SEARCH · Engineering Papers
Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.
Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.
Measuring technique discriminates between jet noise and core noise of jet engine. Results of experimentation confirmed that core noise and jet noise can be separated by examining cross-correlation of far-field microphone signals and that crossover point between core noise and jet noise moves toward higher velocities at higher angles with respect to jet axis.
A method of identification and measurement of core noise and jet noise separately has been developed based on cross-correlation of signals from microphones located at widely separated angles in the far field of a jet. The different coherent properties of core noise and jet noise are used in this method to achieve this separation. Experimental data obtained in a small scale facility is analyzed to demonstrate that this method can be used successfully to separate the mean square pressures of core noise and jet noise.
A review of 34 studies indicates that several factors or variables might be important in providing a psychoacoustic foundation for measurements of the noise from helicopters. These factors are phase relations, tail rotor noise, repetition rate, crest level, and generic differences between conventional aircraft and helicopters. Particular attention was given to the impulsive noise known as blade slap. Analysis of the evidence for and against each factor reveals that, for the present state of scientific knowledge, none of these factors should be regarded as the basis for a significant noise measurement correction due to impulsive blade slap. The current method of measuring effective perceived noise level for conventional aircraft appears to be adequate for measuring helicopter noise as well.
The techniques of helicopter rotor noise prediction attempt to describe precisely the details of the noise field and remove the empiricisms and restrictions inherent in previous methods. These techniques require detailed inputs of the rotor geometry, operating conditions, and blade surface pressure distribution. The Farassat noise prediction techniques was studied, and high speed helicopter noise prediction using more detailed representations of the thickness and loading noise sources was investigated. These predictions were based on the measured blade surface pressures on an AH-1G rotor and compared to the measured sound field. Although refinements in the representation of the thickness and loading noise sources improve the calculation, there are still discrepancies between the measured and predicted sound field. Analysis of the blade surface pressure data indicates shocks on the blades, which are probably responsible for these discrepancies.
Even though there are surveys in which annoyance decreases as the number of events increases above about 150 a day, the available evidence is not considered strong enough to reject the conventional assumption that reactions are related to the logarithm of the number of events. The data do not make it possible to reject the conventional assumption that the effects of the number of events and the peak noise level are additive. It is found that even when equivalent questionnaire items and definitions of noise events could be used, differences between the surveys' estimates of the effect of the number of events remained large. Three explanations are suggested for inconsistent estimates. The first has to do with errors in specifying the values of noise parameters, the second with the effects of unmeasured acoustical and area characteristics that are correlated with noise level or number, and the third with large sampling errors deriving from community differences in response to noise. It is concluded that significant advances in the knowledge about the effects of the number of noise events can be made only if surveys include large numbers of study areas.
The goal of this research is the understanding of helicopter internal noise mechanisms and the development, design, and testing of noise control concepts which will produce significant reductions in the acoustic environment to which passengers are exposed. The Phase 3 effort involved the identification and evaluation of current and advanced treatment concepts, including isolation of structure-borne paths. In addition, a plan was devised for the full-scale evaluation of an isolation concept. Specific objectives were as follows: (1) identification and characterization of various noise control concepts; (2) implementation of noise control concepts within the S-76 SEA (statistical energy analysis) model; (3) definition and evaluation of a preliminary acoustic isolation design to reduce structure-borne transmission of acoustic frequency main gearbox gear clash vibrations into the airframe; (4) formulation of a plan for the full-scale validation of the isolation concept; and (5) prediction of the cabin noise environment with various noise control concepts installed.
