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  • 1
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    Effects of acoustic sources (1987)
    In:  CASI
    Details
    Publication Date: 2013-08-31
    Description: An experiment was conducted to determine the effect of acoustics on the laminar flow on the side of a nacelle. A flight test was designed to meet this goal and a brief review of the purpose is given. A nacelle with a significant length of laminar flow was mounted on the wing of NASA OV-1. Two noise sources are also mounted on the wing: one in the center body of the nacelle; the second in a wing mounted pod outboard of the nacelle. These two noise sources allow for a limited study of the effect of source direction in addition to control of the acoustic level and frequency. To determine the range of Tollmien-Schlichting frequencies, a stability analysis using the pressure coefficient distribution along the side of the nacelle was performed. Then by applying these frequencies and varying the acoustic level, a study of the receptivity of the boundary layer to the acoustic signal, as determined by the shortening of the length of laminar flow, was conducted. Results are briefly discussed.
    Keywords: ACOUSTICS
    Type: Research in Natural Laminar Flow and Laminar-Flow Control, Part 3; p 914-921
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  • 2
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    Evaluation of a multi-point method for determining acoustic impedance (1989)
    In:  Other Sources
    Details
    Publication Date: 2011-08-19
    Description: A multipoint method for determining acoustic impedance was evaluated in comparison with the traditional standing wave and two-microphone methods using 30 test samples covering the reflection factor magnitude range 0.004-0.999. The multipoint method is shown to combine the strengths of the standing wave and two-microphone methods while avoiding some of their inherent weaknesses. In particular, the results obtained suggest that the multipoint method will be less subject to flow induced random error than the two-microphone method in the presence of significant broadband noise levels associated with mean flow.
    Keywords: ACOUSTICS
    Type: Mechanical Systems and Signal Processing (ISSN 0888-3270); 3; 15-35
    Format: text
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  • 3
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    Experimental validation of a two-dimensional shear-flow model for determining acoustic impedance (1987)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Tests were conducted to validate a two-dimensional shear-flow analytical model for determining the acoustic impedance of a liner test specimen in a grazing-incidence, grazing-flow environment. The tests were limited to a test specimen chosen to exhibit minimal effects of grazing flow so that the results obtained by using the shear-flow analytical model would be expected to match those obtained from normal-incidence impedance measurements. Impedances for both downstream and upstream sound propagation were generally consistent with those from normal-incidence measurements. However, sensitivity of the grazing-incidence impedance to small measurement or systematic errors in propagation constant varied dramatically over the range of test frequencies.
    Keywords: ACOUSTICS
    Type: NASA-TP-2679 , L-16203 , NAS 1.60:2679
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  • 4
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    Unsteady pressure loads in a generic high speed engine model (1992)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Unsteady pressure loads were measured along the top interior wall of a generic high-speed engine (GHSE) model undergoing performance tests in the combustion-Heated Scramjet Test Facility at the Langley Research Center. Flow to the model inlet was simulated at 72000 ft and a flight Mach number of 4. The inlet Mach number was 3.5 with a total temperature and pressure of 1640 R and 92 psia. The unsteady pressure loads were measured with 5 piezoresistive gages, recessed into the wall 4 to 12 gage diameters to reduce incident heat flux to the diaphragms, and distributed from the inlet to the combustor. Contributors to the unsteady pressure loads included boundary layer turbulence, combustion noise, and transients generated by unstart loads. Typical turbulent boundary layer rms pressures in the inlet ranged from 133 dB in the inlet to 181 dB in the combustor over the frequency range from 0 to 5 kHz. Downstream of the inlet exist, combustion noise was shown to dominate boundary layer turbulence noise at increased heat release rates. Noise levels in the isolator section increased by 15 dB when the fuel-air ratio was increased from 0.37 to 0.57 of the stoichiometric ratio. Transient pressure disturbances associated with engine unstarts were measured in the inlet and have an upstream propagation speed of about 7 ft/sec and pressure jumps of at least 3 psia.
    Keywords: ACOUSTICS
    Type: NASA-TP-3189 , L-16912 , NAS 1.60:3189
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  • 5
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    A finite element propagation model for extracting normal incidence impedance in nonprogressive acoustic wave fields (1995)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A propagation model method for extracting the normal incidence impedance of an acoustic material installed as a finite length segment in a wall of a duct carrying a nonprogressive wave field is presented. The method recasts the determination of the unknown impedance as the minimization of the normalized wall pressure error function. A finite element propagation model is combined with a coarse/fine grid impedance plane search technique to extract the impedance of the material. Results are presented for three different materials for which the impedance is known. For each material, the input data required for the prediction scheme was computed from modal theory and then contaminated by random error. The finite element method reproduces the known impedance of each material almost exactly for random errors typical of those found in many measurement environments. Thus, the method developed here provides a means for determining the impedance of materials in a nonprogressirve wave environment such as that usually encountered in a commercial aircraft engine and most laboratory settings.
