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This paper explores the secondary surge during compressor surge recovery. Utilizing the test rig at Northwestern Polytechnical University, the surge is induced by tail cone adjustment. A secondary surge follows tail cone retraction post-initial surge. Signal changes are analyzed, and causes are expounded from energy and fluid dynamics. Experimental outcomes confirm the secondary surge is repeatable and observable under set conditions. Its frequency stabilizes at 4.83 Hz across consecutive events and rotor axes, being system-inherent. From energy, with enthalpy as a proxy, potential energy amasses as the compressor interacts with cavity potential shifts from tail cone changes. Exceeding kinetic supply instigates violent airflow oscillations. Peak pressure trends and simulations suggest surge likelihood at ∼3.33 × 105 J enthalpy. In fluid dynamics, post-initial surge, a stall state ensues. Tip clearance vortices from spillage and backflow expand, merge, and fill the passage. Their pre-secondary surge breakdown elevates pipeline pressure, causing gas reflux and surge recurrence. Conclusively, this research enhances compressor secondary surge comprehension, offering support for design optimization and anti-surge strategies. [ABSTRACT FROM AUTHOR]

In order to improve the compressor efficiency, a wavy leading edge design method is introduced. A single objective optimization method is adopted to find the optimal wavy leading edge design and the effect of the wavy leading edge on the suction surface separation of a subsonic compressor stator is investigated by means of numerical simulations and controlling the sinusoidal function. The optimization results indicate that the amplitude has a significant impact on the compressor efficiency. Three amplitude schemes were studied, and the results show that as the amplitude increases, the efficiency of the compressor first increases and then decreases. The optimal schemes can reduce flow loss by 21.8% and increase the peak efficiency of the compressor by 1.46%. The wavy leading edge blades could change the static pressure distribution and induce the formation of leading edge vortex pairs behind the valley to improve the ability of resisting the separation for the flow. [ABSTRACT FROM AUTHOR]

Casing treatment is a highly effective and promising technology for improving compressor stability. The study aims to enrich the casing treatment design method and reveal its mechanism for improving compressor stability and design guidelines. Focusing on Rotor37, we develop a multi-objective optimization platform that combines geometric parameterization, mesh generation, numerical calculations, optimization algorithms, and other relevant components. Five design variables are optimized to maximize stall margin improvement and minimize peak efficiency loss concurrently. The optimal T-shaped multi-groove casing treatment improved the stall margin by 11.49% and reduced peak efficiency loss to 0.13%. Furthermore, the impact of casing treatment on the flow field characteristic is quantified through three kinds of parameters. The quantification analysis demonstrates that the casing treatment enhances stability by improving the axial momentum of the mainstream flow through the radial momentum transport effect and reducing passage blockage. Additionally, Spearman correlation analysis and the Apriori algorithm are employed to investigate the relationships between seven T-shaped muti-groove geometric parameters and five flow field characteristic quantification parameters. The results validate the mechanism of casing treatment for improving compressor stability and obtain nine strongly associated rules. Finally, the linear correlation of the associated rules is evaluated to transform into design guidelines for casing treatment. The guidelines guide the design of T-shaped muti-groove geometry parameters for optimal performance. These guidelines stipulate that a single T-shaped groove should possess a minimum cross-sectional area of 10 mm2 and a depth of not less than 5 mm. Moreover, the cross-sectional area of the straight groove should exceed 4 mm2. [ABSTRACT FROM AUTHOR]

This study addresses the technical challenge that conventional self-recirculating casing treatments predominantly enhance the compressor stability with minimal efficiency gains. Focusing on a transonic axial flow compressor, National Aeronautics and Space Administration Stage35, the research designed an innovative stator suction-type self-circulating casing treatment (SS-SCT), which features the suction slots on the stator instead of the traditional self-circulating casing treatment's suction port. Using multipassage unsteady numerical simulations, this paper investigated the underlying mechanisms by which the SS-SCT affects the compressor performance. The numerical results reveal that the SS-SCT not only increases the compressor stall margin by 3.36% but also enhances the efficiency by 0.13% at the design point and by 0.6% at the near-stall point. The jetting action of the SS-SCT mitigates the detrimental effects of the rotor tip leakage flow and alleviates the blockages in the rotor tip passage at the near-stall point, thereby improving the compressor stability. Concurrently, the suction slots extract the low-speed fluid generated by the boundary layer separation in the stator passage, reducing flow losses, and consequently enhancing the compressor efficiency. [ABSTRACT FROM AUTHOR]

