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An on-the-fly unsteady adjoint-based aerodynamic and aeroacoustic optimization methodology is presented, aiming to achieve practical engineering applications to explore high-efficiency and low-noise design for aerodynamic shapes. Firstly, a novel on-the-fly hybrid CFD-CAA approach is developed with a close integration of unsteady Reynolds-averaged Navier-Stokes equations and a fully viscous time-domain FW-H formulation. Subsequently, an adjoint-based sensitivity analysis method is proposed for unsteady aerodynamic and aeroacoustic problems with either stationary or moving boundaries, wherein a unified architecture for discrete-adjoint sensitivity analysis of both aerodynamics and aeroacoustics is achieved by integrating the on-the-fly hybrid CFD-CAA approach. The on-the-fly approach facilitates direct evaluation of partial derivatives required for solving adjoint equations, eliminating the need for explicitly preprocessing flow and adjoint variables at all time levels in a standalone adjoint CAA solver and consequently substantially reducing memory consumption. The proposed optimization methodology is implemented within an open-source suite SU2. Results show that the proposed on-the-fly adjoint methodology is capable of achieving highly accurate sensitivity derivatives while significantly reducing memory requirements by an order of magnitude, and further demonstrations of single-objective and coupled aerodynamic and aeroacoustic optimizations highlight the potential of the proposed method in exploring high-efficiency and low-noise design for aerodynamic shapes. [ABSTRACT FROM AUTHOR]

Shock waves, characterized by abrupt changes in pressure, temperature, and density, play a significant role in various materials science processes involving fluids. These high-energy phenomena are utilized across multiple fields and applications to achieve unique material properties and facilitate advanced manufacturing techniques. Accurate simulations of these phenomena require numerical schemes that can represent shock waves without spurious oscillations and simultaneously capture acoustic waves for a wide range of wavelength scales. This work suggests a high-order discontinuous Galerkin (DG) method with a finite volume (FV) subcell limiting strategies to achieve better subcell resolution and lower numerical diffusion properties. By switching to the FV discretization on an embedded sub-cell grid, the method displays advantages with respect to both DG accuracy and FV shock-capturing ability. The FV scheme utilizes a class of high-fidelity schemes that are built upon the boundary variation diminishing (BVD) reconstruction paradigm. The method is therefore able to resolve discontinuities and multi-scale structures on the subcell level, while preserving the favorable properties of the high-order DG scheme. We have tested the present DG method up to the 6th-order accuracy for both smooth and discontinuous noise problems. [ABSTRACT FROM AUTHOR]

Computationally extensive fluid flow simulations usually require numerical schemes with a high degree of numerical precision, where high wavenumber and frequency capturing capability is desired in solutions with high physical fidelity. Numerical dispersion and diffusion errors must be reduced to a minimum in order to caputre low amplitude and high frequency flow aspects, such as flow induced sound waves and minor instabilities. Such schemes also must be capable of dealing with flow discontinuities, such as shockwaves and expansion waves in cases where the flow velocity stands near or above sound velocity. In this work the Dispersion Relation Preserving (DRP) optimization is formally extended to the compact Finite Volume (FV) framework where the error is minimized along a given spectral interval. This is done by equaling the resulting algebraic compact spectral optimized Finite Difference (FD) operators with the Finite Volume formulation with the desired numeric characteristics. Numerical tests are carried over linear and nonlinear governing equations models for one and three-dimensional cases, where the travelling Gaussian wave, the transient expansion wave and the smooth to nonsmooth step wave flows are considered. The proposed compact DRP FV schemes surpasses its counterparts in some numerical tests while maintaining comparable capabilities with other more basic schemes in the remaining numerical tests. Excellent broadband content capturing capability is observed, while possessing reduced numerical instabilities generation. [ABSTRACT FROM AUTHOR]

Airfoil self-noise is investigated on a cantilever wing with a supercritical profile with a compressible wall-resolved large-eddy simulation. The Reynolds number based on the chord is 620,000, and the angle of attack is 5°. The aerodynamic results reveal the development of a complex vortex system at the side edge, including primary, secondary, and tertiary vortices that govern aerodynamic noise production. The wall-pressure coefficient and root-mean-square pressure coefficient contours highlight the side-edge shear layer and flow impingement of the primary vortex at the pressure-side edge as important noise-generation mechanisms. Dynamic mode decomposition shows that the dominant pressure modes on the airfoil exist along primary and secondary vortex impingement lines. Far-field acoustic predictions based on both solid and porous Ffowcs-Willliams and Hawkings' analogies demonstrate good agreement with experimental results between 1.5 and 9 kHz. Low-frequency tonal humps observed in the spectra result from duct acoustic modes excited by airfoil self-noise, a consequence of the installation effect. For this configuration, airfoil side-edge noise did not dominate over airfoil surface noise levels, suggesting that at low to moderate angles of attack, airfoil side-edge noise is not a significant contributor to the overall acoustic emissions. [ABSTRACT FROM AUTHOR]

