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All sourcesNASA NTRSDOE OSTI
NASA NTRS1974

Cross-correlation methods for studying near- and far-field noise characteristics of several flow-surface interaction problems

Systematic methods based on cross-correlation techniques are presented for experimental studies of near- and far-field noise characteristics of airflow-surface interaction problems. By cross-correlating any two of three-kinds of microphone measurements: one flush-mounted on the surface, one located in the near field, and one placed in the far field, far-field noise intensities and near-field acoustic energy fluxes can be determined. In both near and far fields, the noise characteristics due to the flow turbulence and surface fluctuating pressure contributions can be evaluated separately. Specific applications to the noise field generated by impinging jets, by surface blowing jets, and by turbulent flow over trailing edges are discussed.

Pan, Y. S.↗
NASA NTRS1975

Cross-correlation methods for studying near- and farfield noise characteristics of flow-surface interactions

Systematic methods based on cross-correlation techniques are presented for experimental studies of near- and farfield noise characteristics in airflow-surface interaction problems. Analyses show that, in near- and farfields, the noise characteristics due to the surface contribution of fluctuating pressures and velocities and due to the volume contribution of the turbulence in the flow can be determined separately. Both farfield noise intensities and nearfield acoustic energy fluxes can be expressed in terms of appropriate cross correlations. These correlations can be obtained by making microphone measurements in the farfield, in the nearfield, and on the surface. Examples of the applications to the noise field generated by impinging jets, by surface blowing jets, and by turbulent flow over trailing edges are presented. Advantages of the present method over conventional methods are also discussed.

Pan, Y. S.↗
NASA NTRS2014

Aerodynamic Measurements of a Gulfstream Aircraft Model With and Without Noise Reduction Concepts

Steady and unsteady aerodynamic measurements of a high-fidelity, semi-span 18% scale Gulfstream aircraft model are presented. The aerodynamic data were collected concurrently with acoustic measurements as part of a larger aeroacoustic study targeting airframe noise associated with main landing gear/flap components, gear-flap interaction noise, and the viability of related noise mitigation technologies. The aeroacoustic tests were conducted in the NASA Langley Research Center 14- by 22-Foot Subsonic Wind Tunnel with the facility in the acoustically treated open-wall (jet) mode. Most of the measurements were obtained with the model in landing configuration with the flap deflected at 39º and the main landing gear on and off. Data were acquired at Mach numbers of 0.16, 0.20, and 0.24. Global forces (lift and drag) and extensive steady and unsteady surface pressure measurements were obtained. Comparison of the present results with those acquired during a previous test shows a significant reduction in the lift experienced by the model. The underlying cause was traced to the likely presence of a much thicker boundary layer on the tunnel floor, which was acoustically treated for the present test. The steady and unsteady pressure fields on the flap, particularly in the regions of predominant noise sources such as the inboard and outboard tips, remained unaffected. It is shown that the changes in lift and drag coefficients for model configurations fitted with gear/flap noise abatement technologies fall within the repeatability of the baseline configuration. Therefore, the noise abatement technologies evaluated in this experiment have no detrimental impact on the aerodynamic performance of the aircraft model.

Neuhart, Dan H.↗
NASA NTRS1990

Acoustic and aerodynamic study of a pusher-propeller aircraft model

An aerodynamic and acoustic study was made of a pusher-propeller aircraft model in the NASA-Ames 7 x 10 ft Wind Tunnel. The test section was changed to operate as an open jet. The 591 mm diameter unswept propeller was operated alone and in the wake of three empennages: an I tail, Y tail, and a V tail. The radiated noise and detailed wake properties were measured. Results indicate that the unsteady blade loading caused by the blade interactions with the wake mean velocity distribution had a strong effect on the harmonics of blade passage noise. The blade passage harmonics above the first were substantially increased in all horizontal directions by the empennage/propeller interaction. Directivity in the plane of the propeller was maximum perpendicular to the blade surface. Increasing the tail loading caused the propeller harmonics to increase 3 to 5 dB for an empennage/propeller spacing of 0.38 mean empennage chords. The interaction noise became weak as empennage propeller spacing was increased beyond 1.0 mean empennage chord lengths. Unlike the mean wake deficit, the wake turbulence had only a small effect on the propeller noise, that effect being a small increase in the broadband noise.

