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This study, conducted at the University of Bristol's aeroacoustic facility, explores the aeroacoustic characteristics of a 10-inch APC propeller in pusher configuration within ground effect conditions. Tested at a rotational speed of 7000 r/min, the propeller's performance was evaluated at varying distances from the ground. Wall pressure fluctuations near the propeller were measured at eight radial points, complemented by far-field acoustic measurements. Consistent with previous research, distinctive thrust and torque behaviours were observed when operating both in and out of ground effect. Notably, a significant noise increase of approximately 5 dB was observed on the ground's reflected side, while a reduction of about 14 dB was detected on the shielded side. Higher-order statistical analysis such as skewness and kurtosis revealed substantial effects of tip vortex impingement in ground effect conditions. This study applies two-point statistics and the Corcos model to analyze propeller-induced wall pressure fluctuations. The coherence function effectively captures the intricate dynamics of pressure variations across different spatial locations and frequencies. This approach offers novel insights into the behaviour of boundary layers and flow-induced pressures in the context of propellers operating in the ground effect. Wavelet decomposition was utilized to differentiate the influences of the boundary layer and the acoustic field. This comprehensive analysis highlights the intricate dynamics of propeller operation in ground effect, contributing valuable insights to the field of aeroacoustic research. [ABSTRACT FROM AUTHOR]

Heavy-duty centrifugal fans account for a significant share of energy consumption in the process and manufacturing industries. As a result, these machines are under increasing pressure to operate at maximum efficiency to reduce costs, pollutants and noise: only combined optimization is considered competitive for future generations of fans. Preliminary studies have shown that applying structured porosity to aerofoil rear parts can lead to a reduction in self noise and trailing edge shedding noise in the mid-to-high frequency range. With this in mind, a porous surface cover is applied to a prototype centrifugal fan to evaluate the aeroacoustic potential in a complex rotating machinery. The optimal geometric characteristics of the perforation are derived from experiments with single aerofoils, while the perimeter of the covered area is varied in eight steps. The centrifugal fan specimen is rapid-prototyped and tested at different fan speeds along the complete characteristic curves, while both aerodynamic and aeroacoustic performances are simultaneously recorded. The results obtained show a significant reduction in overall noise level while aerodynamic performance is maintained. Spectral analysis shows that the noise reduction is due to a broadband effect, where the upper and lower cut-off frequencies are determined by the rotational speed and the location of the applied porosity along the blade chord. However, the maximum noise reduction is obtained as a clear function of the minimum distance between the perforation and the trailing edge of the blade, indicating that the underlying working mechanisms are a combination of broadband dissipation effects due to porosity and destructive interference. [ABSTRACT FROM AUTHOR]

To analyze the noise induced by moving rigid structures in low Mach number flows, acoustic governing equations based on the viscous/acoustic splitting method and the arbitrary Lagrangian–Eulerian method are rigorously derived. In order to resolve the numerical instability generated in a non-uniform mean flow, the modified viscous/acoustic method, based on the filtering method, is developed. The acoustic equations are transformed into the same form as the incompressible flow equations by introducing the acoustic co-velocity and solved based on a collocated grid finite volume method. An approach for solving acoustic equation based on the PIMPLE algorithm is presented and computed in open-source computational fluid dynamics software OpenFOAM, which brings down communication costs and speeds up computing efficiency. Furthermore, the source term decomposition is extended to study the noise generated by each source term in a motion grid. Several examples including stationary and moving meshes have been designed to prove the accuracy of this approach. Finally, the aerodynamic and acoustic properties for the flow past a transversely oscillating cylinder at Re = 200, Ma = 0.2 in lock-in and non-lock-in regions is present. [ABSTRACT FROM AUTHOR]

This review presents an overview of advanced air mobility broadband noise (BBN) prediction and control techniques, highlighting significant advancements in various prediction models. Methods such as the semi-empirical Brooks–Pope–Marcolini (BPM) model, analytical Amiet model, and time-domain models based on the FW-H equation have been extensively studied. Machine learning (ML) shows promise in BBN prediction but requires extensive data training and application to noise source mechanisms. Passive control methods, such as leading and trailing edge serrations and blade tip designs, have been partially successful but often compromise the aerodynamic performance. Active control methods, like suction and blowing control, trim adjustments, and dielectric barrier discharge (DBD) plasma actuators, show great potential, with the latter two being particularly effective for reducing BBN in thin propeller structures. Overall, while progress has been made in understanding and predicting BBN, further research is needed to refine these methods and develop comprehensive noise control strategies. These advancements hold significant promise for effective and efficient noise mitigation in future AAM vehicles. [ABSTRACT FROM AUTHOR]

The accurate prediction of the trailing-edge noise and the determination of its sources are essential to reduce fans and propellers noise. This noise component is due to the scattering of the turbulent boundary layer into acoustic waves by the trailing edge. In this paper, the noise emanating from a CD (Controlled-Diffusion) airfoil is simulated and computed via the hybrid methods of aeroacoustics. In these methods, the aerodynamic and acoustic fields are computed separately. The flow data are obtained using the in-house LES solver SFELES. ACTRAN acoustic solver has been used to solve the acoustics and to provide the near and far fields propagation via Lighthill's analogy. Curle's analogy is applied as well in its integral compact formulation which takes the presence of walls into account. Curle's formulation is applied proposing an approach where the volume and surface integrals have been implemented in SFELES to be calculated simultaneously with the flow in order to avoid the storage of noise sources which requires a huge space. In Lighthill's analogy, sources and near field maps show that the turbulent boundary layer and wake are the more efficient sources and the center of radiation is the trailing edge. The comparison of the numerical results with the experimental measurements, performed by Moreau and Roger and Moreau et al., shows an overall excellent agreement confirming the capability of SFELES (LES sources) combined with ACTRAN (Lighthill's analogy) to predict correctly the noise generated by turbulent flows around airfoils. The acoustic spectrum presents an overprediction up to 5 dB at the frequencies 300 Hz and 550 Hz and an underprediction about 5 dB at the frequencies 1100 Hz and 1750 Hz. The sound pressure level (SPL) obtained using the proposed approach of Curle's analogy matches very well the experimental results. Thus, Curle's analogy can be used to obtain a fast, approximated and acceptable results about the noise radiation of airfoils avoiding the storage of noise sources which requires a huge space and time. [ABSTRACT FROM AUTHOR]