Numerical simulations are performed for flows over a circular cylinder at a Reynolds number ranging from 150 to 5000, and Mach number of 0.2, to assess the predictive capability of URANS and hybrid RANS/LES for acoustic waves generation and propagation. Complementary direct numerical simulations (DNS) are performed to generate validation datasets and to provide more insight into the problem. DNS predictions show that noise induced by the vortex shedding is radiated primarily at a 90-degree angle with respect to the wake direction and dictates the dominant frequency of the sound pressure waves. Turbulence dominates the noise in the near-field wake, resulting in a broadband pressure spectrum. URANS both underpredicts and overpredicts the noise levels in the wake region and in the direction normal to the freestream flow, respectively, which is attributed to its inability to accurately predict the turbulent kinetic energy content. Hybrid RANS/LES computations, using a second-order low-dissipation optimization-based gradient reconstruction scheme on a grid that is three times coarser than the DNS, provide an accurate prediction of the far-field noise levels, except in the wake region, where they are overpredictive. [ABSTRACT FROM AUTHOR]

The conventional sonic boom propagation prediction method widely adopted in supersonic aircraft design involves a two-step procedure. In the first step, a compressible viscous or inviscid Computational Fluid Dynamics analysis is applied to the aircraft geometry at flight conditions to produce a flow field solution near the aircraft. Then in the second step, a one dimensional nonlinear Burgers' equation model is used to propagate sonic boom signature traces from the near field solution at flight altitude to the ground along a ray path. For an accurate ground signature prediction the near field signature must be accurately modeled at a sufficient distance from the aircraft flight path in order to minimize errors in the one dimensional propagation model. This is a very challenging task for general purpose CFD tools in a design environment because the cost of maintaining highly accurate off-body solutions increases dramatically as the radial distance is enlarged in the computational domain. It is also particularly difficult to apply these tools for wave propagation because the algorithms are normally lower order and numerically dissipative and dispersive. In this work a space marching procedure based on an optimized higher order finite difference method is developed and applied in conjunction with a CFD solution concentrated in the close vicinity of the aircraft. This new approach is much more efficient, compared to previous methods, in providing highly accurate near field signatures for full carpet ground predictions. Results indicate that near field signatures retain more waveform shape information farther from the aircraft geometry while reducing the CFD cost significantly. The predicted ground signature is also shown to converge in shape as the radial distance of the near field signature grows, which is indicative of a more ideal initial condition being supplied to the one dimensional wave propagation to ground. This feature is very difficult to replicate in a tractable manner with common CFD approaches used in design. [ABSTRACT FROM AUTHOR]

It has been well-known that the Ingard-Myers impedance condition, while simple to apply, is subject to the hydrodynamic Kelvin-Helmholtz-type instability due to its use of a vortex sheet in modeling the flow at the liner boundary. Recently, in the development of a time domain boundary element method for acoustic scattering by treated surfaces, it was found that by neglecting a certain second-order spatial derivative term in the Ingard-Myers formulation, the hydrodynamic instability can be avoided. The present paper aims to provide further analysis of this modified condition, hereby referred to as the Truncated Ingard-Myers Impedance Boundary Condition (TIMIBC). It will be shown, based on the dispersion relations of linear waves, that the instability intrinsic to the Ingard-Myers condition is eliminated in the proposed new formulation. Quantitative assessments on the accuracy of the TIMIBC for scattering of acoustic waves by lined surfaces are carried out, and its effectiveness is demonstrated by a numerical example. It is found that the TIMIBC provides a good approximation to the original Ingard-Myers condition for flows of low to mid subsonic Mach numbers. As such, the proposed TIMIBC can offer a practical solution for overcoming the intrinsic instability associated with the Ingard-Myers condition. Moreover, time domain implementation of the TIMIBC is also discussed and illustrated with a numerical example using a finite difference scheme. In particular, a minimization procedure for finding the poles and coefficients of a broadband multipole expansion for the impedance function is formulated by which, unlike the commonly used vector-fitting method, passivity of the model is ensured. [ABSTRACT FROM AUTHOR]

This introduction provides a brief personal background and highlights the scientific accomplishments of Professor Christopher Tam. He has maintained extremely high standards of research and an enviable track record for productivity, spanning more than five decades. The depth and breadth of his technical output are truly breathtaking, and he has inspired scores of researchers and graduate students. When the first author was an Acoustics Engineer at Lockheed, Georgia, he started collaborating with Chris in the early 1990s. There were several joint projects when the author moved to Boeing, which encompassed both fundamental research and applications. The second author has an even longer association with Chris, initially as a PhD advisee in the late 1980s and then as a co-researcher. In this special volume we pay tribute to Chris's many achievements, an endeavor that provides us with great satisfaction and which is long overdue. [ABSTRACT FROM AUTHOR]

Toroidal propellers, a new type of drone propellers capable of significantly reducing noise, offer new possibilities for future low-altitude flying platforms. In this study, a numerical model was established to analyze the aerodynamic noise of the toroidal propeller under normal atmospheric conditions. The aerodynamic calculations for the toroidal and benchmark propellers were performed to obtain noise source information using a transient large eddy simulation. The hybrid computational aeroacoustic method was employed to calculate the noise spectrum at different speeds and locations. Moreover, an experimental system for propellers was designed and built within an anechoic chamber to investigate the aerodynamic performance and noise characteristics at various speeds. The noise reduction effect of the toroidal propeller compared to the benchmark propeller was analyzed, taking into account the sound pressure level, lift coefficient, and figure of merit. At the same thrust level, the lift coefficient of the toroidal propeller increased by 187% relative to the benchmark propeller. The radial and axial sound pressure levels decreased by 5.2 dB(A) and 19.6 dB(A), respectively. The toroidal propeller significantly reduced noise while improving aerodynamic performance. This study provides a theoretical basis, experimental methods, and data support for the aerodynamic and noise calculations of toroidal propellers. It has significant engineering implications for the development of low-noise propellers for drones.

