compressor blade loading was found to be improved by 5.9% after incorporating the blade tip surface roughness. The isosurfaces of vorticity contour plotted using the Q-criterion showed the reduction in strength of the tip leakage vortex. The tip leakage trajectory was found to be shifted toward the suction surface of the blade for the blade tip with surface roughness. This positive alteration in the tip leakage flow structure led to the improved performance for the blade tip with surface roughness. [ABSTRACT FROM AUTHOR]

The demands to apply transonic centrifugal compressor have increased in the advanced gas turbine engines. Various techniques are used to increase the aerodynamic performance of the centrifugal compressor. The effects of the inclined leading edges in diffuser vanes of a transonic centrifugal compressor on the flow-field unsteadiness and noise generation are investigated by solving the compressible, three-dimensional, transient Navier–Stokes equations. Diffuser vanes with various inclination angles of the leading edge from shroud-to-hub and hub-to-shroud are numerically modeled. The results show that the hub-to-shroud inclined leading edge improves the compressor performance (2.6%), and the proper inclination angle is effective to increase the stall margin (3.88%). In addition, in this study, the transient pressure variations and radiated noise prediction at the design operating point of the compressor are emphasized. The influences of the inclined leading edges on the pressure waves were captured in time/space domain with different convective velocities. The pressure fluctuation spectra are calculated to investigate the tonal blade passing frequency (BPF) noise, and it is shown that the applied inclination angles in the diffuser blades are effective, not only to improve the aerodynamic performance and stall margin, but also to reduce the BPF noise (7.6 dB sound pressure level reduction). Moreover, it is found that the diffuser vanes with inclination angles could suppress the separation regions and eddy structures inside the passages of the diffuser, which results in reduction of the overall sound pressure level and the broadband noise radiated from the compressor. [ABSTRACT FROM AUTHOR]

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]

In the present study, the numerical analysis of blade tip geometry effect on the performance of a single-stage axial compressor has been the focus of attention. The studied geometries included a rotor with variable tip clearance. For the first model, the tip clearance increases as it moves toward the blade trailing edge. The tip clearance of the second model reduces as it approaches the trailing edge, whereas in the third model, the tip clearance remains constant. The results indicated that the tip clearance of the first sample, as the worst tip clearance case, creates a 10% reduction in the stall margin with respect to the third standard model, and the second sample tip clearance brings about a stall margin reduction of 4% efficiency with respect to the third standard model. Then, the effect of blade tip clearance geometry on outlet flow angle of the rotor was inspected. The results showed that in the first model, the outlet flow angle has the largest deviation than the third standard model and the second model performance places somewhere between the two other tip clearance geometries. Also evident from the result is that taking advantage of the variable blade tip clearance is not an appropriate method for improving compressor performance.

In this study, a spiral ring cavity (SRC) is proposed to improve the operational stability of a centrifugal compressor. Aerodynamic performance of the compressor was analyzed using three-dimensional Reynolds-averaged Navier-Stokes equations with the shear stress transport turbulence model. To prove the superiority of the proposed cavity, adiabatic efficiency and stall margin of an SRC were compared with those of the smooth casing and conventional cavities, i.e., ring cavity and discrete ring cavities. To find the optimum shape of the cavity, a parametric study was performed for the SRC using three geometric parameters, i.e., the axial location, width of the cavity port, and circumferential location of the SRC. The operation stability was sensitive to the width of cavity port and circumferential location. The SRC improved the stall margin considerably compared to other cavities. [ABSTRACT FROM AUTHOR]

The performance and stability of a low-bypass turbofan transonic axial compressor with a nonuniform inlet flow is a significant concern in recent times. In both military and commercial aircraft, serpentine ducts produce significant inlet swirl distortion. Moreover, the nonuniform inlet flow frequently acts upon aircraft gas-turbine engines causing deteriorating effects on the aircraft engine. High circumferential swirl flows and inlet flow angularity decrease the aerodynamic performance and the stall margin and increase the rotor blade loading. The current paper is aimed at the investigation of the flow field in the tip clearance region of low-bypass turbofan transonic compressor rotor under nonuniform circumferential flow conditions through numerical simulation using Ansys CFX. The mathematical models based on 1D Mean Line Code and Dynamic Turbine Engine Compressor Code (DYNTECC) are used to analyze the nonuniform inlet swirl flow of the compressor rotor. The mathematical model is limited to compute the multistage compressor characteristics for the compression system and the combustor of a turbine engine. For the single-stage swirl flow analysis current paper focuses on the CFD based results. The results based on CFD show that co-swirl patterns slightly improve the stability range of the compressor; counter-swirl flows diminish it. [ABSTRACT FROM AUTHOR]

In the current study, the influence of blade tip clearance in different stages of a three-stage compressor is investigated. Performance diagrams of compressor were verified against experiment when there is no change in the tip clearance, after which the effect of tip clearance for the cases, 1, 1.5 and 2 mm, in the first, second and third stages of rotor was studied. The results indicated that the impact of tip clearance increase did not have any effect on choked flow rate value in the first and second stages, and only the change in the third stage tip clearance reduced the choked flow rate. For the same tip clearance value, the highest compressor performance loss occurs in the case of applying the tip clearance in the third stage, which is also the final stage and is highly sensitive to tip clearance changes. Moreover, modifying the tip clearance is effective on the flow angle in the trailing edge and as the tip clearance increases, the same thing happens about the flow angle. The maximum value of flow angle changes by modifying the tip clearance belongs to the third stage of the compressor.

