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Chinese Journal of Aeronautics 2024年第10期37卷 83-102页

作者： Ruoye XIAO Dong SUN Jian YU School of Aeronautic Science and Engineering Beihang UniversityBeijing 100191China State Key Laboratory of Aerodynamics Mianyang 621000China

shock wave/boundary layer interaction in a 24°turning angle of the compression ramp at Mach number 2.9 controlled by steady microjet is investigated using direct numerical *** different jet spacings which are termed as sparse,moderate and dense are considered,and the induced vortex system and shock structures are compared.A moderate jet spacing configuration is found to generate counter-rotating vortex pairs that transport high-momentum fluid towards the vicinity of wall and strengthen the boundary layer to resist separation,reducing the separation *** dense jet spacing configuration creates a larger momentum deficit region,reducing the friction downstream of the *** of pressure and pressure gradient reveals that dense jet spacing configuration reduces the intensity of separation *** impact of varying jet spacings on the turbulent kinetic energy transport mechanism is also investigated by decomposing the budget terms in the transport ***,the spectral characteristics of the separation region are studied using power spectral density and dynamic mode decomposition methods,revealing that moderate jet spacing configuration suppresses low-frequency fluctuations in the separation region.

关键词： shock wave/boundary layer interaction Compression ramp Steady microjet Different jet spacings Direct numerical simulation

Passage shock wave/boundary layer interaction control for transonic compressors using bumps

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Chinese Journal of Aeronautics 2022年第2期35卷 82-97页

作者： Yongzhen LIU Wei ZHAO Qingjun ZHAO Qiang ZHOU Jianzhong XU Institute of Engineering Thermophysics Chinese Academy of SciencesBeijing 100190China School of Aeronautics and Astronautics University of Chinese Academy of SciencesBeijing 100049China Innovation Academy for Light-duty Gas Turbine Chinese Academy of SciencesBeijing 100190China

Flow separation due to shock wave/boundary layer interaction is dominated in blade passage with supersonic relative incoming flow,which always accompanies aerodynamic performance penalties.A loss reduction method for smearing the passage shock foot via shock Control Bump(SCB)located on transonic compressor rotor blade suction side is implemented to shrink the region of boundary layer *** curved windward section of SCB with constant adverse pressure gradient is constructed ahead of passage shock-impingement point at design rotor speed of Rotor 37 to get the improved *** investigations on both two models have been conducted employing Reynolds-Averaged Navier-Stokes(RANS)method to reveal flow physics of *** and analyses on simulation results have also been carried out,showing that passage shock foot of baseline is replaced with a family of compression waves and a weaker shock foot for moderate adverse pressure gradient as well as suppression of boundary layer separations and secondary flow of low-momentum fluid within boundary *** is found that adiabatic efficiency and total pressure ratio of improved blade exceeds those of baseline at 95%-100%design rotor speed,and then slightly worsens with decrease of rotatory speed till both equal below 60%rated *** investigated conclusion implies a potential promise for future practical applications of SCB in both transonic and supersonic compressors.

关键词： Flow separation Passage shock shock Control Bump(SCB) shock wave/boundary layer interaction Transonic compressors

Numerical evaluation of passive control of shock wave/boundary layer interaction on NACA0012 airfoil using jagged wall

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Acta Mechanica Sinica 2016年第5期32卷 792-804页

作者： Mojtaba Dehghan Manshadi Ramin Rabani Mechanical and Aerospace Engineering Department Malek-Ashtar University of TechnologyEsfahan Iran Department of Mechanical Engineering Yazd University

shock formation due to flow compressibility and its interaction with boundary layers has adverse effects on aerodynamic characteristics, such as drag increase and flow separation. The objective of this paper is to appraise the practicability of weakening shock waves and, hence, reducing the wave drag in transonic flight regime using a two-dimensional jagged wall and thereby to gain an appropriate jagged wall shape for future empirical study. Different shapes of the jagged wall, including rectangular, circular, and triangular shapes, were employed. The numerical method was validated by experimental and numerical studies involving transonic flow over the NACA0012 airfoil, and the results presented here closely match previous experimental and numerical results. The impact of parameters, including shape and the length-to-spacing ratio of a jagged wall, was studied on aerodynamic forces and flow field. The results revealed that applying a jagged wall method on the upper surface of an airfoil changes the shock structure significantly and disintegrates it, which in turn leads to a decrease in wave drag. It was also found that the maximum drag coefficient decrease of around 17 % occurs with a triangular shape, while the maximum increase in aerodynamic efficiency（lift-to-drag ratio）of around 10 % happens with a rectangular shape at an angle of attack of 2.26？.

