scholarly journals Investigations of Shock/Boundary-Layer Interaction in a Highly Loaded Compressor Cascade

1995 ◽  
Author(s):  
Ralf M. Bell ◽  
Leonhard Fottner
Keyword(s):  
Boundary Layer ◽  
Reynolds Number ◽  
Mach Number ◽  
Flow Behavior ◽  
Compressor Cascade ◽  
Flow Angle ◽  
Layer Interaction ◽  
Shock Boundary Layer Interaction ◽  
Boundary Layer Interaction ◽  
Highly Loaded

Experimental investigations of the shock/boundary-layer interaction were carried out in a highly loaded compressor cascade under realistic turbomachinery conditions in order to improve the accuracy of semi-empirical flow and loss prediction methods. Different shock positions and strengths were obtained by variations of inlet flow angle and inlet Mach number. The free stream turbulence intensity, depending on the inlet Mach number, changed between 4% and 8%. The influence of the inlet Reynolds number based on blade chord is also examined for two different values (Re1=450000, 900000). Schlieren pictures of the transonic cascade flow reveal an unsteady flow behavior with different shock configurations, depending on the pre-shock Mach number. Wake distributions and boundary-layer measurements with the Laser two-focus velocimetry show that the increase of total pressure loss with increasing inlet Mach number is mainly due to the shock/boundary-layer interaction. The shock interaction with a laminar/transitional boundary-layer causes a wide streamwise pressure diffusion, clearly shown by profile pressure distributions. This has a strong influence on the flow outside of the boundary-layer presented by a quantitative Schlieren image. The transition process, investigated with the analysis of thin-film signals, is induced by the shock-wave and occurs above a separated-flow region. At the higher Reynolds number a shock-induced transition takes place without separation.

Detailed Flow Study of Mach Number 1.6 High Transonic Flow With a Shock Wave in a Pressure Ratio 11 Centrifugal Compressor Impeller

2004 ◽  
Vol 126 (4) ◽  
pp. 473-481 ◽  
Author(s):  
Hirotaka Higashimori ◽  
Kiyoshi Hasagawa ◽  
Kunio Sumida ◽  
Tooru Suita
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Mach Number ◽  
Centrifugal Compressor ◽  
Transonic Flow ◽  
Reverse Flow ◽  
Pressure Ratio ◽  
Layer Interaction ◽  
Shock Boundary Layer Interaction ◽  
Boundary Layer Interaction

Requirements for aeronautical gas turbine engines for helicopters include small size, low weight, high output, and low fuel consumption. In order to achieve these requirements, development work has been carried out on high efficiency and high pressure ratio compressors. As a result, we have developed a single stage centrifugal compressor with a pressure ratio of 11 for a 1000 shp class gas turbine. The centrifugal compressor is a high transonic compressor with an inlet Mach number of about 1.6. In high inlet Mach number compressors, the flow distortion due to the shock wave and the shock boundary layer interaction must have a large effect on the flow in the inducer. In order to ensure the reliability of aerodynamic design technology, the actual supersonic flow phenomena with a shock wave must be ascertained using measurement and Computational Fluid Dynamics (CFD). This report presents the measured results of the high transonic flow at the impeller inlet using Laser Doppler Velocimeter (LDV) and verification of CFD, with respect to the high transonic flow velocity distribution, pressure distribution, and shock boundary layer interaction at the inducer. The impeller inlet tangential velocity is about 460 m/s and the relative Mach number reaches about 1.6. Using a LDV, about 500 m/s relative velocity was measured preceding a steep deceleration of velocity. The following steep deceleration of velocity at the middle of blade pitch clarified the cause as being the pressure rise of a shock wave, through comparison with CFD as well as comparison with the pressure distribution measured using a high frequency pressure transducer. Furthermore, a reverse flow is measured in the vicinity of casing surface. It was clarified by comparison with CFD that the reverse flow is caused by the shock-boundary layer interaction. Generally CFD shows good agreement with the measured velocity distribution at the inducer and splitter inlet, except in the vicinity of the casing surface.

Detailed Flow Study of Mach Number 1.6 High Transonic Flow With a Shock Wave in a Pressure Ratio 11 Centrifugal Compressor Impeller

2004 ◽  
Author(s):  
Hirotaka Higashimori ◽  
Kiyoshi Hasagawa ◽  
Kunio Sumida ◽  
Tooru Suita
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Mach Number ◽  
Centrifugal Compressor ◽  
Transonic Flow ◽  
Reverse Flow ◽  
Pressure Ratio ◽  
Layer Interaction ◽  
Shock Boundary Layer Interaction ◽  
Boundary Layer Interaction

