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Reducing the Levelized Cost of Energy is the main objective of wind turbine industry, in particular for the emerging sector of floating offshore turbines. In this work, a novel Economic Nonlinear Model Predictive Control (ENMPC) strategy is developed to maximize the power production of floating offshore wind turbines. The control problem is solved through an indirect method, which achieves the computational efficiency required to apply it in real world cases. A non-linear Reduced Order Model of the floating turbine predicts aerodynamic power, generator temperature and platform motions inside the controller. A set of constraints, including a bound on the generator temperature, the thrust and platform velocities are imposed. Simulations using the open-source engineering tool OpenFAST on the 5MW NREL wind turbine supported by the OC3 spar buoy platform [1] are performed to validate the turbine model and then to assess the controller performances in realistic wind and sea state conditions. With respect to the standard controller, a 4.3% increase of generated power in rated conditions is achieved with a more stable generator temperature.

This paper deals with the development and application of an algorithm aimed at the identification of open-loop transfer functions from closed-loop data. Particularly suited for helicopter applications, it allows the identification of the machine transfer functions even in the presence of a controller. The identification algorithm is based on the knowledge of the time histories of arbitrary external inputs and corresponding control actuation and responses. It is successfully applied to the AW-09 helicopter for the identification of the transfer functions relating pilot inputs to vehicle attitudes and kinematics. Then, from their rational approximation the time domain representation of the helicopter dynamics is achieved. The capability of the final state-space helicopter model to simulate vehicle dynamics is assessed by comparing the provided helicopter time responses to arbitrary inputs, with those obtained through the high-fidelity nonlinear solver formerly used to obtain the database for the identification of the transfer functions.

The diffusion of electric system on-board has started after 2000 due to the need of fulfilling new regulations regarding environmental impact of aviation and of reducing maintenance costs. This process has led to the more-electric aircraft concept design, for which several hydraulic and pneumatic systems have been replaced by electric drives. The main obstacles to the switch to electric propulsion and energy storage/production (all-electric aircraft) are the unfavorable energy-to-weight and power-to-weight ratios of battery and fuel cell that represent the two viable technological solutions for energy storage and power generation. Helicopters make no exception in the transition to electric systems, although with a significant lag with respect to fixed-wing aircraft. Nevertheless, it is expected that, in the future, the technological progress will make electric propulsion a standard solution also for rotorcraft. This paper presents a review of the steps toward electrified aircraft taken by the aviation industry in the past decades (mainly concerning fixed-wing vehicles) and the conceptual all-electric retrofit of a case study lightweight helicopter. Different batteries and fuel cell solutions are examined, and the performance (range and endurance) and economic aspects (life cost and investment plan) are critically analyzed, considering both the currently available technology and the expected future developments.

This paper presents an original approach to structural health monitoring of helicopter rotors based on strain measurement on the blades. Three algorithms are presented, one in the time domain and two in the frequency domain. They are based on the analysis of the discrepancies between the strains on damaged and undamaged blades. Two damage types are considered: a mass unbalance at the tip, and a localized stiffness reduction. The performance of the proposed methods is assessed by numerical simulation using a multibody dynamic solver for comprehensive aeroelastic analysis of rotorcraft. The numerical investigation has highlighted the capability of all the presented techniques to detect damages in the blades, even of small entity, in both steady-flight and soft maneuvers. The reduced prediction accuracy in aggressive maneuvered flight suggests the use of these methods only in steady or quasi-steady flight conditions.

Among the renewable energy technologies, offshore wind energy is expected to provide a significant contribution for the achievement of the European Renewable Energy (RE) targets for the next future. In this framework, the increase of generated power combined with the alleviation of vibratory loads achieved by application of suitable advanced control systems can lead to a beneficial LCOE (Levelized Cost Of Energy) reduction. This paper defines a control strategy for increasing floating offshore wind turbine lifetime through the reduction of vibratory blade and hub loads. To this purpose a Proportional-Integral (PI) controller based on measured blade-root bending moment feedback provides the blade cyclic pitch to be actuated. The proportional and integral gain matrices are determined by an optimization procedure whose objective is the alleviation of the vibratory loads due to a wind distributed linearly on the rotor disc. This control synthesis process relies on a linear, state-space, reduced-order model of the floating offshore wind turbine derived from aero-hydroelastic simulations provided by the open-source tool OpenFAST. In addition to the validation of the proposed controller, the numerical investigation based on OpenFAST predictions examines also the corresponding control effort, influence on platform dynamics and expected blade lifetime extension. The outcomes show that, as a by-product of the alleviation of the vibratory out-of-plane bending moment at the blade root, significant reductions of both cumulative blade lifetime damage and sway and roll platform motion are achieved, as well. The maximum required control power is less than 1% of the generated power.

