Numerical research on inlet total pressure distortion in a transonic compressor with non-axisymmetric stator

Sun, Peng; Fu, Wenguang; Wang, Hong; Zhong, Jingjun

doi:10.1177/0954410017740385

Cited by 4 publications

(4 citation statements)

References 15 publications

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“…It is worth noting that the structure of non-axisymmetric stator (NAS-CTBS) is consistent with that in Ref. 21. In order to systematically explore the influence of structural parameters of the non-axisymmetric stator on compressor stage performance, it is included in this paper for comparative analysis with other nonaxisymmetric stator schemes.…”

Section: Non-axisymmetric Stator Schemesmentioning

confidence: 71%

“…The test rig and test equipment are described in the Ref. 21. The compressor stage aerodynamic characteristics compared with the numerical simulation results are shown in Figure 6.…”

Section: Numerical Methods Validationmentioning

confidence: 99%

“…In addition, the strength of shock waves in the rotor passage was also weakened. The ensuing researches 21,22 showed that the stability margin of the compressor with non-axisymmetric stator was increased by approximately 12%, from 8.3% to 9.3%. Xu et al 23 found that the non-axisymmetric stator (nine of the stator blades in the distorted region had tip clearance) could ensure stage adiabatic efficiency and total pressure ratio while the surge margin was increased significantly.…”

Section: Introductionmentioning

confidence: 99%

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Effect of three non-axisymmetric stator schemes on compressor performance with distorted inlet

Sun

2021

Proceedings of the Institution of Mechanical Engineers, Part G:

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In the boundary layer ingesting propulsion system, the compressor suffers from a non-uniform flow field. The compressor operating with distorted inflow continuously results in the loss of aerodynamic performance and stability margin. In this paper, three non-axisymmetric configurations are described for the stator of a transonic compressor to match the non-uniform flow field. The flow fields with distorted inflow at near stall condition are obtained and analyzed, the effects of the prototype stator and the three non-axisymmetric stators on aerodynamic performance are compared in detail. Results show that the non-axisymmetric stator schemes can effectively improve the stability margin of the transonic compressor and the maximum stability margin is relatively increased by 22.3% in all the three non-axisymmetric stators. The non-axisymmetric stator design is effective on decreasing the aerodynamic losses and improving the performance of the compressor operating with distorted inflow. Overall, the results show that in the design of the non-axisymmetric stator, the adoption of a curved-twisted blade and the increase of cascade solidity have the potential to reduce loss sources caused by distorted inflow.

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Section: Non-axisymmetric Stator Schemesmentioning

confidence: 71%

“…The test rig and test equipment are described in the Ref. 21. The compressor stage aerodynamic characteristics compared with the numerical simulation results are shown in Figure 6.…”

Section: Numerical Methods Validationmentioning

confidence: 99%

Section: Introductionmentioning

confidence: 99%

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Effect of three non-axisymmetric stator schemes on compressor performance with distorted inlet

Sun

2021

Proceedings of the Institution of Mechanical Engineers, Part G:

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“…The results indicated the VIGV could improve the overall aerodynamic performances of the fan through decreasing the swirls at the inlet of the rotor blade. Furthermore, Sun et al (2019) attempted to apply a non-axisymmetric stator design to a transonic compressor to reduce the negative impacts of circumferential total pressure distortion, and the results validated its effectiveness.…”

Section: Introductionmentioning

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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