An analysis of tone noise propagation through a boundary layer and fuselage scattering effects was derived. This analysis is a three dimensional and the complete wave field is solved by matching analytical expressions for the incident and scattered waves in the outer flow to a numerical solution in the boundary layer flow. The outer wave field is constructed analytically from an incident wave appropriate to the source and a scattered wave in the standard Hankel function form. For the incident wave, an existing function - domain propeller noise radiation theory is used. In the boundary layer region, the wave equation is solved by numerical methods. The theoretical analysis is embodied in a computer program which allows the calculation of correction factors for the fuselage scattering and boundary layer refraction effects. The effects are dependent on boundary layer profile, flight speed, and frequency. Corrections can be derived for any point on the fuselage, including those on the opposite side from the source. The theory was verified using limited cases and by comparing calculations with available measurements from JetStar tests of model prop-fans. For the JetStar model scale, the boundary layer refraction effects produce moderate fuselage pressure reinforcements aft of and near the plane of rotation and significant attenuation forward of the plane of rotation at high flight speeds. At lower flight speeds, the calculated boundary layer effects result in moderate amplification over the fuselage area of interest. Apparent amplification forward of the plane of rotation is a result of effective changes in the source directivity due to boundary layer refraction effects. Full scale effects are calculated to be moderate, providing fuselage pressure amplification of about 5 dB at the peak noise location. Evaluation using available noise measurements was made under high-speed, high-altitude flight conditions. Comparisons of calculations made of free field noise, using a current frequency-domain propeller noise prediction method, and fuselage effects using this new procedure show good agreement with fuselage measurements over a wide range of flight speeds and frequencies. Correction factors for the JetStar measurements made on the fuselage are provided in an Appendix.
The topics covered include the following: community noise sources and noise control issues; noise components for turbine bypass turbojet engine (TBE) turbojet; engine cycle selection and noise; nozzle development schedule; NACA nozzle design; NACA nozzle test results; nearly fully mixed (NFM) nozzle design; noise versus aspiration rate; peak noise test results; nozzle test in the Low Speed Aeroacoustic Facility (LSAF); and Schlieren pictures of NACA nozzle.
An extension of a prior study has been completed to examine the potential reduction of aircraft flyover noise by the method of active noise control (ANC). It is assumed that the ANC system will be designed such that it cancels discrete tones radiating from the engine fan inlet or fan exhaust duct, at least to the extent that they no longer protrude above the surrounding broadband noise levels. Thus, without considering the engineering details of the ANC system design, tone levels am arbitrarily removed from the engine component noise spectrum and the flyover noise EPNL levels are compared with and without the presence of tones. The study was conducted for a range of engine cycles, corresponding to fan pressure ratios of 1.3, 1.45, 1.6, and 1.75. This report is an extension of an effort reported previously. The major conclusions drawn from the prior study, which was restricted to fan pressure ratios of 1.45 and 1.75, are that, for a fan pressure ratio of 1.75, ANC of tones gives about the same suppression as acoustic treatment without ANC. For a fan pressure ratio of 1.45, ANC appears to offer less effectiveness from passive treatment. In the present study, the other two fan pressure ratios are included in a more detailed examination of the benefits of the ANC suppression levels. The key results of this extended study are the following observations: (1) The maximum overall benefit obtained from suppression of BPF alone was 2.5 EPNdB at high fan speeds. The suppression benefit increases with increase in fan pressure ratio (FPR), (2) The maximum overall benefit obtained from suppression of the first three harmonics was 3 EPNdB at high speeds. Suppression benefit increases with increase in FPR, (3) At low FPR, only about 1.0 EPNdB maximum reduction was obtained. Suppression is primarily from reduction of BPF at high FPR values and from the combination of tones at low FPR, (4) The benefit from ANC is about the same as the benefit from passive treatment at fan pressure ratios of 1.75 and 1.60. At the two lower fan pressure ratios, the effectivness of treatment is much greater than that of ANC, and (5) No significant difference in ANC suppression behavior was found from the QCSEE engine database analysis compared to that of the E3 engine database, for the FPR = 1.3 engine cycle. The effects of ANC on EPNL noise reduction are difficult to generalize. It was found that the reduction obtained in any particular case depended upon the frequency of the tones and their shift with rpm, the amount of ANC suppression received by each tone (which depended on its protrusion from the background), and the NOY-value of the tone relative to the NOY-value of other tones and the peak broadband levels, because PNL is determined from the sum of the NOY-values.