This study investigates the influence of forebody configuration on aerodynamic noise generation and radiation in standard squareback vehicles, employing a hybrid computational aeroacoustics approach. Initially, a widely used standard squareback body is employed to establish grid-independent solutions and validate the applied methodology against previously published experimental data. Six distinct configurations are examined, consisting of three bodies with A-pillars and three without A-pillars. Throughout these configurations, the reference area, length, and height remain consistent, while systematic alterations to the forebody are implemented. The findings reveal that changes in the forebody design exert a substantial influence on both the overall aerodynamics and aeroacoustics performance of the vehicle. Notably, bodies without A-pillars exhibit a significant reduction in downforce compared to their A-pillar counterparts. For all configurations, the flow characteristics around the side-view mirror and the side window exhibit an asymmetrical horseshoe vortex with high-intensity pressure fluctuations, primarily within the confines of this vortex and the mirror wake. Side windows on bodies with A-pillars experience more pronounced pressure fluctuations, rendering these configurations distinctly impactful in terms of radiated noise. However, despite forebody-induced variations in pressure fluctuations impacting the side window and side-view mirror, the fundamental structure of the radiated noise remains relatively consistent. The noise pattern transitions from a cardioid-like shape to a monopole-like pattern as the probing distance from the vehicle increases. [ABSTRACT FROM AUTHOR]

The present paper is aimed at providing an updated review of prediction methods for the aerodynamic noise of ducted rotor–stator stages. Indeed, ducted rotating-blade technologies are in continuous evolution and are increasingly used for aeronautical propulsion units, power generation and air conditioning systems. Different needs are faced from the early design stage to the final definition of a machine. Fast-running, approximate analytical approaches and high-fidelity numerical simulations are considered the best-suited tools for each, respectively. Recent advances are discussed, with emphasis on their pros and cons. [ABSTRACT FROM AUTHOR]

In recent years, to predict the performance and the aerodynamic sound of box fans with high accuracy, numerical analysis has been necessary for the transitional boundary layer on the blades. In this study, two types of box fans with a baseline fan and a high-load fan, where the boundary layer that develops on the suction surface of the rotor blade transitions to turbulent flow, were targeted for large eddy simulation (LES) analysis and acoustic analysis using different grid resolutions to predict the far-field sound. The box fan in this study has an impeller of 180 mm outer diameter and five blades, and the Reynolds number based on the diameter and the tip velocity is 3.38 × 105. The decoupled method with LES for incompressible flow and computational aeroacoustics (CAA) based on a dipole sound source was performed. The computational grid resolutions had 25 million grid points (Case 1), 52–56 million grid points (Case 2 and Case 4), and 420–450 million grid points (Case 3 and Case 5). The performance of the fans and the sound pressure level (SPL) were compared with each case and experiment. The performance was in good agreement with the experiment in higher grid resolution cases. Boundary layer transition on the blade surface was predicted as the main source of aerodynamic sound generated from the box fan. Increasing the grid resolution improved the prediction accuracy of the SPL in the frequency range corresponding to the scale of the turbulent eddies captured by the grid resolution.

The results of the numerical investigation of the tone noise of the first booster stage of a high bypass ratio turbofan in the far field in the approach (landing) operational conditions are presented. The study is performed using the frequency domain numerical method of multistage turbomachines tonal noise simulation, developed at the Baranov Central Institute of Aviation Motors (CIAM). A variant of the method suitable for the calculations in a nonlinear setup is used. The calculation is performed for several blade passages in each row. The harmonic balance method for multitonal disturbances with a frequency mapping is implemented. The results of the calculations are compared with the results of the calculations in the time domain, presented in the previous paper (they were also performed in a nonlinear setup) and with the experimental data obtained at the CIAM acoustic test facility. In general, satisfactory correspondence is found between the data in both cases. The results can be treated as an argument for the validity of the proposed nonlinear frequency domain method of the multistage turbomachinery tone noise calculations. [ABSTRACT FROM AUTHOR]