    Keywords: ACOUSTICS
    Type: NASA-TM-110160 , NAS 1.15:110160
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  • 6
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    Evaluation of a multi-point method for determining acoustic impedance (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An investigation was conducted to explore potential improvements provided by a Multi-Point Method (MPM) over the Standing Wave Method (SWM) and Two-Microphone Method (TMM) for determining acoustic impedance. A wave propagation model was developed to model the standing wave pattern in an impedance tube. The acoustic impedance of a test specimen was calculated from a best fit of this standing wave pattern to pressure measurements obtained along the impedance tube centerline. Three measurement spacing distributions were examined: uniform, random, and selective. Calculated standing wave patterns match the point pressure measurement distributions with good agreement for a reflection factor magnitude range of 0.004 to 0.999. Comparisons of results using 2, 3, 6, and 18 measurement points showed that the most consistent results are obtained when using at least 6 evenly spaced pressure measurements per half-wavelength. Also, data were acquired with broadband noise added to the discrete frequency noise and impedances were calculated using the MPM and TMM algorithms. The results indicate that the MPM will be superior to the TMM in the presence of significant broadband noise levels associated with mean flow.
    Keywords: ACOUSTICS
    Type: NASA-TM-100637 , NAS 1.15:100637
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  • 7
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    Airfoil profile drag increase due to acoustic excitation (1989)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: A two-dimensional airfoil (NACA-0009) is subjected to high intensity pure-tone sound over a 1-5 kHz frequency range while immersed in a flow with 240 ft/sec velocity in a quiet flow facility with a Reynolds number of 3 million. Wake dynamic pressures are determined, and the momentum deficit is used to calculate a two-dimensional drag coefficient. Significant increases in drag are observed when the airfoil is subjected to high-intensity sound at critical frequencies. The increased drag is accompanied by movement of the natural transition location. When the transition is fixed by roughness at 10 percent chord, no further transition movement is observed in response to an acoustic Tollmien-Schlichting disturbance. However, a 4 percent increase in the sectional drag coefficient is noted. It is believed to be due to the sound exciting the flow near the airfoil surface (shear layer), thus causing the existing turbulence to become more intense, possess a higher mixing rate (momentum), and increase the skin friction.
    Keywords: ACOUSTICS
    Type: AIAA PAPER 89-1069
    Format: text
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  • 8
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    Acoustic effects on profile drag of a laminar flow airfoil (1987)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A two-dimensional laminar flow airfoil (NLF-0414) was subjected to high-intensity sound (pure tones and white noise) over a frequency range of 2 to 5 kHz, while immersed in a flow of 240 ft/sec (Rn of 3 million) in a quiet flow facility. Using a wake-rake, wake dynamic pressures were determined and the deficit in momentum was used to calculate a two dimensional drag coefficient. Significant increases in drag were observed when the airfoil was subjected to the high intensity sound at critical sound frequencies. However, the increased drag was not accompanied by movement of the transition location.
    Keywords: ACOUSTICS
    Type: NASA-TM-100505 , NAS 1.15:100505
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  • 9
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    Fluctuating pressures measured beneath a high-temperature, turbulent boundary layer on a flat plate at Mach number of 5 (1989)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Fluctuating pressures were measured beneath a Mach 5, turbulent boundary layer on a flat plate with an array of piezoresistive sensors. The data were obtained with a digital signal acquisition system during a test run of 4 seconds. Data sampling rate was such that frequency analysis up to 62.5 kHz could be performed. To assess in situ frequency response of the sensors, a specially designed waveguide calibration system was employed to measure transfer functions of all sensors and related instrumentation. Pressure time histories were approximated well by a Gaussian prohibiting distribution. Pressure spectra were very repeatable over the array span of 76 mm. Total rms pressures ranged from 0.0017 to 0.0046 of the freestream dynamic pressure. Streamwise, space-time correlations exhibited expected decaying behavior of a turbulence generated pressure field. Average convection speed was 0.87 of freestream velocity. The trendless behavior with sensor separation indicated possible systematic errors.
    Keywords: ACOUSTICS
    Type: NASA-TP-2947 , L-16596 , NAS 1.60:2947
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  • 10
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    Measured and calculated acoustic attenuation rates of tuned resonator arrays for two surface impedance distribution models with flow (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An experiment was performed to validate two analytical models for predicting low frequency attenuation of duct liner configurations built from an array of seven resonators that could be individually tuned via adjustable cavity depths. These analytical models had previously been developed for high frequency aero-engine inlet duct liner design. In the low frequency application, the liner surface impedance distribution is unavoidably spatially varying by virtue of available fabrication techniques. The characteristic length of this spatial variation may be a significant fraction of the acoustic wavelength. Comparison of measured and predicted attenuation rates and transmission losses for both modal decomposition and finite element propagation models were in good to excellent agreement for a test frequency range that included the first and second cavity resonance frequencies. This was true for either of two surface impedance distribution modeling procedures used to simplify the impedance boundary conditions. In the presence of mean flow, measurements revealed a fine scale structure of acoustic hot spots in the attenuation and phase profiles. These details were accurately predicted by the finite element model. Since no impedance changes due to mean flow were assumed, it is concluded that this fine scale structure was due to convective effects of the mean flow interacting with the surface impedance nonuniformities.
    Keywords: ACOUSTICS
    Type: NASA-TP-2766 , L-16352 , NAS 1.60:2766
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