Optical crossed-beam investigation of local sound generation in jets
Cross correlation of optical signals from two light beam sound detectors for jet noise source localization
Krause, F. R.↗
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An objective method and measuring equipment for noise control and acoustic diagnostics of motorcars
An objective method and apparatus for noise control and acoustic diagnostics of motorcar engines are reported. The method and apparatus let us know whether the noisiness of the vehicle under test exceeds the admissible threshold levels given by appropriate standards and if so what is the main source of the excessive noise. The method consists in measuring both the overall noise level and the sound pressure levels in definite frequency bands while the engine speed is controlled as well and may be fixed at prescribed values. Whenever the individually adjusted threshold level has been exceeded in any frequency band, a self-sustaining control signal is sent.
Kacprowski, J.↗
Far noise field of a two-dimensional subsonic jet
The acoustical radiation characteristics of a high aspect ratio, cold subsonic, model slot jet have been investigated experimentally. Acoustical measurements were made in both a free-field environment and a reverberation field environment. Mean velocity distribution of the slot jet was determined to provide information on the mixing characteristics of the flow. Free-field results obtained reflected a characteristic asymmetrical directivity pattern of the slot jet noise. The total radiated sound power yielded a typical U to the 7-th power dependence on mean jet velocity. The experimental findings are compared with data from similar studies and also with radiation characteristics of circular cold subsonic jets.
Kouts, C.↗
Noise and static performance characteristics of a STOL aircraft jet flap
Static noise and performance tests were conducted on a 1/4-scale jet flap model with a multilobe nozzle of array area ratio of 2.7. The model nozzle and flap tested were a two-dimensional section of a distributed blowing system similar to previously investigated augmentor wing systems without the upper shroud and intake door. Noise data were measured with the nozzle alone and also during attached flow conditions with the flap at two turning angles representing takeoff and approach conditions. The noise data are scaled to a 200,000-lb TOGW four-engine airplane and are presented in terms of perceived noise level and one-third octave band sound pressure level. Comparisons are made with the noise levels produced by an augmentor wing airplane fitted with a three-element acoustically lined augmentor flap. The static performance is presented in terms of thrust recovery and effective turning angle.
Harkonen, D. L.↗
Annoyance due to simulated blade-slap noise
The effects of several characteristics of blade slap noise on annoyance response were studied. These characteristics or parameters were the sound pressure level of the continuous noise used to simulate helicopter broadband noise, the ratio of impulse peak to broadband noise or crest factor, the number of pressure excursions comprising an impulse event, the rise and fall time of the individual impulses, and the repetition frequency of the impulses. Analyses were conducted to determine the correlation between subjective response and various physical measures for the range of parameters studied. A small but significant improvement in the predictive ability of PNL was provided by an A-weighted crest factor correlation. No significant improvement in predictive ability was provided by a rate correction.
Powell, C. A.↗
The problem of noise reduction with reference to light airplanes
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AIRPLANES, LIGHT↗
Propeller-loudness charts for light airplanes
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CHARTS - PROPELLERS↗
Dynamometer-stand investigation of a group of mufflers
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ENGINES - PRESSURE, BACK↗
The attenuation characteristics of four specially designed mufflers tested on a practical engine setup
Attenuation characteristics of four different resonator mufflers were determined in both cold tests and engine field tests and compared with the theoretical calculations. These mufflers were specifically designed for a helicopter. Engine-exhaust sound pressures, temperatures, and noise levels from the helicopter were measured. The experimental muffler cold tests indicated close a agreement with theory, whereas the engine tests indicated some discrepancies. Test results show the usefulness of the theoretical equation used for predicting muffler attenuation characteristics.
HELICOPTERS↗
Sound-Level Measurements of a Light Airplane Modified to Reduce Noise Reaching the Ground
An Army liaison-type airplane, representative of personal airplanes in the 150 to 200 horsepower class, has been modified to reduce propeller and engine noise according to known principles of airplane-noise reduction. Noise-level measurements demonstrate that, with reference to an observer on the ground, a noisy airplane of this class can be made quiet -- perhaps more quiet than necessary. In order to avoid extreme and unnecessary modifications, acceptable noise levels must be determined.
