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Sanjiva K. Lele - One of the best experts on this subject based on the ideXlab platform.
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Sound generated by instability wave/Shock-cell interaction in supersonic jets
2006Co-Authors: Prasun K. Ray, Sanjiva K. LeleAbstract:Broadband Shock-associated Noise is an important component of the overall Noise generated by modern airplanes. In this study, sound generated by the weakly nonlinear interaction between linear instability waves and the Shock-cell structure in supersonic jets is investigated numerically in order to gain insight into the broadband Shock-Noise problem. The model formulation decomposes the overall flow into a mean flow, linear instability waves, the Shock-cell structure and Shock-Noise. The mean flow is obtained by solving RANS equations with a k - ∈model. Locally parallel stability equations are solved for the Shock structure, and linear parabolized stability equations are solved for the instability waves. Then, source terms representing the instability wave/Shock-cell interaction are assembled and the inhomogeneous linearized Euler equations are solved for the Shock-Noise. Three cases are considered, a cold under-expanded Mj = 1.22 jet, a hot under-expanded M j = 1.22 jet, and a cold over-expanded M j = 1.36 jet. Shock-Noise computations are used to identify and understand significant trends in peak sound amplitudes and radiation angles. The peak sound radiation angles are explained well with the Mach wave model of Tam & Tanna (J. Sound Vib. Vol. 81, 1982, p. 337). The observed reduction of peak sound amplitudes with frequency correlates well with the corresponding reduction of instability wave growth with frequency. However, in order to account for variation of sound amplitude for different azimuthal modes, the radial structure of the instability waves must be considered in addition to streamwise growth. The effect of heating on the M j = 1.22 jet is shown to enhance the sound radiated due to the axisymmetric instability waves while the other modes are relatively unaffected. Solutions to a Lilley-Goldstein equation show that sound generated by 'thermodynamic' source terms is small relative to sound from 'momentum' sources though heating does increase the relative importance of the thermodynamic source. Furthermore, heating preferentially amplifies sound associated with the axisymmetric modes owing to constructive interference between sound from the momentum and thermodynamic sources. However, higher modes show destructive interference between these two sources and are relatively unaffected by heating.
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Shock leakage through an unsteady vortex-laden mixing layer: application to jet screech
Journal of Fluid Mechanics, 2003Co-Authors: Takao Suzuki, Sanjiva K. LeleAbstract:In an under-expanded supersonic jet, the interaction between Shock-cell structure and vortices in a mixing layer generates intense tonal Noise, called jet screech. This Noise generation can be explained as a Shock-leakage process through an unsteady vortex-laden mixing layer. This paper studies the Shock-leakage mechanism based on a geometrical theory and direct numerical simulation (DNS) in two dimensions. In the limit of weak Shocks, the analysis becomes analogous to geometrical acoustics: the eikonal equation demonstrates that Shock waves tend to leak near the saddle points between vortices. Analysing the wavenumber vector, it is shown that the local vorticity behaves as a barrier against Shocks. Using the unsteady DNS data, trajectories of the Shock fronts are computed with the time dependent eikonal equation. Furthermore, the interaction between unsteady vortices and a compression wave is solved using DNS. The geometrical theory shows good agreement with DNS for Shock-front evolution, but the amplitude of the leaked waves agrees only qualitatively. This study also investigates the effects of a temperature difference across the mixing layer. The analysis based on total internal reflection and the numerical results of both geometrical acoustics and DNS indicate that the direction of the radiated Shock Noise tends to rotate downstream as the jet temperature increases.