Soderman, Paul T.↗
NASA NTRS2014

Experiments on Exhaust Noise of Tightly Integrated Propulsion Systems

A wide-ranging series of tests have been completed that seek to map the effects of installation, including jet by jet interaction effects, on exhaust noise from various nozzles in forward flight. The primary data was far-field acoustic spectral directivity. The goals of the test series were (i) to generate enough data for empirical models of the different effects, and (ii) to provide data for advanced computational noise predictions methods applied to simplified yet realistic configurations. Data is presented that demonstrate several checks on data quality and that provide an overview of trends observed to date. Among the findings presented here: (i) Data was repeatable between jet rigs for single nozzles with and without surfaces to within +/- 0.5 dB. (ii) The presence of a second jet caused a strong reduction of the summed noise in the plane of the two plumes and an increase over the expected source doubling in most other azimuthal planes. (iii) The impact of the second jet was reduced when the jets were unheated. (iv) The impact of adding a second isolated rectangular jet was relatively independent of the nozzle aspect ratio up to aspect ratio 8:1. (v) Forward flight had similar impact on a high aspect ratio (8:1) jet as on an axisymmetric jet, except at the peak noise angle where the impact was less. (vi) The effect of adding a second round jet to a tightly integrated nozzle where the nozzle lip was less than a diameter from the surface was very dependent upon the length of the surface downstream of the nozzle. (vii) When the nozzles were rectangular and tightly integrated with the airframe surface the impact of a second jet was very dependent upon how close together the two jets were. This paper serves as an overview of the test; other papers presented in the same conference will give more detailed analysis of the results.

Noise↗
NASA NTRS1999

Summary of HEAT 1 Aeroacoustics Installation Effects

A critical part of the NASA High-Speed Research (HSR) program is the demonstration of satisfactory suppression of the jet noise present at low airspeeds. One scheme for reducing jet exhaust noise generated by a future High-Speed Civil Transport (HSCT) is the use of a mixer/ ejector system which would entrain large quantities of ambient air into the exhaust flow from the powerplant in order to cool and slow the jet exhaust before it leaves the tailpipe. Of the variety of factors which can affect the noise suppression characteristics of the mixer/ejector system, the influence of the wing flow field and high-lift devices is not well understood. The effectiveness of the noise suppression device must be evaluated in the presence of the wing/high-lift system before definitive assessments can be made concerning HSCT noise. Of nearly equal importance is the evaluation of the performance of the high-lift system(s) in the presence of realistic propulsion units which feature high ambient flow entrainment rates and jet thrust coefficients. These noise suppressors must provide the required acoustic attenuation while not overly degrading the thrust efficiency of the propulsion system or the lift enhancement of the high-lift devices on the wing. The overall objective of the NASA High-lift Engine Aeroacoustics Technology program is to demonstrate satisfactory interaction between the jet noise suppressor and the high-lift system at airspeeds and angles of attack consistent with takeoff, climb, approach, and landing. In support of this program, an isolated aeroacoustic test of a 13.5%-scale, candidate mixer/ejector nozzle was performed in the Ames' Research Center 40- by 80-Foot Wind Tunnel. The purpose of the test was to measure the baseline aeroacoustic performance characteristics of this nozzle in isolation from the aerodynamic flowfield induced by an HSCT airframe. The test documented the acoustic signature of the nozzles with treated and hardwall ejector surfaces and with changes in the ratio of ejector-duct-to-jet-area over a wide range range of nozzle pressure ratios and freestream Mach numbers. The test also measured the thrust performance, ambient-flow aspiration ratio, and internal and external static pressures on the nozzles. The isolated aeroacoustic performance data has been compared with results obtained with this nozzle installed on a 13.5% Boeing Reference H HSCT configuration, semi-span model. The semi-span, aeroacoustics integration test documented the first-order effects of the airframe flowfield on the acoustic performance of the nozzle and the effect of the nozzle secondary inlet flows on the aerodynamic performance of the wing high-lift systems. This investigation is critical to understanding the mutual installation effects of mixer/ejector nozzles and wing high-lift systems.