An experimental proof-of-concept test was conducted to demonstrate reduction of rotor-stator interaction noise through the use of rotor-trailing edge blowing. The velocity deficit from the viscous wake of the rotor blades was reduced by injecting air into the wake from a continuous trailing edge slot. Hollow blades with interior guide vanes create flow channels through which externally supplied air flows from the blade root to the trailing edge. A previous paper documented the substantial tonal reductions of this Trailing Edge Rotor Blowing (TERB) fan. This report documents the broadband characteristics of TERB. The Active Noise Control Fan (ANCF), located at the NASA Glenn Research Center, was used as the proof-of-concept test bed. Two-component hotwire data behind the rotor, unsteady surface pressures on the stator vane, and farfield directivity acoustic data were acquired at blowing rates of 1.1, 1.5, and 1.8 percent of the total fan mass flow. The results indicate a substantial reduction in the rotor wake turbulent velocity and in the stator vane unsteady surface pressures. Based on the physics of the noise generation, these indirect measurements indicate the prospect of broadband noise reduction. However, since the broadband noise generated by the ANCF is rotor-dominated, any change in the rotor-stator interaction broadband noise levels is barely distinguishable in the farfield measurements.
Cross-correlation and coherence functions are used to look for periodic acoustic components in turbofan engine combustor time histories, to investigate direct and indirect combustion noise source separation based on signal propagation time delays, and to provide information on combustor acoustics. Using the cross-correlation function, time delays were identified in all cases, clearly indicating the combustor is the source of the noise. In addition, unfiltered and low-pass filtered at 400 Hz signals had a cross-correlation time delay near 90 ms, while the low-pass filtered at less than 400 Hz signals had a cross-correlation time delay longer than 90 ms. Low-pass filtering at frequencies less than 400 Hz partially removes the direct combustion noise signals. The remainder includes the indirect combustion noise signal, which travels more slowly because of the dependence on the entropy convection velocity in the combustor. Source separation of direct and indirect combustion noise is demonstrated by proper use of low-pass filters with the cross-correlation function for a range of operating conditions. The results may lead to a better idea about the acoustics in the combustor and may help develop and validate improved reduced-order physics-based methods for predicting direct and indirect combustion noise.
Aircraft engines have evolved into a highly complex system to meet ever-increasing demands. The evolution of engine technologies has primarily been driven by fuel efficiency, reliability, as well as engine noise concerns. One of the sources of engine noise is pressure fluctuations that are induced on the stator vanes. These local pressure fluctuations, once produced, propagate and coalesce with the pressure waves originating elsewhere on the stator to form a spinning pressure pattern. Depending on the duct geometry, air flow, and frequency of fluctuations, these spinning pressure patterns are self-sustaining and result in noise which eventually radiate to the far-field from engine. To investigate the nature of vane pressure fluctuations and the resulting engine noise, unsteady pressure signatures from an array of embedded acoustic sensors are recorded as a part of vane noise source diagnostics. Output time signatures from these sensors are routed to a control and data processing station adding complexity to the system and cable loss to the measured signal. "Smart" wireless sensors have data processing capability at the sensor locations which further increases the potential of wireless sensors. Smart sensors can process measured data locally and transmit only the important information through wireless communication. The aim of this wireless noise telemetry task was to demonstrate a single acoustic sensor wireless link for unsteady pressure measurement, and thus, establish the feasibility of distributed smart sensors scheme for aircraft engine vane surface unsteady pressure data transmission and characterization.
X-ray micro-tomography systems often suffer from high levels of noise. In particular, severe ring artifacts are common in reconstructed images, caused by defects in the detector, calibration errors, and fluctuations producing streak noise in the raw sinogram data. Furthermore, the projections commonly contain high levels of Poissonian noise arising from the photon-counting detector. This work presents a 3-D multiscale framework for streak attenuation through a purposely designed collaborative filtering of correlated noise in volumetric data. A distinct multiscale denoising step for attenuation of the Poissonian noise is further proposed. By utilizing the volumetric structure of the projection data, the proposed fully automatic procedure offers improved feature preservation compared with 2-D denoising and avoids artifacts which arise from individual filtering of sinograms.