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A NUMERICAL INVESTIGATION OF BROAD-BAND Shock Noise
40th AIAA Aerospace Sciences Meeting & Exhibit, 2002Co-Authors: Calvin Lui, Sanjiva K. LeleAbstract:The interaction between a spatially developing, turbulent shear layer (Mc — 0.6) and an isolated oblique compression wave (AP/P^ = 0.2) is studied by direct numerical simulation. Analysis is performed on three key elements of this problem, namely the oblique compression-expansion wave, the associated acoustic field and the shear-layer turbulence. The acoustic field consists of two separate components: the downstream propagating mixing Noise originating from the transitional region of the shear layer, and a nearly omni-directional Shock Noise component from a region slightly downstream of the interaction location. The Shock Noise component dominates the upstream radiation but its relative importance diminishes in the downstream direction. The frequency spectrum of the Shock Noise has a peak centered around fS^/AU = 1.17, which is slightly higher than that of mixing Noise. In the near field, self-similar turbulence is shown to be established before the interaction. Two-point correlation of the turbulent fluctuations confirms the importance of accounting for the spatial coherence of the turbulence in any Shock Noise model. Minimal oscillations of the compression wave are observed. This is in sharp contrast to the two-dimensional Shock-vortex study by Manning & Lele [8] in which substantial Shock motion was observed.
Zoltán S. Spakovszky - One of the best experts on this subject based on the ideXlab platform.
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Effects of Boundary-Layer Ingestion on the Aero-Acoustics of Transonic Fan Rotors
Journal of Turbomachinery, 2013Co-Authors: Jeff Defoe, Zoltán S. SpakovszkyAbstract:The use of boundary-layer-ingesting, embedded propulsion systems can result in inlet flow distortions where the interaction of the boundary-layer vorticity and the inlet lip causes horseshoe vortex formation and the ingestion of streamwise vortices into the inlet. A previously-developed body-force-based fan modeling approach was used to assess the change in fan rotor Shock Noise generation and propagation in a boundary-layer-ingesting, serpentine inlet. This approach is employed here in a parametric study to assess the effects of inlet geometry parameters (offset-to-diameter ratio and downstream-to-upstream area ratio) on flow distortion and rotor Shock Noise. Mechanisms related to the vortical inlet structures were found to govern changes in the rotor Shock Noise generation and propagation. The vortex whose circulation is in the opposite direction to the fan rotation (counter-swirling vortex) increases incidence angles on the fan blades near the tip, enhancing Noise generation. The vortex with circulation in the direction of fan rotation (co-swirling vortex) creates a region of subsonic relative flow near the blade tip radius that decreases the sound power propagated to the far-field. The parametric study revealed that the overall sound power level at the fan leading edge is set by the ingested streamwise circulation, and that for inlet designs in which the streamwise vortices are displaced away from the duct wall, the sound power at the upstream inlet plane increased by as much as 9 dB. By comparing the far-field Noise results obtained to those for a conventional inlet, it is deduced that the changes in rotor Shock Noise are predominantly due to the ingestion of streamwise vorticity.
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Effects of Boundary Layer Ingestion on the Aero-Acoustics of Transonic Fan Rotors
Volume 8: Turbomachinery Parts A B and C, 2012Co-Authors: Jeff Defoe, Zoltán S. SpakovszkyAbstract:The use of boundary-layer-ingesting, embedded propulsion systems can result in inlet flow distortions where the interaction of the boundary layer vorticity and the inlet lip causes horseshoe vortex formation and the ingestion of streamwise vortices into the inlet. A previously-developed body-force-based fan modeling approach was used to assess the change in fan rotor Shock Noise generation and propagation in a boundary-layer-ingesting, serpentine inlet. This approach is employed here in a parametric study to assess the effects of inlet geometry parameters (offset-to-diameter ratio and downstream-to-upstream area ratio) on flow distortion and rotor Shock Noise.Mechanisms related to the vortical inlet structures were found to govern changes in the rotor Shock Noise generation and propagation. The vortex whose circulation is in the opposite direction to the fan rotation (counter-swirling vortex) increases incidence angles on the fan blades near the tip, enhancing Noise generation. The vortex with circulation in the direction of fan rotation (co-swirling vortex) creates a region of subsonic relative flow near the blade tip radius which decreases the sound power propagated to the far-field.The parametric study revealed that the overall sound power level at the fan leading edge is set by the ingested streamwise circulation, and that for inlet designs in which the streamwise vortices are displaced away from the duct wall, the sound power at the upstream inlet plane increased by as much as 9 dB. By comparing the far-field Noise results obtained to those for a conventional inlet, it is deduced that the changes in rotor Shock Noise are predominantly due to the ingestion of streamwise vorticity.Copyright © 2012 by ASME
James Bridges - One of the best experts on this subject based on the ideXlab platform.
