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Research Papers

A Parametric Study of the Effects of Inlet Distortion on Fan Aerodynamic Stability

[+] Author and Article Information
Wenqiang Zhang

Department of Mechanical Engineering,
Imperial College London,
London SW7 2AZ, UK
e-mail: w.zhang15@imperial.ac.uk

Mehdi Vahdati

Department of Mechanical Engineering,
Imperial College London,
London SW7 2AZ, UK
e-mail: vahdati@imperial.ac.uk

Contributed by the International Gas Turbine Institute (IGTI) of ASME for publication in the JOURNAL OF TURBOMACHINERY. Manuscript received August 19, 2018; final manuscript received August 27, 2018; published online November 29, 2018. Editor: Kenneth Hall.

J. Turbomach 141(1), 011011 (Nov 29, 2018) (11 pages) Paper No: TURBO-18-1212; doi: 10.1115/1.4041376 History: Received August 19, 2018; Revised August 27, 2018

The performance and aerodynamic stability of fan blades operating in a circumferentially nonuniform inlet flow is a key concern in the design of turbofan engines. With the recent trends in the design of civil engines with shorter inlet ducts (such as low-speed fans), or boundary layer ingesting engines, quick and reliable modeling of rotor/distortion interactions is becoming very important. The aim of this paper is to study the effects of inlet distortions on the aerodynamic stability of a fan blade and to identify the parameters that have a major impact on the stability of the blade. NASA rotor 67, for which a significant amount of measured data is available, was used for this study. In the first part of this study, the behavior of the fan with inlet distortion at near stall (NS) condition is analyzed, and it is shown how rotating stall is triggered. In the second part of this study, unsteady simulations with inlet distortion were performed to study how exit duct length influences the stall margin of the blade.

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References

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Figures

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Fig. 1

Isometric view of the computational domain of NASA stage 67 with inlet distortion

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Fig. 2

Initial clean and final operating point for nozzles A and B at 90% corrected speed

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Fig. 3

Physical mass flow history at the rotor outlet for nozzles A and B

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Fig. 4

Instantaneous contours at the stator inlet for nozzles A (left) and B (right) at the sixth revolution

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Fig. 5

Absolute swirl angle and incidence at the rotor inlet, middle span, nozzle A

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Fig. 6

Circumferential contour of the axial velocity for case A, 98% rotor span

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Fig. 7

The vortex in front of the rotor in nozzle A

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Fig. 8

Tip leakage flow of the initial clean flow solution for nozzle A

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Fig. 9

Tip leakage flow for nozzle A at different circumferential positions

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Fig. 10

The circumferential contour of the axial velocity at 98% span of the rotor for nozzle A (left) and nozzle B (right) at different time (legend is the same with Fig. 6)

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Fig. 11

Stall margin loss of NASA stage 67 with DC120 distortion at different rotational speed

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Fig. 12

The stall margin loss for different exit duct length

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Fig. 13

Mass flow history at the rotor outlet for the case with long exit duct and short exit duct

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Fig. 14

The constant radius layer with static pressure contour extracted from the exit duct

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Fig. 15

Static pressure history contour at middle span in the short (left) and long (right) exit duct

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Fig. 16

Schematic plot of the computational domain

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Fig. 17

Mass flow history at the rotor outlet for the stage case with long exit duct and short exit duct

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