This paper describes the proposal for an airframe noise prediction method that can be used in Indonesian aircraft design research and industry. Since the industry is strategic, whether economically or securely, the country should build their own technology in all phases. Noise assessment is one of the sophisticated procedures in designing and manufacturing aircraft. Indonesia has not developed and manufactured engines, therefore airframe noise prediction is a relevant object to study. This method is proposed to be a utility to improve Indonesian aircraft design specifications. Based on the analogy theory of aeroacoustics, the use of Ffowcs Williams and Hawkings (FWH) equation was chosen to represent noise. The noise prediction method was developed by simulating and modeling in OpenFOAM, an open-source software based on the finite volume method. Its reliability has been proven in research and industrial computational fluid dynamics. The proposed method is developed from CFD and is compatible with using a personal computer device. The algorithm based on boundary conditions and hardware specifications has been built. Each step has been designed to support FWH Equation solver uniquely for this study. Computational domains represent the airflow area overrun a solid object. For the reason of far-field radiation, the domain token is 20 times the dimension of the object, and wall functions are applied. The atmosphere condition of low flight level and low to medium subsonic speed is taken to represent various aircraft designs and configurations. The condition represents the existing Indonesian industry of civil and military flying vehicles, and its possibility to develop. [ABSTRACT FROM AUTHOR]

In order to reduce the noise of the axial fan, an uneven blade spacing was used to optimize tonal noise and blade leading edge triangles were added to optimize the broadband noise. First, different angular distribution schemes were designed on the basis of ensuring the dynamic balance of the fan, and the noise of the different schemes was simulated by using the computational aeroacoustics (CAA) numerical simulation method. Then, the best optimal angular distribution design was selected for experimental verification, which confirmed the reliability of the numerical calculation. The test results also show that the uneven blade spacing has little effect on the fan pressure and efficiency. Additionally, triangular leading edges were added to each blade based on the uneven blade spacing design. Combining these two optimization schemes proved to reduce the fan noise by 13.4 dBA. [ABSTRACT FROM AUTHOR]

The effect of aerodynamic noise from high-speed trains on residents living near railway lines is a critical issue. The pantograph cavity is considered to be a major source of aerodynamic noise. To address this problem, this study proposes the application of jetting at the leading edge of the cavity, directly targeting noise reduction at its source. The large eddy simulation approach is used for flow calculations, and the Ffowcs Williams and Hawkings aeroacoustic analogy is adopted for far-field acoustic predictions. Orthogonal designs and the backpropagation algorithm optimized by the genetic algorithm (BP-GA) is used to explore the effects of jetting factors on noise and identify the optimal parameters. Orthogonal design analysis shows that the most influential among the factors is jet orifice diameter, followed by jet velocity and jet angle. The use of the BP-GA algorithm for optimization reveals that the optimal jet parameters are jet velocity of 118.28 m/s, jet angle of 3.17°, and jet orifice diameter of 76.74 mm. The algorithm predicts a minimum noise level of 91.04 dB, which is close to the simulated noise level of 90.74 dB. The jetting process achieves a maximum noise reduction of 4 dB. Results demonstrate that the proposed method for cavity leading edge jetting effectively reduces turbulent kinetic energy and horseshoe-shaped vortices in the cavity, leading to noise reduction. This method also minimizes the effects of aerodynamic noise on distant areas, such as waiting areas and residential buildings. This work provides a theoretical basis for increasing high-speed train speeds. [ABSTRACT FROM AUTHOR]

Jet–flap interaction noise is brought about by the proximity of a turbofan engine to airframe surfaces and arises mainly during take-off and landing. This paper reports a numerical analysis of the flow and acoustic fields generated by an isothermal, subsonic, circular jet close to a simplified geometry of a wing and flap under the intrusive conditions of flap deflection angles of 7 ∘ , 15 ∘ , and 30 ∘ and Mach numbers of 0.5 and 0.7. The jet–flap interaction was seen to deflect the jet stream downwards and increase turbulence intensity levels. Regarding the far-field noise, amplifications with respect to the free-jet noise were verified over the entire frequency range, and noise shielding was predicted at high frequencies. The installation effects at low frequencies were found to be less dominant as the jet velocity was increased. Substantial amplifications were found for a flap angle of 30 ∘ in the high-frequency range, most likely caused by the changes in the jet quadrupole noise due the deformation of the jet flow. The far-field noise was also computed from the pressure field on the wing and flap surfaces, showing that the contributions of both the wing and the flap give rise to a dipole-like directivity pattern perpendicular to the surfaces, at low Strouhal. The noise associated with the flap was found to dominate over that of the wing for almost all polar angles. [ABSTRACT FROM AUTHOR]

Flow and acoustic fields of a rectangular over-expanded supersonic jet interacting with an adjacent parallel plate are investigated using compressible Large Eddy Simulations (LES). The jet exits from a converging diverging rectangular nozzle of aspect ratio 2 with a design Mach number 1.5. Four distances (0 to 3 equivalent diameters) between the plate and the adjacent lip of the rectangular jet in the minor axis plane are studied. The geometry of the nozzle, the positions of the plate, and the exit conditions are identical to the ones of an experimental study. Snapshots and mean velocity fields are presented. Good agreement with the PIV experimental measurements is obtained. Previously, the corresponding free jet has been found to undergo a strong flapping motion in the minor axis plane due to screech. Here, it is shown that the intensity of the screech increases for certain distances from the plate and decreases for others, as compared to the corresponding free jet. Two points space-time cross correlations of the pressure along the jet's shear-layers show, in two cases, an amplification of the aeroacoustic feedback mechanism leading to screech noise in the jet shear-layer closer to the plate. This amplification is due to acoustic waves impinging on the plate, and generating propagating waves back towards the jet, thus exciting the shear-layer at the screech frequency, around the tenth shock cell. Moreover, when the jet develops as a wall jet on the plate, the screech frequency and its associated flapping motion is canceled but a symmetrical oscillation of the jet at a lower frequency becomes dominant and radiates in the near acoustic field. This oscillation mode, as the ones associated with the screech tones for the other cases studied, can be explained by the use of a vortex sheet model of the ideally expanded equivalent planar jet. [ABSTRACT FROM AUTHOR]

