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May 1978

Volume 63, Issue S1, pp. S1-S87

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back to top Session H. Noise I: Generation and Measurement
Contributed Papers
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A possible unsteady thickness noise mechanism for helicopter rotors (A)

D. L. Hawkings

J. Acoust. Soc. Am. Volume 63, Issue S1, pp. S21-S21 (1978); (1 page)

Online Publication Date: 11 Aug 2005

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An alternative theoretical approach to the low‐speed helicopter rotor noise problem is proposed. The traditional treatment of this problem is to model the situation by a system of unsteady thrust and drag dipoles. The dipole strengths are either estimated empirically, or related to the inflow distortion via an unsteady aerodynamic response analysis. An analysis is presented which suggests that an alternative model consists of a system of unsteady lift dipoles, plus a distribution of inplane quadrupoles. The strength of these quadrupoles is related directly to the product of the blade thickness and the distortion inplane velocity components. It is shown how this mechanism is likely to be the cause of the subjectively important higher harmonic rotational noise observed in or near the rotor plane. Some comparisons with experiment are presented.
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Airframe aerodynamic noise—total radiated acoustic power approach (A)

L. L. Shaw and D. R. Houser

J. Acoust. Soc. Am. Volume 63, Issue S1, pp. S22-S22 (1978); (1 page)

Online Publication Date: 11 Aug 2005

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During flight the noise radiated by aircraft is emanating from two distinct types of sources. One source is the propulsion system and the other is the nonpropulsion system noise, or airframe noise, associated with movement of the aircraft through the atmosphere. The purpose of this effort was to study airframe noise using a total radiated acoustic power approach. Methodology was developed to accurately calculate the total acoustic power by using measurements from an array of microphones during aircraft flyover. This methodology was applied to Schweizer 2–32 glider flyovers and it was found that for an aerodynamic configuration (no flaps, wheels, wheel wells, etc.) the total acoustic power can be obtained from one flyover measurement by assuming the directivity is nearly equal in all directions. This assumption was shown to be valid for the glider and is assumed valid for any aircraft in an aerodynamic configuration. The detailed methodology developed is still useful since most commercial aircraft land in a nonaerodynamic configuration and thus their directivity is not equal in all directions. The results from the glider tests were compared to data in the literature and found to agree well. Variation of the total power with aircraft velocity followed a V6 law. The parameter which best normalizes the overall acoustic power from different aircraft was found to be the wing area.
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Fine structure of airfoil tone frequency (A)

M. R. Fink

J. Acoust. Soc. Am. Volume 63, Issue S1, pp. S22-S22 (1978); (1 page)

Online Publication Date: 11 Aug 2005

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A method is described for calculating frequency of tone noise radiated by constant‐chord airfoils in uniform low‐turbulence flow. This noise mechanism is important for small aircraft propellers and helicopter rotors, and for automotive radiator cooling fans. Tones are attributed to a feedback loop between instability waves convected downstream within a laminar boundary layer extending to the trailing edge and sound waves generated at the trailing edge and radiated upstream in the inviscid flow. The upstream extent of the feedback path is the most forward distance at which a laminar fluctuation at the tone frequency is unstable. For small increases of flow velocity, the total number of phase cycles around the feedback loop remains constant and the frequency increases less rapidly than flow velocity, When frequency has increased such that the ratio of chord to half the acoustic wavelength becomes an integer, the number of phase cycles abruptly increases to the largest integer for Which a large‐amplitude feedback cycle can occur. The resulting calculated steplike jumps of tone frequency generally match the available data.
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Relative velocity exponents for jet engine exhaust noise (A)

James R. Stone

J. Acoust. Soc. Am. Volume 63, Issue S1, pp. S22-S22 (1978); (1 page)

Online Publication Date: 11 Aug 2005

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The effect of flight on jet engine exhaust noise has often been presented in terms of a relative velocity exponent n as a function of radiation angle. Prediction methods (e.g., Bushell, 1975) have also been proposed on this basis. The value of n is given by the OASPL reduction due to relative velocity divided by ten times the logarithm of the ratio of relative jet velocity to absolute jet velocity. In such terms, classical jet noise theory (Ffowes Williams, 1963) would predict a value of n = 7 at 90° to the jet axis with n decreasing as the inlet axis is approached and increasing as the jet axis is approached. However, flight tests have shown a wide range of results, including negative values of n in some cases. It is shown in this paper that the exponent n is positive for pure jet mixing noise and varies, in a predictable manner, as a function of jet conditions and flight velocity. It is also shown that when the effects of shock noise and internally generated noise are considered the range of n values predicted is increased, and in some cases negative values of n are predicted. These calculations are based on simple empirical models for jet mixing noise, shock noise and internally generated noise.
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Reduction of fan noise in an anechoic chamber by reducing chamber wall induced inlet flow disturbances (A)

J. H. Dittmar

J. Acoust. Soc. Am. Volume 63, Issue S1, pp. S22-S22 (1978); (1 page)

Online Publication Date: 11 Aug 2005

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The difference between flight and ground static noise data has arisen in the past few years as a significant problem in fan jet engine noise testing. The additional noise for static testing has been attributed to inlet flow disturbances or turbulence interacting with the fan rotor. In an attempt to determine a possible source of inflow disturbances entering fans tested in the Lewis Research Center anechoic chamber the inflow field was studied using potential flow analyses. These potential flow calculations indicated that there was substantial flow over the wall directly behind the fan inlet. This flow near the wall anechoic wedges could produce significant inflow disturbances. Fan noise tests were run with various extensions added to the fan inlet ducting to move the inlet away from this back wall and thereby reduce the inlet flow disturbances. Significant noise reductions were observed with increased inlet length. Over 5‐dB reduction of the blade passage tone sound power level was observed between the shortest and longest inlet at 90% fan speed and the first overtone was reduced 9 dB. High‐frequency reductions in the broadband noise were also observed.
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Measurements of noise of centrifugal fans by flush‐mounted microphones inside the exhaust ducts (A)

