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This paper describes a frequency-domain method for the identification of multi-input control equivalent turbulence input models. Such models can be identified from flight data gathered in specific ...

This paper investigates multi-objective optimization of coordinated patrolling flight of multiple unmanned aerial vehicles in the vicinity of terrain, while respecting their performance parameters....

Accepted: 2021-02-28, 資料番号: PA2120065000

The implementation of interconnect gain from aileron to rudder surface on the majority of the aircraftis to decrease sideslip which is generated because of adverse yaw with the movement of control stick in lateral axis and also enhances the turning rate performance.The Aileron to Rudder Interconnect (ARI)involves significant part to decouple the Dutch roll oscillations from roll rate response to aileron command. ARI is feed-forward gain whichis susceptible to aircraft system uncertainty. Incorrect ARI gain can lead to side slip buildup which can cause aircraft to depart in case of fault scenarios. Four systematic ARI design methods are proposed. One of the proposed methods which use the norm of ARI transfer function at roll damping frequency is suitable for online reconfiguration of control law. The reconfiguration of ARI gain is illustratedwith the simulation responses of fault scenario case of aileron surface damage.

The paper is devoted to the problem of safe flight in vortex hazard conditions. A vortex wake is always produced behind flying aircraft. This vortex wake is invisible to other aircraft following it. Therefore, getting into a vortex wake from a preceding aircraft often becomes unexpected and quite often leads to flight incidents. This is confirmed by the statistics of flight accidents. With the introduction of aircraft weighing more than 500 tons, the problem of vortex safety has only sharpened. The fact is that the old standards that define safe intervals between flying aircraft are still active. These standards are given in this paper. It is also shown that even if these standards are observed, the flight of a medium-class aircraft at the same level as a heavy-class aircraft is unsafe. To study the effect of the vortex wake from the preceding aircraft, a special computational software complex based on the discrete vortex method was developed. This complex has passed the necessary testing and state registration. A number of measures were carried out to validate and verify the developed complex, confirming the operability of the programs included in it and the reliability of the results obtained. On the basis of this computational and software complex studies on the effect of the A380 aircraft vortex wake on MS-21 class aircraft were performed. The studies were undertaken for various altitudes and speeds of the A380 aircraft and for its various flight configurations. As a criterion for the MS-21 aircraft safe flight, the roll moment that occurs due to falling into the vortex wave, in particular, its coefficient was chosen. If the coefficient of the MS-21 aircraft roll moment when it hits the vortex wake exceeds that from the ailerons, such a flight is considered dangerous. The distances behind the A380 that are unsafe for the MS-21 aircraft flight are given.

We describe the Intelligent Autopilot System (IAS), a fully autonomous autopilot capable of piloting large jets such as airliners by learning from experienced human pilots using Artificial Neural Networks. The IAS is capable of autonomously executing the required piloting tasks and handling the different flight phases to fly an aircraft from one airport to another including takeoff, climb, cruise, navigate, descent, approach, and land in simulation. In addition, the IAS is capable of autonomously landing large jets in the presence of extreme weather conditions including severe crosswind, gust, wind shear, and turbulence. The IAS is a potential solution to the limitations and robustness problems of modern autopilots such as the inability to execute complete flights, the inability to handle extreme weather conditions especially during approach and landing where the aircraft’s speed is relatively low, and the uncertainty factor is high, and the pilots shortage problem compared to the increasing aircraft demand. In this paper, we present the work done by collaborating with the aviation industry to provide training data for the IAS to learn from. The training data is used by Artificial Neural Networks to generate control models automatically. The control models imitate the skills of the human pilot when executing all the piloting tasks required to pilot an aircraft between two airports. In addition, we introduce new ANNs trained to control the aircraft’s elevators, elevators’ trim, throttle, flaps, and new ailerons and rudder ANNs to counter the effects of extreme weather conditions and land safely. Experiments show that small datasets containing single demonstrations are sufficient to train the IAS and achieve excellent performance by using clearly separable and traceable neural network modules which eliminate the black-box problem of large Artificial Intelligence methods such as Deep Learning. In addition, experiments show that the IAS can handle landing in extreme weather conditions beyond the capabilities of modern autopilots and even experienced human pilots. The proposed IAS is a novel approach towards achieving full control autonomy of large jets using ANN models that match the skills and abilities of experienced human pilots and beyond.

