Research Papers

Compressible Modal Instability Onset in an Aerodynamically Mistuned Transonic Fan

[+] Author and Article Information
Christoph Brandstetter

Ecole Centrale de Lyon,
Laboratoire de Mecanique des
Fluides et d'Acoustique—UMR CNRS 5509,
Univ Lyon,
Lyon, F-69134, France
e-mail: christoph.brandstetter@ec-lyon.fr

Benoit Paoletti, Xavier Ottavy

Ecole Centrale de Lyon,
Laboratoire de Mecanique des
Fluides et d'Acoustique—UMR CNRS 5509,
Univ Lyon,
Lyon, F-69134, France

Manuscript received July 29, 2018; final manuscript received December 11, 2018; published online January 16, 2019. Assoc. Editor: Rakesh Srivastava.

J. Turbomach 141(3), 031004 (Jan 16, 2019) (11 pages) Paper No: TURBO-18-1175; doi: 10.1115/1.4042310 History: Received July 29, 2018; Revised December 11, 2018

This paper describes observed modal oscillations arising from a feedback mechanism between an acoustic resonance in the exit flow channel and aerodynamic and aeroelastic disturbances in a transonic fan stage. During tests, the fan suffered from rotating stall and surge which were preceded by low frequency pressure fluctuations. Through a range of aerodynamic and aeromechanical instrumentations, it was possible to determine a clear chain of cause and effect, whereby geometrical asymmetries trigger local instabilities and modal oscillations through an interaction with the system acoustics. To the authors knowledge, this is the first time that modal oscillations occurring before stall are attributed to multiphysical interactions, showing that acoustic characteristics of the system can influence the aerodynamic as well as the aeromechanical stability of fans. This bears implications for the stability assessment of fans and compressors because first, the stability margin may be affected by standing waves generated in bypass ducts or combustion chambers, and second, geometrical variations of the rotor blades which are believed to be beneficial for aeromechanical stability may lead to complex coupling phenomena.

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Fig. 1

Test facility schematic

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Fig. 2

Facility core schematic

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Fig. 3

Fan stage in anechoic chamber; installed turbulence control system; rotatable microphone antenna

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Fig. 4

Measured fan stage characteristics for subsonic and transonic speedlines

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Fig. 5

Time history of unsteady wall pressure upstream of the rotor for all measured speedlines, stall onset at t = 0

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Fig. 6

Prestall spectrum and temporal evolution of selected modal amplitudes, sliding fft-window 0.8 s, Hanning, 50% overlap

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Fig. 7

14 Hz mode amplitude in different instrumentations for N105%

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Fig. 8

Acoustic spectrum of exhaust, experimental and numerical results

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Fig. 9

Geometry and parameters of numerical setup

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Fig. 10

Shape of low frequency acoustic modes (modal pressure amplitude); numerical results at design mass flow

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Fig. 11

Rotor SG instrumentation

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Fig. 12

SG spectrogram and prestall maximum; (a) raw signal and (b) signal with subtracted sliding ensemble average of each ten revolutions; N105

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Fig. 13

Low frequency spectrum of rotor SGs; average of last 2 s before rotating stall; N105

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Fig. 14

Bandpass filtered strain-gauge data for N105; normalized with average amplitude of blade mode 1 during rotating stall; envelopes on the right representing the connection of adjacent peaks in the bandpass filtered signal of each sensor

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Fig. 15

SG signals of different rotor blades during stall onset; resulting cross-correlation peaks depending on propagation speed N105%

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Fig. 16

Comparison of propagating structural and aerodynamic disturbances for 105% speedline; propagation speed in steady frame of reference

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Fig. 23

Integration of Helmholtz resonators to reduce acoustic resonance of exhaust system

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Fig. 22

Feedback loop between rotor asymmetry and exhaust

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Fig. 21

Modal cycle hysteresis N105%; black crosses show revolution average for last 10 modal oscillations before rotating stall; final cycle in red

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Fig. 20

Spectrum of prestall shock fluctuation for all transonic speedlines; average value for all blades during 2 s (≈ 500 REV.) before rotating stall

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Fig. 19

Comparison of shock position and blade geometry parameter

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Fig. 18

Shock fluctuation around revolution −290 before rotating stall; N105%

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Fig. 17

Variation of shock position for N105%; detected with unsteady pressure sensor upstream of rotor leading edge



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