Smith, Brian E.↗
NASA NTRS1988

Supersonic jet plume interaction with a flat plate

Supersonic jet plume interaction with a flat plate was studied using a model scaled test apparatus designed to simulate plume/aircraft structure interaction for the cruise configuration. The generic configuration consisted of a rectangular supersonic nozzle of aspect ratio 7, and a large flat plate located beneath the nozzle at various nozzle plate distances; the plate was instrumented to measure surface dynamic pressure and mean wall temperature, with provisions for measurements of acceleration and strain on coupon size panels that could be inserted in the plate. Phase-averaged schlieren measurements revealed the presence of high-intensity acoustic emission from the supersonic plume above the plate, directed upstream; this radiation could be associated with the shock noise generation. Narrow band spectra of surface dynamic pressure show spectral peaks with amplitude levels reaching 1 psi, related to the screech tones. Temperature measurements indicated elevated surface temperatures in regions of high turbulence intensity.

Seiner, John M.↗
NASA NTRS2012

Aeroacoustic Study of a High-Fidelity Aircraft Model: Part 1- Steady Aerodynamic Measurements

In this paper, we present steady aerodynamic measurements for an 18% scale model of a Gulfstream air-craft. The high fidelity and highly-instrumented semi-span model was developed to perform detailed aeroacoustic studies of airframe noise associated with main landing gear/flap components and gear-flap interaction noise, as well as to evaluate novel noise reduction concepts. The aeroacoustic tests, being conducted in the NASA Langley Research Center 14- by 22-Foot Subsonic Tunnel, are split into two entries. The first entry, completed November 2010, was entirely devoted to the detailed mapping of the aerodynamic characteristics of the fabricated model. Flap deflections of 39?, 20?, and 0? with the main landing gear on and off were tested at Mach numbers of 0.16, 0.20, and 0.24. Additionally, for each flap deflection, the model was tested with the tunnel both in the closed-wall and open-wall (jet) modes. During this first entry, global forces (lift and drag) and extensive steady and unsteady surface pressure measurements were obtained. Preliminary analysis of the measured forces indicates that lift, drag, and stall characteristics compare favorably with Gulfstream?s high Reynolds number flight data. The favorable comparison between wind-tunnel and flight data allows the semi-span model to be used as a test bed for developing/evaluating airframe noise reduction concepts under a relevant environment. Moreover, initial comparison of the aerodynamic measurements obtained with the tunnel in the closed- and open-wall configurations shows similar aerodynamic behavior. This permits the acoustic and off-surface flow measurements, planned for the second entry, to be conducted with the tunnel in the open-jet mode.

Khorrami, Mehdi R.↗
NASA NTRS2009

ATK Launch Vehicle (ALV-X1) Liftoff Acoustic Environments: Prediction vs. Measurement