Copyright of Transactions of Nanjing University of Aeronautics & Astronautics is the property of Editorial Department of Journal of Nanjing University of Aeronautics & Astronautics and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

The self-circulating casing treatment can effectively expand the stable working range of the compressor, with little impact on its efficiency. With a single-stage transonic axial flow compressor NASA (National Aeronautics and Space Administration) Stage 35 as the research object, a multi-channel unsteady numerical calculation method was used here to design three types of self-circulating casing treatment structures: 20% Ca (axial chord length of the rotor blade tip), 60% Ca, and 178% Ca (at this time, the bleed position is at the stator channel casing) from the leading edge of the blade tip. The effects of these three bleed positions on the self-circulating stability expansion effect and compressor performance were studied separately. The calculation results indicate that the further the bleed position is from the leading edge of the blade tip, the weaker the expansion ability of the self-circulating casing treatment, and the greater the negative impact on the peak efficiency and design point efficiency of the compressor. This is because the air inlet of the self-circulating casing with an air intake position of 20% Ca is located directly above the core area of the rotor blade top blockage, which can more effectively extract low-energy fluid from the blockage area. Compared to the other two bleed positions, it has the greatest inhibitory effect on the leakage vortex in the rotor blade tip gap and has the strongest ability to improve the blockage at the rotor blade tip. Therefore, 20% Ca from the leading edge of the blade tip has the strongest stability expansion ability, achieving a stall margin improvement of 11.28%. [ABSTRACT FROM AUTHOR]

In this paper, a numerical optimization study of single-groove casing treatment was conducted on a high-speed axial compressor. One of the aims is to find the optimal structure of a single groove that can improve compressor stability with minimal loss in efficiency. Another aim is to explore suitable parameters for rapidly evaluating the compressor stall margin. A design optimization platform has been constructed in this paper, which utilizes NSGA-II and a Radial Basis Function (RBF) neural net model to carry out the optimization. The stall margin of the compressor with A single groove was accurately determined by calculating its entire overall performance line. A Pareto front is obtained through optimization, and the optimal design can be selected from the Pareto front. By considering both stall margin and efficiency loss, one of the optimal designs was found to achieve a 7.49% improvement in stall margin with a 0.24% improvement in peak efficiency. Based on the database, the effect of design parameters of a single groove on compressor stability and performance is analyzed. A series of evaluation parameters of stall margin were compared to their degree of correlation with the real stall margin calculated by the entire overall performance line. As a result, tip blockage and momentum ratio can be used as efficient parameters for quickly evaluating the compressor stall margin without the need to calculate the entire performance curve of the compressor. [ABSTRACT FROM AUTHOR]

The three-dimensional numerical simulation was used to study the effects of self-recirculating casing treatment (SRCT) on the stability and performance of a subsonic axial flow compressor. The control variable method was used to compare three SRCT structures with different axial bleeding positions. The mechanism of the influence of different SRCTs on the compressor stability was revealed. Unsteady calculation results show that as the bleeding position moves downstream, the stall margin obtained by SRCT increases first and then decreases. SRCTs with bleeding positions at 0.4, 0.8, and 1.3 axial chord length (Ca) obtain 19.74%, 22.72%, and 15.97% stall margin improvement, respectively. The internal flow field analysis shows that the 0.8Ca bleeding position is close to the low-speed fluid at the blade tip, and there is a large pressure difference between the bleeding port and the injection port. This structure can suppress the secondary flow in the blade passage, thereby reducing the mixing loss caused by the tip leakage flow and the mainstream. The airflow inside the SRCT with 0.8Ca bleeding position is smooth, and the flow loss caused by the bleeding effect is low. Therefore, the compressor efficiency of this structure is the highest. [ABSTRACT FROM AUTHOR]

