High-speed Impulsive Noise Research Articles

Numerical investigation on the use of various blade tips for the helicopter's main rotor blade in forward flight regarding aerodynamic performance and HSI noise considerations

The goal of this research is to investigate aerodynamic performance and high-speed impulsive (HSI) noise of various blade tips for the helicopter's main rotor in forward flight. For this purpose, three-dimensional compressible flow field around the rotor blades is numerically simulated by the solution of the unsteady Reynolds averaged Navier-Stokes (URANS) equations with the "SST k-ω" turbulence model. Solution is quantified by the calculation of torque, drag, and thrust coefficients of CQ, CD, CT, and L/D ratio. HSI noise on the rotor plane is assessed by the calculation of sound pressure level at certain points of the flow field using the Ffowcs Williams-Hawking (FW-H) equation. Blade tips are eleven and include anhedral, dihedral, Eagle and combinations of them, and also Blue Edge and rectangular. The combinations are inspired by the tip configurations of the fixed wings. In this regard effects of the tip geometry parameters including curvature and height of its parts have been considered as well. All of the tips are examined for two flight conditions of Caradonna and Tung two-blade rotor with fixed pitch angle and time-varying pitch angles. Based on the present results, it is found that blade tips of dihedral family do not improve the aerodynamic performance of the helicopter blade. Also, addition of anhedral part to the blade tip improves the L/D ratio. According to the present acoustic analysis, it was seen that after the Blue Edge, the Eagle tip stands in order of priority. For the condition of Caradonna and Tung two-blade rotor with tip Mach number of 0.8 and advance ratio of 0.2, HSI noise of the Eagle-tip blade decreases by 2.39 % compared to that of the two-blade rotor with rectangular tip. However, the aerodynamics performance of the Eagle tip is much better than that of the Blue Edge.

A new blade tip geometry to improve aerodynamic performance and acoustic noise of helicopter blade in hovering flight

The goal of the present study is to improve aerodynamic performance of the helicopter main rotor by geometric redesign of its blades' tips. In this process, the generated noise is examined as well to ensure that it is not increased by the new tip geometry. Three-dimensional flow field around the blades is numerically simulated by the solution of the compressible unsteady Reynolds averaged Navier-Stokes (URANS) equations to calculate torque and thrust coefficients of CQ and CT, and the figure of merit (FoM) ratio. The turbulence model used in this study is “SST k-ω”. To check both the thickness noise and the high-speed impulsive (HSI) noise on the rotor plane, the Ffowcs Williams-Hawking (FW-H) equation is solved to obtain the sound pressure level at certain points in the flow field. Following the noticeable aerodynamic improvements that anhedral and dihedral winglets made on the fixed-wings, aerodynamic performance of the blade tip geometries composed of anhedral and dihedral winglets or a combination of them are investigated on the helicopter blades in present research. Redesign of blade tips minimizes the geometric changes needed for a new blade design. Flow fields around blades with nine different tip geometries are three-dimensionally simulated and analyzed in detail at crucial rotational speeds of 1250 rpm and 2540 rpm in hovering flight. To determine the amount of improvement, results obtained for different tip geometries are compared with the reference results of the rectangular blade. Among all of the different tip geometries, it was concluded that Eagle and Anhedral-Eagle make maximum improvements in aerodynamic and aeroacoustic of the main rotor blades. The FoM ratios of these two tip geometries have increased even up to 17% and 21%, respectively. Also, their noise levels of them have decreased by about 4.5% and 1%.

Numerical analysis on the high-speed impulsive noise propagation characteristic of helicopter rotor in the presence of strong shear flow

Based on the Linearized Euler Equations (LEE) and the Computational Fluid Dynamics (CFD) technique, the propagation characteristics of helicopter rotor High-Speed Impulsive (HSI) noise in non-uniform flowfield are simulated and analyzed. A Hybrid Computational Aeroacoustics (HCAA) method for solving the sound field of rotor is first developed. The acoustic source region is simulated by solving Navier-Stokes equations, and the acoustic near-field is simulated by solving LEE based on the Runge-Kutta Discontinuous Galerkin (RKDG) method. The numerical analysis method is then validated through comparisons with experimental data, and the aeroacoustics characteristics of the UH-1H rotor are calculated. Afterwards, the propagation characteristics of HSI noise in the wind tunnel are calculated, and the “trap” effect of the shear layer on HSI noise is determined. Finally, the peak negative pressure, time-averaged acoustic energy, and instantaneous perturbed pressure field are analyzed in detail, and some conclusions are drawn. In jet flow, the HSI noise energy is reinforced, and the time of the peak negative pressure is significantly advanced. In still air, the HSI noise energy is attenuated, and the time of the peak negative pressure is slightly delayed. These investigations offer valuable data for developing advanced theoretical correction method to consider the effect of shear layer on rotating acoustic source.