Motivated by the growing concerns for the noise emissions generated by the interaction between the engine jet and the airframe surfaces consequent to the progressive increase of the engine by-pass ratio in commercial aircraft architectures, we study the estimation of the far-field noise radiated by a jet in an installed configuration based on sensor readings in the near field of the jet. We carry out an experimental investigation of the jet-surface interaction phenomena in a simplified set-up where a flat plate is installed parallel to the nozzle axis of a subsonic jet. Real-time estimation is performed using empirical linear transfer functions identified between near-field sensors and far-field observers. The transfer function and the accuracy of the noise estimation performed are characterised as a function of jet flow conditions, type and streamwise position of near-field sensors and radial distance between the jet and the edge. This analysis can guide the positioning of near-field sensors for the implementation of closed-loop control strategies for the jet-surface interaction noise. [ABSTRACT FROM AUTHOR]

This paper experimentally characterises the far-field noise emissions of a rotor operating in reverse non-axial inflow conditions. Specifically, experiments were undertaken at a range of rotor tilting angles and inflow velocities to investigate the effects of negative tilting on rotor acoustics and their correlation with aerodynamic performance. The results show that the forces and moments experienced by the rotor blades change significantly with increasing inflow velocity and increasing negative tilting angle. Correspondingly, distinct modifications to the far-field acoustic spectra are observed for the negatively tilted rotor when compared to the edgewise condition, with the broadband noise content notably increasing. Moreover, for a given tilting angle, the broadband noise component is accentuated with increasing inflow velocity, similar to when the negative tilting angle is increased. With reference to the flow-field surveys conducted in the literature and a preliminary in-house flow measurement, the increase in broadband content can possibly be attributed to the heightened level of ingestion of blade self-turbulence, i.e., the ingestion of turbulent wake generated by the upstream portion of the rotor by the downstream portion. At lower inflow velocities, the magnitude of the blade passing frequency at each of the observer angles is found to change minimally with negative tilt. In contrast, at higher inflow velocities, the directivity pattern and intensity of both the blade passing frequency and the overall sound pressure level are shown to change with increases in magnitude, particularly at downstream observer locations with negative tilt. These findings have important ramifications for the design and suitable operational profile of aerial vehicles for future urban air mobility applications. [ABSTRACT FROM AUTHOR]

The aerodynamic noise test of counter-rotating propellers in wind tunnels has the problems of difficult time coordination and high cost. The aerodynamic noise test system of counter-rotating propellers，which is capable of simulating the driving state of counter-rotating propellers on the ground，is developed by relying on the acoustic environment of the ground. The aerodynamic noise test is carried out on a perforated counter-rotating propeller and a reference propeller to validate the test system and evaluate the noise reduction effect of the perforated counter-rotating propeller. The results show that the perforated counter-rotating propeller can effectively reduce the noise under the prerequisite of guaranteeing the aerodynamic performance in the range of the tested working conditions. In the 90° pointing angle，which has the most significant effect on the aircraft cabin，the noise reduction at the secondorder passing frequency reaches 5 dB. The designed aerodynamic noise test system can provide assistance for the evaluation of aerodynamic noise of the counter-rotating propeller and the design of noise reduction.

This paper describes the proposal for an airframe noise prediction method that can be used in Indonesian aircraft design research and industry. Since the industry is strategic, whether economically or securely, the country should build their own technology in all phases. Noise assessment is one of the sophisticated procedures in designing and manufacturing aircraft. Indonesia has not developed and manufactured engines, therefore airframe noise prediction is a relevant object to study. This method is proposed to be a utility to improve Indonesian aircraft design specifications. Based on the analogy theory of aeroacoustics, the use of Ffowcs Williams and Hawkings (FWH) equation was chosen to represent noise. The noise prediction method was developed by simulating and modeling in OpenFOAM, an open-source software based on the finite volume method. Its reliability has been proven in research and industrial computational fluid dynamics. The proposed method is developed from CFD and is compatible with using a personal computer device. The algorithm based on boundary conditions and hardware specifications has been built. Each step has been designed to support FWH Equation solver uniquely for this study. Computational domains represent the airflow area overrun a solid object. For the reason of far-field radiation, the domain token is 20 times the dimension of the object, and wall functions are applied. The atmosphere condition of low flight level and low to medium subsonic speed is taken to represent various aircraft designs and configurations. The condition represents the existing Indonesian industry of civil and military flying vehicles, and its possibility to develop.

Serrations are commonly employed to mitigate the turbulent boundary layer trailing-edge noise. However, significant discrepancies persist between model predictions and experimental observations. In this paper, we show that this results from the frozen turbulence assumption. A fully developed turbulent boundary layer over a flat plate is first simulated using the large-eddy simulation method, with the turbulence at the inlet generated using the digital filter method. The space–time correlations and spectral characteristics of wall pressure fluctuations are examined. The simulation results demonstrate that the coherence function decays in the streamwise direction, deviating from the constant value of unity assumed in the frozen turbulence assumption. By considering an exponential decay function, we relax the frozen turbulence assumption and develop a prediction model that accounts for the intrinsic non-frozen nature of turbulent boundary layers. To facilitate a direct comparison with frozen models, a correction coefficient is introduced to account for the influence of non-frozen turbulence. The comparison between the new and original models demonstrates that the new model predicts lower noise reductions, aligning more closely with the experimental observations. The physical mechanism underlying the overprediction of the noise model assuming frozen turbulence is discussed. The overprediction is due to the decoherence of the phase variation along the serrated trailing edge. Consequently, the ratio of the serration amplitude to the streamwise frequency-dependent correlation length is identified as a crucial parameter in determining the correct prediction of far-field noise. [ABSTRACT FROM AUTHOR]