In order to investigate the effect of a synthetic jet arranged at the suction surface on the aerodynamic performance of an axial flow compressor, a high-speed subsonic axial flow compressor rotor was numerically simulated in this paper. The results showed that after being excited by the synthetic jet, the flow margin of the compressor decreased by 1.3%, but the peak efficiency increased by 0.47%. After analyzing the internal flow field, it was found that although the alternating blowing and suction effect of the synthetic jet can reduce the separation loss of the suction surface, the radial velocity of the flow is higher when stimulated and converges towards the blade tip, resulting in more severe blockage at the blade tip than in the prototype compressor, leading to a slight decrease in stability margin. To balance the stability margin of the compressor, a coupled flow control numerical simulation combining passive control of the casing treatment and active control of the suction surface synthetic jet was subsequently carried out. The results showed that the coupled flow control could exert their respective advantages, resulting in an increase of 17.84% in the compressor flow margin and a decrease of only 0.2% in peak efficiency. [ABSTRACT FROM AUTHOR]

Purpose: The geometric parameters of the compressor blade have a noteworthy influence on compressor stability, which should be meticulously designed. However, machining inaccuracies cause the blade geometric parameters to deviate from the ideal design, and the geometric deviation exhibits high randomness. Therefore, the purpose of this study is to quantify the uncertainty and analyze the sensitivity of the impact of blade geometric deviation on compressor stability. Design/methodology/approach: In this work, the influence of blade geometric deviation is analyzed based on a subsonic compressor rotor stage, and three-dimensional numerical simulations are used to compute samples with different geometric features. A method of combining Halton sequence and non-intrusive polynomial chaos is adopted to carry out uncertainty quantitative analysis. Sobol' index and Spearman correlation coefficient are used to analysis the sensitivity and correlation between compressor stability and blade geometric deviation, respectively. Findings: The results show that the compressor stability is most sensitive to the tip clearance deviation, whereas deviations in the leading edge radius, trailing edge radius and chord length have minimal impact on the compressor stability. And, the effects of various blade geometric deviations on the compressor stability are basically independent and linearly superimposed. Originality/value: This work provided a new approach for uncertainty quantification in compressor stability analysis. The conclusions obtained in this work provide some reference value for the manufacturing and maintenance of rotor blades. [ABSTRACT FROM AUTHOR]

Copyright of Transactions of Nanjing University of Aeronautics & Astronautics is the property of Editorial Department of Journal of Nanjing University of Aeronautics & Astronautics and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

The blade thickness distribution is an important geometric parameter of a compressor rotor blade. It determines the aerodynamic performance of a rotor, which should be carefully designed. However, the manufacturing inaccuracies cause the blade thickness distribution to deviate from the ideal design. These geometric deviations alter the flow field near the blade surface, which affects the aerodynamic performance of the rotor. Therefore, it is of great significance to quantitatively investigate the effects of blade thickness deviation on the aerodynamic performance. In this paper, the influence of blade thickness deviation on the flow field is analyzed based on a transonic compressor rotor, and an uncertainty quantification process is performed to study the effects of blade thickness deviation on the aerodynamic performance. Cases with different geometry features are checked in the current study using 3-dimensional Reynolds-averaged Navier–Stokes simulations. Results show that the blade thickness deviation leads to changes in the intensities of the expansion wave in the upper part of the blade suction and the passage shock wave, which affects the supercharging process of the rotor. The influence of thickness deviation on the strength of the tip leakage vortex changes the isentropic efficiency of the rotor. The results of the uncertainty quantitative analysis indicate that the linear correlation between the variation of rotor performance and the thickness deviation is strong when the machining accuracy is high, whereas the linear correlation between the two is weakened when the tolerance range is larger. [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]

Casing treatment is an effective passive technology for improving the compressor stability. However, the current design methods for the casing treatment rely excessively on trial and error experiences, presenting significant challenges to actual engineering applications. In this paper, we propose a multi-objective optimization design method based on stall margin evaluation and data mining to enhance the stability of axial compressor rotors. We have developed a multi-objective optimization platform that combines geometric parameterization, mesh generation, numerical calculations, optimization algorithms, and other relevant components. To optimize six design variables and two objective functions, we have implemented two optimization strategies based on direct stall margin calculation and stall margin evaluation. The optimization results revealed that optimal casing treatment structures can be obtained by considering both compressor stability and efficiency. Furthermore, we employed data mining of self-organizing maps to explain the tradeoffs from the optimal solutions. The aerodynamic analysis demonstrated that the casing treatment enhances stability by restricting negative axial momentum of tip leakage flow and reducing passage blockage. Four categories of stall margin evaluation parameters were quantified, and their effectiveness was assessed through a correlation analysis. Finally, we used the axial momentum of the tip leakage flow-related evaluation parameter for the optimization of stall margin evaluation. Compared with direct stall margin calculation-based optimization, the evaluation of the parameter-based optimization method effectively predicted the stability enhancement of casing treatment while revealing the optimal geometric features. It suggests that the stall margin evaluation-based optimization method should be utilized in the initial optimization process of casing treatment due to its advantages in the optimization speed. [ABSTRACT FROM AUTHOR]