Efficient Methodology with Grid Configurations for HSI Noise Prediction in Hover

Complex and unsteady flow patterns occur around rotating rotor blades owing to the generation of shock waves at the blade tip and interference between the blade and tip wake when the blade rotates at high speeds, thereby resulting in specific noise characteristics. Furthermore, these noise characteristics cause noise pollution in private areas and degrade detectability of helicopters in military applications. Therefore, it is important to accurately analyze noise characteristics around rotor blades. In this study, a novel computational methodology with a chimera wake grid was proposed to efficiently and accurately predict high-speed impulsive (HSI) noise due to the shock wave at the blade tip. The proposed method enables accurate analysis of a complex flow region by overlapping a wake grid in the vicinity of shock waves. A 1/7-scaled UH-1H rotor blade was used to compare and verify the accuracy of HSI noise prediction and computational efficiency. The chimera grid method was applied to the present simulation for considering the blade motion and moving effects. A permeable surface for wrapping the surface of the physical blade was constructed to include quadrupole noise source generated from the control volume, thickness noise source, and loading noise source. Furthermore, the permeable Ffowcs Williams and Hawkings (FW–H) equations were used with the Kirchhoff approach to realize efficient far-field noise prediction. The proposed HSI noise analysis via chimera wake grid indicated that the strength of the shock wave is more precisely predicted, thereby resulting in improved prediction of the HSI noise. The proposed chimera methodology allows efficient noise prediction using a much coarser background grid system.

Numerical modelling of shock wave-boundary layer interaction control by passive wall ventilation

The normal shock wave-boundary layer interaction (SBLI) phenomenon is known to constitute a main factor limiting the aerodynamic performance in many aeronautical applications (transonic wings, helicopter rotor blades, compressor and turbine cascades). The interaction process highly disturbs the boundary layer, often causing flow separation and onset of large scale unsteadiness (e.g. airfoil buffet or supersonic inlet buzz). In certain conditions it may also initiate a dramatic increase of acoustic emission levels (e.g. high-speed impulsive noise). To limit the negative impact of the phenomenon various flow control strategies are implemented, here in a form of a passive control system realised by placing a shallow cavity covered by a perforated plate just beneath the shock. Details of the flow structure obtained by this method are studied numerically. Three distinctive experimental set-ups are considered with the interaction taking place: on a flat wall (transonic nozzle, ONERA), on a convex wall (curved duct, University of Karlsruhe), and on an airfoil (NACA 0012, NASA Langley). Depending on the relative cavity length the ventilation process leads to a transformation of the normal shock topology into: a large λ-foot structure (classical, short cavity), a system of oblique waves (extended cavity), or a gradual compression (full-chord perforation). The reference and flow control cases are simulated with the SPARC code (RANS) with Spalart–Allmaras turbulence and Bohning–Doerffer transpiration models. The results are compared with the measurements, emphasizing the streamwise evolution of the boundary layer profiles and integral parameters during the interaction. The prediction capabilities of the solver in terms of the shock wave-boundary layer interaction control by wall ventilation are assessed and presented in details for the investigated range of flow configurations and conditions.