The ATK Launch Vehicle (ALV-X1) provided an opportunity to measure liftoff acoustic noise data. NASA Marshall Space Flight Center (MSFC) engineers were interested in the ALV-X1 launch because the First Stage motor and launch pad conditions, including a relativity short deflector ducting, provide a potential analogue to future Ares I launches. This paper presents the measured liftoff acoustics on the vehicle and tower. Those measured results are compared to predictions based upon the method described in NASA SP-8072 "Acoustic Loads Generated by the Propulsion System" and the Vehicle Acoustic Environment Prediction Program (VAEPP) which was developed by MSFC acoustics engineers. One-third octave band sound pressure levels will be presented. This data is useful for the ALV-X1 in validating the pre-launch environments and loads predictions. Additionally, the ALV-X1 liftoff data can be scaled to define liftoff environments for the NASA Constellation program Ares vehicles. Vehicle liftoff noise is caused by the supersonic jet flow interaction with surrounding atmosphere or more simply, jet noise. As the vehicle's First Stage motor is ignited, an acoustic noise field is generated by the exhaust. This noise field persists due to the supersonic jet noise and reflections from the launch pad and tower, then changes as the vehicle begins to liftoff from the launch pad. Depending on launch pad and adjacent tower configurations, the liftoff noise is generally very high near the nozzle exit and decreases rapidly away from the nozzle. The liftoff acoustic time range of interest is typically 0 to 20 seconds after ignition. The exhaust plume thermo-fluid mechanics generates sound at approx.10 Hz to 20 kHz. Liftoff acoustic noise is usually the most severe dynamic environment for a launch vehicle or payload in the mid to high frequency range (approx.50 to 2000 Hz). This noise environment can induce high-level vibrations along the external surfaces of the vehicle and surrounding launch facility structures. The acoustic pressure fluctuations will induce severe vibrations in relatively large lightweight structures. Consequently, there is the potential for failure of the structure or attached electrical components. Due to these potential failures, the liftoff acoustic noise is one of the noise source inputs used to determine the vibro-acoustic qualification environment for a launch vehicle and its components.

Houston, Janice↗
NASA NTRS2006

Flight Test Results for Uniquely Tailored Propulsion-Airframe Aeroacoustic Chevrons: Shockcell Noise

Azimuthally varying chevrons (AVC) which have been uniquely tailored to account for the asymmetric propulsion-airframe aeroacoustic interactions have recently shown significant reductions in jet-related community noise at low-speed take-off conditions in scale model tests of coaxial nozzles with high bypass ratio. There were indications that such AVCs may also provide shockcell noise reductions at high cruise speeds. This paper describes the flight test results when one such AVC concept, namely, the T-fan chevrons with enhanced mixing near the pylon, was tested at full-scale on a modern large twin-jet aircraft (777-300ER) with focus on shockcell noise at mid-cruise conditions. Shockcell noise is part of the interior cabin noise at cruise conditions and its reduction is useful from the viewpoint of passenger comfort. Noise reduction at the source, in the exhaust jet, especially, at low frequencies, is beneficial from the perspective of reduced fuselage sidewall acoustic lining. Results are shown in terms of unsteady pressure spectra both on the exterior surface of the fuselage at several axial stations and also microphone arrays placed inside the fuselage aft of the engine. The benefits of T-fan chevrons, with and without conventional chevrons on the core nozzle, are shown for several engine operating conditions at cruise involving supersonic fan stream and subsonic or supersonic core stream. The T-fan AVC alone provides up to 5 dB low-frequency noise reduction on the fuselage exterior skin and up to 2 dB reduction inside the cabin. Addition of core chevrons appears to increase the higher frequency noise. This flight test result with the previous model test observation that the T-fan AVCs have hardly any cruise thrust coefficient loss (< 0.05%) make them viable candidates for reducing interior cabin noise in high bypass ratio engines.

Mengle, Vinod G.↗
NASA NTRS2015

Investigating the Feedback Path in a Jet-Surface Resonant Interaction

A resonant interaction between an 8:1 aspect ratio rectangular jet and flat-plates, placed parallel to the jet, is addressed in this study. For certain relative locations of the plates, the resonance takes place with accompanying audible tones. Even when the tone is not audible the sound pressure level spectra is often marked by conspicuous peaks. The frequencies of the spectral peaks, as functions of the streamwise length of the plate and its relative location to the jet as well as the jet Mach number, are explored in an effort of understand the flow mechanism. It is demonstrated that the tones are not due to a simple feedback between the plates trailing edge and the nozzle exit; the leading edge also comes into play in determining the frequency. An acoustic feedback path, involving diffraction from the leading edge, appears to explain the frequencies of some of the spectral peaks.