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Turbulence Associated with Broadband Shock Noise in Hot Jets
2013Co-Authors: James Bridges, Mark P. WernetAbstract:Time-Resolved Particle Image Velocimetry (TRPIV) has been applied to a series of jet flows to measure turbulence statistics associated with broadband Shock associated Noise (BBSN). Data were acquired in jets of Mach numbers 1.05, 1.185, and 1.4 at different temperatures. Both convergent and ideally expanded nozzles were tested, along with a convergent nozzle modified to minimize screech. Key findings include the effect of heat on Shock structure and jet decay, the increase in turbulent velocity when screech is present, and the relative lack of spectral detail associated with the enhanced turbulence.
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PIV Measurements of Supersonic
2012Co-Authors: James Bridges, Mark P. WernetAbstract:Abstract While externally mixed, or separate flow, nozzle systems are most common in high bypass-ratio aircraft, they are not as attractive for use in lower bypass- ratio systems and on aircraft that will fly supersonically. The Noise of such propulsion systems is also dominated by jet Noise, making the study and Noise reduction of these exhaust systems very important, both for military aircraft and future civilian supersonic aircraft. This paper presents particle image velocimetry of internally mixed nozzle with different area ratios between core and bypass, and nozzles that are ideally expanded and convergent . Such configurations independently control the geometry of the internal mixing layer and of the external Shock structure. These allow exploration of the impact of Shocks on the turbulent mixing layers, the impact of bypass ratio on broadband Shock Noise and mixing Noise, and the impact of temperature on the turbulent flow field. At the 2009 AIAA/CEAS Aeroacoustics Conference the authors presented data and analysis from a series of tests that looked at the acoustics of supersonic jets from internally mixed nozzles. In that paper the broadband Shock and mixing Noise components of the jet Noise were independently manipulated by holding Mach number constant while varying bypass ratio and jet temperature. Significant portions of that analysis was predicated on assumptions regarding the flow fields of these jets, both Shock structure and turbulence. In this paper we add to that analysis by presenting particle image velocimetry measurements of the flow fields of many of those jets. In addition, the turbulent velocity data documented here will be very useful for validation of computational flow codes that are being developed to design advanced nozzles for fu ture aircraft.
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Development of Jet Noise Power Spectral Laws
2011Co-Authors: Abbas Khavaran, James BridgesAbstract:High-quality jet Noise spectral data measured at the Aero-Acoustic Propulsion Laboratory (AAPL) at NASA Glenn is used to develop jet Noise scaling laws. A FORTRAN algorithm was written that provides detailed spectral prediction of component jet Noise at user-specified conditions. The model generates quick estimates of the jet mixing Noise and the broadband Shock-associated Noise (BBSN) in single-stream, axis-symmetric jets within a wide range of nozzle operating conditions. Shock Noise is emitted when supersonic jets exit a nozzle at imperfectly expanded conditions. A successful scaling of the BBSN allows for this Noise component to be predicted in both convergent and convergent-divergent nozzles. Configurations considered in this study consisted of convergent and convergent- divergent nozzles. Velocity exponents for the jet mixing Noise were evaluated as a function of observer angle and jet temperature. Similar intensity laws were developed for the broadband Shock-associated Noise in supersonic jets. A computer program called sJet was developed that provides a quick estimate of component Noise in single-stream jets at a wide range of operating conditions. A number of features have been incorporated into the data bank and subsequent scaling in order to improve jet Noise predictions. Measurements have been converted to a lossless format. Set points have been carefully selected to minimize the instability-related Noise at small aft angles. Regression parameters have been scrutinized for error bounds at each angle. Screech-related amplification Noise has been kept to a minimum to ensure that the velocity exponents for the jet mixing Noise remain free of amplifications. A Shock-Noise-intensity scaling has been developed independent of the nozzle design point. The computer program provides detailed narrow-band spectral predictions for component Noise (mixing Noise and Shock associated Noise), as well as the total Noise. Although the methodology is confined to single streams, efforts are underway to generate a data bank and algorithm applicable to dual-stream jets. Shock-associated Noise in high-powered jets such as military aircraft can benefit from these predictions.