New Approach for Aerodynamic and Aeroacoustic Analysis of Actively Controlled Flaps Rotor

A new approach for aerodynamic and aeroacoustic analysis of actively controlled flaps (ACF) rotor is established to predict airloads and noise reduction. The approach is divided into three parts. The first is an unsteady aerodynamic model of the ACF described through a panel method, and the unsteady pressure under interaction of rotor tip vortex and ACF trailing-edge vortex, which are depicted through a viscous vortex particle method, are taken into account. The second is the influence of the ACF on trajectory of the wake and ACF trailing-edge vortex, which is carried out by a new imaged vortex model to capture blade-vortex interaction (BVI). The third is an aeroacoustic model coupled into the aerodynamic model by transmitting the unsteady pressure to Ffowcs Williams and Hawkings (FW-H) equation. The present approach is applied to AH-1G rotor, and the results show that the predicted unsteady airloads, high-speed impulsive (HSI), and BVI noise are compared favorably with experiment and Reynolds-averaged Navier–Stokes (RANS)/FW-H results. Finally, the variation of airload and noise reduction of ONERA Active Blade Concept rotor (ABC) is predicted and compared with experiment. It is shown that the present approach predicts the airload and noise reduction of the ACF with different phases and flap locations with good accuracy.

Helicopter rotor noise prediction using a convected FW-H equation in the frequency domain

In this paper, high-speed impulsive noise radiated from helicopter rotors in hover and forward flight is computed by using a convected FW-H equation in the frequency domain with a rotating permeable data surface. Compared to the time-domain formulations, the frequency-domain formula avoids retarded time computation and flow data interpolation with respect to the retarded time. In addition, the complicated emission surface construction is obviated, while a supersonic rotating data surface is chosen to account for quadrupole sources outside the sonic cylinder. The selection of a rotating data surface close to the rotor blades improves noise source resolution and thus ensures noise prediction accuracy at a low computational cost, compared to a stationary data surface approach. Besides, the use of a rotating data surface for hovering helicopter rotors enables noise source identification by a steady flow simulation. Application of the methodology to hovering and forward-flight helicopter rotors has shown good agreement with experimental data, thus indicating that the approach is an accurate and efficient computational tool for noise prediction of helicopter rotors operating at high tip Mach number. Since the formula is derived in a moving medium frame, the effect of forward-flight speed on the radiated noise can be explicitly taken into account.

Development and validation of a hybrid method for predicting helicopter rotor impulsive noise

Prediction of helicopter rotor impulsive noise is practically a very challenging task. This paper describes a hybrid method to predict rotor impulsive noise for both high-speed impulsive noise and blade–vortex interaction noise. The hybrid solver has been developed by combining the advantages of three different methods: (1) a computational fluid dynamics method based on Reynolds-averaged Navier–Stokes equations to account for the viscous and compressible effects near the blade; (2) a vorticity transport model to predict rotor wake system without artificial dissipation; and (3) an acoustic calculation method, based on Ffowcs-Williams Hawkings equation implemented to a permeable data surface. The developed hybrid solver is validated through available test data, for the cases of UH-1H model rotor, AH-1 Operational Loads Survey rotor, and Helishape 7A rotor. Peak sound pressure level of high-speed impulsive noise is accurately predicted with relative errors less than 7%. Additionally, acoustic waveform of blade–vortex interaction noise is well captured though it is sensitive to small changes in aerodynamic load. It is suggested that present hybrid method is versatile for the prediction of rotor impulsive noise with moderate computational cost.

A time-domain method for prediction of noise radiated from supersonic rotating sources in a moving medium.

This paper presents a time-domain method for noise prediction of supersonic rotating sources in a moving medium. The proposed approach can be interpreted as an extensive time-domain solution for the convected permeable Ffowcs Williams and Hawkings equation, which is capable of avoiding the Doppler singularity. The solution requires special treatment for construction of the emission surface. The derived formula can explicitly and efficiently account for subsonic uniform constant flow effects on radiated noise. Implementation of the methodology is realized through the Isom thickness noise case and high-speed impulsive noise prediction from helicopter rotors.

다중 투과면을 이용한 UH-1H 로터 블레이드의 제자리 비행 시 원방 소음 해석 연구

In this study, aero-acoustics in far-field is computed for the UH-1H rotor blade in hovering flight. A Chimera grid method is applied in this simulation to consider the blade motion and moving effects. Multi-permeable surfaces surrounding a blade are constructed to include the flow noise generated from the shock wave and the tip vortices. A computation is performed with three-dimensional Euler’s equations. An acoustic analogy code using the Kirchhoff approach to Ffowcs Williams and Hawkings equations is applied to efficient the noise prediction caused by volume noise sources in far-field. The high speed impulsive noise on the rotor blade is predicted and validated according to the permeable surface locations. It is confirmed that the noise source caused by the shock waves generated on the blade surface is a dominant factor in the far-field noise prediction.