Jet Noise↗
NASA NTRS1995

A numerical study of fundamental shock noise mechanisms

The results of this thesis demonstrate that direct numerical simulation can predict sound generation in unsteady aerodynamic flows containing shock waves. Shock waves can be significant sources of sound in high speed jet flows, on helicopter blades, and in supersonic combustion inlets. Direct computation of sound permits the prediction of noise levels in the preliminary design stage and can be used as a tool to focus experimental studies, thereby reducing cost and increasing the probability of a successfully quiet product in less time. This thesis reveals and investigates two mechanisms fundamental to sound generation by shocked flows: shock motion and shock deformation. Shock motion is modeled by the interaction of a sound wave with a shock. During the interaction, the shock wave begins to move and the sound pressure is amplified as the wave passes through the shock. The numerical approach presented in this thesis is validated by the comparison of results obtained in a quasi-one dimensional simulation with linear theory. Analysis of the perturbation energy demonstrated for the first time that acoustic energy is generated by the interaction. Shock deformation is investigated by the numerical simulation of a ring vortex interacting with a shock. This interaction models the passage of turbulent structures through the shock wave. The simulation demonstrates that both acoustic waves and contact surfaces are generated downstream during the interaction. Analysis demonstrates that the acoustic wave spreads cylindrically, that the sound intensity is highly directional, and that the sound pressure level increases significantly with increasing shock strength. The effect of shock strength on sound pressure level is consistent with experimental observations of shock noise, indicating that the interaction of a ring vortex with a shock wave correctly models a dominant mechanism of shock noise generation.

Meadows, Kristine R.↗
NASA NTRS2015

Investigating the Feedback Path in a Jet-Surface Resonant Interaction

A resonant interaction between an 8:1 aspect ratio rectangular jet and flat-plates, placed parallel to the jet, is studied experimentally. For certain locations of the plate relative to the jet, the resonance takes place with a loud accompanying tone. The sound pressure level spectra are often marked by multiple peaks. The frequencies of the spectral peaks are studied as a function of the streamwise length of the plate, its relative location to the jet as well as the jet Mach number. It is demonstrated that the tones are not due to a simple feedback between the plate's trailing edge and the nozzle's exit; the leading edge of the plate also comes into play in the frequency selection. With parametric variation, it is found that there is an order in the most energetic spectral peaks; their frequencies cluster in distinct bands. The 'fundamental', i.e., the lowest frequency band is explained by an acoustic feedback involving diffraction at the plate's leading edge.

Resonance↗
NASA NTRS2010

Hybrid Wing Body Aircraft System Noise Assessment with Propulsion Airframe Aeroacoustic Experiments

A system noise assessment of a hybrid wing body configuration was performed using NASA s best available aircraft models, engine model, and system noise assessment method. A propulsion airframe aeroacoustic effects experimental database for key noise sources and interaction effects was used to provide data directly in the noise assessment where prediction methods are inadequate. NASA engine and aircraft system models were created to define the hybrid wing body aircraft concept as a twin engine aircraft with a 7500 nautical mile mission. The engines were modeled as existing technology high bypass ratio turbofans. The baseline hybrid wing body aircraft was assessed at 22 dB cumulative below the FAA Stage 4 certification level. To determine the potential for noise reduction with relatively near term technologies, seven other configurations were assessed beginning with moving the engines two fan nozzle diameters upstream of the trailing edge and then adding technologies for reduction of the highest noise sources. Aft radiated noise was expected to be the most challenging to reduce and, therefore, the experimental database focused on jet nozzle and pylon configurations that could reduce jet noise through a combination of source reduction and shielding effectiveness. The best configuration for reduction of jet noise used state-of-the-art technology chevrons with a pylon above the engine in the crown position. This configuration resulted in jet source noise reduction, favorable azimuthal directivity, and noise source relocation upstream where it is more effectively shielded by the limited airframe surface, and additional fan noise attenuation from acoustic liner on the crown pylon internal surfaces. Vertical and elevon surfaces were also assessed to add shielding area. The elevon deflection above the trailing edge showed some small additional noise reduction whereas vertical surfaces resulted in a slight noise increase. With the effects of the configurations from the database included, the best available noise reduction was 40 dB cumulative. Projected effects from additional technologies were assessed for an advanced noise reduction configuration including landing gear fairings and advanced pylon and chevron nozzles. Incorporating the three additional technology improvements, an aircraft noise is projected of 42.4 dB cumulative below the Stage 4 level.

