This paper introduces a novel concept of visual-haptic display for situational awareness improvement for crowded and low altitude airspace situations. The visual augmentation display that constitutes of Virtual Fences delimiting no-fly zones, and a specific tri-dimensional highlight graphics that enhances visibility of other remotely piloted or autonomous agents, as well as conventional manned aircraft operating in the area is presented first. Then the Shared Control paradigm and the Haptic Force generation mechanism, based on a Proportional-Derivative-like controller applied to repulsive forces generated by the Virtual Fences and other UAVs are introduced and discussed. Simulations with 26 pilots were performed in a photo-realistic synthetic environment showing that the combined use of Visual-haptic feedback outperforms the Visual Display only in helping the pilot keeping a safe distance from no-fly zones and other vehicles.

A rigorous proof of global exponential stability is derived for a magnetic control law that drives a rigid satellite toward a pure-spin condition around a prescribed principal axis of inertia with a desired angular rate. The proof represents an extension and a generalization of a method proposed by two of the authors of the present note for demonstrating global asymptotic stability for a B-dot-like control law that detumbles a spacecraft to rest by means of magnetic actuators only. The proof of stability in the case of acquisition of a non-zero desired angular rate pure spin state is derived in terms of robustness of the global exponential stability of a nominal system by means of generalized exponential asymptotic stability in variations (GEASV) tools. To this aim, the error dynamics equation is first derived in the classical form of a nominal system perturbed by a vanishing perturbation term. Then, after proving the generalized exponential stability for the nominal system, such result is extended to the perturbed system. As a further contribution, an approach for the choice of the control law gain is proposed to the present application, thus allowing to perform the acquisition of the desired pure-spin condition in quasi-minimum time from arbitrary initial tumbling conditions. Stability and performance of the approach are extensively tested by means of numerical simulation.

The use of a hybrid powertrain for a conventional single main rotor helicopter is investigated, with the objective of assessing its feasibility and its potential impact on improving safety, especially for single-engine rotorcraft. The study is focused on the characteristics of the powertrain and required battery pack. It is based on a simple analysis of power required in forward flight and the estimate of the total energy required for a powered landing maneuver after thermal engine failure. Current technologies are considered as well as expected improvements, especially as far as energy density and power density of the battery are concerned. The latter analysis is based on current trends for battery and motors technologies, in order to determine the technological breakthrough limit.

The paper deals with the assessment of the reliability of simplified rotorcraft models in the evaluation of maneuvering potential using inverse simulation. Inverse solutions obtained for the same maneuver from models of different complexity and fidelity are compared with the objective of identifying the most appropriate model for a consistent evaluation of vehicle handling qualities at the minimum computational cost. A total of nine rotor blade models, three main rotor inflow models, and two fuselage aerodynamic databases are combined for deriving 10 different helicopter simulation models analyzed in five maneuvers. Inverse solutions are obtained by means of an integration method for hurdle-hop, slalom, and lateral repositioning maneuvers. Once the maneuvers are solved for the baseline rotorcraft model, the uncertainty associated with the command law identified by means of simpler models is determined. The quantitative evaluation of model reliability from a set of simulation tests is further analyzed on a pop-up-pop-down maneuver and a 180 deg fast turn for validating the approach. The results show that uncertainty intervals are correctly identified, although with some degree of conservativeness when less demanding maneuvers are dealt with.

The paper presents a technique for assessing the reliability of a set of helicopter models in predicting the required control action when executing a given (set of) manoeuvre task(s). An inverse simulation algorithm based on the integration method is used in order to derive the time–history of control commands necessary for following a prescribed flight path. A quantitative comparison between the control laws thus obtained is performed in order to assess the reliability of lower order models with respect to the baseline, most complete one, adopted as a reference for the analysis. Two metrics are developed, one for evaluating a global error level in the definition of the required control law, and a second one for the identification of the uncertainty in the control action when adopting a lower order model. A total of 9 main rotor dynamic models, 3 main rotor inflow models and 3 fuselage aerodynamic databases are combined in order to obtain as many as 13 different helicopter simulation models, analyzed in 3 manoeuvres: a hurdle–hop, a slalom and a lateral reposi- tioning. The evaluation of the uncertainty associated with the command law identified by means of simpler models is thus performed in terms of the considered metrics, the validity of which is then tested on two more manoeuvres: a pop–up–pop–down manoeuvre and a 180 deg fast turn. The results show that most of the times uncertainty intervals are correctly identified, although with some degree of conservativeness, when less demanding manoeuvres are dealt with.

