Browse Topic: Noise
An experimental investigation was conducted to explore the loads, acoustics, and tip vortex trajectories of coaxial counter-rotating (CCR) rotor with unequal upper and lower radii. The upper and lower rotor radii were tested both at the nominal radius of 1.108 m, and also with a lower rotor radius of 90% nominal radius, for a constant rotor speed of 1180 RPM and a constant inter-rotor spacing of z/R = 0.108. Rotors were torque balanced and tested for a range of upper rotor collective pitch from -2◦ to 10◦ . The power required for both CCR systems was within 0.9% for most trim conditions, and equal thrust was produced at upper rotor collectives of 6◦ and 8◦ (within 1.0%). At low loading conditions the unequal radii configuration produced more thrust for the same power due to a reduction in profile drag. The overall sound pressure level (OASPL) was lower for the CCR rotor with shortened lower rotor blades at all angles of elevation. Larger reductions in A-weighted OASPL(A) were observed, due to a larger contribution of broadband noise to the total OASPL(A).
This paper describes the electromagnetic noise mitigation on the Maryland Tiltrotor Rig (MTR) and presents its first hover test results. The primary source of noise was found to be pulse width modulation associated with the motor controller. Due to this noise, testing was limited to unpowered, freewheeling cases. To solve the noise problem and allow powered testing, three hardware filters were integrated into the power and data systems. A complementary digital filter was also used. With the filtering solution in place, hover tests were carried out to high collectives of 30◦and blade loadings of 0.2. The test data was assessed using blade element-momentum theory predictions.
Achieving noise reduction in rotorcraft requires an analysis of various design parameters and flight conditions. However, high-fidelity methods are computationally expensive. To overcome this limitation, reduced order model (ROM)-based surrogate models have been applied to aerodynamics and aeroacoustics prediction. This study proposes a ROM-based surrogate model employing a variational autoencoder (VAE) to predict rotor aerodynamic loads and associated noise. Train and test datasets were generated using reformulated vortex particle method across a wide range of flight conditions. The proposed framework was applied to a single rotor, and its performance was evaluated qualitatively and quantitively in comparison with proper orthogonal decomposition (POD)-based surrogate model. The results show that VAE-based model consistently outperformed the POD model in noise prediction. These results demonstrate that the proposed framework enables accurate rotor noise prediction under various flight conditions and provides a promising approach for low-noise rotorcraft design and operational planning.
This study examines the capability of medium-fidelity comprehensive analysis models to predict the acoustics for manned and unmanned rotorcraft configurations. Using the automated tool NDARC2RCAS developed at DEVCOM Army Research Laboratory, multiple configurations including a single main rotor, tilt rotor, coaxial and pusher, quadcopter, and hexacopter are evaluated at various mission segments including hover, advancing climb, and forward flight. Each configuration and condition is evaluated using a range of aerodynamic models from lower to higher fidelity, including uniform inflow, dynamic inflow, prescribed wake, free wake, and viscous vortex particle method (VVPM). These evaluations are then used with another automated tool, RCAS Acoustics, to predict noise on a Voronoi observer sphere. A comparison of the results for the single main showed good agreement between all of the aerodynamic models except VVPM. For the tilt rotor in forward flight, the higher-fidelity models produced changes in rotor loads due to the interaction with the wing. With prescribed and free wake models, this change in load is sharp and causes noise increases of up to 40 dB in front of and behind the vehicle, while the VVPM model produced a smoother change that results in a smaller, 20 dB increase in noise. The quadcopter and hexacopter show similar in-plane noise levels for all models, with alternating cancellation and amplification patterns due to rotor phasing, while out-of-plane noise is increased on the hexacopter when using the higher fidelity models.
