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

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

This paper presents a comprehensive experimental aeroacoustic investigation of a propeller under turbulence ingestion. Two turbulence-generating passive grids were utilized to quantify the effect of turbulence intensity on the aeroacoustic characteristics of the propeller. A two-component hot-wire anemometry was employed to study the flow field. The flow field results demonstrate a substantial increase in fluctuating velocity components in both axial and radial directions, concentrated at the mid-span of the blade and near the tip, respectively. Energy spectral analysis in the vicinity of the propeller blade shows significantly higher broadband energy levels with multiple haystacking peaks at the harmonics of the blade passage frequency. Far-field noise and load measurement results show that turbulence ingestion has a strong effect on the aerodynamic loading and acoustic response at the blade passage frequency. The directivity of noise radiation at low frequency shows a significant tonal noise contribution. Meanwhile, broadband noise radiation is more dominant at a higher range of frequency, especially when the propeller is operated with turbulence ingestion and at higher advance ratio settings. The far-field noise results revealed the haystacking trends in the low frequency domain of the spectra and are most significant for propellers operating in turbulent inflows. [ABSTRACT FROM AUTHOR]

This paper presents an experimental investigation into the effects of turbulence ingestion on the aerodynamic noise characteristics of rotor blades in edgewise flight. A small-scaled, two-bladed rotor was used in the study. The test utilised two turbulence-generating grids, to generate turbulence inflows with different characteristics, and to compare them to the baseline configuration of the laminar inflow. The experiments were set at forwarding edgewise flight configuration, with freestream inflow velocity ranging from 10 m/s to 22 m/s. Simultaneous measurements of far-field acoustic pressure and load were conducted, along with a separate flow measurement using particle image velocimetry. The acoustic spectra demonstrated a larger contribution to the tonal noise radiation at blade passing frequency, and to the broadband noise radiation at the mid-frequency domain, due to turbulence ingestion. However, the broadband responses in the high-frequency domain were comparable between the tested laminar and turbulence inflow cases, with similar broadband humps featuring in the acoustic spectra. The directivity patterns of the overall sound pressure level showed that the noise radiation was lowest near the plane of rotation, and highest downstream. Turbulence ingestion effects could also be seen in the elevated noise levels throughout the observation positions for the grid inflow cases, particularly at larger advance ratios. [ABSTRACT FROM AUTHOR]

Autorotating samaras such as Sycamore seeds are capable of descending at exceptionally slow speeds and the secret behind this characteristic is attributed to a flow mechanism known as the leading edge vortex (LEV). A stable LEV is known to increase the maximum lift coefficient attainable at high angles of attack and recent studies of revolving and flapping wings have proposed suitable lift and drag coefficient models to characterise the aerodynamic forces of the LEV. For the samara, however, little has been explored to properly test the suitability of these low-order lift and drag coefficient models in describing the aerodynamic forces produced by the samara. Thus, in this paper, we aim to analyse the use of two proposed aerodynamic models, namely, the normal force and Polhamus models, in describing the sectional aerodynamic lift of a samara that is producing a LEV. Additionally, we aim to quantify the aerodynamic parameters that can describe the lift and drag of the samara for a range of wind speed conditions. To achieve this, the study first examined the samara flight data available in the literature, and from it, the profiles of the lift coefficient curves were investigated. Subsequently, a numerical Blade Element-Momentum model (BEM) of the autorotating samara encompassing different lift profiles was developed and validated against a comprehensive set of samara flight data, which were measured from wind tunnel experiments conducted at the University of Bristol for three different Sycamores. The results indicated that both the normal force and Polhamus lift models combined with the normal force drag can be used to describe the two-dimensional lift characteristics of a samara exhibiting an LEV. However, the normal force model appeared to be more suitable, since the Polhamus relied on many assumptions. The results also revealed that the aerodynamic force parameters can vary with windspeed and with the samara wing characteristics, as well as along the span of the samara wing. Values of the lift curve slope, zero-lift drag coefficient, and maximum lift coefficient are predicted and presented for different samaras. The study also showed that the low-order BEM model was able to generate a good agreement with the experimental measurements in the prediction of both rotational speed and thrust. Such a validated BEM model can be used for the initial design of bio-inspired rotors for micro-air vehicles. [ABSTRACT FROM AUTHOR]