This work presents the correlation between experimental measurements and numerical simulations concerning the aerodynamics of a hovering rotor operating in ground effect above parallel and inclined surfaces. The capability of a potential-flow, free-wake aerodynamic solver based on a boundary-element-method formulation to simulate in-ground-effect problems is examined in detail. In particular, two approaches for including ground effect are outlined and compared, namely the bounded domain method and the mirror image method. Analyses of rotor aerodynamic performance and flow-field visualization reveal that the experimental data and the numerical predictions given by the solver based on the mirror-image method are in good agreement, thus proving the suitability of this developed computational tool to simulate rotor aerodynamics in the presence of close ground. Concerning the bounded domain method, it has been proven that it is suitable for the evaluation of rotor thrust and wake shape close to the rotor, but does not provide accurate simulation of the flow-field in proximity to the ground and corresponding inflow over the rotor disk.

This paper presents an investigation of the aerodynamic and aeroacoustic interaction of propellers for distributed electric propulsion applications. The rationale underlying the research is related to the key role that aeroacoustics plays in the establishment of the future commercial aviation scenario. The sustainable development of airborne transportation system is currently constrained by community noise, which limits the operations of existing airports and prevents the building of new ones. In addition, the substantial saturation of the existing noise abatement technologies inhibits the further development of the existing fleet, and imposes the adoption of disruptive configurations in terms of airframe layout and propulsion technology. Simulation-based data may help in clarifying many aspects related to the acoustic impact of such innovative concepts. Blended-wing-body equipped with distributed electric propulsion is one of the most promising, due to the beneficial effect of the substantial shielding induced by its geometry. Nevertheless, the novelty of the layout requires a thorough investigation of specific aspect for which no previous experience is available. Herein, the interaction between propellers is analysed for a fixed propeller geometry, as a function of their mutual distance and compared to the acoustic pattern of the isolated one. The aerodynamic results have been obtained using a boundary integral formulation for unsteady, incompressible, potential flows which accounts for the interaction between free wakes and propellers. For the aeroacoustic analyses, the Farassat 1A boundary integral formulation for the solution of the Ffowcs Williams and Hawkings equation has been used. These results provide an insight into the minimum distance between propellers to avoid aerodynamic/aeroacoustic interaction effects, which is an important starting point for the development of distributed propulsion systems.

The scope of this paper is the presentation of the computational methodologies applied in the comprehensive code for rotorcraft developed in the last years at Roma Tre University, along with the assessment of its prediction capabilities focused on flight conditions characterized by strong blade–vortex interactions. Boundary element method approaches are applied for both potential aerodynamics and aeroacoustics solutions, whereas a harmonic-balance/modal approach is used to integrate the rotor aeroelastic equations. The validation campaign of the comprehensive code has been carried out against the well-known HART II database, which is the outcome of a joint multi-national effort aimed at performing wind tunnel measurements of loads, blade deflection, wake shape and noise concerning a four-bladed model rotor in low-speed descent flight. Comparisons with numerical simulations available in the literature for the same test cases are also presented. It is shown that, with limited computational cost, the results provided by the Roma Tre aero-acousto-elastic solver are in good agreement with the experimental data, with a level of accuracy that is in line with the state-of-the-art predictions. The influence of the vortex core modeling on aerodynamic predictions and the influence of the inclusion of the fuselage shielding effect on aeroacoustic predictions are discussed.

Wake inflow modelling is a crucial issue in the development of efficient and reliable computational tools for flight dynamics and aeroelastic analyses of rotorcraft. The aim of this work is the development of state- space, wake inflow modelling techniques for coaxial rotors perturbed from steady flight conditions, based on simulations provided by high-fidelity aerodynamic solvers. These provide relationships of the coefficients of an approximated linear distribution of wake inflow over upper and lower rotor discs either with rotor controls and helicopter kinematic variables or with thrust and in-plane moments generated by the rotors. A three-step identification procedure is proposed. It consists in: (i) evaluation of wake inflow due to harmonic perturbations of rotor kinematics; (ii) determination of the corresponding inflow coefficients (and rotors loads) transfer functions, followed by (iii) their rational approximation and transformation into time domain to obtain the final differential form. Aerodynamic responses are predicted through a boundary element method tool for potential flows, capable of capturing effects due to wake distortion, multi-body interference (like that in coaxial rotors) and severe blade–vortex interactions. The derived state-space models are validated by correlation with inflow distributions directly calculated by the aerodynamic solver, for a coaxial rotor subject to arbitrary perturbations.