Research is carried out on the prediction of aircraft surface noise during flight with high maneuverability and high overload. The dual hybrid Reynolds-averaged Navier-Stokes/large eddy simulation (RANS/LES) calculation model based on zonal mixing and turbulent scale mixing is modified, and the new hybrid model is established to simulate complex separated flow accurately. A calculation method for simultaneously calculating aerodynamic/motion, internal/external flow, and compressible/incompressible flow is employed. The dual hybrid RANS/LES model combined with six-degree-of-freedom motion equations is used to simulate the maneuvering process of the aircraft and calculate the surface acoustic load. The calculation results reveal that the acoustic load at the inlet of the engine and tail nozzle is large. During positive overload flight, the surface noise is greater than negative overload due to the stronger turbulent fluctuation on the aircraft surface and the impact of jet noise at the tail nozzle. The contribution rate of jet noise to the overall sound pressure level (OASPL) of the fuselage gradually increased from front to rear. In the maneuvering state, the vortex motion on the aircraft surface is enhanced, and the pressure fluctuation is stronger than that in the cruise state, which makes the surface acoustic load greater than that in the cruise state. [ABSTRACT FROM AUTHOR]

Numerical simulations are performed for flows over a circular cylinder at a Reynolds number ranging from 150 to 5000, and Mach number of 0.2, to assess the predictive capability of URANS and hybrid RANS/LES for acoustic waves generation and propagation. Complementary direct numerical simulations (DNS) are performed to generate validation datasets and to provide more insight into the problem. DNS predictions show that noise induced by the vortex shedding is radiated primarily at a 90-degree angle with respect to the wake direction and dictates the dominant frequency of the sound pressure waves. Turbulence dominates the noise in the near-field wake, resulting in a broadband pressure spectrum. URANS both underpredicts and overpredicts the noise levels in the wake region and in the direction normal to the freestream flow, respectively, which is attributed to its inability to accurately predict the turbulent kinetic energy content. Hybrid RANS/LES computations, using a second-order low-dissipation optimization-based gradient reconstruction scheme on a grid that is three times coarser than the DNS, provide an accurate prediction of the far-field noise levels, except in the wake region, where they are overpredictive.

In order to obtain the influence mechanism of the rotor noise and the local aerodynamic variation under different operating conditions, a novel discrete noise analysis strategy for the acoustic source localization is established. The analysis strategy has two aspects including the blade division method and the noise contribution calculation method. Firstly, the body-fitted rotor blade grids are preprocessed and refined at the division position before the flowfield simulation. Then, based on the flowfield result, the blade grids are divided into several blocks in the chordwise and spanwise directions, and the acoustic pressure of each block is predicted. Finally, a discriminant function for the contribution of the block to the rotor noise is proposed, and the acoustic source localization is carried out. The URANS equations and FW-H equations are used to capture high-fidelity flowfield and rotor acoustic pressure. Parameters such as the block position in different direction on the rotor blade and the collective pitches are quantified, and the relationship between air flow and aeroacoustics is discussed in detail. Some conclusions are obtained by analyzing the BO105 model rotor in hover. The acoustic pressure produced by the leading edge of the upper surface could cancel about 50% acoustic pressure of the remainder of these blocks. Increasing the force at this position will be benefit to the noise reduction. Acoustic source distribution is closely linked to the flow separation near the blade tip: the main acoustic source moves toward around 0.9 R section of the blade in the spanwise direction, and it moves from the leading edge towards the trailing edge in the chordwise direction. [ABSTRACT FROM AUTHOR]

The present paper is aimed at providing an updated review of prediction methods for the aerodynamic noise of ducted rotor–stator stages. Indeed, ducted rotating-blade technologies are in continuous evolution and are increasingly used for aeronautical propulsion units, power generation and air conditioning systems. Different needs are faced from the early design stage to the final definition of a machine. Fast-running, approximate analytical approaches and high-fidelity numerical simulations are considered the best-suited tools for each, respectively. Recent advances are discussed, with emphasis on their pros and cons.

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In this study, the concept of a zero mass flow liner is evaluated. The concept enables impedance control by the induction of, acoustically actuated, periodic bias flow through the facing sheet of the liner. By means of the periodic bias flow, the impedance of the liner is adapted to different grazing flow conditions. The equivalent fluid impedance model for perforated plates is modified to account for the effects of periodic bias and grazing flow. A generally applicable optimization routine, using the impedance of the lined surface as a boundary condition in a numeric calculation, is implemented. Based on the results of the optimization, a zero mass flow liner is manufactured and evaluated experimentally. The damping characteristics are assessed in the form of dissipated energy along the lined surface. Prediction and measurements show reasonable agreement. The zero mass flow liner delivers broadband dissipation of high peak value over a range of grazing flow Mach numbers. Under grazing flow, the effect of periodic bias flow is reduced. For a ratio of grazing to bias flow velocities larger than five, no appreciable effect is found. This poses considerable energy requirements on the actuation source for the application in high-Mach-number flow regimes. [ABSTRACT FROM AUTHOR]