This paper investigates the relationship between broadband noise behavior and helical wake structure in coaxial corotating rotors. Experimental measurements were conducted across variations in collective pitch (9.4°, 12.5°, and 15.0°) and rotor speeds (1500–4500 RPM). The inflow ratio (λ) was shown to govern the slope of broadband noise trends mapped in phase offset versus separation distance space, with experimental and theoretical λ values agreeing within 1%. Tip vortex core growth was estimated using the Ramasamy-Leishman model and normalized by the blade tip chord, reflecting the location of tip vortex formation. Across collective pitch variations, initial vortex core radii ranged between 7.5% and 9.1% and across rotor speeds, it ranged between 7.5% to 8.5% of the blade tip chord. When broadband noise trends became less coherent across phase offset angles, the corresponding vortex core radii were observed to approach or exceed 10% of the tip chord. At 4500 and 3500 RPM, vortex cores radii remained near or slightly above 10% at diffusion and depletion of broadband noise coherence. At 1500 RPM, initial vortex core radius is significantly beyond 10% of blade tip chord, leading to early diffusion, loss of organized wake structures, and broadband noise saturation across all phase offsets. If an inboard chord length were used instead for normalization, the specific normalized values would differ, but the overall trends in vortex diffusion and broadband noise coherence would remain similar. These results define a practical RPM threshold below which coherent wake structures and slope-based inflow prediction are no longer valid. Overall, the findings offer a framework for anticipating wake breakdown and broadband noise saturation, providing guidance for rotor design and operation to preserve low-noise performance.
In this study, we employ the Polynomial Chaos Expansion (PCE) and Monte Carlo (MC) methods to quantify the uncertainty of unsteady loading noise generated by a hovering rotor under the presence of vertical gust. The unsteady loading noise is predicted using a frequency-domain approach combined with a quasi-steady Blade Element Momentum Theory, accounting for time-varying aerodynamic forces. A sinusoidal gust is modeled using two parameters: gust length and gust amplitude. Then, the uncertainty quantification (UQ) of the unsteady loading noise is performed using the PCE and MC with these two gust parameters. The UQ analyses show that the largest uncertainty in unsteady loading noise occurs at the rotor axis, and PCE and MC simulations show good agreement. The individual and combined effects of the gust parameters on the acoustic uncertainty are analyzed, and parallel coordinate plots are utilized to visualize combinations of the gust parameters that produce noise outliers. It is found that the gust length leads to larger uncertainty in the unsteady loading noise than does the gust amplitude, especially at fifth harmonic. Furthermore, low magnitude of high harmonic noise is produced even at high gust amplitudes when combined with wide gust lengths.
An extensive test campaign was conducted at the National Full-Scale Aerodynamics Complex 40- by- 80-Foot wind tunnel to acquire performance, loads, and acoustics measurements of the Joby Aviation propeller across a variety of operating conditions. The dataset provided validation of the design methodology as well as verification of computational tools. The Vold-Kalman filter was used to extract the shaft-coherent propeller noise in hover to obtain the residual noise, representing the broadband noise. This data verified broadband noise tip speed scaling laws as well as a low-order empirical model for overall sound pressure level. The OVERFLOW/PSU-WOPWOP method was used to simulate the propeller in pure edgewise flight and shown to accurately predict propeller performance. The low-frequency acoustics were predicted well but the solver underpredicted frequencies above 300 Hz, possibly due to the inability to capture the turbulent component of the blade-wake and blade-vortex interaction. The computational method was used to simulate the propeller at various angles of attack in low-speed edgewise flight and captured trends and spectral content up to 1 kHz. The predictions showed noise sources moving from root to tip as the propeller angle of attack increased. However, the high-frequency content was not captured for all cases. The experimental campaign was successful in characterizing the acoustics of the Joby Aviation propeller, but more research is needed to be able to properly predict and understand the noise sources throughout the flight envelope.
Blade–wake interaction (BWI) is a significant source of broadband noise and is often dominant in rotors with high blade counts. Accurately capturing the resulting unsteady blade loading is computationally expensive and, therefore, drives the cost of BWI noise calculation. To address this challenge, a low-fidelity BWI noise prediction tool was developed using aerodynamic data from the blade element momentum theory (BEMT) and the lattice Boltzmann method (LBM) for a series of rotor configurations with medium to high solidity. Starting from a six-bladed baseline rotor, 13 additional configurations were generated by varying blade twist, taper, root collective, solidity, and blade count. The relationship between vortex miss distance and blade loading unsteadiness was quantified to construct a semi-empirical BWI noise model. The model predicted BWI noise with a root mean square error of 3.9 dBA and a mean absolute percentage error of 1%. It was subsequently integrated into a BEMT framework to produce aerodynamic and acoustic data for training a tandem neural network (TNN) that was employed to optimize two rotor geometries. The optimized designs achieved up to a 7% reduction in BWI noise and a 7% improvement in performance. Additional geometric modifications—including blade tip anhedral, forward sweep, and a mixed configuration—were also assessed using LBM, each demonstrating notable noise reduction.