This paper presents the developments incorporated in a low–order simplified model of helicopter motion in order to devise a Flight and Navigation Procedures Trainer. A minimum–complexity rotorcraft model is used as the baseline for developing a simple and computationally efficient flight simulator, which runs in real–time on relatively inexpensive facilities. In order to fulfill qualification standards for the simulator, a series of features are added to the baseline model so as to deal with all flight conditions possibly encountered during vehicle operation. These additional features include engine model and governor, a simple undercarriage and ground–effect model, the representation of vortex ring state and icing phenomena. The architecture of both hardware and software for the simulator is presented and discussed.

This paper presents a novel approach to the design of low-thrust trajectories, based on a first order approximated analytical solution of Gauss planetary equations. This analy- tical solution is shown to have a better accuracy than a second-order explicit numerical integrator and at a lower computational cost. Hence, it can be employed for the fast propagation of perturbed Keplerian motion when moderate accuracy is required. The analytical solution was integrated in a direct transcription method based on a decomposition of the trajectory into direct finite perturbative elements (DFPET). DFPET were applied to the solution of two-point boundary transfer problems. Furthermore the paper presents an example of the use of DFPET for the solution of a multiobjective trajectory optimisation problem in which both the total DV and transfer time are minimised with respect to departure and arrival dates. Two transfer problems were used as test cases: a direct transfer from Earth to Mars and a spiral from a low Earth orbit to the International Space Station.

The paper discusses the relevance of eccentric reference orbits on the dynamics of a tethered formation, when a massive cable model is included in the analysis of a multi-tethered satellite formation. The formations examined in this study are hub-and-spoke (HAS) and closed-hub-and-spoke (CHAS) configurations for in-plane and Earth-facing spin planes. Stability of the formations is studied by means of numerical simulation, together with the evaluation of the effects of eccentricity on tether elongation, agents relative position, and formation orientation and shape.

This paper discusses the relevance of J2 effects induced by Earth’s oblateness on a tethered formation, when a massive cable model is included in the dynamic analysis of a multi–tethered satellite formation. The formations examined in this study are Hub-And-Spoke (HAS) and Closed-Hub-And-Spoke (CHAS) configurations for In-Plane and Earth-Facing spin planes. Some configurations present a cluster with deputies affected differently by the J2 potential depending on their initial position with respect to the orbital plane. The differences in stability, elongation of tethers and shapes of the formations under the J2 potential are compared with those in absence of the perturbation.

The use of an evolutionary algorithm in the framework of H ∞ control theory is being considered as a means for synthesizing controller gains that minimize a weighted combination of the infinite norm of the sensitivity function (for disturbance attenuation requirements) and complementary sensitivity function (for robust stability requirements) at the same time. The case study deals with a complete full-authority longitudinal control system for an unstable high-performance jet aircraft featuring (i) a stability and control augmentation system and (ii) autopilot functions (speed and altitude hold). Constraints on closed-loop response are enforced, that representing typical requirements on airplane handling qualities, that makes the control law synthesis process more demanding. Gain scheduling is required, in order to obtain satisfactory performance over the whole flight envelope, so that the synthesis is performed at different reference trim conditions, for several values of the dynamic pressure, used as the scheduling parameter. Nonetheless, the dynamic behaviour of the aircraft may exhibit significant variations when flying at different altitudes, even for the same value of the dynamic pressure, so that a trade-off is required between different feasible controllers synthesized at different altitudes for a given equivalent airspeed. A multi-objective search is thus considered for the determination of the best suited solution to be introduced in the scheduling of the control law. The obtained results are then tested on a longitudinal non-linear model of the aircraft.