Vogeley, A W↗
Solutions of the benchmark problems by the dispersion-relation-preserving scheme
The 7-point stencil Dispersion-Relation-Preserving scheme of Tam and Webb is used to solve all the six categories of the CAA benchmark problems. The purpose is to show that the scheme is capable of solving linear, as well as nonlinear aeroacoustics problems accurately. Nonlinearities, inevitably, lead to the generation of spurious short wave length numerical waves. Often, these spurious waves would overwhelm the entire numerical solution. In this work, the spurious waves are removed by the addition of artificial selective damping terms to the discretized equations. Category 3 problems are for testing radiation and outflow boundary conditions. In solving these problems, the radiation and outflow boundary conditions of Tam and Webb are used. These conditions are derived from the asymptotic solutions of the linearized Euler equations. Category 4 problems involved solid walls. Here, the wall boundary conditions for high-order schemes of Tam and Dong are employed. These conditions require the use of one ghost value per boundary point per physical boundary condition. In the second problem of this category, the governing equations, when written in cylindrical coordinates, are singular along the axis of the radial coordinate. The proper boundary conditions at the axis are derived by applying the limiting process of r approaches 0 to the governing equations. The Category 5 problem deals with the numerical noise issue. In the present approach, the time-independent mean flow solution is computed first. Once the residual drops to the machine noise level, the incident sound wave is turned on gradually. The solution is marched in time until a time-periodic state is reached. No exact solution is known for the Category 6 problem. Because of this, the problem is formulated in two totally different ways, first as a scattering problem then as a direct simulation problem. There is good agreement between the two numerical solutions. This offers confidence in the computed results. Both formulations are solved as initial value problems. As such, no Kutta condition is required at the trailing edge of the airfoil.
Tam, Christopher K. W.↗
Numerical Simulation of the Generation of Axisymmetric Mode Jet Screech Tones
An imperfectly expanded supersonic jet, invariably, radiates both broadband noise and discrete frequency sound called screech tones. Screech tones are known to be generated by a feedback loop driven by the large scale instability waves of the jet flow. Inside the jet plume is a quasi-periodic shock cell structure. The interaction of the instability waves and the shock cell structure, as the former propagates through the latter, is responsible for the generation of the tones. Presently, there are formulas that can predict the tone frequency fairly accurately. However, there is no known way to predict the screech tone intensity. In this work, the screech phenomenon of an axisymmetric jet at low supersonic Mach number is reproduced by numerical simulation. The computed mean velocity profiles and the shock cell pressure distribution of the jet are found to be in good agreement with experimental measurements. The same is true with the simulated screech frequency. Calculated screech tone intensity and directivity at selected jet Mach number are reported in this paper. The present results demonstrate that numerical simulation using computational aeroacoustics methods offers not only a reliable way to determine the screech tone intensity and directivity but also an opportunity to study the physics and detailed mechanisms of the phenomenon by an entirely new approach.
Shen, Hao↗
Determination of Jet Noise Radiation Source Locations using a Dual Sideline Cross-Correlation/Spectrum Technique
The goal of our efforts is to extrapolate nearfield jet noise measurements to the geometric far field where the jet noise sources appear to radiate from a single point. To accomplish this, information about the location of noise sources in the jet plume, the radiation patterns of the noise sources and the sound pressure level distribution of the radiated field must be obtained. Since source locations and radiation patterns can not be found with simple single microphone measurements, a more complicated method must be used.
Allen, C. S.↗
A Computational and Experimental Study of Slit Resonators
Computational and experimental studies are carried out to offer validation of the results obtained from direct numerical simulation (DNS) of the flow and acoustic fields of slit resonators. The test cases include slits with 90-degree corners and slits with 45-degree bevel angle housed inside an acoustic impedance tube. Three slit widths are used. Six frequencies from 0.5 to 3.0 kHz are chosen. Good agreement is found between computed and measured reflection factors. In addition, incident sound waves having white noise spectrum and a prescribed pseudo-random noise spectrum are used in subsequent series of tests. The computed broadband results are again found to agree well with experimental data. It is believed the present results provide strong support that DNS can eventually be a useful and accurate prediction tool for liner aeroacoustics. The usage of DNS as a design tool is discussed and illustrated by a simple example.
