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Research Papers

High Resolution RANS Nonlinear Harmonic Study of Stage 67 Tip Injection Physics

[+] Author and Article Information
Allan D. Grosvenor

Ramgen Power Systems, LLC,
Bellevue, WA 98005
e-mail: allan.grosvenor@gmail.com

Gregory S. Rixon, Logan M. Sailer

Ramgen Power Systems, LLC,
Bellevue, WA 98005

Michael A. Matheson

Oak Ridge National Laboratory,
Oak Ridge TN 37831

David P. Gutzwiller, Alain Demeulenaere

NumecaUSA,
San Francisco, CA 94109

Mathieu Gontier

Numeca International,
Brussels B-1170, Belgium

Anthony J. Strazisar

AJS Aero Incorporated,
Chesterland, OH 44026-1722

1Corresponding author.

Contributed by the International Gas Turbine Institute (IGTI) Division of ASME for publication in the JOURNAL OF TURBOMACHINERY. Manuscript received July 29, 2014; final manuscript received August 17, 2014; published online November 18, 2014. Editor: Ronald Bunker.

The United States Government retains, and by accepting the article for publication, the publisher acknowledges that the United States Government retains, a nonexclusive, paid-up, irrevocable, worldwide license to publish or reproduce the published form of this work, or allow others to do so, for United States government purposes.

J. Turbomach 137(5), 051005 (May 01, 2015) (13 pages) Paper No: TURBO-14-1188; doi: 10.1115/1.4028550 History: Received July 29, 2014; Revised August 17, 2014; Online November 18, 2014

Numerical prediction of the Stage 67 transonic fan stage employing wall jet tip injection flow control and study of the physical mechanisms leading to stall suppression and stability enhancement afforded by endwall recirculation/injection is the focus of this paper. Reynolds averaged Navier–Stokes (RANS) computations were used to perform detailed analysis of the Stage 67 configuration experimentally tested at NASA's Glenn Research Center in 2004. Time varying predictions of the stage plus recirculation and injection flowpath were executed utilizing the nonlinear harmonic (NLH) approach. Significantly higher grid resolution per passage was achieved than what has been generally employed in prior reported numerical studies of spike stall phenomena in transonic compressors. This paper focuses on characterizing the physics of spike stall embryonic stage phenomena and the influence of tip injection, resulting in experimentally and numerically demonstrated stall suppression.

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Figures

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Fig. 1

Geometrical configuration of Stage 67 transonic fan stage with tip injection. (a) Labeled stage components, (b) schematic view of rotor shock waves at near-stall, (c) and (d) two magnifications shown, and (e) view of viscous distortion displayed at >90% span, colored by static pressure.

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Fig. 2

Views of Stage 67 computational grid. Level 111 shown here: (a) view of overall grid, only one passage is meshed per row, (b) rotor passage grid displayed at casing, (c) rotor grid resolution surrounding rotor LE at casing, (d) meridional view of grid distribution—passage mesh displayed at periodic boundaries, and (e) magnified view of (d) at injector nozzle intersection with casing, identifying rotor/stator interface location.

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Fig. 8

Interaction zone phenomena at near-stall condition. (a) and (b) Contours of relative Mach number at two spanwise positions, midjet and 90% span—magnified from Fig. 5, (c) λ2 isosurfaces colored by static pressure—magnified from Fig. 7(c), (d) shape factor (or form parameter) at casing in injected and noninjected passages, and (e) 3D shock structures plus λ2 isosurfaces—domain shows only axial portion through rotor, and spanwise portion above 90% span calculated based on axial velocity profiles from 111 near-stall condition.

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Fig. 9

Vortex identification plotted from verge-of-stall calculation. λ2 isosurfaces colored by circumferential component of vorticity. Blue indicates opposite sense of rotation than red.

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Fig. 3

Comparison of predicted performance and injection properties with experiment, for varying grid resolution, and cases with and without injection

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Fig. 4

Circumferential cutting planes displaying jet/TC leakage vortex interaction at the near-stall (111 grid level) and verge-of-stall (222 grid level) conditions via Mach number calculated in absolute reference frame (left), relative total pressure (middle), axial velocity (right)

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Fig. 5

Blade-to-blade views of quantities plotted over spanwise cutting planes. (a) Top: taken at jet-center, bottom: 90% span and (b) shock structures traced at 90% span.

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Fig. 6

Velocity triangles plotted from 222 verge-of-stall calculation. Top: taken at jet-center, and bottom: 95% span.

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Fig. 7

Vortex identification based on λ2 demonstrating injection influence on TC vortex at two operating conditions. Isosurfaces of λ2 are shown, colored by static pressure. View is constrained to enable observation of rotor tip/casing zone by removing everything outside of spanwise 90%-casing and axial range from rotor LE to TE—injector surfaces also shown. Top: 3D viewing direction shown in Fig. 1, demonstrating formation of radial structures, spanwise depth of vortex structures, and rotor tip suction surface separating flow structures (a) near-stall and (b) verge-of-stall. Bottom: view direction looking from centerline outward toward casing (c) near-stall and (d) verge-of-stall.

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Fig. 10

Comparison of predicted flowfields at the near-stall condition, displayed for three different grid resolutions

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