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  • 1
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    Experiments on Exhaust Noise of Tightly Integrated Propulsion Systems (2014)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A wide-ranging series of tests have been completed that seek to map the effects of installation, including jet by jet interaction effects, on exhaust noise from various nozzles in forward flight. The primary data was far-field acoustic spectral directivity. The goals of the test series were (i) to generate enough data for empirical models of the different effects, and (ii) to provide data for advanced computational noise predictions methods applied to simplified yet realistic configurations. Data is presented that demonstrate several checks on data quality and that provide an overview of trends observed to date. Among the findings presented here: (i) Data was repeatable between jet rigs for single nozzles with and without surfaces to within +/- 0.5 dB. (ii) The presence of a second jet caused a strong reduction of the summed noise in the plane of the two plumes and an increase over the expected source doubling in most other azimuthal planes. (iii) The impact of the second jet was reduced when the jets were unheated. (iv) The impact of adding a second isolated rectangular jet was relatively independent of the nozzle aspect ratio up to aspect ratio 8:1. (v) Forward flight had similar impact on a high aspect ratio (8:1) jet as on an axisymmetric jet, except at the peak noise angle where the impact was less. (vi) The effect of adding a second round jet to a tightly integrated nozzle where the nozzle lip was less than a diameter from the surface was very dependent upon the length of the surface downstream of the nozzle. (vii) When the nozzles were rectangular and tightly integrated with the airframe surface the impact of a second jet was very dependent upon how close together the two jets were. This paper serves as an overview of the test; other papers presented in the same conference will give more detailed analysis of the results.
    Keywords: Aircraft Propulsion and Power; Aircraft Design, Testing and Performance; Acoustics
    Type: GRC-E-DAA-TN15171 , AIAA-CEAS Aeroacoustics Conference; Jun 16, 2014 - Jun 20, 2014; Atlanta, GA; United States
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  • 2
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    Background Oriented Schlieren Implementation in a Jet-Surface Interaction Test (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Many current and future aircraft designs rely on the wing or other aircraft surfaces to shield the engine noise from observers on the ground. However the available data regarding how a planar surface interacts with a jet to shield and/or enhance the jet noise are currently limited. Therefore, the Jet-Surface Interaction Tests supported by NASA's Fundamental Aeronautics Program's Fixed Wing Project were undertaken to supply experimental data covering a wide range of surface geometries and positions interacting with high-speed jet flows in order to support the development of noise prediction methods. Phase 1 of the Test was conducted in the Aero-Acoustic Propulsion Laboratory at NASA Glenn Research Center and consisted of validating noise prediction schemes for a round nozzle interacting with a planar surface. Phased array data and far-field acoustic data were collected for both the shielded and reflected sides of the surface. Phase 1 results showed that the broadband shock noise was greatly reduced by the surface when the jet was operated at the over-expanded condition, however, it was unclear whether this reduction was due a change in the shock cell structure by the surface. In the present study, Background Oriented Schlieren is implemented in Phase 2 of the Jet-Surface Interaction Tests to investigate whether the planar surface interacts with the high-speed jet ow to change the shock cell structure. Background Oriented Schlieren data are acquired for under-expanded, ideally-expanded, and over-expanded ow regimes for multiple axial and radial positions of the surface at three different plate lengths. These data are analyzed with far-field noise measurements to relate the shock cell structure to the broadband shock noise produced by a jet near a surface.
    Keywords: Acoustics
    Type: AIAA Paper-2013-0038 , E-18606 , GRC-E-DAA-TN7066 , 51st AIAA Aerospace Sciences Meeting; Jan 07, 2013 - Jan 10, 2013; Grapevine, TX; United States
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  • 3
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    LES of a Jet Excited by the Localized Arc Filament Plasma Actuators (2011)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The fluid dynamics of a high-speed jet are governed by the instability waves that form in the free-shear boundary layer of the jet. Jet excitation manipulates the growth and saturation of particular instability waves to control the unsteady flow structures that characterize the energy cascade in the jet.The results may include jet noise mitigation or a reduction in the infrared signature of the jet. The Localized Arc Filament Plasma Actuators (LAFPA) have demonstrated the ability to excite a high-speed jets in laboratory experiments. Extending and optimizing this excitation technology, however, is a complex process that will require many tests and trials. Computational simulations can play an important role in understanding and optimizing this actuator technology for real-world applications. Previous research has focused on developing a suitable actuator model and coupling it with the appropriate computational fluid dynamics (CFD) methods using two-dimensional spatial flow approximations. This work is now extended to three-dimensions (3-D) in space. The actuator model is adapted to a series of discrete actuators and a 3-D LES simulation of an excited jet is run. The results are used to study the fluid dynamics near the actuator and in the jet plume.
