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Dec 1977

Volume 62, Issue S1, pp. S1-S102

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back to top Session II. Noise VI: Aeroacoustics
Contributed Papers
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Generation of acoustic tones by flow through multiholed plates (A)

F. C. De Metz and David W. Taylor

J. Acoust. Soc. Am. Volume 62, Issue S1, pp. S80-S80 (1977); (1 page)

Online Publication Date: 11 Aug 2005

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The mechanisms and scaling laws for acoustic tone generation by gas and liquid flow through multiholed plates have been studied experimentally. Rigid and compliant multiholed plates were tested in quiet air and water flow facilities to determine hydroacoustic and hydroelastic resonance frequencies as a function of hole diameter, length and spacing, and flow conditions. The resonance frequencies are expressed as a Strouhal number based on hole dimensions and flow velocity and presented as a function of Reynolds number, hole shape, and Mach number. The scaling laws for predicting upper bounds for the tone pressure amplitudes are also presented.
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Suppression of aerodynamically induced cavity pressure oscillations (A)

L. L. Shaw

J. Acoust. Soc. Am. Volume 62, Issue S1, pp. S80-S80 (1977); (1 page)

Online Publication Date: 11 Aug 2005

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A flight test program was performed to gain further insight into the phenomenon of flow‐induced cavity pressure oscillations and to evaluate the effectiveness of numerous suppression concepts in eliminating or reducing the pressure oscillations. The cavities tested were rectangular with approximate dimensions of 17‐in. long, 8‐in. deep, and 9‐in. wide and were instrumented with microphones, static pressure ports, and a thermocouple. The flight speeds ranged from Mach number 0.6 to 1.3 at pressure altitudes of 3000, 20 000, and 30 000 ft. The suppression devices included leading edge spoilers and deflectors and trailing edge ramps and deflectors. Several combinations of these were tested. The results indicate that the flow‐induced pressure oscillations in a cavity, of the dimensions tested and for the speed range tested, can be significantly reduced with leading edge spoilers in conjunction with a trailing edge ramp. A 40‐dB reduction was achieved for the predominant modal frequency. Other combinations of the suppression devices afforded some reduction but the spoiler‐ramp combination proved most effective.
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Transfer matrix method for sound transmission in acoustically lined horns (A)

Roy J. Beckemeyer

J. Acoust. Soc. Am. Volume 62, Issue S1, pp. S80-S81 (1977); (2 pages) | Cited 1 time

Online Publication Date: 11 Aug 2005

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One‐dimensional sound propagation in a waveguide of continuously varying area is governed by Webster's equation. For certain axial variations in area and in locally reacting wall impedance, closed form solutions can be obtained. However, for arbitrary variations in these parameters as well as for more complicated forms of the equation, numerical solution techniques are required. In this paper, modified versions of Webster's equation are derived which allow consideration of horns with axial temperature gradients and with inserts of porous material. Numerical solutions to the equations are effected by means of a transfer matrix technique. The horn is segmented axially and the segment length is used as an expansion parameter with which to perform an asymptotic expansion of the governing differential equation. As a result the segment boundary value problem is replaced by a sequence of initial value problems. These are solved to yield a transfer matrix relating the acoustic pressure and particle velocity on the right end of the segment to those on the left. Transfer matrices are developed for each of the horn configurations for which governing equations are derived. Some sample numerical results are included.
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Noise of deflectors used for flow attachment with STOL‐OTW configurations (A)

U. von Glahn and D. Groesbeck

J. Acoust. Soc. Am. Volume 62, Issue S1, pp. S81-S81 (1977); (1 page)

Online Publication Date: 11 Aug 2005

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For future STOL aircraft utilizing engine‐over‐the‐wing (OTW) installations in which the exhaust nozzles are located above and separated from the upper surface of the wing, an external jet‐flow deflector can be used to provide flow attachment to the wing/flap surfaces for lift augmentation. In the present work, the deflector noise in the flyover plane measured with several model‐scale nozzle/deflector configurations is examined. The OTW configurations consisted of a circular nozzle (5.18‐cm diameter) mounted 0.1 wing chord lengths above and aft of the wing leading edge. Wing chords (flaps retracted) of 33 and 49.5 cm were used. Tests included flap settings of 20° and 60° and jet Mach numbers of 0.6 and 0.8. The data indicate that the effective sound pressure level of the deflector peaks in the forward quadrant. With some deflector/wing combinations, a severe tone/haystack was obtained at a model‐scale frequency near 6‐8000 Hz. The effective sound pressure level is a primary function of deflector immersion in the jet flow and the deflector span. Peak deflector noise levels as much as 12 dB above those of other jet/flap interaction noise sources were obtained.
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Computational experiments on the effects of inlet turbulence and inflow distortion on fan noise (A)