Aileron to Rudder Interconnect (ARI) gain is implemented on most fighter aircraft, primarily to reduce the side slip produced due to adverse yaw from pilot lateral control stick input and to improve the turn rate response. A systematic and non-iterative design procedure for ARI gain is proposed herein based on the evaluation of a transfer function magnitude at the aircraft roll mode frequency. The simplicity of the proposed method makes it useful for real-time flight control law reconfiguration in situations where the aileron control authority is diminished due to damage. This is demonstrated by a simulation example considering an aileron surface damage scenario.

This work aims to model transonic airfoil–gust interaction and the gust response on transonic aileron-buzz problems using high-fidelity computational fluid dynamics (CFD) and the Long Short Term Me...

For a mechatronic system, the lack of robustness is mainly due to an ignorance of variability, which is unavoidable in every designed system at all stages of system development and during its life cycle. In fact, the inherent imperfections of manufacturing processes and operational variations such as material wear or multiphysical effects (thermal, vibrations, etc.) involve parametric variations, which can degrade the proper system function. To ensure a high level of quality and to improve design robustness, the deviations between actual and target definition should be restricted by specified tolerances. Moreover, it is important to understand the relationship between parameter deviations and system behavior. Thus, in this article, a variational approach is proposed to identify performance sensitivity to parameter deviations in a mechatronic system and to determine influential parameters that affect significantly the proper functioning of the system and its behavior. Furthermore, this approach is used to specify the admissible parameter deviations for which the given specifications can be guaranteed. The complete approach is applied to an example in the aeronautic field: an electromechanical actuator driving an aircraft primary flight control surface (aileron).

Variable-camber technology is considered an effective way to adaptively improve the aerodynamic performance of aircraft under various flight conditions. This paper studies the aerodynamic characteristics of the trailing-edge variable-camber technology by means of Computational Fluid Dynamics (CFD) and a drag decomposition method. Trailing-edge variable-camber technology can be simply realized by the continuous deflection of the flaps and ailerons of a wing. A supercritical airfoil is used to study the two-dimensional effect of variable-camber technology, and a wide-body airplane model is used to validate the three-dimensional improvement in the wing's airfoil made by variable-camber technology. An optimization strategy for airfoil that incorporates variable-camber technology is proposed. The optimization results demonstrate that the proposed method can obtain better results than the traditional segregated shape optimization.

In this paper we consider Model Predictive Control (MPC) design for roll angle control for a Fixed-Wing Unmanned Aerial Vehicle (UAV) with Multiple Segmented Control Surfaces. The challenge of roll angle control for a Fixed-Wing UAV consists of switching between inner and outer aileron-pairs with hard constraints due to safety, energy saving and switching actuators. The novelty consists of formulating a hybrid-control problem as a switched linear constrained MPC-QP problem and switched state observer design for Fixed-Wing UAV. A fast novel QP-solver based on the active-set QP-solver Hildreth is developed to meet the real-time implementation target, which is to stay below the sampling time of Ts = 10 ms. The designed MPC controllers are simulated using Matlab. Simulations and the CPU-time from the improved QP-solvers show MPC to be a very good choice for real-time roll angle control for Fixed-Wing UAVs., In this paper we consider model predictive control (MPC) design for roll angle control for a fixed-wing unmanned aerial vehicle (UAV) with multiple segmented control surfaces. The challenge of roll angle control for a fixed-wing UAV consists of switching between inner and outer aileron pairs with hard constraints due to safety, energy saving and switching actuators. The novelty consists of formulating a hybrid control problem as a switched linear constrained MPC-QP problem and switched state observer design for fixed-wing UAV. A fast novel QP-solver based on the active-set QP-solver Hildreth is developed to meet the real-time implementation sampling time of Ts = 10 ms. The designed MPC controllers are simulated using Matlab. Simulations and the CPU-time from the improved QP-solvers show MPC to be a very good solution for real-time roll angle control of fixed-wing UAVs.

