33,986 research outputs found

    Investigation of the Effect of Blade Sweep on Rotor Vibratory Loads

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    The effect of helicopter rotor blade planform sweep on rotor vibratory hub, blade, and control system loads has been analytically investigated. The importance of sweep angle, sweep initiation radius, flap bending stiffness and torsion bending stiffness is discussed. The mechanism by which sweep influences the vibratory hub loads is investigated

    Design of helicopter rotor blades for optimum dynamic characteristics

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    The possibilities and limitations of tailoring blade mass and stiffness distributions to give an optimum blade design in terms of weight, inertia, and dynamic characteristics are discussed. The extent that changes in mass of stiffness distribution can be used to place rotor frequencies at desired locations is determined. Theoretical limits to the amount of frequency shift are established. Realistic constraints on blade properties based on weight, mass, moment of inertia, size, strength, and stability are formulated. The extent that the hub loads can be minimized by proper choice of E1 distribution, and the minimum hub loads which can be approximated by a design for a given set of natural frequencies are determined. Aerodynamic couplings that might affect the optimum blade design, and the relative effectiveness of mass and stiffness distribution on the optimization procedure are investigated

    Hingeless helicopter rotor with improved stability

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    Improved stability was provided in a hingeless helicopter rotor by inclining the principal elastic flexural axes and coupling pitching of the rotor blade with the lead-lag bending of the blade. The primary elastic flex axes were inclined by constructing the blade of materials that display non-uniform stiffness, and the specification described various cross section distributions and the resulting inclined flex axes. Arrangements for varying the pitch of the rotor blade in a predetermined relationship with lead-lag bending of the blade, i.e., bending of the blade in a plane parallel to its plane of rotation were constructed

    Control load envelope shaping by live twist

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    Rotor control systems experience a rapid load growth resulting from retreating blade stall during flight conditions of high blade loading or airspeeds. An investigation was undertaken to determine the effect of changing blade torsional properties over the rotor flight envelope. The results of this study show that reducing the blade stiffness to introduce more blade live twist significantly reduces the large retreating blade control loads, while expanding the flight envelope and reducing retreating blade stall loads

    Results of a parametric aeroelastic stability analysis of a generic X-wing aircraft

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    This paper discusses the trends in longitudinal dynamic aeroelastic stability of a generic x-wing aircraft model with design parameter variations. X-wing rotor blade sweep angle, ratio of blade mass to total vehicle mass, blade structural stiffness cross-coupling and vehicle center-of-gravity location were parameters considered. The typical instability encountered is body-freedom flutter involving a low frequency interaction of the first elastic mode and the aircraft short period mode. Parametric cases with the lowest static margin consistently demonstrated the highest flutter dynamic pressures. As mass ratio was increased, the flutter boundary decreased. The decrease was emphasized as center-of-gravity location was moved forward. As sweep angle varied, it was observed that the resulting increase in forward-swept blade bending amplitude relative to aft blade bending amplitude in the first elastic mode had a stabilizing effect on the flutter boundary. Finally, small amounts of stiffness cross-coupling in the aft blades increased flutter dynamic pressure

    Effect of rotor stiffness and lift offset on the aeroacoustics of a coaxial rotor in level flight

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    The acoustic characteristics of a twin contra-rotating coaxial rotor configuration with significant flapwise stiffness are investigated in steady forward flight. The Vorticity Transport Model is used to simulate the aerodynamics of the rotor system and the acoustic field is determined using the Ffowcs Williams-Hawkings equation implemented using the Farassat-1A formulation. Increasing the hub stiffness alters the strengths of the blade vortex interactions, particularly those between the upper and lower rotors, and affects the intensity and directivity of the blade vortex interaction noise produced by the system. The inter-rotor blade vortex interaction on the advancing side of the lower rotor is the principal source of the most intensively focused noise that is generated by a conventionally articulated coaxial rotor system. For stiffened coaxial rotors, this particular inter-rotor blade vortex interaction is weakened as a result of a broad redistribution in lateral loading, yielding a reduction in the intensity of the noise that is produced by this interaction. The spanwise distribution of loading on the rotors of a stiffened coaxial system can be modified further by altering the lateral partition of lift (or lift offset). It is shown that decreasing the lift offset has the effect of counteracting the redistribution of loading due to flapwise stiffness and hence increases the blade vortex interaction noise as well as the power consumed by the rotor. Conversely, a reduction in both the power consumption and the blade vortex interaction noise is observed if the lift offset is increased, with the maximum benefit of lift offset being achieved at high speed. The computational results suggest that the noise from the dominant inter-rotor blade vortex interaction can be ameliorated through the use of lift offset control on stiffened coaxial systems, to the extent that the noise produced by this interaction can be made to be comparable to that produced by the other, weaker interactions between the two rotors of the system