This paper deals with the development of a noise propagation tool for the prediction of perceived noise due to aircraft/rotorcraft flying in a non-uniform atmosphere over a non-flat ground. The far-field noise is predicted through a direct ray tracing method, once the near-field aeroacoustic solution, used to define the aircraft/rotorcraft equivalent noise source, is known from a near-field aeroacoustic solver. The atmospheric effects are taken into account in the ray tracing algorithm through standard formulations. A new algorithm capable of taking into account the realistic orography and terrain properties handles the interaction between sound rays and ground. The ground orography is automatically generated by exploiting the open-source Application Programming Interface Open-Elevation, whereas the soil composition is defined using the CORINE Land Cover inventory. The new algorithm, applying the ground absorption on each ray independently and moving the waves interference evaluation to a post-processing space–time interpolation process, allows the inclusion of the contributions of multiple-reflecting rays in the evaluation of the receiver noise. The numerical investigation is aimed at validating the proposed tool against a Fast Field Program solver, as well as at assessing the effects of the realistic description of the terrain on the predictions of perceived noise.

This paper deals with the aeroacoustic analysis of helicopter rotors operating in ground effect. Assuming the ground to be a purely reflective flat surface, the scope is twofold: comparison of different techniques to model its presence and assessment of its effects on the emitted noise field. To this purpose, the mirror image method is applied to evaluate the influence of the ground on the aerodynamic sound sources distributed over the rotor blades (indirect acoustic effect), whereas a noise scattering formulation is used to determine its effects as a sound reflecting surface impinged by the pressure field radiated by the rotor (direct acoustic effect). The numerical investigations examine the impact of the rotor height above the ground both on near and far field noise, and provide highlights on potentialities and drawbacks of the different modeling approaches considered.

Wake inflow modeling is a crucial issue in the development of efficient and high-fidelity simulation tools for rotorcraft flight dynamics and aeroelasticity. This paper proposes a space-time accurate, finite-state, dynamic wake inflow modeling suitable for conventional and innovative rotor configurations, based on simulations provided by high-fidelity aerodynamic solvers. It relates the coefficients of a rotor-disc, radial-azimuthal wake inflow distribution to the rotor kinematic variables, and is capable to take into account the intrinsic periodicity of aerodynamic responses of rotors in steady forward flight. The proposed inflow modeling consists of a three-step process: (i) numerical evaluation of wake inflow due to perturbations of rotor kinematic variables, (ii) determination of transfer functions of multi-harmonic components of a suitable set of inflow coordinates, followed by (iii) their rational approximation and transformation into time domain to derive the differential operators governing multi-harmonic dynamics. The numerical investigation concerns the derivation of finite-state inflow models for single and coaxial rotors, through application of an aerodynamic boundary-element-method solver for potential flows. These are successfully validated by comparison with inflows directly calculated by the aerodynamic tool for arbitrary rotor perturbations.

International audience; This paper deals the activities conducted in the GARTEUR Action Group HC/AG-24 to address the noise scattering of helicopter rotors in the presence of the fuselage. The focus of the Action Group is on the development and the validation of numerical prediction methods and establishing an experimental data base for numerical validations. The numerical methods include boundary element method for solving the Helmholtz equation and computational aeroacoustic method for solving linearized Euler equations. The test activities are conducted in the DLR Acoustic Wind Tunnel in Braunschweig. In the current paper, the results from the numerical simulations, the wind tunnel tests as well as numerical and test comparisons are addressed. The test configurations include the sphere, the sphere with sting support and a finite NACA0012 wing (thanks to share data with a NATO STO group). The acoustic predictions of the scattering by spheres will be compared with the analytic solution to address the influence of the sphere support systems at different wind speeds. The acoustic predictions of the scattering by NACA0012 wing will be analyzed and compared to available test results for different source locations and frequencies.

The aim of this work is the development of dynamic, finite-state modelling of wake inflow generated by kinematic perturbations of rotors in steady flight conditions. Extracted from responses of high-fidelity aerodynamic solvers, it is suited for flight dynamics applications. A three-step identification procedure is proposed: (1) evaluation by a high-fidelity solver of the wake inflow due to harmonic perturbations of rotor kinematics, (2) determination of the corresponding inflow coefficient transfer functions, and (3) rational approximation of the transfer functions. Wake inflow models related to rotor loads (like the well-known Pitt–Peters model) are obtained, as well, as by-products of that proposed. Considering aerodynamic simulations provided by a solver based on a boundary element method for potential flows, the numerical investigation presents the validation of the proposed finite-state wake inflow modelling, along with the examination of identified models related to rotor loads, for a rotor in steady flight conditions subject to arbitrary perturbations.