Long-Range and Nonlinear Propagation of Helicopter High-Speed Impulsive Noise

Long-Range and Nonlinear Propagation of Helicopter High-Speed Impulsive Noise

Passive control of rotorcraft high-speed impulsive noise

A strong, normal shock wave, terminating a local supersonic area located at the tip of a helicopter blade, not only limits the aerodynamic performance, but also constitutes an origin of the High-Speed Impulsive (HSI) noise. The application of a passive control device (a shallow cavity covered by a perforated plate) just beneath the interaction region weakens the compression level, thus reducing the main source of the HSI noise. The numerical investigation based on the URANS approach and Bohning/Doerffer (BD) transpiration law (SPARC code) confirms a large potential of the new method. Two exemplary implementations, adapted to model helicopter rotors tested at NASA Ames facility in transonic conditions: Caradonna-Tung (lifting, transonic hover) and Caradonna-Laub-Tung (non-lifting, high-speed forward flight), demonstrate the possible gains in terms of the reduction of acoustic pressure fluctuations in the near-field of the blade tip. The CFD results are validated against the experimental data obtained for the reference configurations (no control), while the analysis of the passive control arrangement is based on a purely numerical research. The normal shock wave is effectively eliminated by the wall ventilation exerting a positive impact on the generated level of the HSI noise.

Optimization for rotor blade-tip planform with low high-speed impulsive noise characteristics in forward flight

Rotor noise is one of the most important issues for helicopter designer, and high-speed impulsive noise is particularly intense among the various rotor noise sources due to compressibility. Based on Computational Fluid Dynamics/Ffowcs Williams and Hawkings equations with Penetrable Data Surface (CFD/FW-Hpds) methods and hybrid optimization technique, a new optimization design procedure for rotor blade planform with low high-speed impulsive noise characteristics is established. First, in order to accurately capture the unsteady aerodynamic characteristics of rotor, based on the moving-embedded grid methodology, a CFD simulation is developed by solving the compressible Reynolds-average Navier–Stokes equations with Baldwin–Lomax turbulence model. The low dissipation Roe-Monotone Upwind-centered Scheme for Conservation Laws (MUSCL) scheme and highly efficient implicit lower–upper symmetric Gauss–Seidel scheme are used for spatial and temporal discretization, respectively. Second, taking the CFD results as sound pressure information input, the high-speed impulsive noise characteristics generated by transonic rotor are analyzed through a robust numerical method based on FW-Hpds equations. Third, the genetic algorithm and surrogated model based on radial basis function are combined as a hybrid optimization technique; during the optimization process, the blade grids are generated by a highly efficient parameterized method. Aiming at the minimization of the sound pressure level of rotor in forward flight, the parametric effect analyses of blade-tip shapes on transonic noise have been conducted first. Then, optimization analyses based on the rotor blade with double-swept and tapered tip have been accomplished with the aerodynamic performance as constraints. Compared with the baseline blade, it shows that the sound pressure level of rotor with optimized blade-tip shape can be decreased obviously at the present calculating condition due to its weaker transonic delocalization phenomenon in the region of blade tip. In the rotor plane, absolute peak value of sound pressure produced by the optimized blade planform is reduced about 59.1% of that by the baseline one, and the reduced value in sound pressure level is up to 5.6 dB.

Unsteady transonic flow control around an airfoil in a channel

Transonic internal flow around an airfoil is associated with self-excited unsteady shock wave oscillation. This unsteady phenomenon generates buffet, high speed impulsive noise, non-synchronous vibration, high cycle fatigue failure and so on. Present study investigates the effectiveness of perforated cavity to control this unsteady flow field. The cavity has been incorporated on the airfoil surface. The degree of perforation of the cavity is kept constant as 30%. However, the number of openings (perforation) at the cavity upper wall has been varied. Results showed that this passive control reduces the strength of shock wave compared to that of baseline airfoil. As a result, the intensity of shock wave/boundary layer interaction and the root mean square (RMS) of pressure oscillation around the airfoil have been reduced with the control method.

