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Research Papers

Numerical Investigation of End Wall Boundary Layer Removal on Highly Loaded Axial Compressor Blade Rows

[+] Author and Article Information
V. Gümmer

Compressor Engineering, Rolls-Royce Deutschland Ltd. & Co. KG, Eschenweg 11, D-15827 Dahlewitz, Germanyvolker.guemmer@rolls-royce.com

M. Goller, M. Swoboda

Compressor Engineering, Rolls-Royce Deutschland Ltd. & Co. KG, Eschenweg 11, D-15827 Dahlewitz, Germany

J. Turbomach 130(1), 011015 (Jan 25, 2008) (9 pages) doi:10.1115/1.2749297 History: Received February 09, 2006; Revised September 22, 2006; Published January 25, 2008

This paper presents results of numerical investigations carried out to explore the benefit of end wall boundary layer removal from critical regions of highly loaded axial compressor blade rows. At the loading level of modern aero engine compressors, the performance is primarily determined by three-dimensional (3D) flow phenomena occurring in the end wall regions. Three-dimensional Navier–Stokes simulations were conducted on both a rotor and a stator test case in order to evaluate the basic effects and the practical value of bleeding air from specific locations at the casing end wall. The results of the numerical survey demonstrated substantial benefits of relatively small bleed rates to the local flow field and to the performance of the two blade rows. On the rotor, the boundary layer fluid was removed from the main flow path through an axisymmetric slot in the casing over the rotor tip. This proved to give some control over the tip leakage vortex flow and the associated loss generation. On the stator, the boundary layer fluid was taken from the flow path through a single bleed hole within the passage. Two alternative off-take configurations were evaluated, revealing a large impact of the bleed hole shape and the location on the cross-passage flow and the suction side corner separation. On both blade rows investigated, rotor and stator, the boundary layer removal resulted in a reduction of the local reverse flow, blockage, and losses in the respective near-casing region. This paper gives insight into changes occurring in the 3D passage flow field near the casing and summarizes the effects on the radial matching and pitchwise-averaged performance parameters, namely loss and deviation of the rotor and stator when suction is active. Primary focus is put on the aerodynamics in the blade rows in the main flow path; details of the internal flow structure within the bleed off-take cavities/ports are not discussed here.

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Copyright © 2008 by American Society of Mechanical Engineers
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Figures

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Figure 1

Low-speed research compressor

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Figure 2

LSRC model grid topology

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Figure 3

Chordwise midgap velocity distributions, operation at DP with bleed off/on

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Figure 4

Contours of axial velocity and relative velocity streamlines at midgap radial height

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Figure 5

Development of near-casing relative total pressure contours through the rotor passage, bleed off

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Figure 6

Development of near-casing relative total pressure contours through the rotor passage, bleed on

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Figure 7

Rotor pitchwise-averaged deviation and loss coefficient

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Figure 8

Embedded stator arrangement

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Figure 9

Stator grid model featuring different bleed geometries: (a) circular OT; (b) tailored OT

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Figure 10

Stator surface Mach number contours and streak lines, stator exit total pressure contours, no OT

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Figure 11

Stator surface Mach number contours and streak lines, stator exit total pressure contours, circular OT

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Figure 12

Stator surface Mach number contours and streak lines, stator exit total pressure contours, tailored OT

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Figure 13

Stator pitchwise-averaged deviation and loss coefficient

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