This work addresses the involuntary interaction between the pilot and the vehicle, with specific reference to helicopters along the vertical (or heave) axis. The problem is described and formulated mathematically. Means of analysis are discussed. Focus is placed on modeling requirements for both the vehicle and the pilot. Open problems and future work lines are sketched.

Offshore wind energy is expected to provide a significant contribution to the achievement of the European Renewable Energy targets. One of the main technological issues affecting floating offshore wind turbines concerns generated power fluctuations and structural fatigue caused by sea-wave/platform interactions. This paper presents a fully-coupled aero/hydro/servo-mechanic model for response and control of floating offshore wind turbines in waves, suitable for preliminary design. The wind-turbine is described by a multibody model consisting of rigid bodies (blades and tower) connected by hinges equipped with springs and dampers (for realistic low-frequency simulation). The aerodynamic loads are evaluated through a sectional aerodynamic approach coupled with a wake inflow model. A spar buoy floating structure supports the wind turbine. The hydrodynamic forces are evaluated through a linear frequency-domain potential solver, with the free surface deformation effects included through a reduced-order, state-space model. An optimal controller is identified and applied for rejection of annoying fluctuations of extracted power and structural loads. The developed comprehensive model has been successfully applied to a floating version of the NREL 5 MW wind turbine for stability analysis, as well as for the analysis of uncontrolled and controlled responses to regular and irregular short-crested sea waves. The proposed controller, based on the combined use of blade pitch and generator torque as control variables and the application of an observer for non-measurable aerodynamic and hydrodynamic states estimation, has been demonstrated to be effective in a wide frequency range for alleviation of both generated power fluctuations and vibratory loads.

Traditional helicopter blades are subject to significant deformations, which influence control forces and moments, as well as the helicopter aeroelastic and aeroacoustic behavior. Thus, the knowledge of rotor elastic states could help improving flight control efficiency, and reducing vibration level and acoustic emissions of next-generation helicopters. This paper presents an original and computationally efficient modal approach aimed at dynamic shape sensing of helicopter rotor blades. It is based on strain measurements in a limited number of points over the blade surface. Although the algorithm is based on the cascaded solution of linear algebraic equations, much like other modal-based algorithms, it is able to reconstruct nonlinear, moderate lag, flap and torsional deflections, which are typical in helicopter structural dynamics. The algorithm is tested on non-rotating and rotating hingeless blades through numerical simulations based upon a multibody dynamics solver for general nonlinear comprehensive aeroelastic analysis. Its capabilities are assessed against those of classical modal approaches. Numerical investigations show that the proposed algorithm is reliable, accurate and robust to measurement noise.

This paper investigates different methodologies for the evaluation of the acoustic disturbance emitted by helicopter’s main rotors during unsteady maneuvers. Nowadays, the simulation of noise emitted by helicopters is of great interest to designers, both for the assessment of the acoustic impact of helicopter flight on communities and for the identification of optimal-noise trajectories. Typically, the numerical predictions consist of the atmospheric propagation of a near-field noise model, extracted from an appropriate database determined through steady-state flight simulations/measurements (quasi-steady approach). In this work, three techniques for maneuvering helicopter noise predictions are compared: one considers a fully unsteady solution process, whereas the others are based on quasi-steady approaches. These methods are based on a three-step solution procedure: first, the main rotor aeroelastic response is evaluated by a nonlinear beam-like rotor blade model coupled with a boundary element method for potential flow aerodynamics; then, the aeroacoustic near field is evaluated through the 1A Farassat formulation; finally, the noise is propagated to the ground by a ray tracing model. Only the main rotor component is examined, although tail rotor contribution might be included as well. The numerical investigation examines the differences among the noise predictions provided by the three techniques, focusing on the assessment of the reliability of the results obtained through the two quasi-steady approaches as compared with those from the fully unsteady aeroacoustic solver.