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Understanding the occurrence of various feedback mechanisms of an under-expanded impinging supersonic jet is a crucial task in research. The presence of several jet modes is examined in this study for the flat and corrugated impinging plate geometries. The behavior of impinging plate configurations during mode switching is investigated by varying the flow state, such as the jet Mach number. The staging behavior at various jet Mach numbers is observed using acoustic spectral plots and schlieren flow visualization. To explore the presence of various types of modes during the jet impingement due to the modification of jet Mach number, ensemble averaging and Proper Orthogonal Decomposition of schlieren images are carried out. In the majority of situations, the corrugated design shows a reduction in tonal noise and overall sound pressure level. In exceptional cases, for the corrugated plates, the enhanced overall sound pressure level is caused by the existence of axisymmetric instability (A1, A2). [ABSTRACT FROM AUTHOR]

Being a direct computational aeroacoustics method, Lattice Boltzmann method (LBM) has great potential and broad application perspective in the field of numerical simulation of aerodynamic noise due to its low dispersion and low dissipation. A series of numerical algorithms and the related improvements based on the standard LBM method are proposed and developed in this paper to adapt to the airfoil noise calculation with complex grid at middle‐high Reynolds number. First, a new mathematical‐boundary‐recognition algorithm based on Green's formula is proposed to deal with complex curved geometric models, which is validated by three‐element airfoil 30P30N benchmark. Then, in order to reduce grid redundancy and improve computing efficiency, the grid refinement technique of domain decomposition model (DDM) is adopted and also improved, which is verified by calculating the flow and sound fields around 2D and 3D cylinders at Reynolds number equal to 90,000. Finally, three different LES turbulence models are combined with the standard MRT‐LBM method, where different finite difference schemes are used to solve Reynolds stress tensor which is different from the traditional one. Through the direct acoustic numerical simulation of NACA0012 airfoil at Reynolds number equal to 200,000, the effects of Smagorinsky models and Wall‐adapting local eddy‐viscosity (WALE) model on aerodynamic noise prediction are compared and analyzed. Overall, the proposed methodology is shown to be appropriate for predicting the aerodynamic noise at low Mach number and can successfully simulate the generation and propagation of far field acoustics. [ABSTRACT FROM AUTHOR]

Currently, the most advanced micromachined microphones on the market are based on a capacitive coupling principle. Capacitive micro-electromechanical-system-based (MEMS) microphones resemble their millimetric counterparts, both in function and in performance. The most advanced MEMS microphones reached a competitive level compared to commonly used measuring microphones in most of the key performance parameters except the acoustic overload point (AOP). In an effort to find a solution for the measurement of high-level acoustic fields, microphones with the piezoelectric coupling principle have been proposed. These novel microphones exploit the piezoelectric effect of a thin layer of aluminum nitride, which is incorporated in their diaphragm structure. In these microphones fabricated with micromachining technology, no fixed electrode is necessary, in contrast to capacitive microphones. This specificity significantly simplifies both the design and the fabrication and opens the door for the improvement of the acoustic overload point, as well as harsh environmental applications. Several variations of piezoelectric structures together with an idea leading to electrically controlled sensitivity of MEMS piezoelectric microphones are discussed in this paper. [ABSTRACT FROM AUTHOR]

From an engineering point of view pulsatile pipe flows, which include the hydrodynamic and acoustic fluctuations, can cause many problems such as vibration, noise and structural fatigue etc. Knowledge of the characteristics of hydrodynamic fluctuations (pseudo sound) and acoustic waves inside pipes particularly those with bends make is crucial to pipeline vibration and noise reduction. Numerical simulation of the pulsatile pipe flows shows challenges especially in curved pipes with different flow modes compounded, and predicting the acoustic fluctuations is even tougher. In this work, the unsteady flows of a straight and curved pipe with a 90° elbow induced by a pulsatile pressure inlet were calculated by Detached Eddy Simulations (DES). The pipe acoustic fields are solved through an acoustic perturbation equation, which takes the time derivative of hydrodynamic pressure as the aeroacoustic source. The characteristics of the flow and sound fields are discussed detailly, and special attention is paid to the influence of the 90° elbow on the curved pipe flow. The results show that the amplitude of hydrodynamic pressure fluctuations decrease almost linearly along the straight pipe, and the 90° elbow has obvious local influence on the flow pressure and velocity especially in the downstream side. The acoustic field is of 1-D standing wave pattern because of the fundamental pulsation frequency lower than the pipe cutoff frequency, and the 90° elbow has only negligible effect on the acoustic field in the flow frequency range. [ABSTRACT FROM AUTHOR]

In this study, the asymptotic solutions of the pressure variations induced by two trains passing each other in a tunnel are theoretically investigated. The one-dimensional inviscid compressible airflow is analysed, and two methods to obtain numerically exact solutions and MH expansion formulas for approximate equations are presented, where MH is the Mach number of the high-speed train. The pressure coefficient, corresponding to the maximum value of the magnitude of the pressure, is expressed as |cp|max = |cp,min| = [(R/(1 - R))(1 + a)2 + (R(1 - R)/(1 - 2R)2)(1 - a)2] + O[MH], where cp,min < 0, a = UL/UH and UL and UH denote the speeds of the low- and high-speed trains, respectively, and R is the cross-sectional area ratio of the train to the tunnel. The theoretical results indicate the dependence of the speeds of the two trains on the pressure distribution and that the maximum magnitude of the asymptotic pressure for a fixed value of MH is obtained for a = 1 and a = 0 when R < Rc and R > Rc, respectively, where Rc denotes the critical blockage ratio. Because the airflow along the side of the low-speed train, induced by the low-speed train, is along the running direction of the high-speed train and reduces the relative velocity of the high-speed train as the two trains pass each other, |cp|max for a = 0 is larger than |cp|max for a = 1 when R > Rc. It is theoretically demonstrated that, as conventional high-speed railway systems satisfy R < Rc, a conservative pressure estimation can be established assuming a = 1. [ABSTRACT FROM AUTHOR]