Carbon/epoxy stiffened panels are being increasingly used in transport rotorcraft. The reduced mass density and high stiffness of carbon/epoxy composites can lead to higher levels of vibration relative to comparable metallic structures, which themselves can have vibrations and interior noise high enough to damage the hearing of crew and passengers. The current investigation explores a method to reduce the vibration of carbon/epoxy stiffened panels by introducing thickness tapers known as acoustic black holes (ABHs). The ABH feature is integrated into either the stiffeners or plate of a representative stiffened panel configuration. A finite element (FE) parametric study was used to guide designs that reduce the vibration of the panel without compromising the compressive buckling capability or mass of the panel. FE studies showed that a 30 ply to 12 ply thickness taper longitudinally oriented in the blade stiffener can reduce vibrations and increase compressive buckling capability. Carbon/epoxy panels were manufactured using a low-cost out-of-autoclave material with simple molding. Experimental testing concluded that integrating the ABH into the stiffeners longitudinally helped to reduce the broadband vibration by 5 dB and increase the buckling load (+4.3%) and collapse load (+16.5%) without increasing the mass greatly compared to a traditional baseline design.
Acoustic flight testing of rotorcraft often involves generating noise source hemispheres to gain an understanding about the aircraft's acoustic emissions. However, aerodynamically complex Urban Air Mobility and Future Vertical Lift vehicles may not maintain a steady aerodynamic state during flight, making source hemispheres measured using traditional linear arrays unreliable or difficult to interpret. To address this challenge, all emission angles need to be measured simultaneously. This has lead to the concept of the two dimensional 'snapshot' array layout. A mathematically defined microphone distribution was utilized to achieve uniform coverage on the source hemisphere. Within the chosen distribution, two lower microphone count distributions are embedded, allowing for a comparison of the effects of number of microphones. The array was deployed as part of a joint Army/NASA acoustic research flight test in July of 2024. Data were collected using an MD530F helicopter as the test vehicle, executing both steady and unsteady flight. Analysis of the array resolution is used to determine adequate channel count and spatially varying sensitivity to array density.
The performance and acoustics of a scaled propeller designed for an eVTOL vehicle were investigated in axial and edgewise flight. The measured performance compared well with BEMT predictions in axial flight conditions. The noise produced by the propeller is dominated by broadband noise sources, where there is evidence of contributions from blade wake interaction noise, turbulent boundary layer trailing edge noise, and laminar boundary layer vortex shedding noise. The directivity of the noise was found to be dependent on the advance ratio. Beamform maps also identified changes in the dominant noise source at different observer locations as a function of advance ratio.
Survivability in the future operating environment is becoming more challenging as threat systems evolve and become more sophisticated. The ability to tailor and manage signatures will be one of the key methods to improve survivability, allowing operators to minimise detection and maximise the effectiveness of countermeasures. This paper presents the findings of an investigation into the application of classical Signal Detection Theory (SDT) to the aural detectability of helicopter noise signatures, considering human auditory capabilities. The paper has thus developed a novel methodology, applied it to both the experimental and numerical helicopter acoustics signatures of an LH platform, and used these results to infer the detectability characteristics of the aircraft, as well as how they are affected by the presence of background noise in different environments.
In the context of developing new rotorcrafts dedicated to Advanced Air Mobility, aeroacoustic simulations of co-axial rotor systems have been conducted using the lattice Boltzmann method with the ProLB code. Eight configurations, spanning from co-rotating rotors to contra-rotating shrouded rotors, were analyzed in stationary conditions through comparisons with experimental data and flow field analysis. This investigation validates our numerical methodology, based on direct noise simulation, and enhances our understanding of noise generation and propagation of such propulsive systems. Our simulations successfully replicate all measured trends in global aerodynamic performance, average noise levels, and noise directivities. Maximum discrepancies were 1.5 N (9%) and 2.1 dB (averaged noise level on the considered microphones). Based on our analysis, the following observations have been made. The open co-rotating rotors is the quietest configuration due to reduced Blade-Vortex Interaction (BVI). A 30° leading rotor angle increases noise by inducing BVI and flow detachment on the downstream rotor. On the other hand, contra-rotation offers the best performance but generates high tonal noise. Finally, the shrouded configuration is the noisiest due to blade-flow interactions and detachments along the duct, resulting in reduced performance and additional noise sources primarily due to non-adapted blades and duct.