The aim of this paper is the evaluation of several pilot models found in the literature, suited for helicopter pilot-assisted and pilot-induced oscillations analyses. Three main topics are discussed: (i) sensitivity of rotorcraft-pilot couplings simulations on the application of the different pilot models available in the literature; (ii) effect of vehicle modeling on active pilot modeling; (iii) effects of interactions between active and passive pilot models. The focus is on hovering flight, where a specific adverse rotorcraft-pilot coupling phenomenon, the vertical bounce, may occur. Pilot models are coupled with a comprehensive aeroservoelastic model of a mid-weight helicopter. The numerical investigations are performed in frequency domain, in terms of eigenanalysis and frequency response analysis.

Rotorcraft–pilot coupling (RPC) denotes inter- play between pilot and helicopter (or tiltrotor) that may give rise to adverse phenomena. These are usually divided into two main classes: pilot-induced oscillations driven by flight dynamics and behavioural processes, and pilot- assisted oscillations (PAO) caused by unintentional actions of pilot on controls, owing to involuntary reaction to seat vibrations. The aim of this paper is the development of mathematical helicopter models suited for analysis of RPC phenomena. In addition to rigid-body dynamics, RPCs (especially PAO) are also related to fuselage structural dynamics and servoelasticity; however, a crucial role is played by main rotor aeroelasticity. In this work, the aeroelastic behaviour of the main rotor is simulated through a novel finite-state modeling that may conve- niently be applied for rotorcraft stability and response analyses, as well as for control synthesis applications. Numerical results, first are focused on the validation of the proposed novel main rotor model, and then present appli- cations of the developed comprehensive rotorcraft model for the RPC analysis of a Bo-105-type helicopter. Specif- ically, these deal with the stability of vertical bouncing, which is a PAO phenomenon caused by coupling of ver- tical pilot seat acceleration with collective control stick, driven by inadvertent pilot actions. Further, considering quasi-steady, airfoil theory with wake inflow correction and three-dimensional, potential-flow, boundary element method approaches, the sensitivity of PAO simulations to different aerodynamic load models applied within the main rotor aeroelastic operator is also investigated.

"This paper discusses the aeroelastic interaction between the helicopter and the pilot called collective bounce. The problem is mostly studied in the time domain, using the multibody system dynamics approach to model the dynamics of the vehicle and the aeroelasticity of the main rotor and a linear or quasilinear transfer function approach for the voluntary and involuntary dynamics of the pilot. Different models are considered for the aerodynamic forces acting on the rotor, ranging from blade-element\/momentum theory to a boundary-element method used independently and in cosimulation with the multibody model. The problem is analyzed in hover and forward flight, highlighting modeling requirements and the sensitivity of the stability results to a variety of parameters of the problem."

Numerical tools aiming at the evaluation of ground acoustic impact of helicopters typically rely on databases given in terms of acoustic disturbance over hemispheres surrounding the helicopter (noise hemispheres). These are evaluated for a discrete number of steady flights falling within the flight envelope. The objective of the present work is the identification of flight parameters to be considered for the characterization of noise hemispheres, particularly when related to unsteady maneuvers. To this purpose, different approaches based on steady flight aeroelastic/aerodynamic/aeroacoustic predictions are examined for assessing their capability of simulating the acoustic impact of helicopters in arbitrary unsteady flight. The numerical investigation demonstrates that at least three parameters, including disk loading, are required to adequately characterize noise hemispheres. Conversely, the similarity of kinematic parameters alone may yield steady flight acoustic predictions poorly correlated with those obtained for unsteady maneuvers.

The aim of the paper is the assessment of a methodology for the identification of a helicopter rotor aeroelastic operator in state-space form relating airframe motion and blade controls perturbations to corresponding hub loads, as extracted from a high-fidelity rotor aeroelastic solver. It is suited for helicopter flight dynamics stability and real-time commands response analyses, as well as for control laws synthesis.The identification method consists of a three-step process, starting with evaluation of responses to small perturbation harmonic inputs, followed by their spectral analysis and rational form approximation of corresponding transfer functions. Considering a Bo-105-type helicopter, numerical investigation is focused on: analysis of critical parameters affecting accuracy and efficiency of transfer functions identification, interpretation of the additional states introduced by rational approximation, validation of rotor finite-state aeroelastic representation.

Boundary Element Method (BEM) is a widely used technique for the aerodynamic analysis of lifting bodies, due to the good compromise between fidelity and computational costs. This is especially true in rotorcraft applications, where the effects of the rotor wake play a crucial role in the evaluation of aerodynamic loads, and CFD solvers suffer for dissipation problems. However, potential formulations are unable to simulate many phenomena, like stall or separation around non-aerodynamic (bluff) bodies. This limits they application to the aerodynamic analysis of complete aircraft/rotorcraft configurations, where massive separations might arise downstream of the fuselage. In this paper, several potential-flow, BEM methodologies are proposed, with the objective of overcoming these limitations, at the cost of receiving information concerning the position of separation point and/or the value of pressure in the separated-flow region. Numerical results concerning the assessment of the proposed BEM techniques are presented: these consist in comparisons with data available in the literature for simple-geometry bodies, and with measurements obtained by experimental tests performed in a semi-open wind tunnel facility on a light-weight helicopter fuselage.