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.

Casing treatment is an effective way to improve the stable working margin of the compressor and plays an important role in ensuring the stable and safe work of the compressor. On the premise that the numerical results are in agreement with the experimental values, the paper study on a new type of the self-circulating casing treatment with double-bleed ports structure was carried out on the NASA rotor 37 by numerical simulation. It is found that compared with the self-circulating casing treatment with single-bleed port structure, the self-circulating casing treatment with double-bleed ports structure is more favorable for improving the stall margin of the rotor, and the improvement of the stall margin obtained at the rotor design speed is as high as 9.57%. Compared with the other two self-circulating casing treatments with single-bleed port structure, the improvement of the stall margin is increased by 37.90 and 49.53%, respectively. The self-circulating casing treatment with double-bleed ports structure has a significant effect on the enhancing stability, but at the same time, it causes a greater efficiency loss, which reduces the design point efficiency of the rotor by 0.34%. Through detailed analysis of the flow field, it is found that the self-circulating casing treatment improves the flow capacity of the blade tip region, weakens the blockage degree of the blade tip region, and reduces the blade tip load, thereby improving the stall margin of rotor. Compared with the self-circulating casing treatment with single-bleed port structure, the self-circulating casing treatment with double-bleed ports structure is more advantageous for delaying the flow separation of the suction surface, delaying the occurrence of the blade tip stall to make the stability of rotor more favorable, and obtained stall margin improvement is the greatest. [ABSTRACT FROM AUTHOR]

In this paper, a computational investigation of circumferential groove casing treatment in a highly-loaded low-reaction transonic compressor rotor is conducted, in which the stage reaction is significantly reduced due to a larger meridional contraction with respect to conventional transonic compressors. Steady computation at near-stall point is performed first to capture the stall inception of the rotor with smooth casing. Detailed observations, which mainly focus on the tip leakage flow behavior, obstruction and vortical structures in the tip region, determine the reason for the compressor stall. There is tip leakage vortex breakdown in the tip region. Moreover, it yields passage obstruction, and finally leads to the compressor stall. Then, attempts are made to investigate how the circumferential grooves can be applied for the compressor's stall margin enhancement without compromising efficiency. Three configurations are obtained and analyzed by changing axial position and the number of the circumferential grooves. The results of computational parametric study indicate the optimal location of the groove is near the leading edge and the downstream grooves combine their influence on the compressor's stabilization and performance in a cumulative manner. The optimal circumferential groove configuration produces an increase of 1% in total pressure ratio at the near-stall point and a gain of 3.7% in stall margin, without any penalty in efficiency. Furthermore, the impact the grooves will exert on the flow mechanisms between the grooves and the main flow is also considered. [ABSTRACT FROM AUTHOR]

Inlet flow non-uniformity, commonly known as inflow distortion, has been a long-standing problem in the history of gas turbine engines. Distortion can be present in the form of total pressure, total temperature or inflow incidence or any combinations of these. The search for better and robust performance requires engines that can sustain a large amount of inlet distortion without considerable loss in the thrust. In the present paper, the effect of total pressure distortion on a single-stage compressor and low bypass ratio fans are studied. Distortion near hub and tip in the form of step radial total pressure profiles is imposed at far upstream of the rotor leading edge. A systematic approach to qualitatively predict the performance maps in the presence of these distortions is discussed. Further, two extents of total pressure distortion are explored for constant inlet distortion intensity. Hub distortion is found to increase the stability margin, whereas tip distortion reduces it. On extending the distortion extent, hub distortion drastically reduces the stability margin, whereas a comparatively lower reduction in stability margin with tip distortion is observed. The critical distortion limit is observed by varying the inlet distortion extent. Also, it is found that downstream ducts in the bypass axial fan do not interact with the upstream fan. This can be exploited to perform independent simulations of the core engine from low bypass ratio fans. Hub distortion is found to drastically affect the duct performance owing to the presence of thicker upstream inlet boundary layer. [ABSTRACT FROM AUTHOR]