• Original expressions are proposed for the gradient of the radiated pressure. • Efficient algorithms for acoustic calculations using integral methods are detailed. • The integration techniques can handle grids in supersonic motion. • The order of accuracy of the acoustic calculations is precisely quantified. Starting from the Ffowcs Williams-Hawkings surface integral formulation for a moving medium, the article proposes rather simple expressions for the radiated pressure and its gradient in the time domain, that are valid for solid or porous, fixed or moving integration surfaces. Moreover, these original expressions allow calculations with integration surfaces in supersonic motion (such as rotating surfaces around propellers or rotors). Versions dedicated to fixed integration surfaces are also proposed. The usual locally compact and the fully non-compact integration techniques are recalled, with, for both, a detailed description of efficient calculation algorithms. Particular attention is devoted to the order of precision of the calculations. The time derivation techniques and integration schemes used in this study lead to a theoretical second order that can easily be increased for the locally compact integration method. The results of these expressions and integration methods are compared to the analytical solution for the case of a fixed monopole and for that of a rotating monopole. They clearly show the benefit of the direct gradient calculation over a calculation by finite differences of the pressure around the observation point, particularly for broadband signals. [ABSTRACT FROM AUTHOR]

The article presents a fairly simple way to take solid bodies into account in acoustic radiation calculations using integral methods, while still using the free-space Green’s function. The approach is based on the Kirchhoff method and on a locally plane reflection assumption. It can be applied to both analytical noise sources and acoustic disturbances provided by numerical simulations, to fixed or mobile noise sources, concentrated or widely spread in a moving medium. The time-domain formulation is an important advantage for periodic signals rich in harmonics (rotors or propellers impulsive noise) and for broadband signals (profile or jet noise). The formulation and calculation algorithm are described in detail. The method’s accuracy and limitations are shown first by comparing the results with analytical solutions for the acoustic scattering of a point source by a sphere, for a fluid at rest. An application example is then given for a wing in a Mach 0.5 flow, and the results are compared with the numerical solution of the linearized Euler equations, in the presence of a mean flow. In addition, the article proposes expressions for direct calculation of the pressure gradient by Kirchhoff and Ffowcs Williams-Hawkings surface formulations.

In this study, three voluteless centrifugal fans are compared for their aeroacoustic performances. The tonal noise is predicted by coupling the IDDES with Formulation 1A of Farassat. The sources of the tonal noise at the blade passing frequency ( B P F ) are identified. It is found that the sources are related to the fan inlet gap, which introduces higher velocity intensities and turbulent fluctuations interacting with the blade leading edge. By redesigning the gap, the tonal noise at the B P F is reduced effectively. [ABSTRACT FROM AUTHOR]

The sound generated by a vortex pair moving close to a rectangular cylinder is studied by high-resolution simulations of the two-dimensional unsteady compressible Navier-Stokes equations at Mach numbers ranging from 0.05 to 0.3. In this study, the Mach number Ma is defined by Ma=Vt /c0, where Vt is the translation velocity of the co-moving vortex pair and c0 denotes the speed of sound. The directly computed far-field sound is compared to the prediction of the acoustic analogy proposed by Lighthill (1952) and Curle (1955). The prediction is in excellent agreement with the simulation, where the two-dimensional form of Curle’s modified solution is adopted for an accurate calculation of the sound waves with a phase difference between various vortices around the cylinder. It could be confirmed that the scaling law in two dimensions holds in this simulation. In this study, the sound generation due to the dipole and quadrupole fields is investigated by calculating the two terms in Curle’s analogy. Examining the surface distribution of dipoles on the cylinder indicates that the force exerted on the fluid by the solid boundaries of the side/upper wall of the rectangular cylinder generates the sound when the co-translating vortices approach /pass the rectangular cylinder, which is the sound due to dipoles corresponding to the directivity of the sound in the polar diagrams. The area distribution of quadrupoles which is the second time derivative of the area integral of Lighthill’s stress tensor indicates that the sound is emitted from a vortex generated at the corner of the rectangular cylinder when the vortex pair approaches the cylinder and where the interaction between many vortices around the cylinder and the co-translating vortices generates the sound when the co-translating vortices pass the cylinder.

This paper proposes the development of a mechanism to minimise the assembly time for a multi-headed gantry and high-speed surface mounting technology machine by determining the component assignment to feeder slots. Since a gantry moves long distances in order to pick components, place them on the board and then return them to the feeder slots, we classified the overall assembly time according to the different movements of a gantry. The overall assembly time is then minimised by presenting a new heuristic for the feeder assignment, consisting of Nearest Component Allocation (NCA) and Globally Updated Assignment (GUA). NCA contains information about how each component type locates closely to others on the board. Using the solution from NCA, the component distance function calculates the most representative distance between component types. Then, GUA is applied to improve the NCA solution. The experiments consist of several printed circuit boards with numbers of component types and points to be placed. Highlights of this paper are that: a classification of the gantry movements is proposed based on the average speed; a heuristic NCA-GUA for feeder assignment is developed by considering the placements on the board; the computational time is greatly reduced by NCA-GUA without degrading the solution quality; and a decision process for nozzle assignment is proposed to minimise the overall assembly time. The results show how NCA and GUA affect the final results, and how this mechanism leads to better performance than a genetic algorithms or 2-opt swap search. This comparison proves that our method provides competitive and effective solutions in terms of minimising the overall assembly time. [ABSTRACT FROM AUTHOR]

