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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.

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.

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.

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.

The present research investigates the effect of the location and the width of single shallow circumferential groove casing treatment on the flow field and the stability improvement of NASA Rotor 37 utilizing the help of computational fluid dynamics. At first, steady state simulation of Rotor37 was presented for smooth casing (without groove). Then, forty five various grooved casing were simulated and compared with the smooth casing. The results indicated that narrow grooves had slight effect on the adiabatic efficiency but as the width of the groove was increased, a decline in efficiency was observed. The investigation on the stall margin revealed that narrow grooves next to the leading edge could improve the stall margin by a reduction in the size of vortex breakdown zone. Medium-width grooves displayed an effective role in delaying the separation- produced by shock wave and boundary layer interaction- on the blade suction side near the casing. This type of grooves could improve the stall margin more than narrow grooves when located on the top of separation zone near the blade suction side. Wide grooves had negative effect on the stall margin and caused a significant drop in the efficiency and the total pressure ratio of the compressor.

Flow separations in the vaned diffuser play a critical role in the stall margin of centrifugal compressors. To extend the stable operating range of a highly loaded centrifugal compressor, a flow control technique of locating a slot near the vaned diffuser's leading edge is proposed. A parametric study of slot geometry variations, in terms of the slotted portion depth and length, is carried out to summarize the trend of compressor performance and to reveal the corresponding flow physics related to the slot geometries. The design methodology is concluded to provide a guideline for stall margin improvement by the slotted diffuser. Application of a well-designed slotted diffuser shows that the flow separations inside the diffuser passages are suppressed at low flow rates, and the stall margin is increased without performance deterioration. The leakage flow passing through the slotted portion intensifies the passage vortex at the hub/suction surface corner, and decreases the flow angle at the diffuser inlet, leading to the suppression of flow separations on both the suction and pressure surfaces of the diffuser, respectively. However, it is found that the compressor performance deteriorates with increasing slotted portion depth and length, even though the stall margin is increased. When the slot depth is increased by up to 12% of the diffuser vane height, the leakage flow is not able to roll-up into a vortex that has the potential to mitigate the flow separation at the hub/suction surface corner. Instead, it enhances the flow separation region near the suction surface due to the negative effect of leakage flow; eventually, the stage performance is deteriorated. In conclusion, the slot depth should not exceed 9% of the diffuser vane height, and the optimum slot length is no greater than 6%–12% of the diffuser chord length, which provides a basic guideline for slotted diffuser design. [ABSTRACT FROM AUTHOR]

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

This paper presents a performance evaluation of non-axisymmetric casing grooves combined with airflow injection in a transonic axial compressor with NASA Rotor 37, using three-dimensional Reynolds-averaged Navier-Stokes equations with the k-ε turbulence model. An axisymmetric casing groove was divided circumferentially into 36 non-axisymmetric grooves. The numerical results for adiabatic efficiency and total pressure ratio were validated with experimental data. A parametric study for stall margin, stable range extension, peak adiabatic efficiency, and total pressure ratio at peak adiabatic efficiency of the compressor was performed using five parameters: the front and rear lengths, the height of the casing groove, the injection mass flow rate, and the injection angle. The non-axisymmetric casing grooves combined with injection improve greatly the stall margin and stable range extension of the transonic axial compressor, but reduce only slightly the peak adiabatic efficiency in all cases, compared to the results for a smooth casing. [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. [ABSTRACT FROM AUTHOR]

The objective of this study is to evaluate the effect of cross-blade slot casing treatment on the stability and performance of an axial flow compressor rotor. The experimental and unsteady calculated results both show that cross-blade slot casing treatment can generate about 22% stall margin improvement, and the compressor peak efficiency is reduced by about 13%. The detailed flow-field analyses indicate that the sucked and injected flow caused by the slots of cross-blade slot casing treatment can restrain the rotor tip passage blockage, which is made by the low energy tip clearance leakage vortex. When cross-blade slot casing treatment is applied, not only the rotor wheel flange work becomes lower in most of the rotor blade span, but also the flow loss in the blade tip passage becomes fairly large due to the strong interaction between the mainstream and the injected flows made by the slots. As a result, the compressor total pressure ratio and efficiency for cross-blade slot casing treatment are reduced obviously. Three kinds of new cross-blade slot casing treatment were designed according to the previous successful experience and investigated in this paper. The numerical results show that the new three cross-blade slot casing treatments both generate about 54% stall margin improvement at the cost of minor peak efficiency. For one new cross-blade slot casing treatment (CSCT2), the compressor peak efficiency is reduced by about 0.3%. The low energy TLV, which is present for cross-blade slot casing treatment, is removed by the strong sucked flow made by CSCT2. Moreover, the interaction between the mainstream and the injected flows caused by CSCT2 becomes weak obviously, and the corresponding flow loss is reduced greatly. Hence, the compressor stability and performance with CSCT2 are higher than those with cross-blade slot casing treatment. [ABSTRACT FROM AUTHOR]