G. Krishnappa

J. Acoust. Soc. Am. Volume 63, Issue S1, pp. S22-S23 (1978); (2 pages)

Online Publication Date: 11 Aug 2005

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There are no existing standards on measurements of noise radiated by fans and blowers inside small exhaust ducts. Noise measurements by in‐duct flush‐mounted microphones are likely to be affected by pressure fluctuations caused by the presence of the turbulent boundary layer along the duct. By using cross‐correlation techniques, the validity of such measurements are examined in this paper. Noise signals produced by several impellers of 203‐mm diameter operating inside two different casing configurations were measured in a centrifugal fan noise test rig containing a long 50‐mm‐diameter exhaust duct, with an anechoic termination. The signals from two flush‐mounted microphones separated by a distance of 1.22 m were cross correlated at several fan speeds and flow velocities. There were good correlations observed between the two microphone signals at low flow velocities for the time interval corresponding to the acoustic velocity. Although at higher flow velocities the correlation between the two overall signals was poor, encouraging correlations were obtained for the filtered signals in octave bands at frequencies above 125 Hz.
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Marching techniques for estimating duct attenuation and source pressure profiles (A)

K. J. Baumeister

J. Acoust. Soc. Am. Volume 63, Issue S1, pp. S23-S23 (1978); (1 page)

Online Publication Date: 11 Aug 2005

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A numerical method is developed to predict the pressure distribution of a ducted source from farfield pressure measurements. Using an initial value formulation, the Helmholtz wave equation (no steady flow) is solved using explicit marching techniques which can readily be extended to three‐dimensional and nonlinear problems. The Von Neumann method is used to develop relationships which describe how sound frequency and grid spacing affect numerical stability. At the present time, stability considerations limit the approach to high‐frequency sound. Example calculations for both hard and soft wall ducts compare favorably to known boundary value solutions. In addition, under the assumption that aerodynamic forces which produce sound dominate (negligible reflections), this initial‐value approach is successfully used to determine the attenuation of a straight soft wall duct. Compared to conventional finite‐difference or finite‐element boundary value approaches, the numerical marching technique is orders of magnitude shorter in computation time and can be easily employed in problems involving high‐frequency sound.
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Prediction of interface wave profiles for a flow‐excited cavity resonator (A)

S. A. Elder

J. Acoust. Soc. Am. Volume 63, Issue S1, pp. S23-S23 (1978); (1 page)

Online Publication Date: 11 Aug 2005

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Experimental and theoretical results are presented for a wall‐mounted cavity in turbulent flow, oscillating at Helmholtz or depth mode resonance. Data for interface waves at the mouth of the cavity, obtained by the CASH method, can be fitted to limiting expressions obtained from linearized instability theory. At typical steady‐state amplitudes, the wavelength of the disturbance agrees well with the predictions of Michalke [A. Michalke, J. Fluid Mech. 23, 521–544 (1965)], using an equivalent laminar flow model based on the mean velocity profile. Simple exponential growth of the driven wave does not occur, however, on account of nonlinear effects. Near the leading edge, the interface displacement is a superposition of a traveling wave and a stationary acoustic contribution. For the portion of the profile near the trailing edge, the traveling wave becomes predominant. Theoretical interface profiles have been used successfully in a root‐locus solution of the resonant frequency lock‐in problem, making it possible to predict Strouhal number and amplitude of cavity oscillation. [Work supported by a GHR contract with the Naval Ship Systems Command, administered by the Naval Ship Research & Development Center, Carderock, MD.]
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Calibration of “half‐inch” condenser microphones: Comparison of selected electrostatic actuator methods with measurement of pressure response in ANSI couplers (A)

V. Nedzelnitsky, W. B. Penzes, and F. P. Lalli

J. Acoust. Soc. Am. Volume 63, Issue S1, pp. S23-S23 (1978); (1 page)

Online Publication Date: 11 Aug 2005

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Coupler techniques based upon reciprocity are generally considered the most accurate method for absolute pressure calibration of laboratory microphones. Electrostatic actuator techniques are convenient and widely used for obtaining the relative frequency‐response levels of such microphones. However, actuator‐determined response levels can differ appreciably from the pressure‐response level because the electrostatically induced motion of a microphone diaphragm is influenced by the relation between its own acoustic impedance and the radiation impedance with which it is loaded. To evaluate the magnitude of this difference in some practical situations, microphones of different nominal equivalent volumes and resonant frequencies were calibrated in standard couplers (ANSI S1.10‐1966, R1971) over the frequency range 50–20 000 Hz. Response levels for these microphones were determined with electrostatic actuators and normalized to coincide, at 700 Hz, with the coupler‐determined response levels. Differences between the normalized actuator‐determined levels and the coupler‐determined levels were usually largest at frequencies above 5 kHz and were dependent upon actuator configuration and microphone type. Differences of less than 0.3 dB could be found for microphones of high nominal resonant frequency and low nominal equivalent volume, but greater differences (e.g., 1 dB) could be observed for microphones of low resonant frequency and high equivalent volume.
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