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Research Papers

A Method for Evaluating the Aerodynamic Stability of Multistage Axial-Flow Compressors

[+] Author and Article Information
Tsuguji Nakano, Andy Breeze-Stringfellow

 GE Aviation, Cincinnati, OH 45215-1988

J. Turbomach 133(3), 031004 (Nov 11, 2010) (11 pages) doi:10.1115/1.4001225 History: Received August 05, 2009; Revised October 05, 2009; Published November 11, 2010; Online November 11, 2010

A new simple engineering parameter to evaluate the stability of multistage axial compressors has been derived. It is based on the stability analysis for a small circumferential disturbance imposed on the steady-state flow field. The analytical model assumes that the flow field is two dimensional and incompressible in the ducts between blade rows although the steady-state density is permitted to change across the blade rows. The resulting stall parameter contains terms that relate to the slope of the pressure rise characteristic of the blade rows and the inertia effects of the fluid in the blade rows and ducts. The parameter leads to the classical stability criteria based on the slope of the overall total to static pressure rise coefficient in the limit where constant density and constant blade rotational speed are assumed across the compressor. The proposed stall parameter has been calculated for three different multistage axial flow compressors, and the results indicate that the parameter has a strong correlation with the measured stability of the compressors. The good correlation with the test data demonstrates that the newly derived stall parameter captures much of the fundamental physics of instability inception in multistage compressors, and that it can be a good guideline for designers and engineers needing to evaluate the stability boundary of multistage machines.

Copyright © 2011 by American Society of Mechanical Engineers
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References

Figures

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Figure 13

Stall parameter without the duct effects

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Figure 14

Stall parameter without the blade inertia and duct effects

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Figure 15

Stall parameter including estimated stage10 pressure rises

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Figure 16

First harmonic disturbance propagation speed

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Figure 17

Propagation speed of higher harmonic disturbances

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Figure 18

Parameter calculated with the fixed rotor exit relative flow angle

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Figure 4

Schematic of the compressor flow field model

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Figure 3

Slopes of the overall pressure rise

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Figure 2

Six-stage compressor overall pressure rise

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Figure 1

Ten-stage compressor overall pressure rise

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Figure 9

Six-stage compressor 97%Nc rotor performances

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Figure 10

Six-stage compressor 92%Nc rotor performances

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Figure 11

Six-stage compressor 80%Nc rotor performances

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Figure 12

Stall parameter (Eq. 22)

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Figure 5

Schematic of the blade-row model

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Figure 6

Ten-stage compressor 100%Nc rotor performances

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Figure 7

Ten-stage compressor 90%Nc rotor performances

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Figure 8

Ten-stage compressor 80%Nc rotor performances

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