Effects of Mach number on upstream influence in sharp fin-induced shock wave turbulent boundary layer interaction

1984 ◽  
Author(s):  
D. DOLLING
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Mach Number ◽  
Turbulent Boundary Layer ◽  
Layer Interaction ◽  
Boundary Layer Interaction

Numerical Investigation on Suppression of 3-Dimensional Shock Wave/Turbulent Boundary Layer Interaction by Passive Cavity

2003 ◽  
Author(s):  
Kazuyuki Toda ◽  
Shinsuke Dambara ◽  
Makoto Yamamoto ◽  
Shinji Honami ◽  
Nobuyuki Akahoshi
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Mach Number ◽  
Turbulent Boundary Layer ◽  
Passive Control ◽  
Interaction Region ◽  
Mean Flow ◽  
3 Dimensional ◽  
Layer Interaction ◽  
Boundary Layer Interaction

Suppression of three-dimensional shock wave/turbulent boundary layer interaction is one of the important subjects on supersonic air intake. In the present study, the passive control of 2- and 3-dimensional shock wave/turbulent boundary layer interactions is considered. First, computations are performed for two-dimensional flow field at freestream Mach number of 2.46 with various passive cavities beneath the interaction region. The results suggest that the parallel blowing from a cavity to the mean flow with a guide plate can highly keep the interaction region narrow. Next, the most suitable cavity shape clarified in the 2-dimensional computations is applied to the 3-dimensional swept shock wave/turbulent boundary layer interaction at Mach number of 3.11. It is exhibited that the blowing direction is important, and the effect of passive cavity is nearly the same as the bleeding.

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.

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