A cooperative flight test campaign between the US Army and NASA was performed. This test sought to characterize the acoustic emissions of a fully instrumented MD530F helicopter using a snapshot array and a phased array of microphones. The snapshot array of microphones aimed to provide even coverage across the surface of a hemisphere, providing an acoustic emission hemisphere in a single 'snapshot' of time. The phased array of microphones was designed to provide enough resolution to determine noise sources from each individual blade as well as perform source separation from main rotor and tail rotor emissions. Test conditions for the characterization effort were chosen using a traditional one-factor-at-a-time approach as well as three design of experiment approaches. Characterization conditions included constant speed level flight, descent, and ascent conditions. Transient maneuver conditions were also captured over the snapshot array. The vehicle instrumentation included measurements of pilot controls, optical sensors to measure blade azimuth locations, pitch link loads, along with strain gauges to measure structural loads, blades and fuselage. This report will provide an overview of the test, document the data acquired, and provide some initial results.
A hybrid RANS/LES simulation of the Ideally Twisted Rotor (ITR) in hover was interrogated to identify bluntness vortex shedding (BVS) and determine the contribution to the predicted rotor broadband self-noise. Three rotor blade stations were extracted to study spanwise variations in the BVS shedding frequency and amplitude. Corresponding 2-D airfoil simulations were performed to evaluate a simplified modeling approach that effectively isolates BVS. The BVS shedding frequencies predicted by the 2-D airfoil simulations differed by less than 2% from the corresponding rotor stations in the 3-D simulation. The increased computational cost incurred by performing 3-D airfoil simulations did not lead to a worthwhile increase in simulation fidelity. Farfield noise was predicted for the three rotor stations and the 2-D airfoil simulations, and trends in frequency agreed well. The 2-D approach overpredicted the 3-D peak amplitudes by 5 - 10 dB. This work demonstrates that 2-D hybrid RANS/LES airfoil simulations can be used to investigate BVS noise trends on the ITR.
This paper explores a significant step forward, regarding the further detailed understanding of the Fenestron®. Since its patent in 1968 – for the Gazelle helicopter –, the shrouded tail rotor has been resized, inclined, modulated, etc. and has thus been continuously enhanced on different rotorcraft. Half a century after its invention, Airbus is once again exploring in more detail the magic of the Fenestron®, with the objective of optimizing it even further, for future helicopter applications. To grasp and observe properly some specific phenomena, a model (scaled to one third) capable of both unprecedented functions and modularities, was developed. The present paper will describe in detail the novel model and the related challenges and solutions. This model is capable of high rotor speed and dynamic pitch inputs, delivering power levels high enough to reach stall effects, while allowing the measurement of propulsive efficiency and to differentiate rotor vs fairing thrust. Furthermore, the model had to provide aerodynamic-shape modularity, both in the shroud, on the covers and on the tail gearbox supports. The first test campaign performed on this model allowed us to define design drivers and aerodynamic preferences. And since one idea often leads to another, another challenge is being addressed on this scaled mock-up: Reduction of the tail rotor noise level.
This study investigates the effects of chord-to-radius ratio (c/R) and blade count on the aerodynamic and aeroacoustic performance of cyclorotors through experimental testing and a low-fidelity streamtube model. Cyclorotors with c/R ratios between 0.3 to 0.75 and blade counts ranging from 5 to 9 were tested across pitch amplitudes up to 51°. For a 5-bladed configuration, the pitch amplitude that maximizes the force-to-power coefficient (CF/CP) increases with c/R from approximately 32° at low c/R to around 51° at high c/R. However, the peak attainable CF/CP decreases with increasing c/R, indicating a trade-off between optimal pitch amplitude and aerodynamic efficiency. Increasing blade count enhances the generated force but reduces efficiency in all cases except for the lowest c/R configuration (0.3). Aeroacoustic analysis shows that tonal noise is primarily driven by pitch amplitude and intensifies with increasing c/R, while additional blades effectively mitigate it. In contrast, broadband noise is less sensitive to variations in pitch amplitude, c/R, and blade count. The streamtube model captures key aerodynamic trends, particularly at moderate pitch amplitudes. Scaling studies identify optimal configurations for a given disk loading, balancing power consumption and noise levels and highlighting key trade-offs critical to urban air mobility applications.
