Improvements of performance and stability of a single-stage transonic axial compressor using a combined flow control approach

Hu, Junfeng; Wang, Rugen; Huang, Danqin

doi:10.1016/j.ast.2018.12.033

Cited by 31 publications

(15 citation statements)

References 13 publications

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“…The results evidently showed a positive effect on compressor performances caused by corner separation reduction. Corner separation was also found to be suppressed by slot technique in the numerical study by (Hu et al, 2019). Therefore, this method, slotted rotor, clearly justifies itself to be an effective way to improve the compressor stability.…”

Section: Introductionmentioning

confidence: 78%

“…Basically, there are two types of flow control techniques divided by external work requirement, which are actuator (requiring additional devices such as synthetic jets and bleed techniques) and geometric configuration (not required external work input such as casing treatment and slotted rotor). While the geometric configurations seems to be most available in use for the industrial in nowadays (Hu et al, 2019).…”

Section: Introductionmentioning

confidence: 99%

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Numerical Investigation on the Performance of Transonic Axial Compressor with Naturally Aspirated Slots on Blade

Aksorn-In¹,

Chen²,

Zhao³

et al. 2019

Proceedings of Global Power &Amp; Propulsion Society

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The stable operating range is one of the key performance parameters on safety for general compressors. In order to improve the stable range, the flow control technique by slotted rotor was computationally studied and investigated based on the case of well-known transonic axial compressor (NASA Rotor 67). The principle of slotted blade is that the rotor blade is manufactured with the aim to have a slot (connecting between pressure side and suction side) at designed location. Then, the fluid is able to flow from pressure side to suction side by the pressure differences. This effect could literally energize the mainstream flow including the retarded flow caused by the aerodynamic losses. As a result, the operating range of compressor would be improved. Throughout this work, the action mechanism of the slot and the impact on compressor performance were analyzed by the prominent commercial software, NUMECA/EURANUS. The slot was located at three significant areas (e.g., leading edge, midchord, and trailing edge). Moreover, it was also located relative to the axial tip chord of the rotor blade. The stability improvement and the modified performance of the investigated compressor (e.g., pressure ratio and efficiency) were evaluated and compared with those of original blade. The location of slot giving the best improvement was finally identified. Afterwards, the optimum location would be selected for the further investigation by orienting the slot to be inclined configuration. In addition, the geometry of slot would be changed to nozzle as well for high velocity flow at the exit in order to gain more energization (more improvement).

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Section: Introductionmentioning

confidence: 78%

Section: Introductionmentioning

confidence: 99%

Numerical Investigation on the Performance of Transonic Axial Compressor with Naturally Aspirated Slots on Blade

Aksorn-In¹,

Chen²,

Zhao³

et al. 2019

Proceedings of Global Power &Amp; Propulsion Society

View full text Add to dashboard Cite

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“…For a transonic axial-flow compressor rotor, flow separations might occur in the blade hub region even for a well designed blade, especially under the inlet total pressure distortion condition. The separation loss is another major source of flow losses in the rotor blade passage and can significantly affect the compressor overall aerodynamic performances (Hu et al, 2019;Wang & Wu, 2019). In this section, the influences of the optimized rotor sweep design on the hub flow separations have been discussed.…”

Section: Effects On the Hub Flow Separationsmentioning

confidence: 99%

Optimization of a transonic axial-flow compressor under inlet total pressure distortion to enhance aerodynamic performance

