A Computation and Comparison With Measurements of Transonic Flow in an Axial Compressor Stage With Shock and Boundary Layer Interaction

1982 ◽  
Vol 104 (2) ◽  
pp. 510-515 ◽  
Author(s):  
U. K. Singh
Keyword(s):  
Boundary Layer ◽  
Mach Number ◽  
Axial Flow ◽  
Surface Boundary ◽  
Compressor Stage ◽  
Blade Surface ◽  
Layer Interaction ◽  
Flow Adjustment ◽  
Displacement Thickness ◽  
Boundary Layer Interaction

The flow field within a transonic axial flow compressor stage has been computed using a three-dimensional time-marching technique. Limited viscous effects are considered by including a calculation of the blade surface boundary layers. The boundary layer calculation forms an integral part of the whole computation scheme, which consists of, respectively: (i) inviscid Mach number calculations, (ii) blade surface boundary layer displacement thickness calculations, (iii) inviscid Mach number calculations with mass flow adjustment (based on the calculated displacement thicknesses) on the blade surfaces. The boundary layer computation is done by using integral calculation methods and has specifically been developed to account for a shock and boundary layer interaction (should one exist). Comparisons are made with measured results obtained with an advanced laser velocimeter. The calculated Mach number contours are in extremely good agreement with the experimental results. It is concluded that the calculation technique is a useful tool in the design of transonic axial flow turbomachines.

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.

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.

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.

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