• Hybrid approach using standard and finite-volume lattice Boltzmann methods. • Efficient aeroacoustic simulation for low-speed turbomachines. • Computations on Cartesian and body-fitted grids with the direct forcing method. • The hybrid method simulated flow and acoustic fields around a crossflow fan. In this study, we developed a hybrid approach using the lattice Boltzmann method (LBM) and finite-volume lattice Boltzmann method (FVLBM) to perform efficient aeroacoustic simulations with moving boundaries. An entire domain, including the flow and acoustic fields, was computed using the standard LBM with a Cartesian grid. Local domains around the moving objects were computed using the FVLBM with body-fitted grids. These simulations were coupled using the direct forcing method to consider moving boundaries. The hybrid method was validated for several problems including turbulent flows and flow-induced sounds under low-Mach-number conditions. These validation problems covered flows with Reynolds numbers up to 2.0 × 10 5 and Mach numbers less than 0.2. In the simulations of the aeolian tone generated from stationary and rotating cylinders, the hybrid method results were consistent with those of the conventional methods. In the simulation of the turbulent flow and broadband sound of the isolated airfoil, the hybrid method was 15 times faster than the standard LBM and produced results consistent with the experimental results. Furthermore, the application of a cross-flow fan demonstrated that the hybrid method successfully simulated the flow and acoustic fields around the rotor. [ABSTRACT FROM AUTHOR]

The co-rotating vortex pair arrangement is a frequently applied test case in computational aeroacoustics. Its flow quantities and far-field radiated sound pressure can be computed analytically, assuming delta-distribution potential vortices. When this configuration is used for verification in a numerical computation, the singular vortex cores cannot be modeled directly; they have to be desingularized, which affects the associated flow and acoustical fields. Closed-form expressions for the aeroacoustic source terms are derived for the Scully vortex model and used for detailed numerical computations of the radiated sound pressure. The effects of desingularization and finite source region size are analyzed in the frequency domain for Lighthill's equation and the perturbed convective wave equation by means of an efficient numerical approach that computes the convolution of the aeroacoustic source terms and the 2D free-field Green's function. It is shown that the deviations of the far-field radiated sound pressure from the idealized analytical solution depend on the Mach number, vortex core radius, and truncation limit of the source region. A significant increase in the effective size of the source region due to the desingularization is demonstrated. Minimal error bounds compared to the analytic far-field pressure solution are established. The presented converged numerical results serve as reference solutions for validating implementations of hybrid computational aeroacoustic workflows. • Analytical expressions for aeroacoustic source terms of the desingularized vortex pair. • Revised, ab initio deduction of the far-field sound pressure of the co-rotating vortex pair. • Efficient numerical strategy for performing the convolution integral with the Green's function. • Convergence of Lighthill's analogy and the PCWE are analyzed and compared. • Lower error bounds are determined for various Mach numbers and vortex radii. [ABSTRACT FROM AUTHOR]

This paper examines the modifications in the flow and sound fields due to the application of a periodic suction-blowing excitation strip on the top and bottom surfaces of a NACA0012 airfoil. The flow and sound fields have been directly computed (D N S approach) by solving unsteady, two-dimensional compressible Navier–Stokes equations using physical dispersion relation preserving schemes. The flow approaches the airfoil at a Reynolds number of 5000 and 5 o angle of attack. In the no-excitation case, the sound is mainly created by the lift and drag dipole equivalent sources associated with vortex shedding. However, the application of suction-blowing excitation results in two additional sound sources: monopole-equivalent sound sources due to fluid movement in and out of the airfoil surface and the drag dipole-equivalent sound sources due to periodic change in the airfoil's effective cross-sectional area. The combined effects of these sound sources result in significant mitigation of sound on the suction side of the airfoil. The excitation amplitude and forcing frequency effects on sound field modifications have been studied in depth. Beats of sound have been observed when the forcing frequency is close to the excitation frequency. • Direct Numerical Simulation approach is used for Navier–Stokes equations. • Suction-blowing excitation induces monopole and drag dipole sound sources. • Excitation provided near the leading edge increases the lift-to-drag ratio. • Bias in sound directivity is observed for excitation near the trailing edge. • Sound beat is observed for frequencies close to vortex shedding frequency. [ABSTRACT FROM AUTHOR]

In the last decades flow simulations have become a routine practice in many industrial fields for the aerodynamic and noise prediction. Moreover, the ever increasing interest in simulating off‐design operating conditions promoted the development of high‐fidelity simulation tools to overcome the modeling and accuracy limits of standard industrial codes in predicting turbulent separated flows. The discontinuous Galerkin (DG) method is well suited for this class of simulations, but today DG‐based CFD and CAA (Computational AeroAcoustics) solvers cannot yet reach the computational efficiency of well‐established commercial codes. As a consequence, the present paper aims at exploiting some attractive strategies, such as the adaptation of the elemental polynomial degree (p‐adaptation) and of the degree of exactness of quadrature rules, to enhance the computational efficiency of an implicit DG platform for CFD and CAA simulations. Moreover, a sponge layer non‐reflecting boundary treatment has been also implemented for CAA. The predicting capabilities of the method have been assessed on classical CAA and CFD test cases. The proposed adaptive strategy guarantees a significant reduction (≈50%) of the computational effort for both CFD and CAA simulations, compared to uniform‐order discretizations, while not spoiling the high accuracy requested to an high‐fidelity simulation tool. [ABSTRACT FROM AUTHOR]

