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Research Papers

Effect of Inlet Distortion Features on Transonic Fan Rotor Stall

[+] Author and Article Information
James H. Page

Whittle Laboratory,
Department of Engineering,
University of Cambridge,
Cambridge CB3 0DY, UK
e-mail: jpage9@hotmail.com

Paul Hield

Fan Systems Engineering,
Rolls-Royce plc, Filton,
Bristol BS34 7QE, UK
e-mail: Paul.Hield@Rolls-Royce.com

Paul G. Tucker

Department of Engineering,
Trumpington St.,
University of Cambridge,
Cambridge CB2 1PZ, UK
e-mail: pgt23@cam.ac.uk

1Corresponding author.

Contributed by the International Gas Turbine Institute (IGTI) of ASME for publication in the JOURNAL OF TURBOMACHINERY. Manuscript received March 5, 2018; final manuscript received March 28, 2018; published online June 14, 2018. Editor: Kenneth Hall.

J. Turbomach 140(7), 071008 (Jun 14, 2018) (11 pages) Paper No: TURBO-18-1051; doi: 10.1115/1.4040030 History: Received March 05, 2018; Revised March 28, 2018

The effect of inlet distortion from curved intake ducts on jet engine fan stability is an important consideration for next-generation passenger aircraft such as the boundary layer ingestion (BLI) “silent aircraft.” Highly complex inlet flows which occur can significantly affect fan stability. Future aircraft designs are likely to feature more severe inlet distortion, pressing the need to understand the important factors influencing design. This paper presents the findings from a large computational fluid dynamics (CFD) investigation into which aspects of inlet distortion cause the most significant reductions in stall margin and, therefore, which flow patterns should be targeted by mitigating technology. The study considers the following aspects of distortion commonly observed in intakes: steady vortical distortion due to secondary flow, unsteady vortical distortion due to vortex shedding and mixing, static pressure distortion due to curved streamlines, and low momentum endwall flow due to thickened boundary layers or separation. Unsteady CFD was used to determine the stall points of a multipassage transonic rotor geometry with each of the inlet distortion patterns applied. Interesting new evidence is provided, which suggests that low momentum flow in the tip region, rather than distortion in the main body of the flow, leads to damaging instability.

Copyright © 2018 by ASME
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References

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Figures

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Fig. 1

Flow solution of an S-duct les. (a) Instantaneous static pressure for flow from left to right. (b) Time-averaged velocity field in cross-sectional plane at exit. (c) Instantaneous velocity field in cross-sectional plane at exit. (d) Instantaneous total pressure for flow from left to right. (e) Time history of exit centerline axial velocity.

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Fig. 2

Fan rotor geometry, multipassage meshes, and intakes

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Fig. 3

Method used to plot a characteristic curve and calculate the stall point

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Fig. 5

Prestall unsteady tip clearance streamlines and casing static pressure

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Fig. 6

Prestall clean unsteady tip clearance flow negative axial velocity

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Fig. 7

Time history of negative axial velocity at a point in the relative frame over each blade in the tip gap

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Fig. 8

Tip clearance flow with a clean inlet near stall—static pressure field (left), streamlines (middle) and q-criterion (right). Each row of images depicts an instant in time in chronology from top to bottom.

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Fig. 9

Q-criterion visualization of LE spillage at the moment of rotating stall inception, colored by axial velocity from blue (negative) to red (positive)

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Fig. 10

Q-criterion visualization just after rotating stall inception, colored by axial velocity from blue (negative) to red (positive)

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Fig. 12

Q-criterion and radial velocity visualization of flow calculation with inlet vortex

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Fig. 13

Characteristic plots for inlet vortices injected at 25%, 50%, and 75% of span with core radii of 10% and normalized maximum local tangential velocities of 0.3. Smallest stable nozzle areas are circled in red.

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Fig. 14

Characteristic plots for inlet vortices injected at 50% span with core radii of 5%, 10%, and 20% and normalized maximum local tangential velocities of 0.3. Smallest stable nozzle areas are circled in red.

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Fig. 15

Characteristic plots for inlet vortices injected at 50% span with core radii of 20% inlet span and doubled normalized maximum local tangential velocities (0.6). Smallest stable nozzle areas are circled in red.

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Fig. 4

Reproduction of Fig. 5 in Ref. [2], originally published by ASME, showing the normalized radial profiles of (a) stagnation pressure and (b) stagnation temperature without VIGVs, with clean inlet flow

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Fig. 11

Schematic diagram of wingtip vortex local tangential velocity profile

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Fig. 16

Tip clearance flow streamlines with a large strong corotating vortex, at a nozzle area 2% less than the clean inlet stalling area. Time progression is from (a) to (f).

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Fig. 17

Azimuthal snapshot of normalized axial velocity with unsteady vortical inlet distortion

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Fig. 18

Characteristic plots for the unsteady S-duct inlet and clean inlet cases

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Fig. 19

Snapshot of normalized axial velocity on four axial slices from upstream to downstream of the rotor with unsteady vortical inlet distortion: (a) upstream, (b) rotor front, (c) rotor middle, and (d) rotor rear

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Fig. 21

Relative Mach plot at 95% span in the rotor at the last stable nozzle area (slip walls on inlet). The positions of the inside and outside of the intake curve are indicated.

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Fig. 22

Pressure ratio characteristic plot comparing the stall points of the clean inlet case and the curved streamline static pressure and velocity distortion case. Smallest stable nozzle areas are circled in red.

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Fig. 23

Curved streamline with boundary layer inlet distortion. Axial slices are downstream of the intake bend and upstream of the rotor.

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Fig. 20

Curved streamline inlet distortion. Axial slices are downstream of the intake bend and upstream of the rotor.

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Fig. 24

Relative Mach plot at 95% span in the rotor, with inlet boundary layer, at the last stable nozzle area (no-slip walls on inlet). The positions of the inside and outside of the intake curve are indicated.

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Fig. 25

Pressure ratio characteristic plot comparing the stall points of curved streamline static pressure and velocity distortion cases with and without boundary layer. Smallest stable nozzle areas are circled in red.

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Fig. 26

Casing static pressure and tip leakage vortex streamlines at four locations around the annulus: (a) 55 deg, (b) 161 deg, (c) 268 deg, and (d) 321 deg

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