Compressed air energy storage systems must promptly adapt to power network demand fluctuations, necessitating a high surge margin in the compression system to ensure safety. It is challenging to completely eliminate blade geometric variations caused by limited machining precision, the important effects of which should be considered during aerodynamic shape design and production inspection. The present paper explores the uncertainty impact of geometric deviations on the stability margin of a multi-stage axial compressor at a low rotational speed. Initially, an adaptive polynomial chaos expansion-based universal Kriging model is introduced, and its superior response performance in addressing high-dimensional uncertainty quantification problems is validated through rigorous analytical and engineering tests. Then, this model is used to statistically evaluate the stability margin improvement (SMI) of the compressor due to the Gaussian and realistic geometric variabilities separately. The results show that the mean and standard deviation of SMI are −0.11% and 0.5% under the Gaussian geometric variability, while those are 0.33% and 0.39% under the realistic variability. For both the geometric variabilities, the stagger angle and maximum thickness deviations of the first-stage rotor are the most influential parameters controlling the uncertainty variations in the stability margin. Finally, the underlying impact mechanism of the influential geometric deviations is investigated. The variation in the stability margin caused by the geometric deviations primarily results from the alteration of inlet incidences, affecting the size of the tip leakage vortex blockage and boundary-layer separation regions near the blade tip of the first-stage rotor. [ABSTRACT FROM AUTHOR]

The impact of geometric deviation due to manufacturing on compressor performance is considerable in engineering practice. To investigate the impact of blade thickness deviation on compressor performance and flow loss at various rotational speeds, a three-dimensional steady numerical simulation on Rotor 37 was conducted. The quantification of uncertainty was accomplished using a non-intrusive polynomial chaos method. The viscous dissipation coefficient was introduced to analyze the uncertain influence of blade thickness deviation on flow loss. Based on the type of loss source, the flow field was divided into six regions, including the blade tip region, blade root region, leading edge region, trailing edge region, blade surface region, and mainstream region. The results indicate that the sensitivity of total pressure ratio to thickness deviation increases significantly with an increase in the rotational speed. Under peak efficiency conditions, the effect of blade thickness deviation on flow dissipation in leading edge region decreases initially and then increases with an increase in the rotational speed. Meanwhile, the impact on flow loss in other regions increases with the increase in the rotational speed. Under near stall conditions, the blade thickness deviation has a great impact on the flow losses in the blade tip region, leading edge region, and mainstream region at 60% design rotational speed. However, the blade tip region and trailing edge region are more noticeably affected at 100% design rotational speed. Furthermore, the quantification of standard deviation of flow losses in various regions under different rotational speeds and conditions reveals that the flow loss fluctuation in the leading edge region and mainstream region varies with changes in operating conditions and rotational speeds, but the fluctuation of flow loss in other regions is independent of the rotational speed. [ABSTRACT FROM AUTHOR]

Leading-edge tubercles are an effective method to improve the stall margin in a compressor. In existing studies on leading-edge tubercles, achieving a better control on all conditions is a huge difficulty and challenge. Hence, a new method of leading-edge tubercles profiling based on sine and attenuation functions is introduced in this paper. First, the wavelength and amplitude of the leading-edge tubercles were varied by sine function to study their effects on compressor performance. The research reveals uniform tubercles with small amplitude and large wavelength can delay stall incidence from 7.9° to 8.8° and increase it by 10% compared to the baseline. A small amplitude is beneficial to reduce the additional loss caused by the leading-edge tubercles near the blade middle, and a large wavelength is conducive to the development of separation vortex. Then, the leading-edge tubercles were further modified and investigated by introducing some attenuation functions. A suitable attenuation function is introduced to the uniform tubercles with small amplitude and large wavelength so that stall incidence is delayed to 9.7° and increased by 21.25% compared with the baseline. Finally, the vorticity transport equation and three-dimensional streamline reveal that the formation and development of leading-edge vortex pairs are mainly related to the axial bending of the circumferential vortex, the axial stretching of the axial vortex, and vortex viscous dissipation. For this paper, the principal purpose is to offer useful design guidelines, study flow control mechanisms, and achieve better aerodynamic performances under all working conditions for the leading-edge tubercles in the compressor. [ABSTRACT FROM AUTHOR]

