Numerical simulation of compressor cascade flow with strong shock-wave boundary layer interaction

1997 ◽  
Author(s):  
B. Kuesters ◽  
H. Schreiber ◽  
B. Kuesters ◽  
H. Schreiber
Keyword(s):  
Numerical Simulation ◽  
Boundary Layer ◽  
Shock Wave ◽  
Strong Shock ◽  
Strong Shock Wave ◽  
Compressor Cascade ◽  
Layer Interaction ◽  
Cascade Flow ◽  
Boundary Layer Interaction ◽  
Wave Boundary

An Investigation of a Strong Shock-Wave Turbulent Boundary Layer Interaction in a Supersonic Compressor Cascade

1992 ◽  
Vol 114 (3) ◽  
pp. 494-503 ◽  
Author(s):  
H. A. Schreiber ◽  
H. Starken
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Mach Number ◽  
Turbulent Boundary Layer ◽  
Strong Shock ◽  
Strong Shock Wave ◽  
Compressor Cascade ◽  
Blade Passage ◽  
Layer Interaction ◽  
Boundary Layer Interaction

Experiments have been performed in a supersonic cascade facility to elucidate the fluid dynamic phenomena and loss mechanism of a strong shock-wave turbulent boundary layer interaction in a compressor cascade. The cascade geometry is typical for a transonic fan tip section that operates with a relative inlet Mach number of 1.5, a flow turning of about 3 deg, and a static pressure ratio of 2.15. The strong oblique and partly normal blade passage shock-wave with a preshock Mach number level of 1.42 to 1.52 induces a turbulent boundary layer separation on the blade suction surface. The free-stream Reynolds number based on chord length was about 2.7 × 106. Cascade overall performance, blade surface pressure distributions, Schlieren photographs, and surface visualizations are presented. Detailed Mach number and flow direction profiles of the interaction region (lambda shock) and the corresponding boundary layer have been determined using a Laser-2-Focus anemometer. The obtained results indicated that the axial blade passage stream sheet contraction (axial velocity density ratio) has a significant influence on the mechanism of strong interaction and the resulting total pressure losses.

scholarly journals An Investigation of a Strong Shock-Wave Turbulent Boundary Layer Interaction in a Supersonic Compressor Cascade

1991 ◽  
Author(s):  
H. A. Schreiber ◽  
H. Starken
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Mach Number ◽  
Turbulent Boundary Layer ◽  
Strong Shock ◽  
Strong Shock Wave ◽  
Compressor Cascade ◽  
Blade Passage ◽  
Layer Interaction ◽  
Boundary Layer Interaction

Experiments have been performed in a Supersonic cascade facility to elucidate the fluid dynamic phenomena and loss mechanism of a strong shock-wave turbulent boundary layer interaction in a compressor cascade. The cascade geometry is typical for a transonic fan tip section that operates with a relative inlet Mach number of 1.5, a flow turning of about 3 degrees, and a static pressure ratio of 2.15. The strong oblique and partly normal blade passage shock-wave with a pre-shock Mach number level of 1.42 to 1.52 induces a turbulent boundary layer separation on the blade suction surface. Freestream Reynolds number based on chord length was about 2.7×106. Cascade overall performance, blade surface pressure distributions, Schlieren photographs, and surface visualisations are presented. Detailed Mach number and flow direction profiles of the interaction region (lambda shock) and the corresponding boundary layer have been determined using a Laser-2-Focus anemometer. The obtained results indicated that the axial blade passage stream sheet contraction (axial velocity density ratio) has a significant influence on the mechanism of strong interaction and the resulting total pressure losses.

Unsteady Numerical Simulation in a Supersonic Compressor Cascade with a Strong Shock Wave

2019 ◽  
Vol 0 (0) ◽  
Author(s):  
Shaowen Chen ◽  
Qinghe Meng ◽  
Yueqi Liu ◽  
Hongyan Liu ◽  
Songtao Wang
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Low Frequency ◽  
Strong Shock ◽  
Oblique Shock Wave ◽  
Compressor Cascade ◽  
Eddy Simulation ◽  
Boundary Layer Interaction ◽  
Unsteady Numerical Simulation ◽  
Wave Boundary

Abstract The most important flow behaviour of supersonic compressor cascades is the shock wave boundary layer interaction (SWBLI). Large eddy simulation (LES) and multiple analysing methods are applied in current study to capture more details of the flow field. It is noted that the LES can catch the dual peaks feature near the SWBLI region with respect to the experimental results. Besides, SWBLI is not only the main losses source in the cascade, but also the most important origin of the unsteadiness behaviour. The high frequency signals correspond to the coherent structure in the boundary layer and dissipate downstream in the cascade, while the low frequency signals relate to the motion of the reflection point of the passage oblique shock wave and dominate the frequency spectrum downstream.

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