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

In 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. [ABSTRACT FROM AUTHOR]

Stall recovery process for performance enhancement of an axial compressor has been experimentally carried out using air injection at its rotor blade row tip region. Twelve air injectors had been mounted evenly spaced around the compressor casing upstream the rotor blade row. Instantaneous flow velocities at various radial and circumferential positions were measured simultaneously utilizing hot wire anemometry. These unsteady results, separately presented in frequency and time domains, provided to distinguish stall inception process and consequent flow induced fluctuations. Time-dependent responses of the flow field properties within the compressor passage and progressive alleviation of stall cells are demonstrated during the tip injection process. Blade tip air injection worked effectively and enhanced the compressor stall margin about 9%. This attractive result occurred for the case where the total mass flow rate passing through the air injectors was as small as 0.8% of the compressor main flow rate. In addition, this augmentation in the stall margin was accompanied by increase in the compressor delivery total pressure. Air injection at the blade row tip region caused its beneficial effects to extend throughout the blade whole span, especially while working at the near-stall conditions. [ABSTRACT FROM AUTHOR]

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

Inflow temperature distortion is an influential factor to determine the stable work margin of aero-engine. In current paper, an overview of temperature distortion introduction mechanism is first reviewed. Fundamental study concerning temperature distortion introduction mechanism is presented. Summarization is given about influence of temperature distortion on propulsion component characteris tics and the aero-engine overall performance. Then comparative study is made based on currently applied temperature distortion simulation devices as well as their operation mechanisms. Progress made in both fundamental and applied research with the aid of temperature distortion simulation devices is introduced.In addition, inflow temperature distortion assessment methodology as well as corresponding standard is introduced with highlight on key techniques concerning temperature distortion test. Finally, techniques to mitigate temperature distortion on aero-engine are summarized and trade study is made upon those techniques. Aero-engine temperature distortion research will provide guidance for engine temperature resistance design and test. [ABSTRACT FROM AUTHOR]

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

The present research investigates the effect of the location and the width of single shallow circumferential groove casing treatment on the flow field and the stability improvement of NASA Rotor 37 utilizing the help of computational fluid dynamics. At first, steady state simulation of Rotor37 was presented for smooth casing (without groove). Then, forty five various grooved casing were simulated and compared with the smooth casing. The results indicated that narrow grooves had slight effect on the adiabatic efficiency but as the width of the groove was increased, a decline in efficiency was observed. The investigation on the stall margin revealed that narrow grooves next to the leading edge could improve the stall margin by a reduction in the size of vortex breakdown zone. Medium-width grooves displayed an effective role in delaying the separation- produced by shock wave and boundary layer interaction- on the blade suction side near the casing. This type of grooves could improve the stall margin more than narrow grooves when located on the top of separation zone near the blade suction side. Wide grooves had negative effect on the stall margin and caused a significant drop in the efficiency and the total pressure ratio of the compressor. [ABSTRACT FROM AUTHOR]

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

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

Parametric studies of recirculating casing treatment were experimentally performed in a subsonic axial flow compressor. The recirculating casing treatment was parameterized with injector throat height, injection position, and circumferential coverage percentage. Eighteen recirculating casing treatments were tested to study the effects on compressor stability and on the compressor overall performance at three blade speeds. The profiles of recirculating casing treatment were optimized to minimize the losses generated by air recirculation. In the experiment, the stalling mass flow rate, total pressure ratio, and adiabatic efficiency of the compressor were measured to study the steady-state effects on the compressor performance of recirculating casing treatments, and static pressure disturbances on the casing wall were monitored to study the influence on stall dynamics. Results indicate that both the compressor stability and overall performance can be improved through recirculating casing treatment with appropriate geometrical parameters for all the test speeds. The influence on stall margin of one geometric parameter often depends on the choice of others, i.e. the interaction effects exist. In general, the recirculating casing treatment with a moderate injector throat and a large circumferential coverage is the optimal choice to enhance compressor stability. The injector of recirculating casing treatment should be placed upstream of the blade tip leading edge and the injector throat height should be lower than four times the rotor tip gap for the benefits of compressor efficiency. At 71% speed, the blade tip loading is decreased through recirculating casing treatment at the operating condition of near peak efficiency and increased near stall. Moreover, the outlet absolute flow angle is reduced in the tip region and enhanced at lower blade spans for both operating conditions. The stall inceptions are not changed with the implementation of recirculating casing treatment for all the test speeds, but the stall patterns are altered at 33% and 53% speeds, i.e. the stall with two cells is detected in the recirculating casing treatment compared with the solid casing with only one stall cell. [ABSTRACT FROM AUTHOR]