An experimental investigation was conducted on a 1.108 m radius coaxial co-rotating (or stacked) rotor and a coaxial counter-rotating (CCR) rotor of identical geometry to compare the acoustics and loads of both rotor configurations in hover. The rotors were operated at a tip Mach number of 0.40, tip Reynolds number of 765,000 and an axial spacing of 1.55 chord lengths, and the index angle between the upper and lower blades of the stacked rotor was varied. The overall sound pressure level (OASPL) was significantly larger for the CCR rotor. For example, total rotor noise at -45◦ angle of elevation was 6 dB greater for the CCR rotor than the stacked rotor at 8◦ collective. These increases in OASPL were driven by large increases in tonal noise for the CCR rotor, of up to 10 dB higher than the stacked rotor at some angles of elevation. This was attributed to additional tonal noise occurring at harmonics of 2Nb/rev, due to vibratory loads from 2Nb/rev blade crossings. The results of the experimental study suggest that, in terms of reduced noise and increased hover efficiency, the stacked rotor is preferable to the CCR rotor for eVTOL vehicle configurations that implement compact or closely spaced designs.
ABSTRACT Small-scale rotorcraft exhibit degraded aerodynamic efficiency, which has been linked to non-ideal losses within the wake. Unique high-frequency, broad-band features have also been observed, without a physical verification of their origin. This work seeks to gather insight into the behavior of the rotor wake structures as a function of Reynolds number (Re), relate this to performance capabilities and the corresponding far-field acoustic signature. Two-component particle image velocimetry (PIV), performance, and acoustic measurements were performed using three small-scale, NACA 0012 rotors operated over a range of low-Reynolds number conditions. Rotor geometry and operational speed (Ω) were varied to obtain the desired Re variation. Span-wise PIV has demonstrated an absence of tip vortex formation as the operational thrust coefficient (CT) is increased, suggesting outboard tip stalling. Phase-locked, chordwise PIV has confirmed this hypothesis, showing the development of large-scale separation and a highly turbulent downstream wake. The wake characteristics at moderate collective angles (θ) show the presence of periodic counter-rotating structures attributed to laminar boundary layer vortex-shedding. Significant broad-band frequency content was measured, and found to correlate well with a physical quantification of the shedding phenomenon.
Airfoil optimization for rotor blades is a critical endeavor aimed at enhancing aerodynamic performance and reducing noise. This paper employs a Kriging surrogate model combined with a multi-objective genetic algorithm to optimize thrust, power, and broadband noise. Three airfoil parameterization methods including ParFoil, PARSEC, and CST are compared when used to generate various airfoil shapes for the surrogate model and optimization process. We utilize low-fidelity aerodynamic tools such as XFOIL and blade element momentum theory for aerodynamics. In addition, acoustic modeling is conducted using Lee's wall pressure spectrum model alongside Amiet's trailing-edge noise model. The paper focuses on small-scale rotor configurations, specifically an ideally twisted rotor using the NACA 0012 airfoil and a modified XV-15 blade. Both blades are used as baseline models for hover optimization. The optimization of the ideally twisted rotor across various parameterization methods demonstrates a significant reduction in A-weighted overall sound pressure level by approximately 4.0 dBA. The primary contributors to this noise reduction are identified as a decrease in the chordwise pressure gradient and wall shear stress, which are key factors that contribute to the wall pressure spectrum and resulting trailing-edge noise. Similarly, the optimization of the modified XV-15 blade using the ParFoil method achieves a notable decrease in noise of 3.5 dBA. Further analysis is extended to axial flight conditions, with axial velocities of 5.0, 10.0, and 20.0 m/s, to simulate vertical climb. The results show that these hover-optimized blades are capable of sustaining noise decreases of 3.0-4.0 dBA compared to the baseline, provided that the thrust coefficient is maintained through adjustments in collective pitch.