Numerical analysis on noise of rotor with unconventional blade tips based on CFD/Kirchhoff method

A solver is developed aiming at efficiently predicting rotor noise in hover and forward flight. In this solver, the nonlinear near-field solutions are calculated by a hybrid approach which includes the Navier–Stokes and Euler equations based on a moving-embedded grid system and adaptive grid methodology. A combination of the third-order upwind scheme and flux-difference splitting scheme, instead of the second-order center-difference scheme which may cause larger wake dissipation, has been employed in the present computational fluid dynamics (CFD) method. The sound pressure data in the near field can be calculated directly by solving the Navier–Stokes equations, and the sound propagation can be predicted by the Kirchhoff method. A harmonic expansion approach is presented for rotor far-field noise prediction, which gives an analytical expression for the integral function in the Kirchhoff formula. As a result, the interpolation process is simplified and the efficiency and accuracy of the interpolation are improved. Then, the high-speed impulsive (HIS) noise of a helicopter rotor at different tip Mach numbers and on different observers is calculated and analyzed in hover and forward flight, which shows a highly directional characteristic of the rotor HIS noise with a maximum value in the rotor plane, and the HSI noise weakens rapidly with the increasing of the directivity angle. In order to investigate the effects of the rotor blade-tip shape on its aeroacoustic characteristics, four kinds of blade tips are designed and their noise characteristics have been simulated. At last, a new unconventional CLOR-II blade tip has been designed, and the noise characteristics of the presented CLOR-II model rotor have been simulated and measured compared to the reference rotors with a rectangular or swept-back platform blade tip. The results demonstrate that the unconventional CLOR-II blade tip can significantly reduce the HSI noise of a rotor.

Design Exploration of Helicopter Blades for HSI Noise and Aerodynamic Performance

In this study, the high-speed impulsive (HSI) noise and aerodynamic performance of helicopter blades were improved using multi-objective design exploration (MODE) comprised of multi-objective shape optimization and data mining. To broaden the design space of the problem, geometry definitions with a high degree of freedom were adopted. As a result, remarkable improvements in both HSI noise and aerodynamic performance were achieved. For data mining, analysis of variance (ANOVA) and self-organizing map (SOM) were used to extract design knowledge of the helicopter blades. The results indicated that tip chord length and blade twist are important factors for HSI noise and aerodynamic performance, respectively. Based on the design knowledge obtained from data mining, an additional design variable (twist) was introduced. The solutions obtained from the shape optimization using variable twist showed better blade loading performance than that from shape optimization using fixed twist.

Application of the passive control of shock wave to the reduction of high-speed impulsive noise

Strong, normal shock wave, terminating a local supersonic area on an airfoil (or helicopter blade), not only limits aerodynamic performance but also becomes a source of a high-speed impulsive (HSI) noise. The application of a passive control system (a cavity covered by a perforated plate) on a rotor blade should reduce the noise created by a moving shock. This article covers details of the numerical implementation of the Bohning/Doerffer transpiration law into the SPARC code and includes a validation against the experimental data obtained for the ONERA transonic nozzle with a flat wall. The passive control device is then applied numerically on a helicopter blade in high-speed transonic hover conditions to weaken the shock wave – the main source of HSI noise.

Improved Algorithm for Nonlinear Sound Propagation with Aircraft and Helicopter Noise Applications

An efficient numerical method to solve nonlinear sound propagation is presented. The frequency-domain Burgers equation, which includes nonlinear steepening and atmospheric absorption, is formulated in the form of the real and imaginary parts of the pressure. The new formulation effectively eliminates possible numerical issues associated with zero amplitude at higher frequencies occurring in a previous frequency-domain algorithm. In addition, to circumvent a high-frequency error that can occur in the truncated higher frequencies, a split algorithm is developed, in which the Burgers equation is solved below a cutoff frequency and a recursive analytic expression is used beyond the cutoff frequency. Finally, the Lanczos smoothing filter is incorporated to remove the Gibbs phenomenon. The new method is found to successfully eliminate high-frequency numerical oscillations and to provide excellent agreement with the exact solution for an initially sinusoidal signal with only a few harmonics. The new method is applied to a broad range of applications with a comparison to other methods to assess the robustness and numerical efficiency of the method. These include sonic boom, broadband supersonic jet engine noise, and helicopter high-speed impulsive noise. It is shown that the new method provides the fastest and most accurate predictions compared to the other methods for all the application problems.

