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

Schreiber, H. A.; Starken, H.

doi:10.1115/1.2929170

Cited by 23 publications

(14 citation statements)

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“…The 4 th stage is from the upstream of the second passage shock to the outlet location of the supersonic cascade. According to the experimental results, 9,22 the second passage shock, especially as standing close to the exit of the cascade passage, tilts or presents as lambda shock near the blade surface because of the strong shock/boundary layer interaction (Figure 1). Hence, for most of the previous shock loss models, [5][6][7] modeling of the second passage shock as an entire normal shock may result in an overpredicted shock loss level.…”

Section: ¼ L Np L O ð1þmentioning

confidence: 99%

A new method for rapid shock loss evaluation and reduction for the optimization design of a supersonic compressor cascade

Liu

Shi

2017

Proceedings of the Institution of Mechanical Engineers, Part G:

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A one-dimensional analytical shock loss prediction method was proposed to tailor the shock system, i.e. the strength of the first and second passage shock, and reduce the shock loss in a supersonic cascade. To develop the one-dimensional analytical model, the shock system in a supersonic cascade was divided into four processes which can be seen in most supersonic compressor cascade, i.e. the flow upstream the extending-external shock, the flow between the extendingexternal shock and the first passage shock, the accelerating flow from the first passage shock to the second passage shock, and the flow downstream of the second passage shock. Based on some flow assumptions and experimental empirical correlations, the complex flows, containing the shock system, in the blade passage of a supersonic compressor cascade could be described with one-dimensional relationships, which can be used to predict shock losses along the flow passage rapidly and determine the shock system improving direction for achieving lower shock loss while keeping the same cascade static pressure ratio. In order to validate the one-dimensional analytical method, the shock system of two supersonic cascades ARL-SL19 and DLR-PAV-1.5 are modified based on the analysis of the model. The modified cascades achieved about 29% and 25% reduction of shock loss at redesign point compared with baseline cascades, respectively.

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Section: ¼ L Np L O ð1þmentioning

confidence: 99%

A new method for rapid shock loss evaluation and reduction for the optimization design of a supersonic compressor cascade

Liu

Shi

2017

Proceedings of the Institution of Mechanical Engineers, Part G:

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“…This is in contrast to strong shock interactions with turbulent boundary-layers (e.g. Schreiber and Starken, 1991). The oil flow visualization also illustrates the importance of the experimental setup for the investigation of shock/boundary-layer interactions under turbomachinery conditions.…”

Section: Description Of Interaction Processmentioning

confidence: 86%

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

Bell

Fottner

1995

Volume 1: Turbomachinery

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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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“…Based on the experimental results, 11 the wedge angle 2 of the upstream branch shock about 8.5 and decelerated from Mach number 1.5 to about 1.19. The wedge angle 2 is calculated by equation (12) in this article 2 % 8:5 ðÀ2Ma 2 2 þ 7M À 5Þ ð 12Þ…”

Section: Shock Loss Of the Terminal Passage Shockmentioning

confidence: 99%

Shock loss model and blade profile optimization design of a supersonic cascade

Sun

Yang

Zhao

2014

Proceedings of the Institution of Mechanical Engineers, Part G:

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With the improvement of single-stage compressor load, the inflow velocity and relative Mach number at blade tip of compressor are further increased. Using supersonic blade profile to design highly loaded compressor is an effective method to satisfy the requirements of the compressor aerodynamic design. In the internal flow of a blade row, detached shock is caused at the leading edge of supersonic cascade, and complex passage shock structure is produced by shock and boundary layer interaction at blade surface. Establishing shock loss calculation method for detached shock and passage shock and optimizing the blade profile according to the shock loss mechanisms are effective means to design high-efficiency compressors. The earlier studies of shock loss calculation model in supersonic compressor cascade have shown that shock loss of the detached shock at leading edge is related to leading edge radius and inlet Mach number. Therefore, several shock loss models for detached shock are established and proved to be effective. In previous studies on passage shock loss, the changes in geometry of passage shock, which is caused by shock and boundary layer interaction, are not fully considered. So, it is necessary to do further study on passage shock loss model. Based on the internal flow characteristics in supersonic cascade, passage shock loss model is proposed. Physics-based passage shock geometry, which is determined by static pressure rise, Mach number before wave fronts and blade profile curvature, is applied for this passage shock loss model. Near the blade surface, passage shock geometry is transformed into lambda-type shock that is composed of two branch oblique shocks, and oblique shock relationships are used to calculate their total pressure losses. Away from the blade surface, passage shock geometry is very close to the normal shock shape, so that normal shock relationships are used for loss calculation.

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

Cited by 23 publications

References 0 publications

A new method for rapid shock loss evaluation and reduction for the optimization design of a supersonic compressor cascade

A new method for rapid shock loss evaluation and reduction for the optimization design of a supersonic compressor cascade

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

Shock loss model and blade profile optimization design of a supersonic cascade

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