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An MDOE Investigation of Chevrons for Supersonic Jet Noise Reduction
16th AIAA CEAS Aeroacoustics Conference, 2010Co-Authors: Brenda Henderson, James BridgesAbstract:The impact of chevron design on the Noise radiated from heated, overexpanded, supersonic jets is presented. The experiments used faceted bi-conic convergent-divergent nozzles with design Mach numbers equal to 1.51 and 1.65. The purpose of the facets was to simulate divergent seals on a military style nozzle . The nozzle throat diameter was equal to 4.5 inches. Modern Design of Experiment (MDOE) techniques were used to investigate the impact of chevron penetration, length, and width on the resulting acoustic radiation. All chevron configurations used 12 chevrons to match the number of facets in the nozzle. Most chevron designs resulted in increased broadband Shock Noise relative to the baseline nozzle. In the peak jet Noise direction, the optimum chevro n design reduced peak sound pressure levels by 4 dB relative to the baseline nozzle. Th e penetration was the parameter having the greatest impact on radiated Noise at all observatio n angles. While increasing chevron penetration decreased acoustic radiation in the pea k jet Noise direction, broadband Shock Noise was adversely impacted. Decreasing chevron length increased Noise at most observation angles. The impact of chevron width on radiated Noise depended on frequency and observation angle. I. Introduction He application of chevrons (serrations applied to a nozzle trailing edge that protrude into the exhaus ting flow) to military aircraft is particularly attractive becaus e existing engines can be retrofitted rather than r edesigned to incorporate these devices. At takeoff, high perfor mance tactical aircraft typically have overexpanded , supersonic jet-exhausts that contain Noise sources not present in the subsonic exhausts of commercial aircraft en gines. As a result, chevrons that have been optimized for Noise reduction in commercial aircraft may not perform a dequately on tactical aircraft. While a reasonably large number of investigations have studied the impact of chevr ons on subsonic jets, similar studies for supersonic flows are limi ted. The present investigation uses a Modern Desig n of Experiments (MDOE) approach to explore the impact of chevron design on the acoustic radiation of overe xpanded supersonic jets. An overexpanded jet resulting from operating a convergent-divergent nozzle at a stagnation pressu re below that corresponding to the nozzle design Mach number contains a quasi-periodic Shock cell structure that can persist for several diameters downstream of the nozzle exit. T he constructive interference of sound waves produce d by the interaction of large-scale jet disturbances with th e Shock waves within the Shock cell structure resul ts in broadband Shock Noise 1,2,3 . Shock Noise can dominate the acoustic spectra at upstream and broadside observation angles relative to the nozzle exit. Additionally, mixing Noise sources are present and are associated with l arge scale jet disturbances (radiating in the downstream direction ) that become very effective Noise sources when the ir phase speeds (relative to the ambient speed of sound) bec ome supersonic 4 , and with fine scale turbulence 5 (radiating in the upstream direction). Mixing Noise sources are also present in subsonic jets but the large-scale distu rbances typically have subsonic phase speeds. In subsonic jets, properly designed chevron no zzles produce lower overall acoustic radiation leve ls than those of a corresponding round nozzle 6,7,8,9 . Experiments have shown that increasing chevron p enetration decreases low frequency Noise and often increases high frequency Noise (sometimes referred to as high frequency cros sover). The number of chevrons also impacts the acoustic radiat ion but not as significantly as the penetration. J et shear velocity (the velocity difference between the inner and oute r jet streams) impacts chevron acoustic performance with increases in shear velocity increasing low frequenc y Noise reduction but sometimes increasing high fre quency Noise * Researcher, Acoustics Branch, MS 54-3, 21000 Brookpark Rd., Cleveland, OH 44135. † Reseacher, Acoustics Branch, MS 54-3, 21000 Brookpark Rd., Cleveland, OH 44135, Associate Fellow.
