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Research Papers

Shock Propagation and MPT Noise From a Transonic Rotor in Nonuniform Flow

[+] Author and Article Information
Jeffrey J. Defoe

Postdoctoral Associate
e-mail: jdefoe@mit.edu

Zoltán S. Spakovszky

Associate Professor of
Aeronautics and Astronautics
Gas Turbine Lab;
Department of Aeronautics and Astronautics,
Massachusetts Institute of Technology,
Cambridge, MA 02139

45 dB was reported in Ref. [10] due to an inadvertent postprocessing error, which has since been corrected.

Though the cut-back Mach number for the SAX-40 is 0.22, the free-stream Mach number of 0.1 is consistent with the experimental R4 wind tunnel data and is thus used throughout this work.

Fan broadband noise is not modeled in the simulations, and the background noise floor is therefore set by numerical noise.

This is consistent with the assumptions used in the body force representation of axisymmetric through-flow for identical blade passages.

This is deemed appropriate as the inlet pressure recovery is 99% at the low cut-back flight Mach number of 0.1.

Contributed by the International Gas Turbine Institute (IGTI) of ASME for publication in the JOURNAL OF TURBOMACHINERY. Manuscript received July 9, 2011; final manuscript received August 24, 2011; published online October 30, 2012. Editor: David Wisler.

J. Turbomach 135(1), 011016 (Oct 30, 2012) (9 pages) Paper No: TURBO-11-1104; doi: 10.1115/1.4006497 History: Received July 09, 2011; Revised August 24, 2011

One of the major challenges in high-speed fan stages used in compact, embedded propulsion systems is inlet distortion noise. A body-force-based approach for the prediction of multiple-pure-tone (MPT) noise was previously introduced and validated. In this paper, it is employed with the objective of quantifying the effects of nonuniform flow on the generation and propagation of MPT noise. First-of-their-kind back-to-back coupled aero-acoustic computations were carried out using the new approach for conventional and serpentine inlets. Both inlets delivered flow to the same NASA/GE R4 fan rotor at equal corrected mass flow rates. Although the source strength at the fan is increased by 38 dB in sound power level due to the nonuniform inflow, far-field noise for the serpentine inlet duct is increased on average by only 3.1 dBA overall sound pressure level in the forward arc. This is due to the redistribution of acoustic energy to frequencies below 11 times the shaft frequency and the apparent cut-off of tones at higher frequencies including blade-passing tones. The circumferential extent of the inlet swirl distortion at the fan was found to be two blade pitches, or 1/11th of the circumference, suggesting a relationship between the circumferential extent of the inlet distortion and the apparent cut-off frequency perceived in the far field. A first-principles-based model of the generation of shock waves from a transonic rotor in nonuniform flow showed that the effects of nonuniform flow on acoustic wave propagation, which cannot be captured by the simplified model, are more dominant than those of inlet flow distortion on source noise. It demonstrated that nonlinear, coupled aerodynamic and aero-acoustic computations, such as those presented in this paper, are necessary to assess the propagation through nonuniform mean flow. A parametric study of serpentine inlet designs is underway to quantify these propagation effects.

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References

Figures

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Fig. 1

MPT noise prediction framework

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Fig. 2

Body force perturbation to generate rotor blade shocks

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Fig. 3

Computational domains

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Fig. 4

Axial and tangential Mach number distributions at rotor leading edge for serpentine inlet

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Fig. 5

Relative Mach number at 92% span from fan (x/RAIP = 0) to AIP/throat (x/RAIP = 0.84)

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Fig. 6

Unsteady pressure at rotor leading edge over mean dynamic pressure at AIP

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Fig. 7

Unsteady pressure at AIP over mean dynamic pressure at AIP

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Fig. 8

Unsteady pressure at serpentine inlet throat over mean dynamic pressure at AIP

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Fig. 9

Linear far-field spectra (dashed lines: computation; solid lines: experimental data)

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Fig. 10

Full-scale linear far-field spectra (dashed lines: conventional inlet; solid lines: serpentine inlet)

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Fig. 11

Control volume analysis for detached shock strength (adapted from Ref. [21])

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Fig. 12

Modulated shock surface model

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Fig. 13

Computed Mach numbers and relative flow angle at 92% span at rotor leading edge for serpentine inlet case

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Fig. 14

Shock strength dependence on relative Mach number; inset: inlet distortion as idealized relative Mach number distribution

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Fig. 15

Rotor-locked sawtooth wave modulated by stationary shock surface in nonuniform flow

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