Requirements for aeronautical gas turbine engines for helicopters include small size, low weight, high output, and low fuel consumption. In order to achieve these requirements, development work has been carried out on high efficiency and high pressure ratio compressors. As a result, we have developed a single stage centrifugal compressor with a pressure ratio of 11 for a 1000 shp class gas turbine. The centrifugal compressor is a high transonic compressor with an inlet Mach number of about 1.6. In high inlet Mach number compressors, the flow distortion due to the shock wave and the shock boundary layer interaction must have a large effect on the flow in the inducer. In order to ensure the reliability of aerodynamic design technology, the actual supersonic flow phenomena with a shock wave must be ascertained using measurement and CFD. This report presents the measured results of the high transonic flow at the impeller inlet using LDV and verification of CFD, with respect to the high transonic flow velocity distribution, pressure distribution and shock boundary layer interaction at the inducer. The impeller inlet tangential velocity is about 460m/s and the relative Mach number reaches about 1.6. Using an LDV, about 500m/s relative velocity was measured preceding a steep deceleration of velocity. The following steep deceleration of velocity at the middle of blade pitch clarified the cause as being the pressure rise of a shock wave, through comparison with CFD as well as comparison with the pressure distribution measured using a high frequency pressure transducer. Furthermore, a reverse flow is measured in the vicinity of casing surface. It was clarified by comparison with CFD that the reverse flow is caused by the shock-boundary layer interaction. Generally CFD shows good agreement with the measured velocity distribution at the inducer and splitter inlet, except in the vicinity of the casing surface.

Role of Shock and Boundary Layer Separation on Unsteady Aerodynamic Characteristics of Oscillating Transonic Cascade

2003 ◽  
Author(s):  
Mizuho Aotsuka ◽  
Toshinori Watanabe ◽  
Yasuo Machina
Keyword(s):  
Boundary Layer ◽  
Pressure Fluctuations ◽  
Boundary Layer Separation ◽  
Aerodynamic Characteristics ◽  
Compressor Cascade ◽  
Layer Interaction ◽  
Unsteady Aerodynamic ◽  
Shock Boundary Layer Interaction ◽  
Boundary Layer Interaction ◽  
Shock Oscillation

The unsteady aerodynamic characteristics of an oscillating compressor cascade composed of Double-Circular-Arc airfoil blades were both experimentally and numerically studied under transonic flow conditions. The study aimed at clarifying the role of shock waves and boundary layer separation due to the shock boundary layer interaction on the vibration characteristics of the blades. The measurement of the unsteady aerodynamic moment on the blades was conducted in a transonic linear cascade tunnel using an influence coefficient method. The cascade was composed of seven DCA blades, the central one of which was an oscillating blade in a pitching mode. The unsteady moment was measured on the central blade as well as the two neighboring blades. The behavior of the shock waves was visualized through a schlieren technique. A quasi-three dimensional Navier-Stokes code was developed for the present numerical simulation of the unsteady flow fields around the oscillating blades. A k-ε turbulence model was utilized to adequately simulate the flow separation phenomena caused by the shock-boundary layer interaction. The experimental and numerical results complemented each other and enabled a detailed understanding of the unsteady aerodynamic behavior of the cascade. It was found that the surface pressure fluctuations induced by the shock oscillation were the governing factor for the unsteady aerodynamic moment acting on the blades. Such pressure fluctuations were primarily induced by the movement of impingement point of the shock on the blade surface. During the shock oscillation the separated region caused by the shock boundary layer interaction also oscillated along the blade surface, and induced additional pressure fluctuations. The shock oscillation and the movement of the separated region were found to play the principal role in the unsteady aerodynamic and vibration characteristics of the transonic compressor cascade.

scholarly journals Effect of wall temperature on separation bubble size in laminar hypersonic shock/boundary layer interaction flows

2019 ◽  
Vol 11 (11) ◽  
pp. 168781401988555 ◽  
Author(s):  
Amjad A Pasha ◽  
Khalid A Juhany
Keyword(s):  
Boundary Layer ◽  
Heat Flux ◽  
Mach Number ◽  
Wall Temperature ◽  
Separation Bubble ◽  
Layer Interaction ◽  
Shock Boundary Layer Interaction ◽  
Boundary Layer Interaction ◽  
Hypersonic Shock ◽  
Bubble Length

At hypersonic speeds, the external wall temperatures of an aerospace vehicle vary significantly. As a result, there is a considerable heat transfer variation between the boundary layer and the wall of the hypersonic vehicle. In this article, numerical computations are performed to investigate the effect of wall temperature on the separation bubble length in laminar hypersonic shock-wave/boundary-layer interaction flows over double-cone configuration at the Mach number of 12.2. The flow field is described in detail in terms of different shocks, expansion fans, shear layer and separation bubble. The variation of the Prandtl number has a negligible effect on the flow field and wall data. A specific heat ratio of less than 1.4 results in the better prediction of wall pressure and heat flux in the shock/boundary-layer interaction region. It is observed that as the wall temperature is increased, the separation bubble size and hence the separation shock length increases. The high firmness of the laminar boundary-layer at a high Mach number shows that the wall temperature in the shock/boundary-layer interaction region has little effect. The peak wall pressure and heat flux decrease with an increase in wall temperature. An estimation is developed between separation bubble length and wall temperature based on the computed results.

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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