Despite current research advances in aircraft dynamics and increased interest in the slowed rotor concept for high-speed compound helicopters, the stability of autogyro rotors remains partially understood, particularly at lightly loaded conditions and high advance ratios. In autorotation, the periodic behavior of a rotor blade is a complex nonlinear phenomenon, further complicated by the fact that the rotor speed is not held constant. The aim of the analysis presented in this article is to investigate the underlying mechanisms that can lead to rotation-flap blade instability at high advance ratios for a teetering autorotating rotor. The stability analysis was conducted via wind tunnel tests of a scaled autogyro model combined with numerical continuation and bifurcation analysis. The investigation assessed the effect of varying the flow speed, blade pitch angle, and rotor shaft tilt relative to the flow on the rotor performance and blade stability. The results revealed that rotor instability in autorotation is associated with the existence of fold bifurcations, which bound the control-input and design parameter space within which the rotor can autorotate. This instability occurs at a lightly loaded condition and at advance ratios close to 1 for the scaled model. Finally, it was also revealed that the rotor inability to autorotate was driven by blade stall. [ABSTRACT FROM AUTHOR]

Recent aircraft designs have considered higher-aspect-ratio wings to reduce induced drag for improved fuel efficiency; however, to remain compliant with airport gate requirements, folding wingtips have been introduced as a solution to the increased wingspan. Recent numerical studies suggest that a folding wingtip solution may be incorporated with spring devices to provide an additional gust loads alleviation ability in flight as well. In this work, a series of low-speed steady and dynamic wind tunnel tests was conducted using a prototype of such a concept. It was found that a folding wingtip with a nonzero relative angle of the folding hinge axis to the streamwise direction could provide gust loads alleviation. The level of load alleviation varied with hinge spring stiffness and lifting condition, with the best performance achieving a 56% reduction in peak loading. [ABSTRACT FROM AUTHOR]

• A rotating propeller mounted on a flexible support is modelled and tested. • A two-bladed propeller results in structural asymmetry and time-dependency. • Frequency splits are observed with increasing rotational speed. • An unstable speed region due to frequency lock-in phenomenon is identified. • The model shows a good agreement with the experiment in time and frequency domains. This work is motivated by the rapid developments of a broad range of electrically powered vertical take-off and landing vehicles which frequently feature two-bladed thrust generating rotor propulsion units. The interaction between a two-bladed rotor and the surrounding structure causes an emergence of the complex dynamics specific to the systems with the time-varying characteristics. This paper introduces a new rotor-structure test rig aimed to support the focused analyses of such systems. The test rig is realized as an idealised configuration consisting of a flexible slender cantilevered beam-like support structure which accommodates a single brushless electrical motor and a two-bladed propeller at its free end. Whilst retaining relative simplicity and scalability, this system allows analysis of complex flexible structure-propeller coupled phenomena, some of which are evidenced in this study. To form a compact and fully defined low-order model suitable for the computationally efficient in-vacuo analysis presented here, a Lagrange-based energy formulation is used to derive the equations of motion. The frequency and time domain techniques, including the complex modal analysis of the linear time-periodic systems, frequency response function and Udwadia-Kalaba method, are used to describe and explain the observed dynamics. The predicted dynamic changes, their impact on the modal interactions and the validity of the finite-sized time-invariant approximating system are successfully demonstrated through correlation with the experiment in the frequency range which encompasses the first four modal families and the rotor speed range spanning up to the first observed experimental instability. Specifically, the study confirms the occurrence of the modal veering and lock-in phenomena, modal splits as well as the strong dominance of the 0th order frequency modulation clusters in the structural response. The results presented in this work show that the proposed experimental platform represents a useful minimal dynamic system with the behavioral characteristics of high significance for the design of novel electrically powered and propeller-driven aircrafts. [ABSTRACT FROM AUTHOR]