John F. Groeneweg

J. Acoust. Soc. Am. Volume 62, Issue S1, pp. S81-S81 (1977); (1 page)

Online Publication Date: 11 Aug 2005

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Existing solutions to the problem of sound production by an axial‐flow fan interacting with inlet turbulence and inflow distortion (Marvin E. Goldstein et al., NASA TN D‐7676, May 1974 and NASA TN D‐7667, June 1974) were evaluated over a range of inflow parameters. Sound power in the fan fundamental and second harmonic tones was computed and the modal power distributions contributing to the total tone power were identified. Comparisons and contrasts between the relations of input parameters to sound power produced were made for the two forms of inflow disturbance. It was found that the bulk of the modal power was carried by modes near cutoff for both forms of inflow disturbance when described by parameter values typical of those expected in practical fan test situations. The implications of the numerical results for the design of passive inlet flow control devices proposed to simulate flight conditions during static fan tests are discussed.
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Effectiveness of an inlet flow turbulence control device to simulate flight fan noise in an anechoic chamber (A)

R. P. Woodward, J. A. Wazyniak, L. M. Shaw, and M. J. MacKinnon

J. Acoust. Soc. Am. Volume 62, Issue S1, pp. S81-S81 (1977); (1 page)

Online Publication Date: 11 Aug 2005

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A hemispherical inlet flow control device was tested on a 50.8‐cm (20‐in.) diameter fan stage in the NASA‐Lewis anechoic chamber. The control device used honeycomb and wire mesh to reduce turbulence intensities entering the fan. Farfield acoustic power level results showed about a 5‐dB reduction in blade passing tone noise at 90% design fan speed with the inlet device in place. The device resulted in about a 10‐dB reduction in multiple pure tone generation at this fan speed. Hot film cross probes were inserted in the inlet to obtain data for two components of the turbulence at 65% and 90% design fan speed. Without the flow control device the axial intensities were below 1.0% while the transverse intensities were almost twice this value. The inflow control device reduced the circumferential turbulence intensities by a factor of three and also reduced the axial length scale.
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Noise‐related turbulence measurements for various surface geometries (A)

W. A. Olsen and D. Boldman

J. Acoust. Soc. Am. Volume 62, Issue S1, pp. S81-S81 (1977); (1 page)

Online Publication Date: 11 Aug 2005

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Fundamental theories for noise generated by flow over surfaces exist for only a few simple surface shapes. The role of turbulence in noise generation by complex surfaces should be essentially the same as for simple surfaces. Examination of simple‐surface theories indicates that the spatial distributions of the mean velocity and turbulence (intensity, spectra, and integral scale length) are sufficient to define the noise emission. Measurements of these flow properties were made for a number of simple and complex (STOL aircraft blown flaps) surfaces. The configurations were selected because their acoustic characteristics (i.e., shape of radiation pattern and velocity power law) are quite different. The spatial distribution of the flow properties around the complex surfaces and approximate theory are used to qualitatively explain the varied acoustic characteristics. The data properties possess some simplifications, which can be used in the theories. The major simplification is that the turbulence spectra are essentially described by the isotropic turbulence relationship, independent of the surface configuration and of the location in the flow field. This result is used to partially explain why the shape of acoustic spectra vary little and why acoustic model data accurately scale to larger size.
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Empirical model for inverted‐velocity‐profile jet noise prediction (A)

James R. Stone

J. Acoust. Soc. Am. Volume 62, Issue S1, pp. S81-S81 (1977); (1 page)

Online Publication Date: 11 Aug 2005

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An empirical model for predicting the noise from inverted velocity‐profile coaxial or coannular jets is presented and compared with small‐scale static and simulated‐flight data. The model considers the combined contributions of as many as four uncorrelated constituent sources: the premerged‐jet/ambient mixing region, the merged‐jet/ambient mixing region, outer‐stream shock/turbulence interaction, and inner‐stream shock/turbulence interaction. The model for both mixing regions is developed from the NASA interim prediction method for jet noise. The noise from the merged region occurs at relatively low frequency and is modeled as the contribution of a circular jet at merged conditions (between inner and outer streams) and total exhaust area, with the high frequencies attenuated (since the high‐frequency region of this fictitious jet does not exist). The noise from the premerged region occurs at higher frequency and is modeled as the contribution of an equivalent plug nozzle at outer stream conditions, with the low frequencies attenuated (since the outer jet is broken up rapidly, before much low frequency is generated). The shock noise for each supersonic stream is calculated from a modification of the Harper‐Bourne and Fisher (1973) model.
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Space‐time correlations of airfoil noise (A)