Presented herein is a comprehensive workflow for the aeroacoustic analysis of a tilt-rotor air taxi during cruise and hover flight using high-fidelity numerical methods. Employing a hybrid approach, the near-field flow is resolved by an unsteady Reynolds-Averaged Navier-Stokes (URANS) solver, which is paired with a Ffowcs Williams-Hawkings (FW-H) acoustic solver to compute the far-field noise. Both impermeable and permeable FW-H integration surface approaches are incorporated. To balance computational resources and accuracy, the flow domain is halved, while the acoustic data is processed to reflect the full vehicle acoustics. Isolated acoustic contributions of each rotor are extracted, allowing the investigation into the impact of phase shifts on the acoustic signature of the vehicle. Spectral analysis, directivity maps, and noise hemispheres reveal the resolution of fundamental open rotor characteristics and amplified tonal interaction noise that can be linked to aeropropulsive phenomena. The introduction of phase shifts can lead to reduced noise levels and an unsymmetrical acoustic field development. A direct comparison of the results based on impermeable and permeable FW-H source surfaces shows general agreement on the directivity prediction but significant differences in noise levels, potentially due to numerical dissipation, grid settings and the symmetry boundary.
This paper describes a mathematical framework for determining the optimal sensor set location for adequately capturing the sound generated by rotors. The approach leverages the gappy-POD method proposed by Everson and Sirovich [J. Opt. Soc. Am., Vol. 12, 1995, pp. 1657-1664], which first identifies the various mode constituents that make up the first few rotor blade-pass frequency harmonics of the sound-field. The algorithm is developed using a covariance matrix for the POD problem comprising auto- and cross-spectral densities of spatially and temporally resolved sound waves captured by an array of microphones oriented parallel to the axis of a laboratory-scale hovering rotor. Three different forms of the technique are developed and compared. These comprise a homogeneous form and two heterogeneous forms; the heterogeneous forms are referred to as XX-topos and XX-chronos and depends on which term in the error minimization equation is assigned the gappy sensor set. A greedy algorithm is then employed to determine the optimal location of the limited sensor set. The findings are analyzed for different combinations of POD modes and blade-pass frequency harmonics of the sound generated by the hovering rotor.
Aeroacoustic characterization of multirotor aircraft is a challenging task, especially due to the variability in the rotational speed of rotors. This problem is exacerbated by the use of variable RPM control, because changes in RPM change the noise sources of the aircraft in a time-dynamic way. As a consequence of the constant variations in noise sources, an accurate assessment of the acoustic characteristics of the aircraft can be difficult to obtain. A possible solution to this problem is to separate the noise sources of the aircraft on a rotor-to-rotor basis and understand the variations based on each rotor's operating states. To achieve this, a source separation process based on the Vold-Kalman filtering approach is applied to separate the contributions of the individual rotors from ground-based acoustic measurements of a hexacopter. The source separation process was applied to separate the tonal noise of each rotor based on the rotor speed measurement and the results were analyzed for a level flight flyover case. The rotor speed has a strong correlation to the A-weighted SPL throughout the flyover and the variations in RPM directly affect the tonal noise levels, as expected from theory.
ABSTRACT The paper investigates the influence of a dynamic active twist control on BVI noise in descent flight for a Mach-scaled Bo 105 model rotor. Therefore, a numerical study has been carried out with DLR's aeromechanics rotor code S4. In addition, sound pressure levels have been computed with DLR's acoustic code APSIM. The intensity of BVI noise is determined by different vortex parameters as well as the blade-vortex miss-distance and the BVI location. In this paper, all parameters that influence the intensity of the radiated BVI noise are analyzed. The analysis is carried out for a rotor shaft angle of attack of αᴿₒ = 4° for which the most intensive BVI noise was identified for the passive rotor blade. The study lays emphasis on the effect of different control amplitudes and phases but concentrates on a control frequency of the third multiple of the rotor rotational frequency for which highest reductions in BVI noise were identified.