Recent studies have considered the use of wings incorporating flared folding wingtips (FFWTs) to enable higher aspect ratios (reducing overall induced drag) while also reducing gust loading and meeting airport operational requirements. This paper presents the first experimental research into the nonlinear dynamic behavior of a wing incorporating an FFWT. Wind-tunnel tests were conducted at a range of velocities below and beyond the linear flutter boundary. The experimental findings are compared with results obtained from continuation and bifurcation analyses on a representative low-fidelity numerical model. The results show that beyond the linear flutter boundary, stable limit cycle oscillations form, which is dependent on the flare angle, are bounded by either geometric or aerodynamic nonlinearities. Also presented is the effect of a wingtip trim tab on the stability boundary of a wing incorporating FFWTs. It is found that the tab angle can significantly alter the stability boundary of the system, indicating that the choice of camber is an important parameter when considering the stability boundary of FFWTs and that a moveable control surface on an FFWT could be used "in flight" to extend the stability boundary of an aircraft. [ABSTRACT FROM AUTHOR]

The dynamics of rotary wing systems are complex and typically feature highly nonlinear and often unsteady aerodynamics, as well as aeroelastic influences. In ongoing efforts to reduce noise and vibration, active devices such as trailing edge flaps on the rotor blades are being studied and these devices can introduce further nonlinearities. Therefore, it is important to be able to evaluate the stability of the overall system with a proper understanding of the global nonlinear behavior. Numerical continuation and bifurcation analysis is well suited to this need, and this paper presents evidence of the technique providing a deeper insight into the stability of helicopter rotor systems than the methods typically adopted in the industry. We first investigate the aeroelastic stability of rotor blades of a medium-sized helicopter in hover and the periodically forced forward flight condition, in both trimmed and untrimmed cases. Then, bifurcation analysis is used to predict the nonlinear stability of a single degree-of-freedom trailing edge flap added to the aeroelastic system, over a range of design parameters. The approach is novel in the context of real-world aeroelastic rotor models, and the emphasis here is on the potential for revealing important multiple-attractor dynamics rather than the study of a particalar system. The results presented highlight the advantages of the approach, both in terms of generating an understanding of local and more global stability, and in the efficiency in obtaining relevant results as parameters vary. [ABSTRACT FROM AUTHOR]

High-aspect-ratio wings are of particular interest to modern aircraft design due to the inherent reduction in induced drag that they provide. However, such wing configurations often come with problems such as increased structural weight and oversized wingspans for existing airport facilities. Unlike conventional folding wingtips, as used on the 777-X, this paper demonstrates the use of semi-aeroelastic hinge devices that enable aircraft incorporating high-aspect-ratio wings not only to fit into airport gates, but also to alleviate aerodynamic loads by allowing floating wingtips to be used in-flight. This study establishes a preliminary design process for such a wing configuration and undertakes a comprehensive sizing process to investigate the impact of the device on wing weight and aircraft performance. For the cases considered, a reduction in wing weight of approximately 25% can be achieved by utilizing the semi-aeroelastic hinge, which can lead to more than 5% improvement in aircraft range. [ABSTRACT FROM AUTHOR]

Recent studies have considered the use of flared folding wingtips (FFWTs) to enable higher aspect ratios--reducing overall induced drag--while reducing gust loading and meeting airport operational requirements. The majority of these analyses have employed linear assumptions despite the presence of large wingtip deformations. In this paper the effect of geometric nonlinearities introduced by an FFWT on the static and dynamic aeroelastic response of a wing is assessed. A geometrically exact expression was formulated to describe the change in the local angle of attack in the chordwise direction across all fold angles. This expression highlighted that the aerodynamic stiffness of an FFWT and, therefore, quantities such as the linear flutter speed are a function of the fold angle and, hence, the attitude of the wing. This effect was verified using both a wind tunnel model of a flexible semispan wing and a numerical model utilizing MSC Nastran, which linearized the model about the equilibrium position of the wingtip. These experiments showed that the geometric nonlinearities introduced due to the large deformations of FFWTs can significantly affect the dynamics of the system, with flutter speeds varying by over 28%, simply by changing the angle of attack of the model. [ABSTRACT FROM AUTHOR]