The paper proposes a simulation approach to evaluate the power required by a rotorcraft in standard flight missions and in emergency landing maneuvers, and the corresponding fuel consumption, in order to compare the feasibility and potential fuel savings for different hybrid power systems. More in detail, three options are analyzed, namely electrification of the tail rotor, fully hybrid electric propulsion and electric emergency landing. Weight penalty and potential fuel saving for the proposed hybridization schemes are evaluated for an Agusta-Westland A109 twin engine helicopter model. Nonetheless the discussed methods of analysis have general validity for single main rotor helicopter configurations. Two different scenarios are considered in this investigation: current technologies for batteries and motors and improved electrical components, with performance projections as of 2040. According to this analysis, electrification of the tail rotor and parallel hybridization are feasible with available technology, whereas a fully electrical power system for emergency landing could be developed only in the future. Finally, a parallel hybrid electric power system is sized according to the analysis of power request over four different missions. Fuel savings are evaluated for different energy management strategies. According to the results of this investigation, the parallel hybrid electric power system with present-day and future technologies can save fuel up to 5% and 12%, respectively, with an appropriate energy management strategy.

This paper discusses a method for solving the so called low-thrust Lambert’ s problem. The problem is formulated as a two-points boundary value one, where the initial and final positions are provided in terms of equinoctial variables. A first-order perturbative method is used for investigate the development of orbital elements generated by the low-thrust propulsion system, which acts as a perturbative parameter with respect to the zero-order Keplerian motion. An implicit formulation is thus obtained which allows for the determination of the low-thrust transfer trajectory driving the equinoctial parameters from the initial to their final values in a prescribed time. Furthermore the paper presents three test cases, which demonstrate the flexibility of the method for different missions: the first example is a spiral multi-revolution transfer from low Earth orbit to the International Space Station, the second is a interplanetary transfer from Earth to Mars, while the third is a geostationary transfer orbit to a geostationary orbit.

An inverse simulation algorithm is applied to the determination of the control laws for tracking desired maneuvers by means of an unconventional quad-rotor configuration. This novel configuration features four rotors that are allowed to tilt around the axis of the support, thus changing the thrust vector direction in order to achieve unprecedented maneuvering capabilities. The novel configuration is compared with a conventional quad-rotor, where control moments are generated by variations of rotor rotation rates. A 90 deg roll maneuver in forward flight is also presented, to highlight the increased maneuvering capabilities of the new configuration.

In this paper, a control law that detumbles a spacecraft with magnetic actuators only is developed. A rigorous mathematical proof of global asymptotic convergence from arbitrary initial tumbling conditions to zero angular velocity in a time-varying magnetic field is derived. Furthermore, a simple criterion for determining a reasonable value of the control gain is developed. The selected gain results in a quasi-minimum time detumbling for different initial conditions, in the presence of magnetic coil saturation. Performance of the proposed detumbling control law is demonstrated by numerical simulations on a large set of test cases using a Monte Carlo approach.

The problem of attitude control can be expressed in terms of Euler axis/angle, where a nominal Euler axis rotation directly takes the spacecraft from an initial attitude to a desired final one. When three independent control torques are not available, e.g. when magnetic torquers are the only attitude effectors, the nominal rotation cannot be performed. Nonetheless a feasible rotation around a 'non-nominal' axis allows exact pointing of a given body-fixed axis towards any prescribed target direction. This geometric property results into a control law that enables magnetically actuated spacecraft to align and spin around a desired direction. When, in addition, the activation of a momentum wheel along the spinning axis allows complete alignment with the desired reference, the problem of saturation arises, for example in the presence of drag disturbance torque. A second magnetic law enables the wheel to never saturate, avoiding the classical periodic desaturation manoeuvres.