Tam, C. K. W.↗
Acoustic-Modal Testing of the Ares I Launch Abort System Attitude Control Motor Valve
The Attitude Control Motor (ACM) is being developed for use in the Launch Abort System (LAS) of NASA's Ares I launch vehicle. The ACM consists of a small solid rocket motor and eight actuated pintle valves that directionally allocate.thrust_- 1t.has-been- predicted-that significant unsteady. pressure.fluctuations.will.exist. inside the-valves during operation. The dominant frequencies of these oscillations correspond to the lowest several acoustic natural frequencies of the individual valves. An acoustic finite element model of the fluid volume inside the valve has been critical to the prediction of these frequencies and their associated mode shapes. This work describes an effort to experimentally validate the acoustic finite model of the valve with an acoustic modal test. The modal test involved instrumenting a flight-like valve with six microphones and then exciting the enclosed air with a loudspeaker. The loudspeaker was configured to deliver broadband noise at relatively high sound pressure levels. The aquired microphone signals were post-processed and compared to results generated from the acoustic finite element model. Initial comparisons between the test data and the model results revealed that additional model refinement was necessary. Specifically, the model was updated to implement a complex impedance boundary condition at the entrance to the valve supply tube. This boundary condition models the frequency-dependent impedance that an acoustic wave will encounter as it reaches the end of the supply tube. Upon invoking this boundary condition, significantly improved agreement between the test data and the model was realized.
Davis, R. Benjamin↗
Analog Binaural Circuits for Detecting and Locating Leaks
Very-large-scale integrated (VLSI) analog binaural signal-processing circuits have been proposed for use in detecting and locating leaks that emit noise in the ultrasonic frequency range. These circuits would be designed to function even in the presence of intense lower-frequency background noise that could include sounds associated with flow and pumping. Each of the proposed circuits would include the approximate electronic equivalent of a right and a left cochlea plus correlator circuits. A pair of transducers (microphones or accelerometers), corresponding to right and left ears, would provide the inputs to their respective cochleas from different locations (e.g., from different positions along a pipe). The correlation circuits plus some additional external circuits would determine the difference between the times of arrival of a common leak sound at the two transducers. Then the distance along the pipe from either transducer to the leak could be estimated from the time difference and the speed of sound along the pipe. If three or more pairs of transducers and cochlear/correlator circuits were available and could suitably be positioned, it should be possible to locate a leak in three dimensions by use of sound propagating through air.
Hartley, Frank T.↗
Aeroacoustic Evaluation of Flap and Landing Gear Noise Reduction Concepts
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 open-wall (jet) mode. Most of the information was acquired with the model in a landing configuration with the flap deflected 39 deg 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 56 deg and 116 deg, with 90 deg 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.
Khorrami, Mehdi R.↗
An Initial Assessment of Variable Depth Liner Optimization for Ducted Proprotor Applications
The rise of the Urban Air Mobility market has spurred the design of a new generation of novel aircraft. To aid industry and researchers interested in these types of aircraft, the Revolutionary Vertical Lift Technology project at NASA has developed a fleet of reference vehicles for system studies. A new six-passenger reference vehicle has recently been added to the research portfolio that has ducted proprotors for propulsors. The ducts present an opportunity to apply acoustic treatment to the interior of the duct that could target both tonal and broadband noise, representative of the sound produced by this type of propulsor. In this paper, design methodologies are presented to design a variable depth liner for this application. An optimizer is used to design multiple liners with variable chamber depths. Experimental results from normal impedance testing are compared to numerical predictions using the optimizer model and a finite element model. Results show that an optimizer can be used to design an acoustic liner with favorable performance for a broad range of frequencies, which could be appropriate for a ducted proprotor application.
Matthew B Galles↗