Casing treatment has been extensively used as a prudent passive flow control means to enhance the stall margin with a little decrease in efficiency. The shape of various grooves has a substantial influence on the controlling effect. In this study, research on the impact of rectangular, fillet and chamfer types rectangular grooves with different tip clearances gap on the performance of casing treatment is carried out using numerical simulation. The performance of different models of grooves with different tip clearance gaps on NASA rotor 37 is investigated by discretizing 3D RANS equations based on the finite volume method. A few numbers of casing grooves models are designed with various tip clearances gap and numerically tested. Rectangular circumferential grooves casing treatment (CGCT) shape and smooth wall casing performances are analyzed. Furthermore, the highest adiabatic efficiencies and stall margins of smooth wall casing, rectangular grooves and fillet, and chamfer types rectangular grooves are compared to evaluate the effect of shapes of grooves on the stability and performance of axial flow compressor with different tip gaps. The adiabatic efficiency of the models 1 ~ 6 decreased by 0.9%, 1.0%, 1.05%, 1.1%, 1.2%, and 1.15% respectively. While the stall margin is increased by 4.4%, 2.5 %, 2.2%, 1.75%, 1.7% and 2.1% for models 1 ~ 6 respectively. [ABSTRACT FROM AUTHOR]

Centrifugal compressors often suffer relatively low efficiency and a terrible operating range particularly due to the complex flow structure and intense impeller/diffuser interaction. Numerous studies have focused on improving the centrifugal compressor performance using many innovative ideas, such as the tandem impeller, which has become increasingly attractive due to its ability to achieve the flow control with no additional air supply configurations and control costs in compressor. However, few studies that attempted to the investigation of tandem impeller have been published until now and the results are always contradictory. To explore the potential of the tandem impeller to enhance the compressor performance and the underlying mechanism of the flow phenomena in the tandem impellers, this paper numerically investigated a high-pressure-ratio centrifugal compressor with several tandem impellers at off-design operating speeds. The results encouragingly demonstrate that the tandem impeller can achieve a performance enhancement over a wide range of operating conditions. Approximately 1.8% maximum enhancement in isentropic efficiency and 5.0% maximum enhancement in operating range are achieved with the inducer/exducer circumferential displacement of λ s = 25% and 50%, respectively. The observed stage performance gain of the tandem impellers decreases when the operating speed increases due to the increased inducer shock, increased wake losses, and deteriorated tandem impeller discharge flow uniformity. In addition, the tandem impeller can extend the impeller operating range particularly at low rotation speeds, which is found to be a result from the suppression of the low-momentum fluid radial movement. The results also indicate that the maximum flux capacity of the tandem impeller decreases due to the restriction of the inducer airfoil Kutta–Joukowsky condition. [ABSTRACT FROM AUTHOR]

The stability range of the gas turbine engine compressors is being challenged in the modern days due to the intention of increasing per stage maximum loading. Casing treatment has been widely adopted as a realistic passive flow control means to improve the stall margin with a slight decrease of efficiency at the same time by various grooves of which shape (location, angles and so on) has a significant influence on controlling effect. However, the influence of some details in grooves is ignored in most of the case that may impact the specific flow field in the groove. A research on the impact of chamfer and fillet corners on the performance of casing treatment was proceeded by numerical simulation in this paper. The performance of different models of grooves on NASA Rotor 37 was investigated by discretizing 3D RANS based on finite volume method. Firstly, steady simulations were performed on NASA rotor 37 for validation. The CFD results and experimental data for adiabatic efficiency and pressure ratio were in good agreement. According to convergence criteria, the initiation of the stall was predicted. Few numbers of circumferential grooves casing treatment (CGCT) models have been proposed and tested numerically. Rectangular CGCT shape and smooth wall casing performances were analyzed. Moreover, the highest adiabatic efficiencies and stall margins of smooth wall casing, rectangular grooves and rectangular grooves with chamfer and fillet corners shapes models were compared to evaluate the influence of the shape of grooves on the stability and performance on axial flow compressor. The rectangular circumferential casing grooves and rectangular grooves with chamfer and fillet corners enhanced significantly stall margin and an operational range of the transonic axial flow compressor but adiabatic efficiency was slightly decreased. [ABSTRACT FROM AUTHOR]

The stable operating range of the centrifugal compressor is limited by stall or surge at lower mass flow rates and by choke at higher mass flow rates. The type of diffuser used has a strong influence on the performance of a centrifugal compressor. This paper presents a numerical investigation of the effect of the leaned vanes in the low solidity vaned diffuser on the stable operating range of a centrifugal compressor. The vanes in the low solidity vaned diffuser were leaned in the direction of impeller rotation by reducing their inlet angle from hub to shroud uniformly. Four cases of leaned vaned diffusers were configured downstream of the same impeller with different lean magnitude. The compressor stage with 6° leaned vaned diffuser gave an improvement of 19.3 % and 12.2 % in the stall margin and the operating range, respectively, over the stage with low solidity vaned diffuser. [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.

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.

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.