Purpose A standard automatic flight control system – autopilot – will become required equipment of the future aircraft, operating in the common sky. For a specific group of aircraft, they are too expensive and too energy-consuming solutions. This paper aims to present the concept of an automatic flight control system that overcomes those limitations. Design/methodology/approach The proposed automatic flight control system uses the trim tabs in all prime flight controlling surfaces: elevator, ailerons and rudder, for stabilizing and controlling the steady flights of an aircraft. Findings The results of an aeroplane flight controlled with the use of trim tabs simulation tests and remarks have been presented and discussed. The simulation was conducted in real-time hardware in the loop environment. The stabilization of the flight was achieved in performed test scenarios. Originality/value The possibility to control an aircraft with coordinated deflections of the trimming surfaces is a beneficial alternate to those currently used and can be recommended for use in the next-generation aircraft.

Extensive research in recent years has focused on improving the current loss-of-control prevention systems and developing new strategies for safe path planning of the impaired aircraft. Success in developing such systems requires a comprehensive perception of the influence of damage on the aircraft’s dynamic behaviour and performance, and the effect of various failure degrees on the flight envelope confinement and the remaining safe maneuvers. This paper comprehensively describes the effects of lateral control surface failure on the NASA Generic Transport Model (GTM) flight envelope, defined by a set of attainable steady-state maneuvers herein referred to as trim points. The study utilises a large database of high-fidelity maneuvering flight envelopes computed for the unimpaired case and wide ranges of the aileron and rudder failure cases at different flight conditions. Flight envelope boundary is rigorously investigated, and the key parameters confining the trim points at different boundary sections are identified. Trend analyses of the impaired flight envelopes and the corresponding limiting factors demonstrate the effect of various failure degrees on the remaining feasible trim points. Results can be employed in emergency path planning with potential uses in the development of aircraft resilient control and upset recovery systems.

Aiming at the problem of energy collection and consumption of solar-powered UAV under different maneuvers, an energy collection model combining solar-powered UAV attitude and solar radiation was established. Firstly, the ASHRAE model was used as the solar radiation estimation model in the flight environment. Secondly, the energy harvesting model of the solar-powered UAV is deduced by the relationship between the attitude of the UAV, the relative position of the sun and the earth. The experimental data was used to establish the energy consumption model of solar-powered UAV in different flight states. Finally, the flight time of the same mission path under different operating modes is obtained through simulation. Simulation results show that when the solar-powered UAV is hovering at a height of 3 000 meters and a radius of 50 meters, the flight time of aileron operation is 7.19% longer than power differential operation; when hovering at a radius of 170 meters, the flight time of the two operations is almost equal; when hovering at a radius of 600 meters and flying in a line, the flight time of the power differential operation is about 1.08% longer than that of the aileron operation.

Based on mechanism research of dynamic coupling and inertial coupling of flying wing UAV at high attack angle or stall maneuver flight, a decoupling compensator is designed to realize the longitude and lateral channels, the aileron and the rudder channel. The maneuver control law is designed based on the robust servo method, which solves the problem of rapid maneuvering of the flying wing. The maneuvering boundary protection controller effectively reduces the risk of boundary during the maneuver flying. At last, the flying tests verified the engineering feasibility of decoupling controller and robust maneuvering controller.

Nixus is the world’s first fly-by-wire (FBW) sailplane, with a custom 92-foot span, 53.3 aspect ratio wing. With the second-largest-aspect-ratio wing ever built for a crewed airplane, Nixus faces unique challenges for aeroelastic design, requiring detailed study and special considerations for safe operation. The use of an FBW system for the wing control surfaces allows the exploration of new strategies for automatic flap positioning, tailored aileron deflections, load alleviation, and, in the future, aeroelasticity control. This paper briefly describes some aspects of this sailplane’s fabrication and covers in detail the ground vibration test (GVT) and use of the GVT results in the flight envelope expansion campaign. The preparation and execution of the GVT are presented, including aspects related to the involvement of Cal Poly’s Aerospace Engineering undergraduate students in this process as part of an extracurricular activity. Results of the GVT, comparison with finite element analysis (FEA), and flutter predictions from the test-verified FEA model are provided. Additionally, the execution of flutter flight tests that used the FBW system to excite the wing is presented.