This study examines the acoustics in hover for manned-size, multi-rotor, eVTOL aircraft in a quadcopter configuration. The rotors on such larger aircraft could have collective pitch control allowing them to operate at a fixed rotational speed. This paper seeks to explore how the relative phasing between the rotors affects the acoustics. Quadcopters with three different rotors are considered: a baseline solidity σ rotor with number of blades N = 2, a 3σ rotor with number of blades N = 2, and a 3σ rotor with number of blades N = 5. The simulations use the Rensselaer Multicopter Analysis Code (RMAC) for the aerodynamic loads on the blades, coupled to an acoustic propagation code for noise predictions at observers in the plane of the quadcopter and at elevations of 30 deg and 60 deg (below the quadcopter). The starting phase of rotors 2, 3, and 4 are varied relative to rotor 1, resulting in 216 total phasing cases for each rotor. From the simulation results in this study, the range of variation in tonal noise (due to thickness and loading) was between 21-30 dB in overall sound pressure level (OASPL). If there is phase locking between rotors, for 2-bladed rotors orthogonal phasing was generally observed to produce low average noise, while tip-to-tip phasing produced higher averaged noise, but this observation did not hold for 5-bladed rotors. For high-solidity 2-bladed rotors, the unweighted OASPL from tonal noise (thickness and loading) is greater than that from broadband noise, especially in-plane and at low elevation angles. But for high-solidity 5-bladed rotors, OASPL from broadband noise was observed to be higher than from tonal noise.
This study examines the acoustics of a wing operating in the wake of a propeller. The propeller wing system is simulated at 24 knots cruise and 8° wing angle of attack. The propeller is simulated using an actuator line model, while the wing is simulated using two different turbulence models: a DDES turbulence model and a higher fidelity LES model. Chordwise compact loads, on-wing pressure surfaces, and pressure surfaces at distances of 2.34% and 10% thickness around the wing surface are used as inputs to PSU-WOPWOP to predict noise at an observer below the wing. Using on-wing surface pressures, the LES broadband noise predictions are 13.5 dB higher than DDES. Chordwise compact loads result in lower noise predictions than on-wing surface pressures, by 11.3 dB for LES and 2.3 dB for DDES. Using off-body pressure surfaces, DDES results remain similar to noise predictions from on-wing pressure surfaces, but with LES the broadband noise predictions are about 2.5 dB lower.
This paper presents activities performed in the frame of MOTUS, a DGAC-funded research project, to better understand and reduce annoyance of helicopter operations. It focuses on the operational context of La Réunion island where local authorities intend to define concrete measures to answer multiple complaints from the population. In parallel with ongoing research towards a better understanding of short- and long-term annoyance thanks to both laboratory and field studies, the paper presents an in-depth analysis of helicopter operations in the area. Furthermore, specific recommendations on low noise operations are proposed to local operators in order to reduce their noise footprint and improve helicopter acceptance.
Carbon fiber reinforced epoxy composite stiffened panels are increasingly being used for structural components in large transport rotorcraft. However, problems are arising with high levels of vibration and interior noise due to the increased stiffness-to-density ratio of composites. The current investigation explores the potential of reducing vibrations in carbon/epoxy stiffened panels with the integration of acoustic black holes (ABH), namely features that incorporate a power law thickness taper. The proposed approach involves designing a taper into the thickness of the blade stiffeners as well as the thin plate. Integration of ABHs into the fuselage structure has the potential to reduce broadband vibrations. Multiple parametric studies with either an ABH integrated into the blade stiffener or a grid of ABHs integrated into the plate were conducted, and the tradeoffs between vibration amplitudes, panel mass, and compressive buckling load were examined. Carbon/epoxy panels were fabricated using vacuum-bag-oven processing with out-of-autoclave prepreg and verified to be of good quality. The integrated velocity response, a proxy for the radiated noise from a panel, and compressive buckling were simulated using finite elements. Comparisons were made to experimentally measured data from modal testing and compression buckling testing. Experimental results indicated that when an ABH is integrated into the blade stiffener and 15 ABHs are integrated into the plate in a grid configuration, the panel mass was unchanged, the integrated velocity response decreased by 2.82 dB, and the buckling load increased by 2.9% compared to a baseline non-tapered design.