Thomas, Russell H.↗
NASA NTRS2011

High-Temperature Smart Structures for Engine Noise Reduction and Performance Enhancement

One of key NASA goals is to develop and integrate noise reduction technology to enable unrestricted air transportation service to all communities. One of the technical priorities of this activity has been to account for and reduce noise via propulsion/airframe interactions, identifying advanced concepts to be integrated with the airframe to mitigate these noise-producing mechanisms. An adaptive geometry chevron using embedded smart structures technology offers the possibility of maximizing engine performance while retaining and possibly enhancing the favorable noise characteristics of current designs. New high-temperature shape memory alloy (HTSMA) materials technology enables the devices to operate in both low-temperature (fan) and high-temperature (core) exhaust flows. Chevron-equipped engines have demonstrated reduced noise in testing and operational use. It is desirable to have the noise benefits of chevrons in takeoff/landing conditions, but have them deployed into a minimum drag position for cruise flight. The central feature of the innovation was building on rapidly maturing HTSMA technology to implement a next-generation aircraft noise mitigation system centered on adaptive chevron flow control surfaces. In general, SMA-actuated devices have the potential to enhance the demonstrated noise reduction effectiveness of chevron systems while eliminating the associated performance penalty. The use of structurally integrated smart devices will minimize the mechanical and subsystem complexity of this implementation. The central innovations of the effort entail the modification of prior chevron designs to include a small cut that relaxes structural stiffness without compromising the desired flow characteristics over the surface; the reorientation of SMA actuation devices to apply forces to deflect the chevron tip, exploiting this relaxed stiffness; and the use of high-temperature SMA (HTSMA) materials to enable operation in the demanding core chevron environment. The overall conclusion of these design studies was that the cut chevron concept is a critical enabling step in bringing the variable geometry core chevron within reach. The presence of the cut may be aerodynamically undesirable in some respects, but it is present only when the chevron is not immersed in the core jet exhaust. When deployed, the gap closes as the chevron tip enters the high-speed, high-temperature core stream. Aeroacoustic testing and flow visualization support the contention that this cut is inconsequential to chevron performance.

Quackenbush, Todd R.↗
NASA NTRS1997

Hierarchical Modeling and Robust Synthesis for the Preliminary Design of Large Scale Complex Systems