The interest in electrically-driven propeller airplanes has been steadily increasing over the last 15 years with applications ranging from small electric remotely piloted vehicles up to large high-altitude/long-endurance aircraft, passing through general aviation size airplanes powered by means of hydrogen fuel cells. Unfortunately, the extension of the results valid for conventional configurations to electrically powered aircraft is not always straightforward. As a matter of fact, the determination of the best range flight condition for piston props as the minimum drag airspeed is carried out by maximizing the specific range, taking into account weight variations induced by fuel consumption. On the converse, the weight of electrically powered aircraft is constant during cruise. This is especially true for battery operated aircraft, but an approximation of constant weight also holds when hydrogen fuel cells are employed, inasmuch as only negligible variations of aircraft weight are expected, since the mass fraction of hydrogen being negligibly small.

A novel method for deriving simplified models of flexible aircraft dynamics by means of a mixed Newtonian–Lagrangian approach is proposed. Lagrange equations are used for deriving the dynamic model for flexible degrees of freedom, discretized by means of a standard Gal ̈erkin method, while the evolution of transport degrees of freedom (translation and attitude variable) is obtained by means of second Newton’s Law and generalized Euler equations, suitable for describing the dynamics of deformable bodies. The method allows for easily highlighting those terms less important in the dynamic response of the aircraft, thus making it possible to simplify the model as much as possible, keeping its complexity down to a minimum level. This feature is particularly interesting when the dynamic model is used for the derivation of control laws and flexibility effects need to be accounted for only if coupling with manoeuvre and/or closed loop control tasks is relevant. Numerical simulation demonstrates the viability of the modeling approach, showing that flexibility effects are correctly represented.

A method for deriving low-order models for flexible aircraft by means of a mixed Newtonian-Lagrangian approach is proposed. Lagrange equations are used for flexible degrees of freedom, discretized by means of the Galërkin method. The evolution of transport degrees of freedom (position and attitude variables) is obtained by means of Newton's second law and a generalized Euler equation. A strong link with conventional rigid aircraft equations of motion is maintained, which allows to highlight those terms less relevant for aircraft response. When negligible, these terms are removed, and a minimum-complexity flexible aircraft model is derived, suitable for realtime simulation and control law synthesis. Numerical results demonstrate how the approach correctly represents flexibility effects on aircraft response for a large transport aircraft model.

A novel inverse simulation scheme is proposed for applications to rotorcraft dynamic models. The algorithm adopts an architecture that closely resembles that of a model predictive control scheme, where the controlled plant is represented by a high-order helicopter model.Afast solution of the inverse simulation step is obtained on the basis of a lower-order, simplified model. The resulting control action is then propagated forward in time using the more complex one. The algorithm compensates for discrepancies between the models by updating initial conditions for the inverse simulation step and introducing a simple guidance scheme in the definition of the tracked output variables. The proposed approach allows for the assessment of handling quality potential on the basis of the most sophisticated model, while keeping model complexity to a minimum for the computationally more demanding inverse simulation algorithm. The reported results, for an articulated blade, single main rotor helicopter model, demonstrate the validity of the approach.

A novel inverse simulation scheme is proposed for application to rotorcraft dynamic models. The algorithm is based on a model predictive control scheme, that allows for a faster solution of the inverse simulation step, working on a lower–order, simplified helicopter model. The control action is then propagated forward in time on a more complete model. The algorithm compensates for discrepancies between the models by means of a simple guidance scheme. The proposed approach allows for the assessment of handling quality potential on the basis of the most sophisticated model, adopted for the forward simulation, while keeping model complexity to a minimum level for the computationally more demanding inverse simulation algorithm. This allows for a faster solution of the inverse problem, if compared with the computational time necessary for solving the same problem on the basis of the full–order, more complex model. At the same time, the results are not affected by modeling approximations at the basis of the simplified one. The reported results, for an articulated blade, single main rotor helicopter model demonstrate the validity of the approach.