This paper deals with the use of photogrammetry for the experimental identification of structural and inertial properties of helicopter rotor blades4. The identification procedure is based upon theoretical/numerical algorithms for the evaluation of mass and flexural stiffness distributions which are an extension of those proposed in the past by Larsen, whereas the torsional properties (stiffness and shear center position) are determined through the Euler–Bernoulli beam theory. The identification algorithms require the knowledge of the blade displacement field produced by known steady loads. These data are experimentally obtained through photogrammetric detection technique, which allows the identification of 3D coordinates of labeled points (markers) on the structure through the correlation of 2D digital photos. Indeed, the displacement field is simply evaluated by comparing the markers positions on the loaded configuration with those on the reference one. The proposed identification procedure, numerically and experimentally validated in the past by the authors, has been here applied to the structural characterization of two main rotor blades, designed for ultra-light helicopters. Strain gauges measurements have been used to assess the accuracy of the identified properties through natural frequencies comparison as well as to evaluate the blades damping characteristics.

An optimal procedure for the design of rotor blade that generates low vibratory hub loads in nonaxial flow conditions is presented and applied to a helicopter rotor in forward flight, a condition where vibrations and noise become severe. Blade shape and structural properties are the design parameters to be identified within a binary genetic optimization algorithm under aeroelastic stability constraint. The process exploits an aeroelastic solver that is based on a nonlinear, beam-like model, suited for the analysis of arbitrary curved-elastic-axis blades, with the introduction of a surrogate wake inflow model for the analysis of sectional aerodynamic loads. Numerical results are presented to demonstrate the capability of the proposed approach to identify low vibratory hub loads rotor blades as well as to assess the robustness of solution at off-design operating conditions. Further, the aeroacoustic assessment of the rotor configurations determined is carried out in order to examine the impact of low-vibration blade design on the emitted noise field.

Wake inflow modelling is a crucial issue in the development of efficient and reliable computational tools for flight mechanic and aeroelastic analysis of rotorcraft. The aim of this work is the development of a finite-state, dynamic wake inflow modelling for rotors in steady flight conditions, based on the simulations provided by aerodynamic solvers of arbitrary accuracy. It is related to the well-known Pitt-Peters model, which relates the coefficients of an approximated linear distribution of the wake inflow over the rotor disc to thrust and roll and pitch moments. A three-step identification procedure is proposed. It consists in: (i) evaluation of wake inflow due to harmonic perturbations of rotor kinematics, (ii) determination of the corresponding inflow coefficient transfer functions, and (iii) their rational approximation. Wake inflow models related to both flight dynamics statespace variables and rotor loads are presented and discussed. Wake inflow models predicted through aerodynamic solutions provided by a boundary element method for potential flows are compared with the Pitt-Peters one, and validated by correlation with the inflow directly calculated by the aerodynamic solver, for a rotor subject to arbitrary perturbations.

This paper describes an innovative use of photogrammetric detection techniques to experimentally estimate structural/inertial properties of helicopter rotor blades. The identification algorithms for the evaluation of mass and flexural stiffness distributions are an extension of the ones proposed by Larsen, whereas the procedure for torsional properties determination (stiffness and shear center position) is based on the Euler-Prandtl beam theory. These algorithms rely on measurements performed through photogrammetric detection, which requires the collection of digital photos allowing the identification of 3D coordinates of labeled points (markers) on the structure through the correlation of 2D pictures. The displacements are evaluated by comparing the positions of markers in loaded and reference configuration. Being the applied loads known, the structural characteristics can be directly obtained from the measured displacements. The accuracy of the proposed identification algorithms has been firstly verified by comparison with numerical and experimental data, and then applied to the structural characterization of two main rotor blades, designed for ultra-light helicopter applications.

