Research Papers

An Improved Mixing-Plane Method for Analyzing Steady Flow Through Multiple-Blade-Row Turbomachines

[+] Author and Article Information
Ding Xi Wang

Siemens Industrial Turbomachinery Ltd.,
Waterside South,
Lincoln LN5 7FD, UK
e-mail: dingxi.wang@siemens.com

Contributed by the International Gas Turbine Institute (IGTI) of ASME for publication in the JOURNAL OF TURBOMACHINERY. Manuscript received May 28, 2013; final manuscript received November 19, 2013; published online January 31, 2014. Assoc. Editor: John Clark.

J. Turbomach 136(8), 081003 (Jan 31, 2014) (9 pages) Paper No: TURBO-13-1085; doi: 10.1115/1.4026170 History: Received May 28, 2013; Revised November 19, 2013

Presented in this paper is an improved method for dealing with a mixing plane that exists between computational domains of two adjacent blade rows of a multiple-blade-row turbomachine. The method makes use of the semidiscrete flow equation updating scheme to convert flux differences across an interrow interface to conservative flow variable incrementals then to characteristic variable perturbations. Therefore, the proposed method bears more physics and is much more robust than any known method of its kind. As a result, reverse flow can be accommodated by this method naturally without any special treatment. Two existing methods are also included to provide a clear illustration of the differences and advantages of the new method. Two numerical test cases using a transonic compressor stage are presented to investigate the robustness of the new method and its influence on solution convergence, accuracy, and time cost in comparison with the two existing methods.

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Fig. 1

Schematic of an artificial interface between a rotor and a stator (left) and the virtual control volume formed by displacing two adjacent domains (right)

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Fig. 2

Blade to blade view of the computational domain at the midspan

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Fig. 3

Meridional view of the computational domain

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Fig. 4

Average residual history of the energy equation at a near peak efficiency

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Fig. 5

Convergence history of inlet mass flow rate and the compressor adiabatic efficiency at a near peak efficiency point

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Fig. 6

Streamwise distribution of mass averaged entropy at a near peak efficiency

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Fig. 7

Mach number contours at the midspan of the compressor (solution from the new method, and the other two methods give identical results that are not provided to avoid redundancy)

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Fig. 8

Pressure ratio variation against mass flow rate at the compressor design speed

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Fig. 9

Adiabatic efficiency variation against pressure ratio at the compressor design speed

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Fig. 10

Average residual history of the energy equation for the reverse flow case

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Fig. 11

Convergence history of inlet/outlet mass flow rate and adiabatic efficiency for the reverse flow case

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Fig. 12

Axial velocity contours and streamlines on a meridional plane for the reverse flow case

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Fig. 13

Total temperature contours at the stator inlet plane for the reverse flow case

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Fig. 14

Total temperature (in the stationary frame of reference) contours at 98% span for the reverse flow case

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Fig. 15

Entropy contours at 98% span for the reverse flow case




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