Multi-rotor autonomous aerial vehicles have been proposed for several missions where the capability of controlled flight in small areas is required. Unfortunately, small-size electrically powered rotary wings UAVs have usually very short endurance, because of inherent limitation in battery energy density and constraints on maximum weight of the battery pack. At the same time, maneuverability can be improved by a variation of rotor thrust orientation. This paper presents a novel quad-rotor configuration where the use of a combustion engine and variable pitch propeller will allow for increasing endurance performance, whereas tilting of rotor disks improves maneuverability. A preliminary design of simple control laws is also discussed for this particular configuration, tested on a few basic maneuvers.

The multibody analysis of the system formed by an entry vehicle and a parachute is the subject of the present paper. In particular, two different models of the system made of bridles (connecting the vehicle to the suspension point) and riser (connecting the suspension point to the suspension lines of the parachute) are considered. This allows one to highlight if and how the simplifying assumptions at the basis of many techniques adopted for sizing the suspension system are reasonable and if they affect the most important factors that characterize the final phase of an entry trajectory, such as peak load and maximum deceleration at parachute inflation and landing site dispersion footprint. A Monte Carlo analysis is performed to account for uncertainties on initial conditions and system parameters.

This study is concerned with the activation energy threshold of bistable composite plates in order to tailor a bistable system for specific aeronautical applications. The aim is to explore potential configurations of the bistable plates and their dynamic behavior for designing novel morphing structure suitable for aerodynamic surfaces and, as a possible further application, for power harvesters. Bistable laminates have two stable mechanical shapes that can withstand aerodynamic loads without additional constraint forces or locking mechanisms. This kind of structures, when properly loaded, snap-through from one stable configuration to another, causing large strains that can also be used for power harvesting scopes. The transition between the stable states of the composite laminate can be triggered, in principle, simply by aerodynamic loads (pilot, disturbance or passive inputs) without the need of servo-activated control systems. Both numerical simulations based on Finite Element models and experimental testing based on different activating forcing spectra are used to validate this concept. The results show that dynamic activation of bistable plates depend on different parameters that need to be carefully managed for their use as aircraft passive wing flaps.

This paper presents an analytical and experimental framework for investigating the performance of fixed-wing electrically-driven propeller aircraft. After the experimental derivation of a novel constant power discharge model for lithium-polymer (Li-Po) batteries, closed-form expressions for flight endurance and range are derived by equating power required in steady level flight with power supplied by the battery pack. Best endurance and best range airspeed are obtained. A methodology is also described to optimally size the battery capacity for a given set of battery and aircraft characteristics. The proposed approach is validated by application to a test case relative to a small-size unmanned aerial vehicle.

In underactuated conditions, satellite attitude actuators can deliver a control torque with two components only. The paper develops a general solution for the problem of attitude maneuver planning for underactuated spacecraft by means of a sequence of N admissible rotations, where a rotation is considered to be admissible if it takes place around an axis which lies on the plane where the actuator system can deliver a control torque. An exact solution for the sequence of two feasible rotations that provides a desired reorientation is derived, where the minimum angular- path sequence can also be identified analytically. The optimal sequence for a higher number of steps is obtained by means of a numerical optimization procedure. The features of the optimal sequences for increasing values of N are investigated in order to obtain a better understanding of their characteristics.

This paper deals with the investigation of performance and sta- bility characteristics of a two-seater, two-bladed, lightweight he- licopter developed in the framework of VLR certification spec- ifications. The main rotor features a gimballed hub with elas- tomeric bearings equipped with a Bell-Hiller to improve stabil- ity, while the fixed pitch, rpm controlled, five-bladed tail rotor is a fenestron design. The main technical drivers of the novel design are to reduce the high level of 2/rev vibrations occurring in teetering rotors, to retain adequate control power in low–g maneuvering and to improve handling qualities using the stabilizing bar to increase roll and pitch damping. A specific aspect of the gimballed rotor is the presence of a sustained wobbling motion of the hub, even in steady–state conditions. A nonlinear model of the vehicle is derived that includes, among other aspects, a detailed model of main rotor, nonlin- ear, quasi–static blade aerodynamics, inflow dynamics, a simple fuselage aerodynamic model and a tail rotor model derived from experimental wind–tunnel tests. Periodic trim conditions are evaluated using a shooting method in order to assess the impact of rotor wobbling mo- tion on helicopter steady–states. Results on performance and controllability are presented and discussed. Finally, the stability characteristics of the vehicle are assessed in order to gain some preliminary insight on the handling qualities of the helicopter.