A numerical model for investigating the aerodynamic sound generated by the interaction between flow and bluff body is developed, and then applied to the computation of near-field noise induced by the turbulent airflow passing through a generic side mirror. The flow field is simulated by employing the viscous vortex method. Then the sound sources within the computational domain are extracted from the simulated results with a vortex sound equation. The sound waves, sum of radiated sound and scattered sound, are determined using a time-domain boundary element method combined with the convolution quadrature method for improving the stability of the time marching algorithm. Further, the fast multipole method is adopted to enhance the computational efficiency. The computed pressure coefficients and surface pressure fluctuations match the measurements and simulations very well, and the obtained spectra of near-field aerodynamic sound are also close to the measured results. The comparisons of computed results of two additional cases with the previous studies demonstrate convincingly that the proposed model can effectively predict the flow-induced near-field noise. [ABSTRACT FROM AUTHOR]

The noise created when a firearm is fired has a lot of adverse effects on humans and the environment, so analyzing and attenuating this noise is essential. This study aimed to examine propellant flow and the sound generated by this flow using a hybrid computational fluid dynamics and computational aeroacoustics method. The compatibility of this numerical study was also validated by comparing it with the experimental result. The impact of critical parameters, such as the shape and number of baffles of the suppressor, was studied. Finally, the overpressure and acoustics results of different suppressors were compared with each other and with the unsuppressed condition. According to the result, the suppressor with curved baffles shows a better performance. When the exit pressure at the tip of suppressor compared to the initial inlet pressure to the suppressor the percentage drop in overpressure was 76.01%, 78.79% and 81.3% for the suppressor with one, three and five curved baffles, respectively. For condition without suppressor, 169.49 dB of SPL was recorded. When using a suppressor without a baffle, this value was reduced to 162.13 dB. For the suppressor with one, three and five curved baffles, the SPL value was 160.234 dB, 159.43 dB and 158.11 dB, respectively. Generally, this study shows that the suppressor's efficiency is directly proportional to its shape and number of baffles. [ABSTRACT FROM AUTHOR]

For noise reduction at a low-to-moderate Reynolds number, airfoil trailing-edge tonal noise has multiple prominent tones. Among these tones, secondary tones are greatly influenced by external disturbances such as oscillations commonly in the environment. In previous experiments, the spatial movement of sources was found to be related to an inherent high-frequency oscillation. Therefore, the spatial influence of external low-frequency oscillations was investigated in this study. By using tripping tapes to construct different symmetry source regions on the pressure side with side secondary tones, a transient spatial analysis of an NACA0012 airfoil at 2 degrees was performed by microphone arrays when a 10 Hz pressure oscillation was significant at 24 m/s. Temporally, this 10 Hz periodic strength change became more intense at a broader frequency bandwidth for a longer source region. Furthermore, a substantial time delay, significantly larger than the sound propagating time difference between microphones, was observed exclusively along the spanwise direction. This delay led to a periodic directivity pattern, particularly when two 0.2 m source regions were separated by a 0.2 m or 0.4 m tripping region. This low-frequency oscillation introduces an asymmetric transient switching pattern for symmetric spanwise source regions. Consequently, the response of airfoils to external oscillations in field tests should be considered. [ABSTRACT FROM AUTHOR]

The impedance matching is essential in producing high-intensity acoustic waves. This paper studies the impedance calculation methods of modulators and double-chamber Helmholtz resonators based on fluid mechanics, aeroacoustics, and electro-acoustic analogy principles. The results: ① The impedance expression of the resonators is derived. ② The impedance variation law of the chamber volume, the stator size, the air supply pressure, the air supply flow, and the working frequency is calculated. ③ The impedance-matching design scheme of the two-chamber Helmholtz low frequency and high acoustic intensity generator is given. Based on the experimental results of the high acoustic intensity generating device that the author has manufactured in the closed space, the correctness of the impedance calculation and matching design scheme in this paper is verified. This method has guiding significance for the design of high-intensity acoustic generating devices in confined space. [ABSTRACT FROM AUTHOR]

It is a distinct privilege to produce this article to honor Professor Tam, the foremost authority on jet aeroacoustics. The fundamental characteristics of jet noise have been studied for 70 years, since the pioneering work of Lighthill in the 1950s. The acoustic analogy, with many variants, has served as the leading theory for nearly 50 years. Many leading researchers in the 1970s formulated theories to interpret the measured trends from subsonic and supersonic jets, using acoustic analogy and flow features as the framework. Quadrupoles, dipoles and monopoles were believed to constitute the sources of noise. The discovery of large-scale organized structures in free shear layers and jets sparked a different avenue of thinking about their importance for noise generation. Tam was the first to clearly demonstrate that these structures are efficient generators of noise and constitute the dominant noise sources, especially in the downstream direction. Now, two schools of thought emerged on the sources of jet noise. Experimental measurements showed that the mean flow as well as the turbulence statistics exhibit a self-similarity in the mixing layer and another similarity in the fully developed jet. Based on these observations, Tam proposed that since noise is generated by the turbulence of the jet, the noise spectra generated by fine-scale and large-scale turbulence should also exhibit self-similarity. By examining a large set of supersonic jet noise data acquired at NASA Langley, Tam offered evidence that the turbulent mixing noise of high-speed jets does consist of two independent self-similar components. In this paper, experimental evidence is compiled from an extensive database that quantifies the effect of several parameters that affect jet spectra. A new scaling method is developed and extended to noise predictions for realistic dual-stream nozzle geometries. The objectives of this paper are to serve as a synthesis of noise characteristics and to focus on application to real-world problems. Results from five different experimental measurements are examined: (1) farfield spectral characteristics; (2) azimuthal and polar correlations in the farfield; (3) correlations of jet turbulence fluctuations and farfield sound; (4) measurement of source distributions with an elliptic mirror; and (5) space-time correlation measurements in the nearfield with a cage array, and nearfield-farfield correlations. Two distinctly different trends are observed in the angular ranges of 50° – ∼120° and ∼120° – 165° for all the parameters investigated with the above five approaches. The salient observations are mutually supporting, and the cumulative weight lends credence to the proposition that there are two distinct sources of turbulent mixing noise. [ABSTRACT FROM AUTHOR]

