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.

© 2013 by ASME
Your Session has timed out. Please sign back in to continue.


Sargeant, M., 2007, “Boundary Layer Ingestion for Advanced Airframes,” Ph.D. thesis, University of Cambridge, Cambridge, UK.
Berrier, B., and Allan, B., 2004, “Experimental and Computational Evaluation of Flush-Mounted S-Duct Inlets,” AIAA Paper No. 2004-0764.
Gliebe, P., Mani, R., Shin, H., Mitchell, B., Ashford, G., Salamah, S., and Connell, S., 2000, “Aeroacoustic Prediction Codes,” NASA Technical Report No. CR-2000-210244.
Mathews, D., and Nagel, R., 1973, “Inlet Geometry and Axial Mach Number Effects on Fan Noise Propagation,” AIAA Paper No. 1973-1022.
Prasad, D., and Feng, J., 2005, “Propagation and Decay of Shock Waves in Turbofan Engine Inlets,” ASME J. Turbomach., 127(1), pp. 118–127. [CrossRef]
Coupland, J., Wilson, A., Pollard, N., Uellenberg, S., Breard, C., and Diamond, J., 2007, “Demonstration of a CFD-CAA Methodology to Predict Buzz-Saw Noise Propagation to the Aircraft,” AIAA Paper No. 2007-3517.
Brambley, E., and Peake, N., 2007, “Sound in Curved Intakes,” AIAA Paper No. 2007-3552.
Brambley, E., and Peake, N., 2008, “Sound Transmission in Strongly Curved Slowly Varying Clylindrical Ducts With Flow,” J. Fluid Mech., 596, pp. 387–412. [CrossRef]
Defoe, J., Narkaj, A., and Spakovszky, Z., 2010, “A Body-Force-Based Method for Prediction of Multiple-Pure-Tone Noise: Validation,” AIAA Paper No. 2010-3747.
Defoe, J., and Spakovszky, Z., 2011, “Shock Propagation and MPT Noise From a Transonic Rotor in Non-Uniform Flow,” ASME Paper No. GT2011-45645. [CrossRef]
Defoe, J., Narkaj, A., and Spakovszky, Z., 2009, “A Novel MPT Noise Methodology for Highly-Integrated Propulsion Systems With Inlet Flow Distortion,” AIAA Paper No. 2009-3366.
Podboy, G., Krupar, M., Helland, S., and Hughes, C., 2002, “Steady and Unsteady Flow Field Measurements Within a NASA 22 Inch Fan Model,” AIAA Paper No. 2002-1033.
Hughes, C., Jeracki, R., Woodward, R., and Miller, C., 2002, “Fan Noise Source Diagnostic Test—Rotor Alone Aerodynamic Performance Results,” AIAA Paper No. 2002-2426.
Woodward, R., Hughes, C., Jeracki, R., and Miller, C., 2002, “Fan Noise Source Diagnostic Test—Far-Field Acoustic Results,” AIAA Paper No. 2002-2427.
Heidelberg, L., 2002, “Fan Noise Source Diagnostic Test—Tone Modal Structure Results,” AIAA Paper No. 2002-2428.
Madani, V., and Hynes, T., 2009, “Boundary Layer Ingesting Intakes: Design and Optimization,” ISABE Paper No. 2009-1346.
Hileman, J., Spakovszky, Z., and Drela, M., 2010, “Airframe Design for Silent Fuel-Efficient Aircraft,” J. Aircr., 47(3), pp. 956–969. [CrossRef]
Huttl, T., Kahl, G., Kennepohl, F., and Heinig, K., 2001, “Resolution Requirements for the Numerical Computation of Tonal Noise in Compressors and Turbines of Aeroengines,” Proceedings of the RTO AVT Symposium on Ageing Mechanisms and Control: Part A—Developments in Computational Aero- and Hydro-Acoustics, Manchester, UK, October 8–11.
Freund, J., 1997, “Proposed Inflow/Outflow Boundary Condition for Direct Computation of Aerodynamic Sound,” AIAA J., 35(4), pp. 740–742. [CrossRef]
Sutliff, D., 2005, “Rotating Rake Turbofan Duct Mode Measurement System,” NASA Technical Report No. NASA/TM-2005-2138328.
Freeman, C., and Cumpsty, N., 1992, “A Method for the Prediction of Supersonic Compressor Blade Performance,” J. Propul. Power, 8(1), pp. 199–208. [CrossRef]


Grahic Jump Location
Fig. 1

MPT noise prediction framework

Grahic Jump Location
Fig. 2

Body force perturbation to generate rotor blade shocks

Grahic Jump Location
Fig. 3

Computational domains

Grahic Jump Location
Fig. 4

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

Grahic Jump Location
Fig. 5

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

Grahic Jump Location
Fig. 6

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

Grahic Jump Location
Fig. 7

Unsteady pressure at AIP over mean dynamic pressure at AIP

Grahic Jump Location
Fig. 8

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

Grahic Jump Location
Fig. 9

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

Grahic Jump Location
Fig. 10

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

Grahic Jump Location
Fig. 11

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

Grahic Jump Location
Fig. 12

Modulated shock surface model

Grahic Jump Location
Fig. 13

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

Grahic Jump Location
Fig. 14

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

Grahic Jump Location
Fig. 15

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



Some tools below are only available to our subscribers or users with an online account.

Related Content

Customize your page view by dragging and repositioning the boxes below.

Related Journal Articles
Related eBook Content
Topic Collections

Sorry! You do not have access to this content. For assistance or to subscribe, please contact us:

  • TELEPHONE: 1-800-843-2763 (Toll-free in the USA)
  • EMAIL: asmedigitalcollection@asme.org
Sign In