This work is related to the investigation of innovative stator designs aiming to reduce the dominant interaction noise in aeroengines. The study of turbulent structures definition is crucial for the accurate prediction of broadband noise radiation from passive treatments, as leading-edge serrations studied here. A modified Fourier modes-based methodology is proposed to obtain a fully three-dimensional incompressible turbulence field, while taking into account periodic and wall-boundary conditions. A low-noise geometry is examined along with the reference profile on a rectilinear seven-vane cascade rig using a hybrid computational fluid dynamics/computational aeroacoustics method. Numerically assessed noise reductions from the serrated airfoils are favorably compared with an analytical solution and a semi-empirical law. An overall sound power-level reduction around 4 to 6 dB is obtained at three acoustic certification points. Finally, the aerodynamic performances are also evaluated through Reynolds-averaged Navier-Stokes computations, and an improved variant of the initial treatment is proposed, allowing for acceptable penalties at the aerodynamic design point. [ABSTRACT FROM AUTHOR]

A linear lattice Boltzmann flux solver (LLBFS) is developed based on the high-precision finite-volume method (FVM) and is used to simulate problems involving acoustic propagation. The least-squares-based finite difference (LSFD) is used to construct the high-precision FVM, which discretizes the linear Navier–Stokes equations (LNSE). In terms of the flux calculation for the interface of the control cell, the LLBFS is derived by comparing the macroscopic and microscopic variables through the Chapman–Enskog (C-E) multi-scale analysis, which establishes the relationship between the linear lattice Boltzmann equation (LBE) and the LNSE. The algorithm has the following advantages. First, the proposed LLBFS inherits the advantages of the lattice Boltzmann flux solver (LBFS), i.e., the local solution of the lattice Boltzmann equation is used to calculate the fluxes; handling the inviscid and viscous fluxes simultaneously makes for a more consistent algorithm and it can be expanded to the multi-dimensional scheme without approximation. Second, solving the LNSE using the high-precision FVM on unstructured grids, which enables problems involving acoustic waves interacting with complex geometries can be simulated conveniently. To validate the accuracy and robustness of the proposed method, several cases are simulated, including acoustic waves propagating in an inviscid and viscous mean flow and interacting with complex geometries. [ABSTRACT FROM AUTHOR]

The installation of muzzle brake changes the flow direction of propellant gas, which leads to a decline in the recoil force but an increase in the muzzle flow noise of the side rear. Therefore, the coordination between high efficiency and low harm of muzzle brake is one of the factors limiting its optimization. To address this problem, a numerical method for evaluating the comprehensive performance of muzzle brake targeting muzzle brake efficiency and impulse noise is proposed in the present study. The two performance indexes are calculated using a computational fluid dynamics (CFD)-computational aeroacoustics (CAA) coupled method. Afterwards, a corresponding experiment is conducted to verify its feasibility. Furthermore, the comprehensive performance of muzzle brakes with three different structures are analysed based on the two indexes obtained by the proposed numerical method. The results reveal that both indexes can be influenced by the structure of muzzle brakes. The increased braking efficiency is almost accompanied by increased impulse noise, but the functional relation is not linear dependence. Thus, the comprehensive performance can be improved by optimization of the structure. The evaluation method proposed in this paper has theoretical significance for optimizing muzzle brakes, which is of great engineering value. [ABSTRACT FROM AUTHOR]

The integral formulation of the Ffowcs-Williams and Hawkings (FW--H) analogy, developed by Farassat (known as Farassat's formulation 1A), is implemented to study the sound generation and propagation of rotating slender bodies. The general post-processing numerical code utilizes the linear acoustic theory to predict the thickness and loading noise terms for bodies in subsonic motion. The developed numerical code is validated for elementary acoustic sources (rotating monopole and dipole) against analytical solutions. The validated code is then applied for prediction of lowfrequency in-plane harmonic noise (LF-IPH) of a model helicopter rotor of Sargent and Schmitz in a low-thrust hover with full-scale tip Mach number. The required loading distribution of the rotor blade is obtained with CFD (RANS) and Blade Element Momentum Theory (BEMT) methods and also validated against literature data. The developed acoustic code, supplemented by CFD and BEMT loading analyses, allows for a detailed comparison (thickness and loading, near- and far-field, etc.) of the LFIPH noise of a helicopter rotor in both, time and frequency domains. The predicted (FW--H) acoustic signals are compared not only with the reference code solutions, but also with the experimental data. Moreover, the paper quantifies the impact of computational grid density and time-step size (used by CFD and FW--H codes) on the final solution accuracy. Additionally, a simplified analytical code is developed (based on elementary dipole solutions, compact chord assumption and BEMT method) allowing for the initial loading noise analysis with highly reduced computational resources. The acquired results are fully compatible with the classical FW--H analysis in terms of the impact of the in-plane and out-of-plane forces on the generated noise. The FW--H code predictions of the acoustic pressure and its components are in satisfactory agreement with the reference and experimental data of Sargent and Schmitz. [ABSTRACT FROM AUTHOR]

