Passive control of oblique shock/boundary layer interaction at high Mach number

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.

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

Volume 5: Turbo Expo 2004, Parts A and B ◽

10.1115/gt2004-53435 ◽

2004 ◽

Cited By ~ 4

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.

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

Volume 2: Symposia, Parts A, B, and C ◽

10.1115/fedsm2003-45636 ◽

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.

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