W. R. Miller, M. P. Bucha, and W. C. Meeecham

J. Acoust. Soc. Am. Volume 62, Issue S1, pp. S81-S82 (1977); (2 pages)

Online Publication Date: 11 Aug 2005

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Space‐time correlation measurements of surface pressure fluctuations on airfoils, immersed in a cold air jet, with their farfield sound were made over the exit velocity range of 68–190 m/sec (Mach 0.2–0.6). Measurements of surface pressure fluctuations were taken both at the one‐third chord and the trailing edge of a 1‐in. chord airfoil. The position of the farfield microphone was varied throughout the hemispherical region downstream of the jet exit. Analysis of the cross‐correlation and delay time measurements on a 1‐in. chord airfoil indicate that the dominant source region for the radiated sound is near the trailing edge. However, for the probe microphone mounted 1 in. from the leading edge of a second, 3‐in. chord airfoil, two correlation peaks were recorded. The first corresponds to direct radiation from the probe location, and the second is due to acoustic transmission upstream from the trailing edge. The normalized cross‐correlation coefficients were in a ratio of 1:5, respectively. This indicates that the sound from midchord is down 14 dB from the trailing edge sound; theory would indicate even less midchord sound. [Work partially supported by NASA and the TRW Foundation.]
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Combustor fluctuating pressure measurements in engine and in a component test facility — a preliminary comparison (A)

Meyer Reshotko and Allen Karchmer

J. Acoust. Soc. Am. Volume 62, Issue S1, pp. S82-S82 (1977); (1 page)

Online Publication Date: 11 Aug 2005

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As part of a program to investigate combustor noise, simultaneous measurements were made with a YF‐102 engine of combustor internal fluctuating pressure and farfield noise. As the dependence of farfield noise on engine internal measurement becomes understood, an important question becomes the connection between combustor internal measurements obtained in an engine and those obtained in a component test facility. To explore this question, a YF‐102 combustor, instrumented identically with that of the engine tests, was operated in a component test facility over a range of conditions encompassing engine operation, and the fluctuating internal measurements were repeated. A comparison of the directly measured spectra at corresponding locations in the two tests shows significant differences. However, the results of two‐point signal analyses within each combustor such as coherence function, transfer function, and phase relationships are similar for both tests. This indicates that the internal dynamics of the combustor as an acoustic source are preserved in a component test facility.
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Water‐flow induced cavity resonances (A)

T. M. Farabee and F. C. DeMetz

J. Acoust. Soc. Am. Volume 62, Issue S1, pp. S82-S82 (1977); (1 page)

Online Publication Date: 11 Aug 2005

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The mechanisms of water flow‐excited acoustic resonances in simple cylindrical cavity systems with circular and slot openings were investigated experimentally. The resonance frequencies of the internal fluctuating pressure field were determined as a function of turbulent boundary‐layer thickness, free‐stream flow velocity, cavity structural parameters, and opening dimensions for flush‐mounted cavities on a large self‐propelled streamlined body of revolution. The scaling laws of the cavity resonance frequencies and peak forcing function amplitudes are considered in terms of feedback mechanisms between the shear layer instabilities in the cavity opening and the vibrations of the cavity structures. Predicted and measured results are reported for the amplitude of the pressure field radiated from the cavity openings.
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Measurements of the low wave number components of turbulent boundary layer wall pressure fluctuations with zero and adverse pressure gradients (A)

M. Moeller, P. Leehey, and N. C. Martin

J. Acoust. Soc. Am. Volume 62, Issue S1, pp. S82-S82 (1977); (1 page)

Online Publication Date: 11 Aug 2005

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Two rectangular plates with approximately clamped boundary conditions were used as spatial filters to measure the low wave number components of the wall pressure fluctuations beneath a turbulent boundary layer. The plates were designed to provide low wave number measurements at higher frequencies and wave numbers than previous plate experiments in order to provide data comparable to measurements made using microphone arrays. Measurements were made in the lower test section wall of the MIT low turbulence subsonic wind tunnel under zero pressure gradient and adverse pressure gradient conditions. The zero pressure gradient data lies in the wave number frequency range 2.5 <ωδ∗/U∞<10 and 0.5 <k1δ∗<1.8 while the adverse gradient data lies in the range 4.5 <ωδ∗/U∞<23 and 0.9 <k1δ∗<4.6. Momentum thickness Reynolds numbers ranged from 6000 to 11 000 for the zero pressure gradient case and 11 000 to 16 000 for the adverse pressure gradient case. When normalized on outer variables the adverse gradient data do not exhibit increased levels from those of the zero pressure gradient data. [Work supported by the Sonar Technology Program, Office of Naval Research.]
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