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SHJAR Jet Noise Data and Power Spectral Laws
2009Co-Authors: Abbas Khavaran, James BridgesAbstract:High quality jet Noise spectral data measured at the Aeroacoustic Propulsion Laboratory at the NASA Glenn Research Center is used to examine a number of jet Noise scaling laws. Configurations considered in the present study consist of convergent and convergent-divergent axisymmetric nozzles. The measured spectral data are shown in narrow band and cover 8193 equally spaced points in a typical Strouhal number range of 0.0 to 10.0. The measured data are reported as lossless (i.e., atmospheric attenuation is added to measurements), and at 24 equally spaced angles (50deg to 165deg) on a 100-diameter (200-in.) arc. Following the work of Viswanathan, velocity power factors are evaluated using a least squares fit on spectral power density as a function of jet temperature and observer angle. The goodness of the fit and the confidence margins for the two regression parameters are studied at each angle, and alternative relationships are proposed to improve the spectral collapse when certain conditions are met. As an immediate application of the velocity power laws, spectral density in Shockcontaining jets are decomposed into components attributed to jet mixing Noise and Shock Noise. From this analysis, jet Noise prediction tools can be developed with different spectral components derived from different physics.
Dilip Prasad - One of the best experts on this subject based on the ideXlab platform.
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Unsteady Aerodynamics and Aeroacoustics of a High-Bypass Ratio Fan Stage
Journal of Turbomachinery, 2005Co-Authors: Anil Prasad, Dilip PrasadAbstract:A numerical investigation of the unsteady aerodynamics of a fan stage comprised of a transonic rotor, swept fan exit guide vane (FEGV), and low-pressure compressor inlet guide vane (IGV) is described, with emphasis on acoustics. It is shown that the effects of the two downstream stator rows on the time-mean blade flow field are negligible, permitting its investigation using isolated rotor calculations. Simulations of this type are carried out along the engine operating line to quantify the acoustic sources associated with the upstream Shock field and wake turbulence-stator interaction. The Shock Noise achieves its maximum value near the flyover acoustic certification condition, while the wake turbulence is least at this condition owing to its proximity to the design point. The behavior of these Noise sources is explained physically by carrying out a detailed examination of the rotor flow field. The unsteady interaction between the rotor and stator rows at a high-power setting is investigated next. It is shown that the time-mean IGV flow is significantly affected by this interaction. Moreover, the unsteady loading on the IGV is found to be large. The behavior of the upstream-propagating acoustic field generated by rotor-IGV interaction is examined. The interaction between the rotor and FEGV is found to be linear in nature. The FEGV surface unsteady pressure and far-field acoustic field behavior are investigated.
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Unsteady Aerodynamics and Aeroacoustics of a High-Bypass Ratio Fan Stage
Volume 5: Turbo Expo 2004 Parts A and B, 2004Co-Authors: Anil Prasad, Dilip PrasadAbstract:A numerical investigation of the unsteady aerodynamics of a fan stage comprised of a transonic rotor, swept fan exit guide vane (FEGV) and low-pressure compressor inlet guide vane (IGV) is described, with emphasis on acoustics. It is shown that the effects of the two downstream stator rows on the time-mean blade flow field are negligible, permitting its investigation using isolated rotor calculations. Simulations of this type are carried out along the engine operating line to quantify the acoustic sources associated with the upstream Shock field and wake turbulence-stator interaction. The Shock Noise achieves its maximum value near the flyover acoustic certification condition, while the wake turbulence is least at this condition owing to its proximity to the design point. The behavior of these Noise sources is explained physically by carrying out a detailed examination of the rotor flow field. The unsteady interaction between the rotor and stator rows at a high-power setting is investigated next. It is shown that the time-mean IGV flow is significantly affected by this interaction. Moreover, the unsteady loading on the IGV is found to be large. The behavior of the upstream-propagating acoustic field generated by rotor-IGV interaction is examined. The interaction between the rotor and FEGV is found to be linear in nature. The FEGV surface unsteady pressure and far-field acoustic field behavior are investigated.Copyright © 2004 by ASME
Jeff Defoe - One of the best experts on this subject based on the ideXlab platform.
