Calibrations of the NASA Langley 14- by 22-Foot Subsonic Tunnel in Acoustic Configuration

Spalt, Taylor B.; Brooks, Thomas F.; Bahr, Christopher J.; Becker, Lawrence E.; Stead, Daniel J.; Plassman, Gerald E.

doi:10.2514/6.2014-2344

Cited by 10 publications

(10 citation statements)

References 15 publications

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“…The purpose of the detonations was to identify any significant spurious reflection paths from the location of anticipated noise sources on the model to the array (e.g., unwanted reflections from the floor, sidewalls, tunnel flow collector, etc.). Details of the acoustic analysis procedure for the pyrotechnic runs are described by Spalt et al 23 Because of the extensive acoustic treatment applied to the test section floor, sidewalls, and ceiling, analysis of the data revealed no appreciable reflection paths that would interfere with array operation.…”

Section: B 14 X 22 Tunnel Background Noisementioning

confidence: 99%

Aeroacoustic Evaluation of Flap and Landing Gear Noise Reduction Concepts

Khorrami

Humphreys

Lockard

et al. 2014

20th AIAA/CEAS Aeroacoustics Conference

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Aeroacoustic measurements for a semi-span, 18% scale, high-fidelity Gulfstream aircraft model are presented. The model was used as a test bed to conduct detailed studies of flap and main landing gear noise sources and to determine the effectiveness of numerous noise mitigation concepts. Using a traversing microphone array in the flyover direction, an extensive set of acoustic data was obtained in the NASA Langley Research Center 14-by 22-Foot Subsonic Tunnel with the facility in the acoustically treated openwall (jet) mode. Most of the information was acquired with the model in a landing configuration with the flap deflected 39º and the main landing gear alternately installed and removed. Data were obtained at Mach numbers of 0.16, 0.20, and 0.24 over directivity angles between 56º and 116º, with 90º representing the overhead direction. Measured acoustic spectra showed that several of the tested flap noise reduction concepts decrease the sound pressure levels by 2 -4 dB over the entire frequency range at all directivity angles. Slightly lower levels of noise reduction from the main landing gear were obtained through the simultaneous application of various gear devices. Measured aerodynamic forces indicated that the tested gear/flap noise abatement technologies have a negligible impact on the aerodynamic performance of the aircraft model.

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Section: B 14 X 22 Tunnel Background Noisementioning

confidence: 99%

Aeroacoustic Evaluation of Flap and Landing Gear Noise Reduction Concepts

Khorrami

Humphreys

Lockard

et al. 2014

20th AIAA/CEAS Aeroacoustics Conference

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“…Heath et al 9,10 described in greater detail the preparations and facility upgrades required to support the HWB acoustic test. While the current manuscript describes the jet noise investigations, colleagues involved in other aspects of the test have provided companion papers describing the microphone phased array development, 11 acoustic calibration procedures, 12 acoustic data processing, 13 shielding characteristics of a mode-generating source, 14 shielding of broadband turbomachinery noise, 15 airframe noise, 16 and community noise metrics. 17 The current study provides an overview of the jet noise portion of the HWB acoustic test and focuses on the shielding benefits observed in conjunction with both the baseline axisymmetric nozzle system and the low noise nozzle system as the engines are translated relative to the fuselage trailing edge.…”

Section: Introductionmentioning

confidence: 99%

Jet noise shielding provided by a hybrid wing body aircraft

Doty

Brooks

Burley

et al. 2018

International Journal of Aeroacoustics

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One approach toward achieving NASA's aggressive N+2 noise goal of 42 EPNdB cumulative margin below Stage 4 is through the use of novel vehicle configurations like the hybrid wing body. Jet noise measurements from a hybrid wing body acoustic test in the NASA Langley 14-by 22-Foot Subsonic Tunnel are described. Two dual-stream, heated Compact Jet Engine Simulator units are mounted underneath the inverted hybrid wing body model on a traversable support to permit measurement of varying levels of shielding provided by the fuselage. Both an axisymmetric and low noise chevron nozzle set are investigated in the context of shielding. The unshielded chevron nozzle set shows 1-2 dB of source noise reduction (relative to the unshielded axisymmetric nozzle set) with some penalties at higher frequencies. Shielding of the axisymmetric nozzles shows up to 6.5 dB of reduction at high frequency. The combination of shielding and low noise chevrons shows benefits beyond the expected additive benefits of the two, up to 10 dB, due to the effective migration of the jet source peak noise location upstream for increased shielding effectiveness. Jet noise source maps from phased array results processed with the deconvolution approach for the mapping of acoustic sources algorithm reinforce these observations.

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“…Multi-positional jet engines and broadband engine noise simulators were also developed by NASA for shielding testing. Aspects of the test, reported in this and companion papers [5][6][7][8][9][10][11][12] , are the examination of noise shielding parameters (such as engine location, vertical and nozzle configurations) with regard to noise emission, and the determination of the noise spectra, levels, and directivity of the base vehicle and its components. Finally, the results from this test are used to support the noise assessment 7 of this HWB design and to validate the "low noise" characteristics of this aircraft concept.…”

Section: Introductionmentioning

confidence: 99%

Shielding of Turbomachinery Broadband Noise from a Hybrid wing Body Aircraft Configuration

Hutcheson

Brooks

Burley

et al. 2014

20th AIAA/CEAS Aeroacoustics Conference

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The results of an experimental study on the effects of engine placement and vertical tail configuration on shielding of exhaust broadband noise radiation are presented. This study is part of the high fidelity aeroacoustic test of a 5.8% scale Hybrid Wing Body (HWB) aircraft configuration performed in the 14-by 22-Foot Subsonic Tunnel at NASA Langley Research Center. Broadband Engine Noise Simulators (BENS) were used to determine insertion loss due to shielding by the HWB airframe of the broadband component of turbomachinery noise for different airframe configurations and flight conditions. Acoustics data were obtained from flyover and sideline microphones traversed to predefined streamwise stations. Noise measurements performed for different engine locations clearly show the noise benefit associated with positioning the engine nacelles further upstream on the HWB centerbody. Positioning the engine exhaust 2.5 nozzle diameters upstream (compared to 0.5 nozzle diameters downstream) of the HWB trailing edge was found of particular benefit in this study. Analysis of the shielding performance obtained with and without tunnel flow show that the effectiveness of the fuselage shielding of the exhaust noise, although still significant, is greatly reduced by the presence of the free stream flow compared to static conditions. This loss of shielding is due to the turbulence in the model near-wake/boundary layer flow. A comparison of shielding obtained with alternate vertical tail configurations shows limited differences in level; nevertheless, overall trends regarding the effect of cant angle and vertical location are revealed. Finally, it is shown that the vertical tails provide a clear shielding benefit towards the sideline while causing a slight increase in noise below the aircraft.

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Calibrations of the NASA Langley 14- by 22-Foot Subsonic Tunnel in Acoustic Configuration

Cited by 10 publications

References 15 publications

Aeroacoustic Evaluation of Flap and Landing Gear Noise Reduction Concepts

Aeroacoustic Evaluation of Flap and Landing Gear Noise Reduction Concepts

Jet noise shielding provided by a hybrid wing body aircraft

Shielding of Turbomachinery Broadband Noise from a Hybrid wing Body Aircraft Configuration

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