    Keywords: Acoustics
    Type: NASA/TM-2011-216984 , AIAA Paper 2010-3872 , E-17625 , 16th Aeroacoustics Conference; Jun 07, 2010 - Jun 09, 2010; Stockholm; Sweden
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  • 4
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    Hybrid Wing Body Shielding Studies Using an Ultrasonic Configurable Fan Artificial Noise Source Generating Simple Modes (2012)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: An Ultrasonic Configurable Fan Artificial Noise Source (UCFANS) was designed, built, and tested in support of the Langley Research Center s 14- by 22-Foot wind tunnel test of the Hybrid Wing Body (HWB) full three-dimensional 5.8 percent scale model. The UCFANS is a 5.8 percent rapid prototype scale model of a high-bypass turbofan engine that can generate the tonal signature of candidate engines using artificial sources (no flow). The purpose of the test was to provide an estimate of the acoustic shielding benefits possible from mounting the engine on the upper surface of an HWB aircraft and to provide a database for shielding code validation. A range of frequencies, and a parametric study of modes were generated from exhaust and inlet nacelle configurations. Radiated acoustic data were acquired from a traversing linear array of 13 microphones, spanning 36 in. Two planes perpendicular to the axis of the nacelle (in its 0 orientation) and three planes parallel were acquired from the array sweep. In each plane the linear array traversed five sweeps, for a total span of 160 in. acquired. The resolution of the sweep is variable, so that points closer to the model are taken at a higher resolution. Contour plots of Sound Pressure Level, and integrated Power Levels are presented in this paper; as well as the in-duct modal structure.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/TM-2012-217685 , AIAA Paper-2012-2076 , E-18638 , 2012 Aeroacoustic Conference; Jun 04, 2012 - Jun 06, 2012; Colorado Springs, CO; United States
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  • 5
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    Acoustic Measurements of an Uninstalled Spacecraft Cabin Ventilation Fan Prototype (2012)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Sound pressure measurements were recorded for a prototype of a spacecraft cabin ventilation fan in a test in the NASA Glenn Acoustical Testing Laboratory. The axial fan is approximately 0.089 m (3.50 in.) in diameter and 0.223 m (9.00 in.) long and has nine rotor blades and eleven stator vanes. At design point of 12,000 rpm, the fan was predicted to produce a flow rate of 0.709 cu m/s (150 cfm) and a total pressure rise of 925 Pa (3.72 in. of water) at 12,000 rpm. While the fan was designed to be part of a ducted atmospheric revitalization system, no attempt was made to throttle the flow or simulate the installed configuration during this test. The fan was operated at six speeds from 6,000 to 13,500 rpm. A 13-microphone traversing array was used to collect sound pressure measurements along two horizontal planes parallel to the flow direction, two vertical planes upstream of the fan inlet and two vertical planes downstream of the fan exhaust. Measurements indicate that sound at blade passing frequency harmonics contribute significantly to the overall audible noise produced by the fan at free delivery conditions.
    Keywords: Acoustics
    Type: NASA/TM-2012-217692 , E-18375 , 12_158 , Inter-Noise 2012, The 41st International Congress and Exposition on Noise Control Engineering; Aug 19, 2012 - Aug 22, 2012; New York, NY; United States
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  • 6
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    Hybrid Wing Body Shielding Studies Using an Ultrasonic Configurable Fan Artificial Noise Source Generating Typical Turbofan Modes (2014)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: An Ultrasonic Configurable Fan Artificial Noise Source (UCFANS) was designed, built, and tested in support of the NASA Langley Research Center's 14- by 22-ft wind tunnel test of the Hybrid Wing Body (HWB) full 3-D 5.8 percent scale model. The UCFANS is a 5.8 percent rapid prototype scale model of a high-bypass turbofan engine that can generate the tonal signature of proposed engines using artificial sources (no flow). The purpose of the test was to provide an estimate of the acoustic shielding benefits possible from mounting the engine on the upper surface of an HWB aircraft using the projected signature of the engine currently proposed for the HWB. The modal structures at the rating points were generated from inlet and exhaust nacelle configurations--a flat plate model was used as the shielding surface and vertical control surfaces with correct plan form shapes were also tested to determine their additional impact on shielding. Radiated acoustic data were acquired from a traversing linear array of 13 microphones, spanning 36 in. Two planes perpendicular, and two planes parallel, to the axis of the nacelle were acquired from the array sweep. In each plane the linear array traversed four sweeps, for a total span of 168 in. acquired. The resolution of the sweep is variable, so that points closer to the model are taken at a higher resolution. Contour plots of Sound Pressure Levels, and integrated Power Levels, from nacelle alone and shielded configurations are presented in this paper; as well as the in-duct mode power levels
    Keywords: Research and Support Facilities (Air); Acoustics; Aircraft Propulsion and Power
    Type: NASA/TM-2014-218119 , AIAA Paper 2014-0256 , E-18861 , SciTech 2014; Jan 13, 2014 - Jan 17, 2014; Baltimore, MD; United States
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  • 7
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    Jet-Surface Interaction Test: Far-Field Noise Results (2012)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Many configurations proposed for the next generation of aircraft rely on the wing or other aircraft surfaces to shield the engine noise from the observers on the ground. However, the ability to predict the shielding effect and any new noise sources that arise from the high-speed jet flow interacting with a hard surface is currently limited. Furthermore, quality experimental data from jets with surfaces nearby suitable for developing and validating noise prediction methods are usually tied to a particular vehicle concept and, therefore, very complicated. The Jet/Surface Interaction Test was intended to supply a high quality set of data covering a wide range of surface geometries and positions and jet flows to researchers developing aircraft noise prediction tools. During phase one, the goal was to measure the noise of a jet near a simple planar surface while varying the surface length and location in order to: (1) validate noise prediction schemes when the surface is acting only as a jet noise shield and when the jet/surface interaction is creating additional noise, and (2) determine regions of interest for more detailed tests in phase two. To meet these phase one objectives, a flat plate was mounted on a two-axis traverse in two distinct configurations: (1) as a shield between the jet and the observer (microphone array) and (2) as a reflecting surface on the opposite side of the jet from the observer.