The Aeroacoustics Laboratory in the Department of Aerospace Engineering at Penn State currently serves several areas of experimental research. Most notable in the laboratory is the Anechoic Chamber of which the Aeroacoustics Wind Tunnel is an integral part. In recent years the dominant areas of research have included model aircraft exhaust jets and model rotors for helicopter applications involving detailed noise experiments. In the jet noise area, the applications have included aircraft in both flyover and take-off configurations. In the flyover application the air flow surrounding the exhaust jet has a significant effect on the radiated noise and must be simulated to produce experimental results that will be most useful in preliminary design of noise suppression applications. The Aeroacoustics Wind Tunnel at Penn State serves the purpose for this simulation. During aircraft take-off, the aircraft airspeed is in a range exceeding 250 ft per second, or Mach number of over M = 0.22. In the years preceding 2022, the velocity of the free jet flow of the Penn State facility was limited to less than 200 ft/sec. This paper reports on a project to design upgrades to the existing wind tunnel to improve the test section flow velocity to values closer to the 250 ft/sec target. The approach began with a preliminary design of the return flow ducting to convert the open jet, open return wind tunnel to a closed return tunnel. The concept is to make use of the flow energy in the exhaust of the tunnel to boost the input energy to the inlet fan that drives the flow to the test section. An integral part of the preliminary design involved making use of an upgraded flow analysis computer code to predict the gain in test section velocity of the new facility. For the available horsepower of two fans in the facility, the configurations of the existing and upgraded wind tunnel were entered into this analysis code based on a quasi-one-dimensional formulation. The code included empirical data formulated to include the various shapes of the components of the tunnels. The relative flow velocities in each section used simple incompressible continuity to calculate all velocities in the component sections. The average dynamic pressure in each section is calculated from the flow (constant) density and the velocity squared. Each section had a non-dimensional pressure loss parameter calculated from empirical data (from the relevant literature). The basic code (using the Excel algorithms) was based on a predictor-corrector concept in which the calculation was initialized with the estimated velocity and pressure of the flow at the entrance to the test section. The loss in each following section is estimated to determine the loss in the total pressure and the cumulated loss compared to the gains made at the two fans in the circuit. The initial guess is adjusted based on the total pressure value at the end of the circuit. During the related experiments, measurements of the total and dynamic pressure were performed at several joints between sections and compared to the values predicted in the analysis code. Various settings of the horsepower of the drive fans were used and the results were compared to flow measurements for the tunnel configurations (predominantly for the open return and closed return tunnel setups available before and after the upgrade project.) Results of the analysis and experiments are presented and analyzed together with the remaining activities planned for the coming months. [ABSTRACT FROM AUTHOR]

Interaction between turbulence and an airfoil is a significant aerodynamic noise source for many engineering applications when turbulence in the wake of upstream blades interacts with the leading edge of downstream blades. Modeling the oncoming turbulence as harmonic gusts is a common approach to study the noise generated by turbulence–airfoil interaction. In former studies, it has been found that the sound reduces with the increase in airfoil thickness. However, this is not always true for high-reduced frequency cases. This paper studies airfoil thickness effects using a computational aeroacoustics approach based on the spectral/hp element method. It is found that the sound pressure decreases with the increase in thickness for low-reduced frequency gusts. However, the sound pressure increases in the upstream direction and declines downstream with the increase in thickness for high-reduced frequency cases. To reveal its mechanism, a semi-analytical method and the convective Ffowcs Williams–Hawkings equation are used for low- and high-reduced frequency, respectively, to analyze the radiated noise. For low-reduced frequency cases, the sound reduction due to thickness is caused by the decrease in amplitude in the leading edge region and the increase in phase difference along the airfoil. For high-reduced frequency cases, the phase difference in the upstream observer is much more significant than the downstream observer's due to the convective effect. The increase in phase difference results in a cancellation effect, which leads to a different behavior upstream. [ABSTRACT FROM AUTHOR]

The paper presents the results of an experimental study of a subsonic turbulent jet noise. The cases of an isolated jet and a jet in the presence of a closely spaced plate are considered. The studies were carried out for the conical and equivalent-sized chevron and lobed nozzles. The work demonstrates the possibility of reducing both broadband jet noise and low-frequency jet–plate interaction noise using a lobed nozzle. [ABSTRACT FROM AUTHOR]

Investigations into the aerodynamic properties of vertical sound barriers exposed to high-speed operations employ computational fluid dynamics. The primary focus of this research is to evaluate the influence of train speed and the distance (D) from the track centerline under various operating conditions. The findings elucidate a marked elevation in the aerodynamic effect amplitude on sound barriers as train speeds increase. In single-train passages, the aerodynamic effect amplitude manifests a direct relationship with the square of the train speed. When two trains pass each other, the aerodynamic amplitude intensifies due to an additional aerodynamic increment on the sound barrier. This increment exhibits an approximate quadratic correlation with the retrograde train speed. Notably, the impact of high-speed trains on sound barrier aerodynamics surpasses that of low-speed trains, and this discrepancy amplifies with larger speed differentials between trains. Moreover, the train-induced aerodynamic effect diminishes significantly with greater distance (D), with occurrences of pressure coefficient (CP) exceeding the standard thresholds during dual-train passages. This study culminates in the formulation of universal equations for quantifying the influence of train speed and distance (D) on sound barrier aerodynamic characteristics across various operational scenarios. [ABSTRACT FROM AUTHOR]