关键词： Jagged wall Passive flow control shock wave/boundary layer interaction Aerodynamic efficiency

Unsteadiness of shock wave/boundary layer interaction in Supersonic Cascade

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Journal of Thermal Science 1996年第4期5卷 243-247页

作者： Hong－JuanWang Ke－JunCai 等 InstituteofEngieeringThermophysics ChineseAcademyofSciencesBeijing100080China TsinghuaUniversity Be

An investigation of the passive control of shock wave/boundary laper interaction for reducing the amplitude of the shock oscillation was conducted on the circular arc-wedge (CW) profile cascade in a 220×290 mm transonic compressor cascade wind tunnel. A perforated surface with a cavity beneath it was positioned on the suction surface of the blade at the location of shock impingement. The Schlieren and high-speed photographs for flow over perforated blade are presented and compared with the results for solid blades. With the perforated surface, the high-speed photographs indicated an significant suppression of shock oscillation.

关键词： shock wave/boundary layer interaction passive control of shock wave/boundary layerinteraction (PSBLI) λ shock wave shock oscillation.

Effects of Number of Bleed Holes on shock-wave/boundary-layer interactions in a Transonic Compressor Stator

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Journal of Thermal Science 2024年第2期33卷 611-624页

作者： LI Bai ZHOU Xun LUO Lei DU Wei School of Energy Science and Engineering Harbin Institute of TechnologyHarbin150001China

An extensive numerical investigation is conducted to characterize the flow separation control in a transonic compressor cascade with a porous *** bleed holes are arranged on the suction surface in a single row,two staggered rows and three staggered *** each bleed scheme,five bleed pressure ratios are examined at an inlet Mach number of *** results indicate that the aerodynamic performance of the cascade is significantly improved by the porous *** the single-row scheme,the maximum reduction in total pressure losses is 57%.For the two-staggered-row and three-staggered-row schemes,there is an optimal bleed pressure ratio of 1.0,and the maximum reductions in total pressure loss are 68% and 75%,*** low loss in the cascade is due to the well-controlled boundary *** new local supersonic region created by the bleed hole is the key reason for the improved boundary *** vortex induced by side bleeding provides another mechanism for delaying flow *** the bleed holes could create multiple local supersonic regions,which reduce the range of the adverse pressure gradient that the boundary layer needs to *** is the reason why cascades with more bleed holes perform better.

关键词： transonic compressor stator shock wave/boundary layer interaction porous bleed number of bleed holes

STUDY OF SWEPT shock wave AND boundary layer interactionS

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Chinese Journal of Aeronautics 1998年第4期11卷 2-10页

作者：邓学蓥 Institute of Fluid Mechanics Beijing Univ. Aero. and Astronaut. Beijing 100083 China

This paper presents briefly the recent progress on study of swept shock wave/boundary layer interactions with emphasis on application of zonal analysis and correlation analysis to them. Based on the zonal analysis an overall framework of complicated interaction flow structure including both surface flowfield and space flowfield is discussed. Based on correlation analysis the conical interactions induced by four families of shock wave generators have been discussed in detail. Some control parameter and physical mechanism of conical interaction have been revealed. Finally some aspects of the problem and the prospects for future work are suggested.

关键词： swept shock wave shock wave/boundary layer interaction zonal analysis correlation analysis

Flow Control Effect of Spanwise Distributed Pulsed Arc Discharge Plasma Actuation on Supersonic Compressor Cascade Flow

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Journal of Thermal Science 2022年第5期31卷 1723-1733页

作者： SHENG Jiaming WU Yun ZHANG Haideng WANG Yizhou TANG Mengxiao Science and Technology on Plasma Dynamic Laboratory Air Force Engineering UniversityXi’an710038China Institute of Aero-engine School of Mechanical EngineeringXi’an Jiaotong UniversityXi’an710049China

To achieve efficient control of supersonic compressor cascade flow,a type of spanwise distributed pulsed arc discharge plasma actuation(PADPA)was *** simulate the influences of PADPA on the flow field,a phenomenological model was ***,the flow control effects of PADPA on supersonic compressor cascade flow were researched *** results show that under low static pressure ratio condition,the compressive wave induced by PADPA reduced the intensity of the passage shock wave,which eventually reduced shock wave *** was also found that PADPA produced an adverse pressure gradient(pre-compression effect)around the actuation location,which reduced the strength of the high adverse pressure gradient induced by the passage shock *** airflow on both sides of the actuation location was accelerated by PADPA owing to the spanwise distributed ***,it improved the ability of the boundary layer to resist the effect of the adverse pressure gradient and reduced the separation ***,the total pressure loss was reduced by 6.8%.Under high pressure ratio condition,the effect of PADPA on the suction side controlling the large separation of the boundary layer was *** total pressure loss also increased slightly.

关键词： plasma flow control supersonic cascade shock wave/boundary layer interaction numerical simulation

ZONAL AND CORRELATION ANALYSIS IN SWEPT shock/boundary layer interactionS

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ZONAL AND CORRELATION ANALYSIS IN SWEPT SHOCK/BOUNDARY LAYER...