To reveal the mechanisms underlying the effect of self-circulating casing treatment with different circumferential coverage ratios on the stability of the axial compressor, a three-dimensional unsteady numerical was hereby conducted on Rotor 35. The circumferential coverage ratios of self-circulating casing of 20%, 40%, 60%, and 80% were designed, respectively. The calculated results point out that all the schemes effectively expand the stable working range of the compressor and that the expansion effect is positively correlated with the circumferential coverage ratio. The self-circulating casing with an 80% circumferential coverage ratio exhibits the highest stall margin improvement at 14.83%. The internal flow field analysis shows that the underlying mechanism for the compressor stability increasing with the increase in the circumferential coverage ratio is that after the flows with a higher circumferential speed component enter the self-circulating casing suction port, sufficient circumferential space is required to complete the transformation in the flow direction, so that the flows can smoothly enter the self-circulating casing and subsequent development can be carried out. The larger circumferential size of the self-circulating casing creates favorable conditions for more airflows to enter the self-circulating casing. With the increase in the circumferential coverage ratio, the suction effect of self-circulating casing on low-speed fluid at the blade tip and the bleeding mass flow rate is larger, and a better compressor expansion effect is thereby achieved. [ABSTRACT FROM AUTHOR]

In this paper, an adaptive sparse arbitrary polynomial chaos expansion (PCE) is first proposed to quantify the performance impact of realistic multi-dimensional manufacturing uncertainties. The Stieltjes algorithm is employed to generate the PCE basis functions concerning geometric variations with arbitrary distributions. The basis-adaptive Bayesian compressive sensing algorithm is introduced to retain a small number of significant PCE basis functions, requiring fewer model training samples while preserving fitting accuracy. Second, several benchmark tests are used to verify the computational efficiency and accuracy of the proposed method. Eventually, the coexistence effects of six typical machining deviations on the aerodynamic performance and flow fields of a controlled diffusion compressor cascade are investigated. The probability distributions of the machining deviations are approximated by limited measurement data using kernel density estimation. By uncertainty quantification, it can be learned that the mean performance seriously deteriorates with increasing incidences, while the performance at negative incidences is more dispersed. By global sensitivity analysis, the leading-edge profile error should be given high priority when working at negative incidences, and the inlet metal angle error would be carefully inspected first when the cascade works at high positive incidences. Furthermore, controlling the manufacturing accuracy of the suction surface profile error can play a certain role in improving the robustness of aerodynamic performance in off-design conditions. Through flow field analysis, it further proves that actual leading-edge errors are the most important ones to aerodynamics and reveals how the effects of leading-edge errors propagate in the cascade passage, thus affecting the aerodynamic loss. [ABSTRACT FROM AUTHOR]

The aim of this research was to explore the mechanisms underlying the effect of self-circulating casing treatment with different circumferential coverage ratios on the aerodynamic performance of a transonic centrifugal compressor. A three-dimensional unsteady numerical simulation was carried out on a Krain impeller. The circumferential coverage ratios of the self-circulating casings were set to 36%, 54%, 72% and 90%, respectively. The numerical results showed that the Stall Margin Improvement (SMI) increased with the increase in circumferential coverage ratios. The self-circulating casing with a 90% circumferential coverage ratio exhibited the highest SMI at 20.22%. Internal flow field analysis showed that the self-circulating casing treatment improved the compressor stability by sucking the low-speed flow in the blade tip passage and restraining the leakage vortexes breaking, which caused flow blockage. The compressor performance was improved at most of the operating points, and the improvement increased with increase in circumferential coverage ratio. The improvement in compressor performance was mainly attributed to reduction in the area of the high relative total pressure loss in the blade tip passage and significant decrease in the flow loss by the self-circulating casings. [ABSTRACT FROM AUTHOR]