Over the past two decades, there has been a renewed interest in the development of a new generation of supersonic aircraft for civil purposes that could potentially succeed Concorde. However, the noise annoyance is still considered one of the hampering factors to meet public consensus. This paper aims at revealing the potential of numerical simulations to predict sonic boom signature in Near Field at early design stages. In particular, the paper further demonstrates the applicability of the numerical approach proposed by NASA and other partners during the Sonic Boom Prediction Workshops held between 2014 and 2021, to compute the pressure signature of aircraft in the zone close to it. The results highlight the suitability of the approach (1) to capture the impact of aircraft flight condition variations on the sonic boom signature, (2) to enable the characterization of novel aircraft layout, including Mach 5 waverider configuration, (3) to provide near-field shock wave noise predictions that can be used to evaluate shock propagation, on-ground signature analyses, and annoyance assessment. [ABSTRACT FROM AUTHOR]

This paper conducts simulations to analyze the flow control characteristics of cylindrical flow over porous plates in the subcritical regime, utilising large-eddy simulation (LES) coupled with the Ffowcs Williams-Hawkings (FW-H) method. Three Reynolds numbers (Re) were studied (3900, 4.7 × 104, and 1.4 × 105) with two configurations: bare cylinder and porous plate attached cylinder (PPAC). The research findings suggest that, at Re of 1.4 × 105, PPAC showed significant reductions in drag (28.45%) and lift fluctuation (97.53%). The drag reduction effect is nearly four times higher than that observed at Re of 3900. The porous plate filters out high-frequency irregular oscillations in the wake flow. At high Res, the dominance of high-frequency modes in the wake of a bare cylinder, along with a smaller recirculation region, enhances the control effectiveness of PPAC. The dominant mode of the Kármán vortex street reappearing in the far wake region is enhanced. Aeroacoustic analysis reveals that, due to the superior noise suppression effect of the PPAC across the full frequency range, particularly within the 102–103 frequency range, on average, the noise surrounding the cylinder experiences a reduction of 27.85 dB, and it exhibits more uniform directionality at Re of 1.4 × 105.

The boundary element method is widely used in practical engineering problems, especially in the field of acoustics. For flow‐induced noise, the main target of acoustic calculations is to solve the wave equation with the flow field information. However, the sound field distribution of noncompact structures in half‐space is especially complex because of the strong scattering effect, while the object surface boundary integration often brings a large workload and generates numerical singularities. In this paper, an improved boundary element method for predicting the aeroacoustic noise of noncompact structures is proposed, which can consider the characteristic distribution of sound field induced by complex structures in half‐space. The smooth permeable boundary surrounding the object is used as the integration boundary, while the scattering effect of the ground boundary is investigated by combining the mirror Green's function method, and the numerical prediction of aeroacoustic noise is carried out for the dipole source and NACA0012 airfoil in half‐space. Numerical results show that the far‐field noise obtained by using the permeable surface is consistent with that obtained by integrating the direct object boundary under the influence of ground boundary scattering. The mirror image Green's function method is able to finely capture the ground scattering effect, which has a significant effect on the sound field as the frequency increases. [ABSTRACT FROM AUTHOR]

The twin-jet configuration allows two different scenarios to close the screech feedback. For each jet, there is one loop involving disturbances which originate in that jet and arrive at its own receptivity point in phase (self-excitation). The other loop is associated with free-stream acoustic waves that radiate from the other jet, reinforcing the self-excited screech (cross-excitation). In this work, the role of the free-stream acoustic mode and the guided-jet mode as a closure mechanism for twin rectangular jet screech is explored by identifying eligible points of return for each path, where upstream waves propagating from such a point arrive at the receptivity location with an appropriate phase relation. Screech tones generated by these jets are found to be intermittent with an out-of-phase coupling as a dominant coupling mode. The instantaneous phase difference between the twin jets computed by the Hilbert transform suggests that a competition between out-of-phase and in-phase coupling is responsible for the intermittency. To model wave components of the screech feedback while ensuring perfect phase locking, an ensemble average of leading spectral proper orthogonal decomposition modes is obtained from several segments of large-eddy simulation data that correspond to periods of invariant phase difference between the two jets. Each mode is then extracted by retaining relevant wavenumber components produced via a streamwise Fourier transform. Spatial cross-correlation analysis of the resulting modes shows that most of the identified points of return for the cross-excitation are synchronised with the guided jet mode self-excitation, supporting that it is preferred in closing rectangular twin-jet screech coupling. [ABSTRACT FROM AUTHOR]

Aerodynamic noise emitted by small wind turbines is a concern due to their proximity to urban environments. Broadband airfoil self-noise has been found to be the major source, and several studies have discussed techniques to reduce airfoil leading-edge and trailing-edge noises. Reduction mechanisms inspired by owl wings and their airfoil sections were found to be most effective. However, their effect/s on the tip vortex noise remain underexplored. Therefore, this paper investigates the effects of implementing an owl airfoil design on the tip vortex noise generated by the National Renewable Energy Laboratory (NREL) Phase VI wind turbine to gain an understanding of the relationship, if any, between airfoil design and the tip vortex noise mechanism. Numerical prediction of aeroacoustics is employed using the Ansys Fluent Broadband Noise Sources function for airfoil self-noise radiation. Detailed comparisons and evaluations of the generated acoustic power levels (APLs) for two distinguished inlet velocities were made with no loss in torque. Although the owl airfoil design increased the maximum generated APL by the baseline model from 105 dB to 110 dB at the lower inlet velocity, it significantly reduced the surface area generating the noise, and reduced the maximum APL generated by the baseline model by 4 dB as the inlet velocity increased. The ability of the owl airfoil to mitigate the velocity effects along the span of the blade was found to be its main noise reduction mechanism. [ABSTRACT FROM AUTHOR]

The ability to predict the mobility of rock avalanches is necessary when designing strategies to mitigate the risks they pose. A popular mobility indicator of the flow front is the Heim's apparent friction coefficient μ H . In the field, μ H shows a decrease in value as flow volume V increases. But this correlation has been a mystery as to whether it is due to a causal relationship between V and mobility since: (1) field data of μ H do not collapse onto a single curve because typically widely scattered and (2) laboratory experiments have shown an opposite volume effect on the center of mass mobility of miniature flows. My numerical simulations confirm for the first time the existence of a functional relationship of scaling parameters where μ H decreases as V increases in unsteady and nonuniform 3D flows. Data scatter is caused by μ H that is affected by numerous other variables besides V. The interplay of these variables produces different granular regimes with opposite volume effects. In particular, μ H decreases as V increases in the regime characterized by a relatively rough subsurface. The relationship holds for large-scale flows that, like rock avalanches, consist of a very large number of fine clasts traveling in wide channels. In these dense flows, flow front mobility increases as flow volume increases, as channel width increases, as grain size decreases, as basal friction decreases and as flow scale increases. Larger-scale flows are more mobile because they have larger Froude number values. [ABSTRACT FROM AUTHOR]