Yang

et al. 2020

Engineering Applications of Computational Fluid Mechanics

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In a boundary layer ingestion (BLI) propulsion system, the compressor rotor operates under the condition of a severe inflow distortion because the boundary layers on the surface of aircraft have been ingested. In this situation, the low-momentum fluid would inevitably raise the blade load and aggravate the local aerodynamic losses, thus severely deteriorating the compressor overall aerodynamic performance. The paper has applied a sweep optimization design based on surrogate model and genetic algorithm to a transonic axial-flow compressor at the condition of total pressure distorted inflow. The optimization objective is to improve the compressor efficiency. The sweep design has been implemented through controlling the shape of the blade stacking line and the time-saving steady numerical simulations are mainly utilized in the optimization process. Then, the full-annulus unsteady CFD calculations have been conducted for validating the optimized design and further analyzing the flow mechanisms of performance improvements. The computational results show that the inflow distortion has led to non-uniform distributions of the tip leakage loss and the hub flow separation loss along the circumferential direction. It is found that the rotor blade has suffered from the highest tip leakage losses and hub separation losses during the blade leaving the distorted region. After rotor sweep optimization, both tip leakage and hub separation losses have been notably reduced along the full circumferential direction. Moreover, for different pitchwise locations, the optimized rotor sweep design has achieved more aerodynamic performance improvements in the process of the rotor blade leaving the distorted region than those in the processes of the rotor blade moving out of, moving towards and entering the distorted region. Overall, the optimized design performs better than the baseline design in terms of adiabatic efficiency in the entire operating range. In the meanwhile, the total pressure ratio has been maintained. Specifically, an approximately 0.5% increment of efficiency has been achieved by the optimized design at the optimization point. For this paper, the major purpose is to offer useful guidelines for fan/compressor design strategy to achieve high aerodynamic performances under distorted inflow working condition.

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“…Hub separations were successfully suppressed and an improvement of 0.97 per cent for the adiabatic efficiency was obtained without any penalty of total pressure ratio. Hu et al (2019) also decreased the rotor hub corner separations through a part-span slot configuration and improvements of both adiabatic efficiency and total pressure ratio were achieved. From the above results, it can be seen that the re-design of the rotor blade or the end-wall can reduce the end-wall secondary losses and achieve aerodynamic performance improvements for axial-flow compressors.…”

Section: Introductionmentioning

confidence: 96%

Analysis and application of shroud wall optimization to an axial compressor with upstream boundary layer to improve aerodynamic performance

Pan

et al. 2019

HFF

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Purpose For an axial-flow compressor rotor, the upstream inflow conditions will vary as the aircraft faces harsh flight conditions (such as taking off, landing or maneuvering) or the whole compressor operates at off-design conditions. With the increase of upstream boundary layer thickness, the rotor blade tip will be loaded and the increased blade load will deteriorate the shock/boundary layer interaction and tip leakage flows, resulting in high aerodynamic losses in the tip region. The purpose of this paper is to achieve a better flow control for tip secondary flows and provide a probable design strategy for high-load compressors to tolerate complex upstream inflow conditions. Design/methodology/approach This paper presents an analysis and application of shroud wall optimization to a typical transonic axial-flow compressor rotor by considering the inlet boundary layer (IBL). The design variables are selected to shape the shroud wall profile at the tip region with the purpose of controlling the tip leakage loss and the shock/boundary layer interaction loss. The objectives are to improve the compressor efficiency at the inlet-boundary-layer condition while keeping its aerodynamic performance at the uniform condition. Findings After the optimization of shroud wall contour, aerodynamic benefits are achieved mainly on two aspects. On the one hand, the shroud wall optimization has reduced the intensity of the tip leakage flow and the interaction between the leakage and main flows, thereby decreasing the leakage loss. On the other hand, the optimized shroud design changes the shock structure and redistributes the shock intensity in the spanwise direction, especially weakening the shock near the tip. In this situation, the shock/boundary layer interaction and the associated flow separations and wakes are also eliminated. On the whole, at the inlet-boundary-layer condition, the compressor with optimized shroud design has achieved a 0.8 per cent improvement of peak efficiency over that with baseline shroud design without sacrificing the total pressure ratio. Moreover, the re-designed compressor also maintains the aerodynamic performance at the uniform condition. The results indicate that the shroud wall profile has significant influences on the rotor tip losses and could be properly designed to enhance the compressor aerodynamic performance against the negative impacts of the IBL. Originality/value The originality of this paper lies in developing a shroud wall contour optimization design strategy to control the tip leakage loss and the shock/boundary layer interaction loss in a transonic compressor rotor. The obtained results could be beneficial for transonic compressors to tolerate the complex upstream inflow conditions.

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Improvements of performance and stability of a single-stage transonic axial compressor using a combined flow control approach

Cited by 31 publications

References 13 publications

Numerical Investigation on the Performance of Transonic Axial Compressor with Naturally Aspirated Slots on Blade

Numerical Investigation on the Performance of Transonic Axial Compressor with Naturally Aspirated Slots on Blade

Optimization of a transonic axial-flow compressor under inlet total pressure distortion to enhance aerodynamic performance

Analysis and application of shroud wall optimization to an axial compressor with upstream boundary layer to improve aerodynamic performance

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