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

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

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

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

Casing treatment is an effective approach for improving compressor stall margin in axial compressors. A plenum chamber, which connects slots in the circumferential direction, can occasionally be found in slot-type casing treatments. This study investigates the effect of a plenum chamber on the compressor overall performance and stall margin via time-accurate three-dimensional numerical simulations in a transonic compressor NASA Rotor 37 with a slot-type casing treatment. Four configurations with variable height of plenum chamber are numerically investigated. The unsteady influence of the casing treatment on compressor flow fields is studied using fast Fourier transforms to gain a deeper insight into the function of the plenum chamber. The results indicate that the plenum chamber can reduce the penalty on the compressor’s total pressure ratio and adiabatic efficiency induced by the casing treatment effect at the design operating condition. A detuning effect of the plenum chamber is found and proven to be responsible for the decreased compressor performance penalty through a detailed analysis in both the blade passage and casing treatments. The detuning effect dampens the unsteady influence of the casing treatment on the compressor passage flow in terms of shock oscillations and mass flow fluctuations. Thus, the radial scope and degree of the influence induced by the casing treatment is distinctly restrained by the plenum chamber. Slot circulation, which is composed of a bleed flow from the passage into the slots and injection flow from the slots into the passage, is closely related to the stall margin improvement as the casing treatment is implemented. The plenum chamber has no significant impact on the slot circulation and thus on compressor stall margin. However, shallow slots decrease the slot circulation efficiency, which leads to a much lower stall margin enhancement. [ABSTRACT FROM PUBLISHER]

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

A multiobjective optimization to enhance the stall margin and peak adiabatic efficiency of an axial compressor with a casing groove combined with injection was performed in this work. The geometry of the casing groove was optimized using a hybrid multiobjective evolutionary algorithm coupled with a surrogate model. Reynolds-averaged Navier–Stokes equations with a k–ɛ turbulence model were discretized by finite volume approximations and solved on hexahedral grids to analyze the flow in the compressor. The stall margin and peak adiabatic efficiency were selected as objective functions with three design variables related to the geometry of the casing groove. Latin hypercube sampling was applied as a design-of-experiment technique to generate 25 design points within the design space. Response surface approximation models as the surrogate models for the objectives were constructed based on the objective function values at the design points. A fast nondominated sorting genetic algorithm for the local search coupled with the surrogate models was applied to determine the global Pareto-optimal solutions. The tradeoff between the two objectives was determined and is described with respect to the Pareto-optimal solutions. The multiobjective optimization results showed that the stall margin and peak adiabatic efficiency of the axial compressor with an optimized casing groove combined with injection were simultaneously improved compared to the compressor with a smooth casing. [ABSTRACT FROM PUBLISHER]

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]

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

The paper investigates the effect of a single circumferential groove casing treatment (CGCT) on a transonic compressor rotor numerically. In particular, the effect of the groove at different axial locations on the flow field is studied in detail and stall margin improvement is also discussed. The present results show that the groove close to the leading edge plays a crucial role in stabilizing the near stall flow structures and, hence, improves the stall margin. The groove at the mid-chord-section of the blade can help exchange and transfer momentums between different directions, and suppress the flow unsteadiness, leading to increased efficiency in rotor performance and extended operation range. The groove located near the blade trailing edge has limited effects on stall margin improvement and may cause additional penalty in efficiency. Through comparison with the recent work on CGCT, some common flow physics can be observed. [ABSTRACT FROM AUTHOR]

Centrifugal compressors used for turbochargers are required to achieve high performance within wide flow ranges. However, their operability is limited at low mass flow rates by the flow instabilities leading to rotating stall and surge, which could be delayed by a suitable casing treatment. This article presents a numerical investigation aimed at studying the effects of a type of casing treatment on the flow structure through a radial compressor and the subsequent improvement of aerodynamic stability. The considered casing treatment consists of an oblique slot recess inserted in the shroud near the impeller throat and a large annular cavity connecting this slot to the bell-mouth. The flow computations were carried out by means of the code ANSYS-CFX, considering the shear stress transport (k–ω) turbulence model, in order to investigate the effects of the cavity dimensions on improving the stability limit of the compressor. The analysis of flow structure near the stall limit has revealed that the large vortex appearing near the blade tip is being controlled by such a type of casing treatment; hence the stall onset is delayed. As consequence, the operating range and the stall margin are improved by reducing the aerodynamic blockage. [ABSTRACT FROM AUTHOR]