This study aimed to investigate the influence of the bleeding position of a self-circulating casing on the aerodynamic performance of a transonic centrifugal compressor. Three types of self-circulating structures with the bleeding positions of 11% Ca (the axial chord length of the blade tip), 14% Ca and 20% Ca from the leading edge of the blade were studied by using the numerical simulation method, with the Krain impeller taken as the research object. It was found that all three types of self-recirculating casing treatments can expand the stable operating range of the impeller, and that at medium and small flow rates, the total pressure ratio and efficiency of the impeller increase gradually with the backward movement of the bleeding position. The self-circulating casing treatment can restrain the development of tip leakage vortex, reduce the blockage area, and improve the stability of the impeller by sucking low-energy fluid. The farther back the bleeding position is, the greater the bleeding mass flow rate of the self-circulating casing for the low-energy fluid in the blade-tip passage becomes. Additionally, a greater inhibition effect on the tip leakage vortex, and a better effect of improving the performance and stability of the impeller, can be obtained. The best air bleeding position is 20% Ca, but it is not directly above the blade-tip blockage center of the solid wall casing passage. Instead, it is downstream of the blockage area. [ABSTRACT FROM AUTHOR]

The aerodynamic sound from bluff bodies is a practically important problem in various engineering applications. To control the aerodynamic noise, the sound emitted from a circular cylinder with and without a splitter plate in a Reynolds number (Re) of 3 × 1 0 4 is studied using Ffowcs Williams and Hawkings (FW-H) acoustic analogy. The flow field is simulated by Detached Eddy Simulation (DES) approach to investigate the mechanism of the sound reduction using a splitter plate in a three-dimensional calculation. The predicted sound of the circular cylinder is compared with the experimental data from the literature, and a good agreement is achieved. The results from this study show that lift and drag fluctuations of the circular cylinder with the splitter plate are smaller than those of the no-splitter case. The Strouhal number related to vortex shedding with the splitter plate is slightly reduced compared to the unmodified circular cylinder due to the stretched shear layers. The pressure fluctuations in the wake are decreased by the splitter plate, resulting from the suppression of vortex shedding. The application of the splitter plate reduces the lift dipole which is the main sound source. It leads to a sound reduction of 13 dB. [ABSTRACT FROM AUTHOR]

This paper presents numerical and experimental investigations about grooved casing treatment with the help of a high-speed small-scale compressor rotor First, the numerical investigation seeks to offer a contribution of understanding the working mechanism by which circumferential grooves improve stall margin. It is found that stall margin gain due to the presence of circumferential grooves arises from the suction-injection effect and the near-tip unloading effect. Based on that, the philosophy of design of experiment is then set up. Finally, parametric studies are carried out through systematical experiments. It is found that the orthogonal experiment and the factorial analyses are successful in identifying the 'best casing configuration' in terms of stall margin improvement. However, the ineffectiveness of the deduction from simulations suggests that the secondary flow circulations on stall margin gain should not be neglected, and the overall contribution of each groove to stall margin gain depends on its unloading effect and the compound effect of suction-injection. Further numerical investigation will focus on how to set up quantitative criteria to evaluate the compound effect of suction-injection and the unloading effect on stall margin gain respectively in each groove. [DOI: 10.1115/1.4003000]

Flow variations at the inlet boundary due to the compressor operational condition changes and geometric variations of the realistic compressor blades due to the manufacturing variability cannot be absolutely avoided, the global and local performance impact of which requires to be considered in the mechanism study of performance change and the aerodynamic shape design. In this paper, a method to analyze the simultaneous impact of the inflow Mach number, inlet incidence and geometric uncertainties was proposed. To make the uncertainty modeling of geometric variations faster and closer to engineering practice, a parametric mathematical model based on scanning points on the blade was introduced to describing the profile and torsion errors in the method specially. Meanwhile, a sparse grid-based Non-Intrusive Polynomial Chaos (NIPC) was used for uncertainty quantification and uncertainty sensitivity analysis to alleviate the computational burden. Then, the method was combined with a loss source calculation method to estimate the global and local aerodynamic loss changes of a controlled diffusion compressor blade in a reference flow state of high inflow Mach number and large positive incidence, and the response performance of sparse grid-based NIPC was verified. The results show that inlet incidence and torsion error have a significance uncertainty effect on the boundary-layer separation above the suction surface, which is main reason for the fluctuation of global aerodynamic loss. The uncertainty effect of profile error on the boundary-layer separation is relatively weak, but profile error could have a certain uncertainty effect on the leading edge separation. The boundary-layer separation is insensitive to Inflow Mach number. Furthermore, a stochastic aerodynamic analysis in different reference inflow states was investigated, which reveals some laws that the uncertainty of the aerodynamic losses and the sensitivity of the inflow and geometric uncertainties change with reference inflow states. [ABSTRACT FROM AUTHOR]