The Advanced Noise Control Fan (neé Active Noise Control Fan - ANCF) was utilized in the design, test, and evaluation for technical risk mitigation of most of the innovative fan noise reduction technologies developed by NASA over the past 20 years. The ANCF is a low-speed, ducted fan, testbed for measuring and understanding fan-generated aeroacoustics, duct propagation, and radiation to the farfield. It is considered a low Technology Readiness Level testbed. The international aeroacoustics research community employed the ANCF to facilitate advancement of multiple noise reduction and measurement technologies, and for code validation. From 1994 to 2016, it was located in the NASA Glenn Research Center’s Aero-Acoustic Propulsion Laboratory. In 2016 the ANCF was transferred to the University of Notre Dame where it is expected to continue to positively impact ducted fan aeroacoustic research and provide STEM support. This paper summarizes the capabilities and contributions of the ANCF to the field by documenting its history. Limited data is presented, focusing on a description of the configurations, goals, and objectives of representative ANCF tests. This provides an overview of the progress of aeroacosustic research as implemented on the ANCF, as well as a background for its continued usage.
The Advanced Noise Control Fan (Active Noise Control Fan - ANCF) was utilized in the design, test, and evaluation for technical risk mitigation of most of the innovative fan noise reduction technologies developed by NASA over the past 20 years. The ANCF is a low-speed ducted fan test bed for measuring and understanding fan-generated aeroacoustics, duct propagation, and radiation to the farfield. It is considered a low Technology Readiness Level testbed. The international aeroacoustics research community employed the ANCF to facilitate advancement of multiple noise reduction and measurement technologies, and for code validation. From 1994 to 2016, it was located in the NASA Glenn Research Center's Applied Aero Propulsion Laboratory. In 2016 the ANCF was transferred to the University of Notre Dame where it is expected to continue to positively impact ducted fan aeroacoustic research. This paper summarizes the capabilities and contributions of the ANCF to the field by documenting its history.
A study has been conducted of the future noise environment of Patric Henry Airport and its neighboring communities projected for the year 1990. An assessment was made of the impact of advanced noise reduction technologies which are currently being considered. These advanced technologies include a two-segment landing approach procedure and aircraft hardware modifications or retrofits which would add sound absorbent material in the nacelles of the engines or which would replace the present two- and three-stage fans with a single-stage fan of larger diameter. Noise Exposure Forecast (NEF) contours were computed for the baseline (nonretrofitted) aircraft for the projected traffic volume and fleet mix for the year 1990. These NEF contours are presented along with contours for a variety of retrofit options. Comparisons of the baseline with the noise reduction options are given in terms of total land area exposed to 30 and 40 NEF levels. Results are also presented of the effects on noise exposure area of the total number of daily operations.
Measurements were conducted in an acoustic wind tunnel to determine vortex noise of nonrotating circular cylinders and NACA 0012 airfoils. Both constant-width and spanwise tapered models were tested at a low turbulence level. The constant-diameter cylinder and constant-chord airfoil also were tested in the turbulent wake generated by an upstream cylinder or airfoil. Vortex noise radiation from nonrotating circular cylinders at Reynolds numbers matching those of the rotating-blade tests were found to be strongly dependent on surface conditions and Reynolds number. Vortex noise of rotating circular cylinder blades, operating with and without the shed wake blown downstream, could be predicted using data for nonrotating circular cylinders as functions of Reynolds number. Vortex noise of nonrotating airfoils was found to be trailing-edge noise at a time frequence equal to that predicted for maximum-amplitude Tollmein-Schlichting instability waves at the trailing edge.
The maxima of amplitudes of thickness and loading noise harmonics are established when the radial distribution of blade chord, thickness ratio, and lift coefficient is specified. It is first shown that only airfoils with thickness distribution and chordwise loading distributions which are symmetric with respect to midchord need be considered for finding the absolute maxima of thickness and loading noise. The resulting chordwise thickness and load distributions for these maximum noise conditions require infinite slope at some points along the chord but otherwise are uniform. It is shown that sweeping the blades reduces the thickness and loading noise, but there is no optimum sweep which generates the lowest noise.