The terminology ‘rotorcraft-pilot coupling’ denotes phenomena arising from interaction between pilot and rotorcraft. Among these, the present work deals with ‘pilot-assisted oscillations’ that derive from unintentional pilot actions on controls due to seat vibrations, and are strictly related to rotor-aeroelasticity/airframe-structural-dynamics coupling, with involvement of blade control actuator dynamics. Focusing the attention on helicopters, a comprehensive rotorcraft model is developed and applied, with main rotor unsteady aerodynamics described in state-space form. This makes it particularly suited for stability and frequency-response analysis, as well as control applications. Numerical investigations address two critical rotorcraft-pilot coupling aeroelastic issues: stability of vertical bouncing and gust response in hovering. Results from main rotor unsteady aerodynamics modelling are compared with widely-used quasi-steady aerodynamics predictions. These suggest that, for accurate RPC/PAO phenomena predictions, mathematical modelling should include the three-dimensional, unsteady-flow effects, and that the pilot-in-the-loop passive behaviour produces a beneficial effect on the load factor generated by gust encountering.

Understanding, predicting and supressing the inadvertent aircraft oscillations caused by Aircraft/Rotorcraft Pilot Couplings (A/RPC) is a challenging problem for designers. These are potential instabilities that arise from the effort of controlling aircraft with high response actuation systems. The present paper reviews, updates and discusses desirable practices to be used during the design process for unmasking A/RPC phenomena. These practices are stemming from the European Commission project ARISTOTEL Aircraft and Rotorcraft Pilot Couplings – Tools and Techniques for Alleviation and Detection (2010–2013) and are mainly related to aerodynamic and structural modelling of the aircraft/rotorcraft, pilot modelling and A/RPC prediction criteria. The paper proposes new methodologies for precluding adverse A/RPCs events taking into account the aeroelasticity of the structure and pilot biodynamic interaction. It is demonstrated that high-frequency accelerations due to structural elasticity cause negative effects on pilot control, since they lead to involuntary body and limb-manipulator system displacements and interfere with pilot's deliberate control activity (biodynamic interaction) and, finally, worsen handling quality ratings.

This paper discusses the impact of different models of aerodynamic loads on rotorcraft-pilot couplings stability using a robust stability analysis approach. The aeroelasticity of the main rotor of a helicopter is formulated using aerodynamic models based on the blade ele- ment/momentum theory and boundary element method coupled to a finite element model of the blade. The resulting linearized models are used to determine stability limits according to the generalized Nyquist criterion, associated with the accelerations of the pilot’s seat caused by the involuntary action of the pilot on the control in- ceptors. The resulting stability curves are discussed considering examples of involuntary pilot transfer functions from the literature.

This paper deals with a computational aeroelastic tool aimed at the analysis of the response of rotary wings in arbitrary steady motion. It has been developed by coupling a nonlinear beam model for blades structural dynamics with a potential-flow boundary integral equation solver for the prediction of unsteady aerodynamic loads around three-dimensional, lifting bodies. The Galerkin method is used for the spatial integration of the resulting differential aeroelastic system, whereas the periodic blade response is determined by a harmonic balance approach. This aeroelastic model yields a unified approach for aeroelastic response and blade pressure prediction, that may conveniently be used for aeroacoustic purposes. It is able to examine configurations where blade–vortex interactions occur. Numerical results show the capability of the aeroelastic tool to evaluate blade response and vibratory hub loads for a helicopter main rotor in level and descent flight conditions, and examine the efficiency and robustness of the different numerical solution algorithms that may be applied in the developed aeroelastic solver. Comparisons among aeroelastic predictions based on different aerodynamic models are also presented.

"Fixed and rotary wing pilots alike are familiar with potential instabilities or with annoying limit cycle. oscillations that arise from the effort of controlling aircraft with high response actuation systems. Understanding, predicting and suppressing these inadvertent and sustained aircraft oscillations, known as aircraft (rotorcraft)-pilot couplings (A\/RPCs) is a challenging problem for the designers. The goal of the. present paper is to give an overview on the state-of-the-art in RPC problem, underlining the future. challenges in this field. It is shown that, exactly as in the case of fixed wing APCs, RPCs existed from the beginning of rotorcraft development and that the problem of eliminating them is not yet solved: the current rotorcraft modelling for RPC analysis is rather limited to the particular case analysed and there is. a lack of quantitative pilot behavioural models to analyse RPCs. The paper underlines the importance of involuntary pilot control actions, generally attributed to biodynamic couplings in predicting RPCs in rotorcraft. It is also shown that recent experiences demonstrate that modern rotorcraft seem to embed. tendencies predisposing the flight control system FCS system towards dangerous RPCs. As the level of automation is likely to increase in future designs, extending to smaller aircraft and to different kinds of. operation, the consequences of the pilot ‘fighting’ the FCS system and inducing A\/RPCs needs to be eradicated. In Europe, the ARISTOTEL project (2010–2013) has been launched with the aim of understanding and predicting modern aircraft's susceptibility to A\/RPC. The present paper gives an overview of future challenges to be solved for RPC-free design and some new solutions herein.. "