The research on geometric parameters of casing treatment (CT) has always been a hot topic, yet the multi-parameter interaction is rarely studied. In order to gain better knowledge of the interaction mechanism of geometric parameters of CT, the experimental and numerical study based on the response surface method has been carried out in an axial compressor rotor. First of all, the statistical analysis based on the database of experimental and numerical results is presented to summarize the influence law of varied parameters on the stability enhancement. It was found that axial overlap, open area ratio, and their interaction had the most significant influence on the stability enhancement of CT. Subsequently, the interaction between axial overlap and open area ratio was analyzed by visualization flow field in details, which provided a deeper insight into stability enhancement mechanism of CT. It indicated that the mass flow and momentum dominated by injection and the suction effect played a key role for extending stability. With smaller open area ratio, it was difficult for the slots to manipulate and control the tip leakage flow or secondary tip leakage flow, resulting in the weak effect of CT on compressor performance. Finally, the underlying flow physics in the tip region and dominant region of CT has also been discussed to penetrate the essential reason of multi-parameter interaction. [ABSTRACT FROM AUTHOR]

High-load axial compressor is the mainstream of current compressor design and development. In order to improve the aerodynamic performance of high-load axial compressor, an active flow control method in which a synthetic jet is applied to the endwall is proposed. Taking the transonic axial compressor NASA Rotor 35 as the research object, using a single factor analysis method, the influence of five different excitation positions, three different excitation frequencies, and three different jet peak velocities on the aerodynamic performance of the compressor was studied in turn, and obtained the influence law of the endwall synthetic jet excitation parameters. The results show that all three parameters have important effects on the performance of the compressor. Among the excitation parameters studied in this paper, there is an optimal excitation position of 25% Ca. When excited at this position, the flow margin of the compressor is expanded the most. On the basis of maintaining the optimal excitation position and the maximum jet peak velocity, the calculation results found that the jet frequency has little effect on the compressor's near stall flow rate, but has a great impact on the total pressure ratio and efficiency. The pressure ratio and efficiency increase with the increase of the excitation frequency. However, there seems to be a threshold of the excitation frequency. Only when the excitation frequency is greater than the threshold can the total pressure ratio and efficiency be higher than the prototype compressor. The jet peak velocity has the smallest impact on the compressor performance. Based on the optimal excitation position and the excitation frequency exceeding the threshold, even if the jet peak velocity is small, the compressor can obtain a higher flow margin, total pressure ratio, and efficiency than the prototype compressor. As the jet peak velocity increases, the performance of compressor can be further improved. [ABSTRACT FROM AUTHOR]

In order to prolong the service life of multistage axial compressors, it is increasingly important to study the influence of blade surface roughness on the compressor performance. In this paper, a five-stage axial compressor of a real aero-engine was selected as the research object, and an equivalent gravel roughness model was used to model the roughness based on measured blade surface roughness data. Furthermore, the impact of blade surface roughness on the performance at design rotational speed was studied by full three-dimensional numerical simulation, and the mechanism of performance variation caused by the roughness was discussed combined with quantitative and flow field analyses. The results show that, when the blade surface roughness of all blades increases, the peak total efficiency decreases by approximately 0.4%, the blocking mass-flow decreases by approximately 0.3%, and the stable working range changes little. When the surface roughness of all rotor blades increases, the performance decline is close to that of all rotor and stator blades, and the variation in stator blade roughness has little effect on the compressor performance. Regarding the variation in roughness, the performance of the latter stage is more sensitive than that of the previous stage, and the decline in the performance of the fifth stage contributes the most to the total performance degradation of the compressor. Once the surface roughness of the fifth-stage rotor blade increases, the flow in the middle of the rotor blade deteriorates and the stage performance decreases obviously, which is the main reason for the decline in the overall performance. [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]

To further control corner separation in high-load axial compressors, this study proposes a new end wall contouring method. It defines multiple standard "surface units" with particular flow control effects and then applies a linear combination, finally forming the geometry of the end wall surface. Based on design experiences, three different end wall contouring cases are generated and calculated on a high-load compressor cascade in the first step. The results show that the new method achieves a clear and intuitive influence on the end wall geometry, with a proper number of design variables, and can effectively combine variables with the development of secondary flow. In the second step, the new method was applied to an axial compressor, with an improvement in the design variables. Although the end wall contouring only improved the efficiency of the compressor stage on the right part of its operating map, the experimental results of the flow field show that the corner separation and end wall loss are suppressed at multiple inflow conditions. The results thus verified the practical effect of the newly developed end wall contouring method. [ABSTRACT FROM AUTHOR]