The dynamics of a multi-tethered satellite formation is considered here, in order to derive a minimum complexity model that successfully represents its dynamics under the action of gravity gradient force and tether tension, where tethers are modeled by means of a sequence of point-masses and massless springs (bead model). The results thus obtained are compared with those derived for a simple, massless tether model, in order to highlight the effects of tether mass on the resulting open-loop behavior. The analysis is performed by means of numerical simulation of the considered models. Once the dynamic behavior of the formation is identified for the nominal values of tether characteristics, the problems raised by the presence of tether mass are highlighted and possible solutions are outlined, when possible. A parametric analysis with respect to the tether linear mass, damping and stiffness is also discussed.

A reduced-order short-period model is derived, which allows for representing the effects of structural flexibility on aircraft response to pilot and disturbance inputs on the basis of a minimum set of relevant information on aerodynamic and structural configuration. The model is derived from first principles by means of a Lagrangian approach, featuring vertical (heave) and pitch degrees of freedom, together with deformation variables for flexural deformation of wing and after portion of the fuselage. As a byproduct, an explicit formulation for stability derivatives with respect to deformation variables is derived. Numerical results are reported for a configuration representative of a modern commercial jet aircraft.

The shape function is defined on the basis of a few parameters that can be modified by means of an optimization algorithm to improve transfer performance in some sense and guarantee its feasibility. The final solution does not need to be the optimal one, but it should provide a reasonable first guess for more sophisticated optimization procedures. Most of the applications of the shape-based approach are proposed in the framework of the two-body problem, where the spacecraft moves under the action of low thrust and a single primary body. This approach is reasonable for most mission scenarios, during preliminary steps of mission design, when approximate solutions are sufficient and a single primary mass is considered during each mission segment. However sometimes it is necessary to account for gravity pull from more than a single primary mass.

The use of magnetic torquers on satellites flying inclined Low Earth Orbits (LEO) arises challenging problems when dealing with control strategies: since the available torque is perpendicular to the local magnetic field, the Attitude Control System (ACS) results to be inherently underactuated. In this paper a rigorous proof of global asymptotic stability is given about a control law that leads the satellite to a desired pure spin condition around a principal axis of inertia. Besides, a heuristic control contribution to this law, inspired by quaternion feedback control and a geometric result by Avanzini and Giulietti is proposed in order to point the spinning axis along the normal to the orbital plane.

A strategy for attitude path planning that performs single-axis pointing in the presence of obstacles along the angular path and constraints on admissible rotation axes is presented. The proposed kinematic solution is based on the identification of the crossing condition for a baseline single-step maneuver with an undesired set of pointing, seen as a conical obstacle along the attitude path. If crossing with the forbidden region occurs, an alternative two-step maneuver is selected, where both rotation axes are admissible and the cones spanned by the sensor boresight are tangent to the prohibited cone. Given the small number of admissible two-step maneuvers (eight at most), it is possible to select the shortest overall angular path with limited computational effort. The test case discussed demonstrates the effectiveness of the approach and provides further insight into the geometry of the cones.