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Effects of Boundary-Layer Ingestion on the Aero-Acoustics of Transonic Fan Rotors
Journal of Turbomachinery, 2013Co-Authors: Jeff Defoe, Zoltán S. SpakovszkyAbstract:The use of boundary-layer-ingesting, embedded propulsion systems can result in inlet flow distortions where the interaction of the boundary-layer vorticity and the inlet lip causes horseshoe vortex formation and the ingestion of streamwise vortices into the inlet. A previously-developed body-force-based fan modeling approach was used to assess the change in fan rotor Shock Noise generation and propagation in a boundary-layer-ingesting, serpentine inlet. This approach is employed here in a parametric study to assess the effects of inlet geometry parameters (offset-to-diameter ratio and downstream-to-upstream area ratio) on flow distortion and rotor Shock Noise. Mechanisms related to the vortical inlet structures were found to govern changes in the rotor Shock Noise generation and propagation. The vortex whose circulation is in the opposite direction to the fan rotation (counter-swirling vortex) increases incidence angles on the fan blades near the tip, enhancing Noise generation. The vortex with circulation in the direction of fan rotation (co-swirling vortex) creates a region of subsonic relative flow near the blade tip radius that decreases the sound power propagated to the far-field. The parametric study revealed that the overall sound power level at the fan leading edge is set by the ingested streamwise circulation, and that for inlet designs in which the streamwise vortices are displaced away from the duct wall, the sound power at the upstream inlet plane increased by as much as 9 dB. By comparing the far-field Noise results obtained to those for a conventional inlet, it is deduced that the changes in rotor Shock Noise are predominantly due to the ingestion of streamwise vorticity.
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Effects of Boundary Layer Ingestion on the Aero-Acoustics of Transonic Fan Rotors
Volume 8: Turbomachinery Parts A B and C, 2012Co-Authors: Jeff Defoe, Zoltán S. SpakovszkyAbstract:The use of boundary-layer-ingesting, embedded propulsion systems can result in inlet flow distortions where the interaction of the boundary layer vorticity and the inlet lip causes horseshoe vortex formation and the ingestion of streamwise vortices into the inlet. A previously-developed body-force-based fan modeling approach was used to assess the change in fan rotor Shock Noise generation and propagation in a boundary-layer-ingesting, serpentine inlet. This approach is employed here in a parametric study to assess the effects of inlet geometry parameters (offset-to-diameter ratio and downstream-to-upstream area ratio) on flow distortion and rotor Shock Noise.Mechanisms related to the vortical inlet structures were found to govern changes in the rotor Shock Noise generation and propagation. The vortex whose circulation is in the opposite direction to the fan rotation (counter-swirling vortex) increases incidence angles on the fan blades near the tip, enhancing Noise generation. The vortex with circulation in the direction of fan rotation (co-swirling vortex) creates a region of subsonic relative flow near the blade tip radius which decreases the sound power propagated to the far-field.The parametric study revealed that the overall sound power level at the fan leading edge is set by the ingested streamwise circulation, and that for inlet designs in which the streamwise vortices are displaced away from the duct wall, the sound power at the upstream inlet plane increased by as much as 9 dB. By comparing the far-field Noise results obtained to those for a conventional inlet, it is deduced that the changes in rotor Shock Noise are predominantly due to the ingestion of streamwise vorticity.Copyright © 2012 by ASME