It is essential for engineering manufacture and robust design to evaluate the impact of manufacturing variability on the aerodynamics of compressor blades efficiently and accurately. In the paper, a novel quadratic curve approximation method based on the scanning points of blade design profiles was introduced and combined with Karhunen–Loève expansion, a mathematical dimensionality reduction method for modeling manufacturing variability as truncated Normal process was proposed. Subsequently, Sparse Approximation of Moment-based Arbitrary Polynomial Chaos (SAMBA PC) and computational fluid dynamics (CFD) were applied to build a computational framework for stochastic aerodynamic analysis considering manufacturing variability. Finally, the framework was adopted to evaluate the aerodynamic variations of a high subsonic compressor cascade under the design incidence. The results illustrate that the SAMBA PC method is more efficient than the traditional methods such as Monte Carlo simulation (MCS) for stochastic aerodynamic analysis. Through uncertainty quantification, the impact of manufacturing variability on the global aerodynamic performance is primarily reflected in the fluctuation of aerodynamic losses, and the fluctuation of the total losses is mainly contributed by the fluctuation of the separation loss after the suction peak (a negative pressure spike near the leading edge (LE)) and the boundary-layer loss on the suction surface (SS). With sensitivity analysis, the most important geometric modes to aerodynamics can be revealed, which provides a useful reference for manufacturing inspection process and helps reduce computational cost in robust design. [ABSTRACT FROM AUTHOR]

To deal with corner separation in high-load axial compressors, this paper proposes a new end wall contouring method aimed at controlling the end wall secondary flow in more than one local area, generating a geometry with fewer control variables that is applicable for multiple working conditions. The new method defines more than one surface unit function, with different effects on end wall secondary flow. Then, the geometry of these surface unit functions will be superposed to generate the end wall contouring, to combine their flow control effects. After applying the new method to a bi-objective optimization design process, with 15 design variables aimed at minimizing the loss of cascade at 0° and 4° incidence, the optimal design reduces the total pressure loss of the high-load cascade by 5% under the former incidence and by 3% under the latter. The most effective design rule is constructing an end wall surface with the rising suction side and sinking pressure side in the blade channel, while locally raising the SS corner with a gentle upstream slope. According to the analysis, the design variables of the new method show an intuitive influence on the variation of end wall geometry and the movement of secondary flow. The corner separation has been effectively suppressed, with fewer control variables than before. It, thus, indicates the advantage of the newly developed end wall contouring method compared with previous studies. [ABSTRACT FROM AUTHOR]

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]

Previous work has shown that performance and internal flow characteristics of a centrifugal pump can be significantly improved with grooved volute casing (GVC). However, it has been found that the selection of the design parameters of the groove structure also has a direct impact on the performance output, internal flow pressure pulsation and erosion wear characteristics of the overflow components of centrifugal pump, so it is necessary to further analyze the design rules of the groove structure parameters. In this study, we first investigated the influence of the number of grooves on the head, efficiency and unsteady pressure pulsation characteristics of the internal flow field of the centrifugal pump, and on this basis, the correlation between different particle parameters and the erosion wear of key overflow components under the conditions of solid–liquid two-phase flow were also studied, and the erosion wear characteristics of the inner wall of the volute casing of centrifugal pump with GVC and original volute casing (OVC) structures were compared. This research leads to the conclusion that when the number of grooves is 3, the groove structure has the least influence on the performance of the centrifugal pump, and the overall change of the performance curve is more stable. Additionally, the pressure pulsation at each monitoring point of the GVC under the same flow condition is smaller, and when the number of grooves increases, the pressure pulsation amplitude also decreases. When the number of grooves is 3, the GVC shows a more significant flow improvement effect under all flow conditions. Based on the improvement of the groove structure on the flow stability, the particle motion behavior can be affected at the same time, so that the pump with GVC can mitigate the erosion wear of the inner wall of the volute casing under the solid–liquid two-phase flow conditions, which improves the critical performance and service life of the key overflow components of the pump. [ABSTRACT FROM AUTHOR]