This paper details the development of a high redundancy Unmanned Aerial Vehicle (UAV) platform with a sliding mode fault tolerant control system. The high redundancy UAV platform is of a Tilt Wing design, utilising four wings, each with a propeller. Conventional fixed wing aircraft control surfaces are also implemented i.e. ailerons, elevators and rudders. The benefits of such a platform is it combines the efficient forward flight of fixed wing aircraft, with the agility and Vertical Take Off and Landing (VTOL) capability of a conventional multi-rotor aircraft. Results show good forward flight performance despite the presence of actuator faults/failures.

This paper presents autonomous landing controller design of unmanned aerial vehicles (UAVs). Feedback linearization approach with anti windup scheme is designed as autonomous landing controller. Anti windup basically controls the integrator component accumulation in the autopilot design and restricts the output within saturation limits. Autonomous landing of a fixed wing UAV consists of approach, glideslope and flare phases. During approach, aerial vehicle aligns itself with runway and reduces its lateral deviations with respect to runway central line. In glide slope phase, aerial vehicle maintains a fixed flight path angle and descends with a constant sink rate. In flare phase, aerial vehicle follows an exponential trajectory and descends with a lower sink rate which keeps reducing further as it goes to lower altitudes. Flare controller is designed with integrator and antiwindup scheme is used to handle the controller output within saturation limits. Landing is primarily a longitudinal mode operation but due to disturbances and coupling lateral and directional modes also gets activated. In this paper pitch angle, roll angle and yaw angle nonlinear dynamic equations are linearized to obtain the control commands in terms of elevator deflection, aileron deflection and rudder deflections. Similarly total velocity of aerial vehicle being an important parameter is controlled using thrust command. A first order linearized model of velocity is used to obtain thrust control command. Autonomous landing of UAV with feedback linearization controller and anti windup scheme is simulated with Six-DOF model of AE2 UAV. The algorithm is implemented with wind disturbances to show the autonomous landing performance of UAV.

A tri-copter unmanned aerial vehicle (UAV) with individually tilted main wings for various flight maneuvers is presented. In contrast to conventional tilt wing UAVs, thrust vectoring and control of the directions of main wing surfaces are attained by tilting main wings individually. Such individually tilted main wings allow more efficient maneuvers of the UAV. Even though the UAV is a tri-copter, it can compensate the reaction torque created by the tail motor by the individual control of the main wings. Few publications have appeared in the literature, dealing with tri-copter UAV with individually tilted main wings. Dynamic model of the UAV is established via the Newton-Euler formulation and effect of individual wing control is examined. The PID controller and PI controller for control of flight maneuvers are implemented and used for simulations and experiments. Results of simulations and experiments are presented for validation of flight performance of the UAV, including the roll rate 80% larger than that of the conventional UAV with ailerons as control surfaces.

Actuation failure is one of the causes of loss of control in flight accidents. Aircraft usually have multiple redundant actuators to mitigate failures, and Failure Detection and Isolation Systems (FDIS) are used to diagnose failures and reconfigure software/hardware to enhance safety. However, the large number of redundant actuators interferes with the FDIS. To detect and isolate failures in fixed-wing aircraft with redundant actuators, this work proposes the combined use of two different strategies of the Two-Stage Kalman Filter. A Supervisory Loop is included using heuristics and statistics to diagnose the actuators, and a Feed-Forward Differential is implemented to improve the isolation process without interfering with the aircraft flight. The solution is evaluated in the detection of an aileron failure in a Boeing 747 simulator.

In this paper, a multi-input/multi-output servomechanism primary flight controller is presented. Also discussed are flight data related to the system performance and pilot assessments of handling q...