There has been a growing interest in utilizing flared folding wingtips as an in-flight load alleviation device to enable increased wing spans that meet airport gate limits but with little increase in wing weight. The semi-aeroelastic hinge (SAH) concept is implemented in high-aspect-ratio wings to enable wingtips to be released during severe load cases such as maneuvers and gusts to alleviate the bending moments while maintaining optimum aerodynamic shape for the rest of the flight. In this paper, scaling methods for wings incorporating the SAH are explored, allowing for the development of equivalent scaled unmanned aerial vehicles or wind tunnel models with similar aeroelastic behavior as full-size aircraft. Three scaling approaches are considered in this study, namely, Iso-Froude, Iso-Frequency, and Iso-Strain, where a set of governing nondimensional quantities and scaling factors are determined. Despite the significant nonlinearities resulting from large wingtip fold angles, it is shown that a linear scaling approach can be appropriate for such a wing configuration. Furthermore, the aeroelastic properties of each scaled model are compared to those of the full-scale model, where the best match was obtained from the Iso-Strain model, although it is challenging to meet the required operational conditions. [ABSTRACT FROM AUTHOR]

Recent studies have shown that semi-aeroelastic hinge devices can enable larger aircraft wingspans. Such a device would be folded on the ground to meet airport width restrictions, locked during cruise for optimal aerodynamic performance, and released during maneuvers to alleviate flight loads. In contrast, this paper uses a wind tunnel experiment to study the aeroelastic behavior of floating wingtip fuel tanks. This device consists of a freely floating wingtip with an additional mass attached in the form of a liquid-filled fuel tank. The static aeroelastic results show that altering the fuel tank's filling level and position allows the wingtip to float at an optimal angle for aerodynamic efficiency across various angles of attack and fuel masses. Additionally, this paper shows that, with careful selection of the mass distribution of the wingtip, dynamic load alleviation comparable to the semi-aeroelastic hinge concept can be achieved during turbulence and one-minus-cosine encounters. Furthermore, the effect of fluid motion is shown to reduce incremental loads during random turbulence encounters by up to 10%; however, it has a negligible impact on the response to one-minus-cosine encounters. Such results are also confirmed by a numerical model incorporating a simple reduced-order fluid sloshing model. [ABSTRACT FROM AUTHOR]

Future aircraft designs look set to use longer wingspans to increase the aspect ratio and therefore overall aerodynamic efficiency of the airframe. Such larger wingspans also reduce roll rates and require increased control surface area to achieve the roll maneuver requirements for certification. In this work, the effect of using flared folding wingtips (FFWTs) on the roll performance of simple aircraft wings is investigated numerically and experimentally. A unique rolling rig is designed, manufactured, and tested, with a series of steady roll and transient tests performed for different wingspans, with and without folding wingtips. It is shown that the use of FFWTs on aircraft wings can enable improved aerodynamic performance due to the increased span while also significantly reducing the aerodynamic damping due to roll, such that the roll performance of a wing incorporating FFWTs is comparable to that of one without the additional span. [ABSTRACT FROM AUTHOR]

Using their light wing and through the use of a leading-edge vortex (LEV), autorotating samaras can generate high lift while descending at extremely low speeds. But the flight performance of the samara, with respect to the wide design envelope, is still not well understood. Therefore, this paper aims to experimentally assess how the flight performance of three natural samara wings varies with differing wind speeds and flight conditions. The tests were conducted within a vertical wind tunnel and a novel rig was devised to effectively measure the vertical thrust and rotational rate of the autorotating samara at near frictionless conditions. Furthermore, a bespoke hub was implemented to control the coning and pitch angles of the samara wing. The tests generated a novel and comprehensive set of experimental data of autorotating samaras with changing wind speed, coning and pitch angles. The results also revealed that coning angles between 5 and 15 degrees can increase the vertical thrust produced by the samara by up to a maximum of 9.6 % . Additionally, it was found that samaras operate at extremely low pitch angles between -0.7 and -2.6 degrees to maximise their thrust, even though the conditions are close to the autorotational stability boundary., Competing Interests: We declare we have no competing interests., (© 2024 The Author(s).)