Rotorcraft-pilot couplings denote interactions between pilot and helicopter (tiltrotors) that may become adverse. They are usually divided into two main classes of phenomena: that including Pilot Induced Oscillations (PIO) phenomena driven by flight dynamics and behavioural processes, and the one conventionally named Pilot Assisted Oscillations (PAO) which is caused by unintentional actions of pilot on controls, due to his involuntary reaction to seat vibrations. The aim of this paper is the development of mathematical helicopter models suited for analysis of PAO phenomena. PAO are strictly related to the structural dynamics of the fuselage and to aervoelasticity, but a crucial role is played by main rotor aeroelasticity. This paper presents a finite-state model of main rotor aeroelastic behavior that may conveniently be applied for PAO stability and response analysis, as well as for control applications aimed at PAO alleviation. It is validated, and its sensitivity to the aerodynamic modeling used within the aeroelastic operator is examined. Further, it is coupled with fuselage dynamics, servo-elastic and pilot models in order to carry out a numerical investigation concerning the stability analysis of vertical bouncing, which is a type of PAO instability which might be caused by the coupling, through the pilot, of vertical acceleration of pilot seat with collective control stick.

This paper deals with the computational analysis of acoustic fields generated by helicopter rotors when Blade-Vortex Interactions (BVI) occur. The prediction procedure starts from the determination of the steady periodic blade deformations. Then, the BVI-affected, unsteady aerodynamics solution is obtained by a potential-flow boundary integral formulation suited for aeronautical configurations experiencing blade-wake impingements. It is applicable to blades with arbitrary shape and motion and evaluates both wake distortion and blade pressure field. Finally, the noise field radiated by the rotor is computed through an aeroacoustic tool based on the Ffowcs Williams and Hawkings equation. The numerical investigation examines the sensitivity of BVI noise prediction on the aeroelastic model applied for the calculation of blade deformations, and assesses the accuracy of the results through correlation with experimental data concerning a helicopter main rotor in descent flight. Noise predicted is examined in terms of both acoustic pressure signatures and noise radiation characteristics.

The acoustic field generated by a helicopter main rotor experiencing blade-vortex interaction (BVI) during a descent flight path is examined. The prediction procedure starts from the determination of the aeroelastic steady periodic solution. Then, a boundary integral formulation for the velocity potential suited for configurations where stong wake/blade impingement occurs is applied. It is fully three-dimensional, can be applied to blades with arbitrary shape and motion and performs the calculation of both wake shape and blade pressure field. Finally, the noise field generated by the helicopter rotor is evaluated through an aeroacoustic tool based on the Ffowcs Williams and Hawkings equation. The numerical investigation discusses the sensitivity of BVI noise prediction on the aeroelastic models applied for the calculation of blade steady periodic deformations. The effects of the different blade deformations given by the aeroelastic solvers considered are examined both in terms of local acoustic signatures and in terms of noise radiation characteristics.

This paper discusses the impact of different aerodynamics models on rotorcraft-pilot couplings stability using a robust stability analysis approach. Models based on blade element/momentum theory and boundary element method, coupled to a finite element model of the blade, are used to formulate the aeroelasticity of the main rotor of a helicopter. The main rotor is coupled to a dynamic model of the airframe that also includes the dynamics of the control system. The resulting linearized models are used to determine stability limits according to the generalized Nyquist criterion, associated with the accelerations of the pilot's seat caused by the involuntary motion of the control inceptors. The resulting stability curves are discussed considering examples of involuntary pilot transfer functions from the literature.

The alleviation of gusts effects on a tiltrotor in aeroplane and helicopter operation modes obtained by an optimal control methodology based on the actuation of elevators, wing flaperons and swashplate is examined. An optimal observer for state estimate is included in the compensator synthesis, with the Kalman-Bucy filter applied in the presence of stochastic noise. Tiltrotor dynamics is simulated through an aeroelastic model that couples rigid-body motion with wing and proprotor structural dynamics. An extensive numerical investigation examines effectiveness and robustness of the applied control procedure, taking into account the action of both deterministic and stochastic vertical gusts. In addition, a passive pilot model is included in the aeroelastic loop and the corresponding effects on uncontrolled and controlled gust response are analysed.