Extensive research in recent years has focused on developing flight envelope estimation methods to improve current loss of control prevention and recovery systems. Such methods are practically efficient only if they are able to evaluate in real time the new flight envelope of damaged aircraft based on the altered dynamics. Due to nonlinear dynamics of aircraft, common approaches to estimate the entire flight envelope of high-fidelity models are numerically intensive and their real time implementation is computationally impossible. So current methods are based on reduced complexity models or flight envelopes are determined locally. This paper presents a novel method to estimate the global flight envelope of impaired aircraft in real-time for any unknown failure degree. In the proposed method, first, numerous flight envelopes are evaluated using a high fidelity model at various failure degrees and different flight conditions and prepared as training data. Then multiple feedforward neural networks are trained offline by a Bayesian regularization backpropagation algorithm. Finally, the trained networks are used to estimate flight envelopes in real time. The method is applied to rudder and aileron failure cases of the NASA Generic Transport Model. Results show that the estimated flight envelopes are good approximations of the high fidelity global flight envelopes.

In this paper, the problem of aircraft spin recovery is solved as a trajectory optimization problem using direct multiple shooting method with time and altitude-loss as cost functions to be minimized. A stable oscillatory spin state is chosen as the initial condition and the optimal control inputs required to transfer the aircraft to a steady level-flight trim state are determined. Optimal spin recovery simulations are carried out with scaled bounds of control inputs to determine their effectiveness and feasibility for recovery in the case of total control failures. It is shown that spin recovery is feasible in the case of aileron failure, and only arresting of yaw-rate is possible in the case of rudder and throttle failures. Optimal spin recovery simulations are also carried out to determine the effects of wind and aerodynamic forces. It is found that spin recovery solutions are sensitive to initial conditions in the presence of wind, and increase of lift and decrease of drag reduce the altitude-loss of the maneuver.

This paper presents the design and manufacturing of a new morphing wing system carried out at the Laboratory of Applied Research in Active Controls, Avionics and AeroServoElasticity (LARCASE) at the ETS in Montréal. This first version of a morphing wing allows the deformation of its trailing edge, denote by Morphing Trailing Edge (MTE). In order to characterize the technical impact of this deformation, we compare its performance with that of a rigid aileron by testing in the LARCASE’s price—Païdoussis subsonic wind tunnel. The first set of results shows that it is possible to replace an aileron by a MTE on a wing, as an improvement was observed for the MTE aerodynamic performances with respect to the aileron aerodynamic performances. The improvement consisted in the fact that the drag coefficient was smaller, and the lift-to-drag ratio was higher for the same lift coefficient. Keywords: Aerodynamic performances, Aileron, Morphing wing, Trailing edge, Wing tunnel tests

This paper deals with the actuation system design of a full-scale morphing aileron for regional aircraft. The aileron is allowed to smoothly change its geometrical configuration and perform the in-flight transition from a baseline shape to a set of optimal morphed ones pre-defined on the basis of aerodynamic requirements. The design of such innovative aileron is aimed not only at substituting the conventional aileron installed on a real aircraft but also to provide additional functionality. The aileron is free to rotate around its main hinge axis and it is also allowed to smoothly modify camber with two independent actuation systems. In such manner it can be used also during cruise with a symmetric deflection between the two half wings in order to reduce drag in off design condition. To accomplish variable aileron shape, a rigid-body mechanism was designed. The proposed aileron architecture is characterised by segmented adaptive ribs rigidly linked each other with spanwise reinforcements such as spars and stringers in a multi-box arrangement. Each rib is split into two movable plates connected by means of rotational hinges in a finger-like mechanism. The mechanism is driven by a load-bearing actuator by means of a kinematic chain opportunely tied based on the structural requirements in terms of shape to be matched and load to be withstood. The proposed device is an innovative arrangement of the quick-return mechanism composed of a beam leverage, commercial linear guides and a crank. The actuator shaft is directly inserted in the crank, which transmits the rotation to the linear guide that slide along a rail moving upward or downward the beam thus resulting in a camber variation. The entire aileron is moved by three leverages internally contained and distributed along the first two bays while the most external ribs are considered passive and their movement slaved. Two actuation layouts are analytically and numerically studied, the analytical theory is presented and validated by means of a multi-body simulation. Moreover, a linear static analysis was carried out under the hypothesis of glued contact between linear guides components simulating a jamming condition. This assumption has been formulated because it represents the most severe condition that envelop all the operative loads to which the actuation system is subjected. The analyses conducted are preliminarily aimed to verify that no failure occur under the imposed loads. In this first design loop, the vertical static force acting on the linear carriage exceeded allowable value and then a new configuration with double-sided linear guides was then investigated.