This experimental study investigates the effect of blade phase angle on noise attenuation in two adjacent, electronically synchronized propellers. Acoustic measurements were performed in an aeroacoustic wind tunnel with a distributed electric propulsion system that involved the adjustment of relative phase angles of 2-bladed propellers between Δψ = 0° and 90°. Ranges of advance ratios (J = 0–0.73) were investigated at a fixed propeller rotation speed of 5000 rpm. The investigation explored the impact on noise directivity and frequency characteristics. The findings reveal significant reductions in noise directivity and tonal noise at the blade pass frequency (BPF). A relative phase angle of Δψ = 90° demonstrated the maximum noise reduction, with an 8 dB decrease at the first BPF and a 2 dB reduction in overall sound pressure level at J = 0. For in-flow conditions (J > 0), a relative phase angle of Δψ = 90° resulted in significant noise reductions of about 24 dB in the first BPF and 6 dB in overall sound pressure level, compared to Δψ = 0°. These observations offer critical insights into the use of the propeller's relative phase angle as an effective noise control method in the distributed electric propulsion system. [ABSTRACT FROM AUTHOR]

This study focuses on the implementation of the spectral difference (SD) method on hexahedral elements to NVIDIA graphics processing units (GPUs) using the Compute Unified Device Architecture (CUDA) for aeroacoustic problems. Three problems were addressed in the implementation of this study: thread parallelism strategy optimization within the GPU, data access patterns management, and multi-GPU parallelization implementation. Computational speed testing showed that the three factors significantly affect the efficiency of the code on the GPU. The implemented GPU solver was validated using an inviscid problem and a viscous problem. The numerical results show that the GPU solver achieves the same level of accuracy as the CPU program, with remarkable speed improvements. Specifically, compared with a single CPU core with a turbo boost frequency of 3.2 GHz (Intel Xeon Silver 4210), the inviscid case tested on an RTX 2070 Super GPU achieved acceleration of 122.4× , and the viscous case conducted on an RTX 3090 GPU achieved acceleration of 229.7×. Additionally, the GPU solver exhibits a parallel efficiency exceeding 93% when performing parallel computing on a platform with multiple RTX 3090 GPU cards. Furthermore, the GPU-accelerated computational aeroacoustics solver was applied to compute the noise from a low-speed propeller. The computed results were compared with experimental data, and the excellent agreement demonstrated the effectiveness and feasibility of the GPU implementation of the SD solver. [ABSTRACT FROM AUTHOR]

The impeller of a multi-disc fan was made of rotating smooth discs spaced close together. The concept of such a machine came from Nikola Tesla in 1909. The fluid, based on the principle of friction against the rotating surfaces of the discs, flows outwards from the disc pack into an outlet. According to the literature, such a design is characterized by low noise emissions. Based on calculations, a prototype of a multi-disc fan was designed and then manufactured. Tests were carried out for the sound power level emitted from the fan's outlet. The examination was performed in a reverberation chamber according to the ISO 3741:2010 standard. The noise source was visualized using an acoustic camera. The aerodynamic characteristics were also determined experimentally according to the ISO 5801:2017 standard. The aim of this study is provide insight into the acoustic and aerodynamic characteristics of a multi-disc fan. [ABSTRACT FROM AUTHOR]

The influence of the boundary layer (BL) thickness on the tonal noise generated by subsonic impinging jets is investigated. For that, initially laminar jets at Mach numbers $0.6$ and $0.9$ with BL thicknesses $0.05r_0$ , $0.1r_0$ and $0.2r_0$ , where $r_0$ is the pipe-nozzle radius, impinging on a plate at $6r_0$ from the nozzle, are simulated. For Mach number $0.9$ , acoustic tones due to feedback loops of axisymmetric nature between the nozzle and the plate are generated at frequencies that do not vary with the BL thickness. The two dominant tones are, however, $17$ and $26\ \mathrm {dB}$ stronger for the thickest BL compared with the thinnest one. For Mach number $0.6$ , for the thinnest BL no acoustic peaks appear, as observed in the experiments of the literature, but narrow peaks resulting from axisymmetric feedback loops emerge for thicker BLs. Therefore, low subsonic impinging jets can be resonant for specific nozzle-exit conditions. The increase in tone amplitude for Mach number $0.9$ , and the establishment of feedback loops for Mach number $0.6$ with increasing BL thickness, are found to result from two changes in the jet flow. The first change is that the shear-layer laminar–turbulent transition occurs farther downstream for a thicker BL, leading to a greater predominance of the axisymmetric aerodynamic fluctuations near the plate. The second change is that the amplification of the flow fluctuations between the nozzle and the plate at the tone frequencies is stronger for thicker BLs. [ABSTRACT FROM AUTHOR]

The aeroacoustics of a boundary layer ingesting (BLI) ducted fan is investigated experimentally. The study examines a ducted fan immersed in an adverse streamwise pressure gradient turbulent boundary layer developed over a curved wall. Aeroacoustics measurements indicate that the noise from the BLI ducted fan results from a complex interaction among the fan, duct and the incoming boundary layer. The fundamental mechanisms of noise generation are explained using a general source separation strategy. A detailed noise comparison is made at varying fan rotational speeds and across a wide range of axial inflow velocities. In a low thrust regime, the noise is found to be driven by the fan loading, coupled with duct acoustics and the haystacking phenomenon. In a high thrust regime, the contribution from duct acoustics diminishes, and the noise is predominantly driven by the fan loading coupled with the haystacking phenomenon. [ABSTRACT FROM AUTHOR]