Helicopter Rotor Shape Optimization for the Improvement of Aeroacoustic Performance in Hover

A helicopter rotor is optimally designed for aeroacoustic performance improvement. As shown in previous reports, the blade shapes can be designed to minimize high-speed impulsive noise but tend to have excessively high tapers and be swept back. Since an overly short chord length around the blade-tip region may cause structural problems and safety issues in autorotation, an autorotation index has been introduced to keep the tip region from having excessive taper ratios. In addition, the changes in thickness and camber of the airfoils can also be taken into account to better represent realistic rotor shapes. Aeroacoustic analysis is performed using Kirchhoff's method coupled with computational fluid dynamics analysis, and the optimization is performed using the kriging-model-based genetic algorithm method. Optimization results are presented that show that the designed blades have improved aerodynamic performance and reduced high-speed impulsive noise characteristics. It is found that a more practical blade shape can be obtained by using airfoil transitions and an autorotation index. The results of the analysis of variance and self-organization map indicate that the taper ratios, the swept back, the tip chord length, the protrusion shape, the camber, and the thickness of the root airfoil are the prominent features affecting the aeroacoustic performance of the rotor.

Prediction of Hovering Rotor Noise Based on Reynolds-Averaged Navier-Stokes Simulation

O VER the past decade, the hybrid method [coupling computational fluid dynamics (CFD) techniques with advanced analytic methods based on acoustic analogy, such as the Ffowcs Williams–Hawkings equation with penetrable data surface (FW–Hpds) method] has been successfully applied to predict the complicated acoustic field of helicopter rotors. The Ffowcs Williams–Hawkings (FW–H) equation [1], a rearrangement of Navier–Stokes equations by using generalfunction theory, provides an accurate theoretical model for describing the propagation of noise from a moving surface to the far field. The Farassat 1Amethod for solving the linear part of the FW–H equation was developed by Farassat and Succi [2]. It has been successfully applied in linear noise prediction [2,3] for more than 20 years. The Farassat 1A method predicts discrete frequency noise quite well, but it would run into complication when predicting nonlinear quadrupole noise of helicopter rotors, because the data surface is the blade itself and nonlinear effects are not included in the surface integral. To calculate the nonlinear noise [e.g., high-speed impulsive (HSI) noise], Farassat and Myers [4] derived the general form of the Kirchhoff equation and its solution (known as the Kirchhoff formulation) to describe the noise radiation from amoving surface. The data surface of the Kirchhoff formulation is fictitious and penetrable. The main benefit of the Kirchhoff method is that the nonlinear effect is accounted for by performing the integral on the data surface covering the nonlinear flow region. The Kirchhoff method coupled with the near-field CFD solution (called the CFD/ Kirchhoff method) has proven to be accurate and efficient when predicting impulsive noise. More recently, a new form of FW–H equation with a penetrable surface (called the FW–Hpds equation) was proposed by Crighton et al. [5] to improve the efficiency of solving the quadrupole noise. The method using a penetrable data surface for solving the FW–H equation with the Euler solution as input data was first implemented by di Francescantonio [6] for prediction of far-field noise from transonic helicopter rotors in hover. Brentner and Farassat [7] conducted an analytical comparison of the FW–Hpds method with the Kirchhoff method and concluded that the FW–Hpds method is more accurate and robust than the Kirchhoff method when the data surface is located in the nonlinear flow region. The FW–Hpds method rapidly showed promise in the studywork of a few researchers [7–10] when it was used for predicting the noise generated by helicopter rotors in hover and forward flight. More recently, Farassat and Casper [11] emphasized the role of analytical methods in computational aeroacoustics and recommended FW–Hpds as a very promising method for noise prediction of a complicated flowfield. To predict nonlinear noise generated by transonic rotors in hover, three-dimensional Euler equations were commonly used to consider the nonlinear effect related to shockwaves. To consider the influence of viscous effect in the near field and get more accurate information about noise sources, this paper uses Reynolds-Averaged Navier– Stokes (RANS) equations to model the nonlinear viscous flowfield near the rotor blades. The far-field noise is calculated by a retardedtime integral formula solving the FW–Hpds equation, with the solution of the RANS equations taken as input data.