    Analysis of potential helicopter vibration reduction concepts

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    Results of analytical investigations to develop, understand, and evaluate potential helicopter vibration reduction concepts are presented in the following areas: identification of the fundamental sources of vibratory loads, blade design for low vibration, application of design optimization techniques, active higher harmonic control, blade appended aeromechanical devices, and the prediction of vibratory airloads. Primary sources of vibration are identified for a selected four-bladed articulated rotor operating in high speed level flight. The application of analytical design procedures and optimization techniques are shown to have the potential for establishing reduced vibration blade designs through variations in blade mass and stiffness distributions, and chordwise center-of-gravity location

    Development of an aeroelastic stability boundary for a rotor in autorotation

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    <p>For the present study, a mathematical model AMRA was created to simulate the aeroelastic behaviour of a rotor during autorotation. Our model: Aeroelastic Model of a Rotor in Autorotation (AMRA) captures transverse bending and teeter, torsional twist and lag-wise motion of the rotor blade and hence it is used to investigate couplings between blade flapping, torsion and rotor speed. Lagrange’s method was used for the modelling of blade flapping and chord-wise bending. Torsional twist of the rotor blade was modelled with the aid of finite element method (FEM), and blade transverse bending could also be modelled in FEM. The model can switch between using a full FEM model for bending and torsion, or a FEM model for torsion and simple blade teeter, depending on the complexity that the user requires.</p> <p>The AMRA model was verified against experimental data obtained during a CAA sponsored flight test programme of the G-UNIV autogyro. Published results of modal analysis of helicopter rotor blades and other data published in open literature were used to validate the FEM model of the rotor blade. The first torsional natural frequency of the ’McCutcheon’ rotor blades was measured with the aid of high-speed camera and used for validation of the FEM model of blade torsional twist. As a further verification of the modelling method, Aérospatiale Puma helicopter rotor blade data were compared on a Southwell plot showing comparison between experimental results and AMRA estimation.</p> <p>The aeromechanical behaviour of the rotor during both axial flight and forward flight in autorotation was investigated. A significant part of the research was focused on investigation of the effect of different values of torsional and flexural stiffness, and the relative positions of blade shear centre/elastic axis and centre of mass of the blade on stability during the autorotation.</p> <p>The results obtained with the aid of the model demonstrate the interesting, and unique, characteristics of the autorotative regime - with instabilities possible in bending and torsion, but also in rotorspeed. Coupled rotor speed/flap/twist oscillations (flutter and divergence) occur if the torsional stiffness of the blade is lower than a critical value, or if the blade centre of mass is significantly aft of the blade twisting axis, as is the case in helicopter pitch-flap flutter. The instability shown here, however, is specific to the autogyro, or autorotating rotor, as it is coupled with rotorspeed, and so differs from both helicopter rotor flutter and fixed-wing flutter. The coupling with rotorspeed allows a combined flutter and divergence instability, where the rotor begins to flutter in rotorspeed, teeter angle and torsional twist and, once the rotorspeed had dropped below a critical value, then moves into divergence in flap and rotorspeed. It was found that the aeroelastic behaviour of a rotor in autorotation is significantly affected by the strong coupling of blade bending stiffness and teeter angle with rotorspeed, and the strong coupling between blade aeroelastic twist and rotor torque.</p&gt

    An experimental investigation of the flap-lag stability of a hingeless rotor with comparable levels of hub and blade stiffness in hovering flight

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    An experimental investigation of the flap-lag stability of a hingeless rotor in hovering flight is presented and discussed. The rotor blade and hub configuration were selected such that the hub and blade had comparable levels of bending stiffness. Experimental measurements of the lag damping were made for various values of rotor rotational speed and blade pitch angle. Specifically at a blade pitch angle of 8 deg at three-quarters radius, the lag damping was determined over a range of rotational speeds from 200 RPM to 320 RPM and also over a range of blade pitch angles from 0 deg to 8 deg
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