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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]

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 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]