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Research Papers

The Present Challenge of Transonic Compressor Blade Design

[+] Author and Article Information
Alexander Hergt

German Aerospace Center (DLR),
Institute of Propulsion Technology,
51147 Cologne, Germany
e-mail: alexander.hergt@dlr.de

J. Klinner

German Aerospace Center (DLR),
Institute of Propulsion Technology,
51147 Cologne, Germany
e-mail: joachim.klinner@dlr.de

J. Wellner

German Aerospace Center (DLR),
Institute of Propulsion Technology,
51147 Cologne, Germany
e-mail: jens.wellner@dlr.de

C. Willert

German Aerospace Center (DLR),
Institute of Propulsion Technology,
51147 Cologne, Germany
e-mail: chris.willert@dlr.de

S. Grund

German Aerospace Center (DLR),
Institute of Propulsion Technology,
51147 Cologne, Germany
e-mail: sebastian.grund@dlr.de

W. Steinert

German Aerospace Center (DLR),
Institute of Propulsion Technology,
51147 Cologne, Germany
e-mail: Wolfgang.Steinert@dlr.de

M. Beversdorff

German Aerospace Center (DLR),
Institute of Propulsion Technology,
51147 Cologne, Germany
e-mail: manfred.beversdorff@dlr.de

1Corresponding author.

Contributed by the International Gas Turbine Institute (IGTI) of ASME for publication in the Journal of Turbomachinery. Manuscript received September 11, 2018; final manuscript received March 26, 2019; published online May 30, 2019. Assoc. Editor: Kenneth Hall.

J. Turbomach 141(9), 091004 (May 30, 2019) (12 pages) Paper No: TURBO-18-1244; doi: 10.1115/1.4043329 History: Received September 11, 2018; Accepted March 27, 2019

The flow through a transonic compressor cascade shows a very complex structure due to the occurring shock waves. In addition, the interaction of these shock waves with the blade boundary layer inherently leads to a very unsteady flow behavior. The aim of the current investigation is to quantify this behavior and its influence on the cascade performance as well as to describe the occurring transonic flow phenomena in detail. Therefore, an extensive experimental investigation of the flow in a transonic compressor cascade has been conducted within the transonic cascade wind tunnel of DLR Institute of Propulsion Technology at Cologne. In this process, the flow phenomena were thoroughly examined for an inflow Mach number of 1.21. The experiments investigate both the laminar and the turbulent shock wave boundary layer interaction within the blade passage and the resulting unsteady behavior. The experiments show a fluctuation range of the passage shock wave of about 10% chord for both cases, which is directly linked with a change of the inflow angle and of the operating point of the cascade. Thereafter, Reynolds-averaged Navier–Stokes (RANS) simulations have been performed aiming at the verification of the reproducibility of the experimentally examined flow behavior. Here, it is observed that the dominant flow effects are not reproduced by a steady numerical simulation. Therefore, a further unsteady simulation has been carried out to capture the unsteady flow behavior. The results from this simulation show that the fluctuation of the passage shock wave can be reproduced but not in the correct magnitude. This leads to a remaining weak point within the design process of transonic compressor blades because the working range will be overpredicted. The resulting conclusion of this study is that the use of scale-resolving methods such as LES or the application of DNS is necessary to correctly predict unsteadiness of the transonic cascade flow and its impact on the cascade performance.

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Figures

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Fig. 1

Transonic cascade with planar endwall

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Fig. 2

Cross-sectional drawing of the test section

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Fig. 3

Definition of the measurement planes and cascade parameters (left) and design of the endwall boundary layer suction slots (right)

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Fig. 4

Schematic diagram of the shock system and locations of PIV measurement regions: region A, distribution of the passage shock position; region B, measurement of SBLI and separation region

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Fig. 5

Single PIV shots of the SBLI for the turbulent case (vector at u > 200 m/s are clipped); nearly attached flow and weak oblique shock

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Fig. 6

Single PIV shots of the SBLI for the turbulent case (vector at u > 200 m/s are clipped); large flow separation with strong oblique shock and transitional separation bubble including reverse flow

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Fig. 7

Inflow angle distribution of the laminar and turbulent case (L2F-measurement plane at midspan)

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Fig. 8

Isentropic Mach number distribution at midspan of the laminar and turbulent case

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Fig. 9

Liquid crystal measurement on the blade suction side of the laminar (top) and turbulent (bottom) case

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Fig. 10

Total pressure ratio distribution in the wake behind the cascade of the laminar and turbulent case (midspan)

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Fig. 11

Schlieren pattern (top) and schematic diagram of the shock structure (bottom) of the laminar case

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Fig. 12

Schlieren pattern (top) and schematic diagram of the shock structure (bottom) of the turbulent case

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Fig. 13

Statistical analysis of the shock movement at PIV region A

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Fig. 14

Conditional averaged PIV results at ROI B, front shock position (top), mid shock position (middle), and rear shock position (bottom)

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Fig. 15

Conditional averaged PIV results at ROI B, front shock position (top), mid shock position (middle), and rear shock position (bottom)

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Fig. 16

Spectrum of the shock wave oscillation of the laminar and turbulent case

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Fig. 17

Principle sketch of the numerical and experimental shock positions at different operation points: (a) numerical aerodynamic design point, (b) numerical stall onset, (c) definition of numerical stall margin, (d) experimental aerodynamic design point, (e) experimental stall onset, and (f) comparison of the numerical and experimental stall margin

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Fig. 18

Mach number distribution of one instantaneous time with the fine mesh; black line indicates the position of the recorded data

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Fig. 19

Experimental (top) and numerical (bottom) Schlieren patterns of the laminar case

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Fig. 20

Temporal evolution of the Mach number: coarse grid configuration

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Fig. 21

Temporal evolution of the Mach number: fine grid configuration

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