The paper proposes a novel approach for single-axis pointing of an underactuated spacecraft using only two reaction wheels (RWs), which is based on a simple yet effective wheel rate command. The control law can be used for aiming the line of sight of a sensor, a nozzle, or an antenna toward a target direction, or solar panels toward the sun, after failure of one wheel for a nonredundant control system hardware or, in the case of multiple failures, for redundant systems. The proposed control methodology represents the practical, dynamic implementation of a kinematic planning scheme proposed by the same authors, developed under the assumptions of zero overall angular momentum and triaxial inertia tensor. Under the zero angular momentum hypothesis, the control law also provides three-axis stabilization at zero angular speed. Only rest-to-rest maneuvers are thus dealt with, as a steady residual rotation rate around an arbitrary body-fixed axis cannot be attained in general unless the axis is the principal of inertia and aligned with the spin axis of one of the two active reaction wheels. The limited computational demand of the control law also makes it a practical solution in the case of small-size satellites, where the computational budget is severely limited by the available CPU processing capabilities. The effects of a nonzero residual angular momentum and control axes not aligned with the principal axes of inertia is also investigated, highlighting limitations on pointing precision and convergence performance.

A method for solving the so-called low-thrust Lambert problem is proposed. After formulating it as a two-point boundary value problem, where initial and final positions are provided in terms of equinoctial variables, a first-order perturbative approach is used for investigating the variation of orbital elements generated by the low-thrust propulsion system, which acts as a perturbing parameter with respect to the zero-order Keplerian motion. An implicit algebraic problem is obtained, which allows for the determination of the low-thrust transfer trajectory that drives the equinoctial parameters from the initial to the final values in a prescribed time. Three test cases are presented, which demonstrate the flexibility of the method for different mission scenarios: an interplanetary transfer from Earth to Mars, a spiral multirevolution transfer from low Earth orbit to the International Space Station, and a maneuver to a highly elliptical orbit with large plane change.

Attitude regulation proves to be a challenging problem, when magnetic actuators alone are used as attitude effectors, since they do not provide three independent control torque components at each time instant. In this paper a rigorous proof of global exponential stability is derived for a magnetic control law that leads the satellite to a desired spin condition around a principal axis of inertia, pointing the spin axis toward a prescribed direction in the inertial frame. The technique is demonstrated by means of numerical simulation of a few example maneuvers. An extensive Monte Carlo simulation is performed for random initial conditions, in order to investigate the effect of changes in control law gains.

This paper aims to analyze the characteristics of a two-bladed gimballed rotor featuring a homokinetic joint between driving shaft and rotor yoke and a fly-bar with paddles. Blades are connected to the yoke by coning hinges. A pitch--coning link is introduced for gust load alleviation. The design and testing of this rotor configuration is part of the development program of a lightweight helicopter in the VLR rotorcraft certification framework. The rotor is designed with the main objective of solving some of the negative issues that affect the use of teetering rotors on light helicopters, such as strong 2/rev oscillatory loads, poor response at low $g$'s and a pronounced sensitivity to gusts and/or large pilot inputs. The work is focused on stability issues related to the presence of coning hinges and their effects on rotor response and loads transmitted to the hub, affected by the variation of mass properties associated to (possibly non--symmetric) coning rotations. To this end a dynamic model is developed, that captures the most relevant aspects of the mechanical interactions between blades and yoke. Numerical simulation and stability analysis are performed to assess possible advantages of the configuration with respect to a conventional teetering rotor.

The paper presents a set of visual aids for enhancing remote pilot perception of potential violations of allowed fly areas or conflicts with conventional air traffic when operating remotely piloted aerial vehicles. Assuming a video stream from an on-board camera is available to the remote pilot, visual aids are provided in a head-up display modality by means of reality augmentation techniques. The main visual element consists of a dynamic set of fences allowing for a safe aircraft separation away from no-fly zones and from neighboring vehicles. The shape of the fences is varied according to aircraft current speed and altitude, in order to allow for a sufficient safety margin, also in case of a loss-of-control accident. As a further visual aid, the predicted future position of the aircraft is determined and fence color is changed in order to highlight potential violations of the allowed operational area. The proposed methodology is validated by means of simulations in a reference operational scenario. Results demonstrate the effectiveness of the proposed approach in improving pilot awareness.