In the context of developing new rotorcraft dedicated to Advanced Air Mobility, the use of non-conventional propulsive systems such as coaxial counter-rotating rotors, equipped or not with a shroud, is often addressed. The present numerical study falls into this framework. We actually propose to analyse the noise generated by such configuration in stationary conditions with a direct noise simulation realized thanks to a lattice Boltzmann Method. In a first step, the predictions are validated by comparisons with measurements realized in an anechoic chamber for the unshrouded case. A good agreement is observed on both tonal and broadband components of the noise radiated by the propulsive system. Then, the effects of using a shroud around the rotors are investigated numerically. It appears that the use of the shroud significantly increases the relative importance of the broadband noise because of interactions between turbulent wakes with the rear propeller and the duct surface. One can consequently conclude that employing a shroud is not necessarily a good option and this architectural choice must be made carefully. Finally, the impact of the rotor-rotor distance is analysed in both cases with and without shroud. As expected, increasing the distance allows reducing the noise levels but this effect seems to be limited to the tonal noise with almost no impact on the broadband component. Those observations are still valid when the duct is used, but placing the front rotor closer to the shroud inlet makes the shrouded configuration even louder. In a near future, this study will be completed by a complementary test campaign that will allow more validations of the present simulations and confirm or not the resulting conclusions.
Small multirotor vehicles, for example, designed for package delivery, are expected to operate in close proximity to populated areas, raising concerns about noise pollution. This study utilizes acoustic flight tests and computational modeling of an instrumented research hexacopter developed at Penn State to investigate noise generation during takeoff and landing maneuvers, considering varying flight path angles and vehicle speeds. Flight tests were conducted at Mid-State Regional Airport and corresponding predictions were made using the Penn State Noise Prediction System. The predicted vehicle states and noise levels are first validated against the flight test data. The validated model and flight test data are then utilized to study the noise emissions of the aircraft. Measurements and predictions of the acoustic characteristics of the vehicle are analyzed using conventional noise metrics, frequency content, and directivity features. Descent maneuvers are found to be noisier than climb maneuvers. Noise generation decreases with an increase in forward speed during both climb and descent and also with an increase in vertical speed during climb. However, during descent, noise generation increases with an increase in sink rate.
The paper describes activities currently run in the frame of the MOTUS project. The developed methodology allows to correct an acoustic measurement database from simulation results in order to model the noise emis-sion of a new helicopter (H/C), accounting for the introduction of low-noise technology bricks (e.g. rotational speed modification, alternate main rotor blade or Fenestron™ design, engine noise control treatment...). The resulting hemisphere database can then be used to simulate various use-cases, from predicting certification noise levels to computing noise footprints on realistic scenarii, including also Low Noise Procedures. All these results can then be auralized in a subsequent step, to address advanced metrics related to noise annoyance.
Helicopters in high-speed forward flight often generate High-Speed Impulse (HSI) noise, presenting a major challenge for noise control and narrowing the range of helicopter use. This paper proposes a novel method for active noise reduction by adjusting the rotor diameter length, effectively delaying HSI noise onset and reducing HSI noise impact. Utilizing the CLORNS solver and the Ffowcs Williams-Hawkings (FW-H) equation, this approach was tested on the AH-1G rotor through simulation analysis. The study simulated the rotor's dynamic diameter length changes, analyzing the effect of crucial parameters on the sound field. Results indicate that this method significantly controls the production of rotor high-speed pulse noise, achieving a noise reduction of up to 2dB at critical operational points. This research aids in formulating specific rotor noise control laws and expands the range of scenarios for helicopter usage.
Over 4 decades of research works on the nutating, now pericyclic, mechanical transmission have studied its capability to achieve high power density, low noise, and amplified single-stage reduction ratios of up to 100:1. These analytical efforts have culminated into the fabrication of a 50 HP and 32:1 reduction ratio pericyclic transmission prototype. This work introduces the prototype with highlights of the assembly and alignment procedures validated by static testing evaluation. Then, discussion of the dynamic test stand integration, instrumentation, and lubrication components lay out the framework of the high-speed testing plan. Power transmission data validated the pericyclic reduction ratio model. Accelerometer data demonstrated the transmission's capability to operate at low vibration, with peak amplitudes of 1.2 and 2.5 inches per second on the pericyclic gear train and output shaft respectively. Acoustic emission data captured the first 5 harmonics of the shaft speed as well as gear mesh frequencies. The thermal profile showed the shaft bearings remained below 180°F throughout testing, implying safe operating conditions. Finally, strain and vibration data showed the pericyclic gear train maintained load sharing throughout the entire operating envelope, further validating static testing and assembly procedures. The results of the experiments demonstrate the technological readiness of pericyclic mechanical transmission.
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