Large-scale complex systems are characterized by multiple interacting subsystems and the analysis of multiple disciplines. The design and development of such systems inevitably requires the resolution of multiple conflicting objectives. The size of complex systems, however, prohibits the development of comprehensive system models, and thus these systems must be partitioned into their constituent parts. Because simultaneous solution of individual subsystem models is often not manageable iteration is inevitable and often excessive. In this dissertation these issues are addressed through the development of a method for hierarchical robust preliminary design exploration to facilitate concurrent system and subsystem design exploration, for the concurrent generation of robust system and subsystem specifications for the preliminary design of multi-level, multi-objective, large-scale complex systems. This method is developed through the integration and expansion of current design techniques: Hierarchical partitioning and modeling techniques for partitioning large-scale complex systems into more tractable parts, and allowing integration of subproblems for system synthesis; Statistical experimentation and approximation techniques for increasing both the efficiency and the comprehensiveness of preliminary design exploration; and Noise modeling techniques for implementing robust preliminary design when approximate models are employed. Hierarchical partitioning and modeling techniques including intermediate responses, linking variables, and compatibility constraints are incorporated within a hierarchical compromise decision support problem formulation for synthesizing subproblem solutions for a partitioned system. Experimentation and approximation techniques are employed for concurrent investigations and modeling of partitioned subproblems. A modified composite experiment is introduced for fitting better predictive models across the ranges of the factors, and an approach for constructing partitioned response surfaces is developed to reduce the computational expense of experimentation for fitting models in a large number of factors. Noise modeling techniques are compared and recommendations are offered for the implementation of robust design when approximate models are sought. These techniques, approaches, and recommendations are incorporated within the method developed for hierarchical robust preliminary design exploration. This method as well as the associated approaches are illustrated through their application to the preliminary design of a commercial turbofan turbine propulsion system. The case study is developed in collaboration with Allison Engine Company, Rolls Royce Aerospace, and is based on the Allison AE3007 existing engine designed for midsize commercial, regional business jets. For this case study, the turbofan system-level problem is partitioned into engine cycle design and configuration design and a compressor modules integrated for more detailed subsystem-level design exploration, improving system evaluation. The fan and low pressure turbine subsystems are also modeled, but in less detail. Given the defined partitioning, these subproblems are investigated independently and concurrently, and response surface models are constructed to approximate the responses of each. These response models are then incorporated within a commercial turbofan hierarchical compromise decision support problem formulation. Five design scenarios are investigated, and robust solutions are identified. The method and solutions identified are verified by comparison with the AE3007 engine. The solutions obtained are similar to the AE3007 cycle and configuration, but are better with respect to many of the requirements.

Koch, Patrick N.↗
NASA NTRS1995

Aerodynamic sound of flow past an airfoil

The long term objective of this project is to develop a computational method for predicting the noise of turbulence-airfoil interactions, particularly at the trailing edge. We seek to obtain the energy-containing features of the turbulent boundary layers and the near-wake using Navier-Stokes Simulation (LES or DNS), and then to calculate the far-field acoustic characteristics by means of acoustic analogy theories, using the simulation data as acoustic source functions. Two distinct types of noise can be emitted from airfoil trailing edges. The first, a tonal or narrowband sound caused by vortex shedding, is normally associated with blunt trailing edges, high angles of attack, or laminar flow airfoils. The second source is of broadband nature arising from the aeroacoustic scattering of turbulent eddies by the trailing edge. Due to its importance to airframe noise, rotor and propeller noise, etc., trailing edge noise has been the subject of extensive theoretical (e.g. Crighton & Leppington 1971; Howe 1978) as well as experimental investigations (e.g. Brooks & Hodgson 1981; Blake & Gershfeld 1988). A number of challenges exist concerning acoustic analogy based noise computations. These include the elimination of spurious sound caused by vortices crossing permeable computational boundaries in the wake, the treatment of noncompact source regions, and the accurate description of wave reflection by the solid surface and scattering near the edge. In addition, accurate turbulence statistics in the flow field are required for the evaluation of acoustic source functions. Major efforts to date have been focused on the first two challenges. To this end, a paradigm problem of laminar vortex shedding, generated by a two dimensional, uniform stream past a NACA0012 airfoil, is used to address the relevant numerical issues. Under the low Mach number approximation, the near-field flow quantities are obtained by solving the incompressible Navier-Stokes equations numerically at chord Reynolds number of 104. The far-field noise is computed using Curle's extension to the Lighthill analogy (Curle 1955). An effective method for separating the physical noise source from spurious boundary contributions is developed. This allows an accurate evaluation of the Reynolds stress volume quadrupoles, in addition to the more readily computable surface dipoles due to the unsteady lift and drag. The effect of noncompact source distribution on the far-field sound is assessed using an efficient integration scheme for the Curle integral, with full account of retarded-time variations. The numerical results confirm in quantitative terms that the far-field sound is dominated by the surface pressure dipoles at low Mach number. The techniques developed are applicable to a wide range of flows, including jets and mixing layers, where the Reynolds stress quadrupoles play a prominent or even dominant role in the overall sound generation.

Wang, Meng↗