Comparing acoustic simulations against experimental data is an essential step in order to prove the correctness of numerical tools. This can be done with wind tunnel experiments where the environmental conditions can be adjusted very accurately. Ultimately, the tools must be capable of predicting real-word scenarios like aircraft flyovers. However, obtaining precise data from flyover experiments is challenging and often important input data is missing. The current paper shows, that by extracting the shielding effect of a small detail, a deflecting flap of an aircraft with rear-mounted engines, it is possible to reproduce flyover measurements with a boundary element method, even when only little engine information is known. The boundary element method can only take a constant mean flow into account, but by additionally evaluating results of a volume-resolved discontinuous Galerkin method more insights into the expected effects of a realistic mean flow is given. [ABSTRACT FROM AUTHOR]

The paper presents validation results of a hybrid simulation method for aeroacoustics in turbulent flows at low Mach numbers. The hybrid method implemented in the Ansys Fluent® CFD package applies a scale-resolving turbulence model to compute the noise sources in an incompressible flow, while the noise propagation is modeled by a wave equation formulated for the acoustic potential. The selected test case deals with a flow and a sound field around a generic side view mirror of a car. The SBES model by Menter, which belongs to the class of the RANS-LES models, is used for the flow simulation. It switches to the Large Eddy Simulation (LES) mode in separated mixing layers and recirculation zone behind the mirror as well as in the following wake, where flow develops intensive turbulence and dominating noise sources. The acoustics wave equation is formulated in the model form of Kaltenbacher et al. and is applied in the time domain. The overall calculation is performed as a transient co-simulation on the same mesh using the finite-volume discretization method for both the flow and the acoustics parts. The wave equation is advanced in time using the HHT-α method. Obtained distribution of the mean wall pressure over the mirror surface closely matches the experimental one. Rich content of the resolved turbulent vortices in the separation zone and good agreement of the calculated and measured wall pressure spectra at sensor locations downstream the mirror evidence a proper LES resolution quality of noise sources. Comparison of the computed noise spectra at the remote microphones with the experimental data demonstrates the sound propagation accuracy and validates the overall aeroacoustics simulation method. [ABSTRACT FROM AUTHOR]

The present work deals with the modeling of the aerodynamic sound generated by the propellers of small-size drones, taking into account the effects of horizontal forward flight with negative pitch and of installation on supporting struts. Analytical aeroacoustic formulations are used, dedicated to the loading noise. The fluctuating lift forces on the blades are expanded as circular distributions of acoustic dipoles, the radiated field of which is calculated by using the free-space Green’s function. This provides descriptions of the sound field, valid in the entire space. The stationary mean-flow distortions responsible for the lift fluctuations and at the origin of the sound are estimated from existing numerical flow simulations and from ad hoc models. Installation and forward-flight effects are found to generate much more sound than the steady loading on the blades associated with thrust. Therefore, the models are believed reliable fast-running tools that could be used for preliminary low-noise design through repeated parametric calculations, or for noise-impact estimates corresponding to prescribed urban traffic.

Optimization design was widely used in the high-lift device design process, and the aeroacoustic reduction characteristic is an important objective of the optimization. The aerodynamic and aeroacoustic study on the baseline wing and flap configuration was performed numerically. In the current study, the three-dimensional Large Eddy Simulation (LES) equations coupled with dynamic Smagorinsky subgrid model and Ffowcs–William and Hawkings (FW–H) equation are employed to simulate the flow fields and carry out acoustic analogy. The numerical results show reasonable agreement with the experimental data. Further, the particle swarm optimization algorithm coupled with the Kriging surrogate model was employed to determine optimum location of the flap deposition. The Latin hypercube method is used for the generation of initial samples for optimization. In addition, the relationship between the design variables and the objective functions are obtained using the optimization sample points. The optimized maximum overall sound pressure level (OASPL) of far-field noise decreases by 3.99 dB with a loss of lift-drag ratio (L/D) of less than 1%. Meanwhile, the optimized performances are in good and reasonable agreement with the numerical predictions. The findings provide suggestions for the low-noise and high-lift configuration design and application in high-lift devices. [ABSTRACT FROM AUTHOR]

A method for calculating the propagation of unsteady disturbances along the duct of a turbofan engine, taking into account nonlinear effects, is presented. It is based on finding a solution in the form of a finite set of circumferential modes and performing calculations for the complex amplitudes of these modes using the harmonic balance method. The use of the method can speed up the calculation of the propagation of disturbances through the air intake and the nozzle of a turbofan engine in the case when the influence of nonlinear effects cannot be neglected. The method under consideration is implemented in the Central Institute of Aviation Motors (CIAM) 3 Dimensional Acoustics Solver (3DAS) software package. It is tested on the problem of calculating the nonlinear interaction of acoustic modes in a 2D cylindrical channel. The calculation results using this method are in close agreement with the results of the calculation performed in the time domain. [ABSTRACT FROM AUTHOR]