    Keywords: Acoustics
    Type: GT2012-69639 , ASME/Turbo Expo 2012; Jun 11, 2012 - Jun 15, 2012; Copenhagen; Denmark
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  • 8
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    Developing an Empirical Model for Jet-Surface Interaction Noise (2014)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The process of developing an empirical model for jet-surface interaction noise is described and the resulting model evaluated. Jet-surface interaction noise is generated when the high-speed engine exhaust from modern tightly integrated or conventional high-bypass ratio engine aircraft strikes or flows over the airframe surfaces. An empirical model based on an existing experimental database is developed for use in preliminary design system level studies where computation speed and range of configurations is valued over absolute accuracy to select the most promising (or eliminate the worst) possible designs. The model developed assumes that the jet-surface interaction noise spectra can be separated from the jet mixing noise and described as a parabolic function with three coefficients: peak amplitude, spectral width, and peak frequency. These coefficients are fit to functions of surface length and distance from the jet lipline to form a characteristic spectra which is then adjusted for changes in jet velocity and/or observer angle using scaling laws from published theoretical and experimental work. The resulting model is then evaluated for its ability to reproduce the characteristic spectra and then for reproducing spectra measured at other jet velocities and observer angles; successes and limitations are discussed considering the complexity of the jet-surface interaction noise versus the desire for a model that is simple to implement and quick to execute.
    Keywords: Aircraft Design, Testing and Performance
    Type: NASA/TM-2014-218120 , AIAA Paper 2014-0878 , E-18862 , SciTech 2014; Jan 13, 2014 - Jan 17, 2014; National Harbor, Maryland; United States
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  • 9
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    Data Quality Assurance for Supersonic Jet Noise Measurements (2010)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The noise created by a supersonic aircraft is a primary concern in the design of future high-speed planes. The jet noise reduction technologies required on these aircraft will be developed using scale-models mounted to experimental jet rigs designed to simulate the exhaust gases from a full-scale jet engine. The jet noise data collected in these experiments must accurately predict the noise levels produced by the full-scale hardware in order to be a useful development tool. A methodology has been adopted at the NASA Glenn Research Center s Aero-Acoustic Propulsion Laboratory to insure the quality of the supersonic jet noise data acquired from the facility s High Flow Jet Exit Rig so that it can be used to develop future nozzle technologies that reduce supersonic jet noise. The methodology relies on mitigating extraneous noise sources, examining the impact of measurement location on the acoustic results, and investigating the facility independence of the measurements. The methodology is documented here as a basis for validating future improvements and its limitations are noted so that they do not affect the data analysis. Maintaining a high quality jet noise laboratory is an ongoing process. By carefully examining the data produced and continually following this methodology, data quality can be maintained and improved over time.
    Keywords: Acoustics
    Type: NASA/TM-2010-216767 , GT2010-22545 , E-17392 , Turbo Expo 2010; Jun 14, 2010 - Jun 18, 2010; Glasgow, Scotland; United Kingdom
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  • 10
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    Effect of Turbulence Modeling on an Excited Jet (2010)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The flow dynamics in a high-speed jet are dominated by unsteady turbulent flow structures in the plume. Jet excitation seeks to control these flow structures through the natural instabilities present in the initial shear layer of the jet. Understanding and optimizing the excitation input, for jet noise reduction or plume mixing enhancement, requires many trials that may be done experimentally or computationally at a significant cost savings. Numerical simulations, which model various parts of the unsteady dynamics to reduce the computational expense of the simulation, must adequately capture the unsteady flow dynamics in the excited jet for the results are to be used. Four CFD methods are considered for use in an excited jet problem, including two turbulence models with an Unsteady Reynolds Averaged Navier-Stokes (URANS) solver, one Large Eddy Simulation (LES) solver, and one URANS/LES hybrid method. Each method is used to simulate a simplified excited jet and the results are evaluated based on the flow data, computation time, and numerical stability. The knowledge gained about the effect of turbulence modeling and CFD methods from these basic simulations will guide and assist future three-dimensional (3-D) simulations that will be used to understand and optimize a realistic excited jet for a particular application.
    Keywords: Acoustics
    Type: NASA/TM-2010-216768 , AIAA Paper 2009-3191 , E-17393 , 15th Aeroacoustics Conference (30th AIAA Aeroacoustics Conference); May 11, 2009 - May 13, 2009; Miami, FL; United States
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