A parallelized flight simulator for the dynamic analysis of airborne wind energy (AWE) systems for ground- and fly-generation configurations is presented. The mechanical system comprises a kite or fixed-wing drone equipped with rotors and linked to the ground by a flexible tether. The time-dependent control vector of the simulator mimics real AWE systems and it includes the length of the main tether, the geometry of the bridle, the torque of the motor controllers of the rotors, and the deflections of ailerons, rudder and elevator. The use of a lagrangian formulation with a minimal coordinate approach and discretizing the main tether as a chain of inelastic straight rods linked by ideal (dissipative-less) rotational joints, yielded a non-stiff set of ordinary differential equations free of algebraic constraints. Several verification tests, including a reel-in maneuver that admits an analytical solution, are presented. The efficiency of the parallelization with the number of tether segments, and trade-off analysis of the lagrangian and hamiltonian formulations are also considered. The versatility of the simulator is highlighted by analyzing two maneuvers that are relevant for AWE scenarios. First, the simulator is used to compute periodic figure-of-eight trajectories with an open-loop control law that varies the geometry of the kite's bridle, as frequently done in ground-generation AWE systems. Second, an unstable equilibrium state of a tethered drone equipped with two rotors for energy harvesting is stabilized by implementing a closed-loop control strategy for the deflection of the control aerodynamic surfaces. This work was supported by the Ministerio de Economía, Industria y Competitividad of Spain and the European Regional Development Fund under the project ENE2015-69937-R (MINECO/FEDER, UE). GSA work is supported by the Ministerio de Economía, Industria y Competitividad of Spain under the Grant RYC-2014-15357. RS was partially supported by the EU projects AWESCO (H2020-ITN-642682) and REACH (H2020-FTIPilot-691173). Publicado

A steady growth of aviation transportation (4-5% per year) causes excessive saturation at numerous major airports. As a result, many flights are delayed. One of the ways to deal with this growing problem is to transfer regional propeller aircraft maintenance to suburban airports. It will require both a modernization of local airports and the design of a new generation of regional aircraft with short takeoff and landing (STOL). The aircraft ability to operate from short runways depends not only on wing unit loading and on high-lift capacities but also it is determined by the control surfaces efficiency. The latter often becomes one of the major reasons for limitation of the amount of lift used in STOL configuration. Thus, the successful application of high-lift devices stipulates the necessity for both the efficiency increasing of existing aircraft control surfaces and the development of some alternative form of lateral control not requiring a significant wingspan proportion. The forms of a lateral control, this article considers, include the interceptor, drooped ailerons, ailerons fitted with mini-flap and one of the alternative forms which uses differential flap section deflection. Several mini-flaps with a various chord are also considered to increase the available rudder yawing moment. The efficiency of the above-mentioned control surfaces has been studied in TsAGI low speed wind tunnel on a model of a twin-engine light aircraft with an enhanced level of lifting capacity on take-off and landing configurations. The tests were conducted at a Reynolds number of 1.0×10 6 and Mach number of 0.15.

PurposeThe purpose of the paper is to analyze the active suppression of the aeroelastic vibrations of ailerons with strongly nonlinear characteristics by neural network/reinforcement learning (NN/RL) control method and comparing it with the classic robust methods of suppression.Design/methodology/approachThe flexible wing and aileron with hysteresis nonlinearity is treated as a plant-controller system and NN/RL and robust controller are used to suppress the nonlinear aeroelastic vibrations of aileron. The simulation approach is used for analyzing the efficiency of both types of methods in suppressing of such vibrations.FindingsThe analysis shows that the NN/RL controller is able to suppress the nonlinear vibrations of aileron much better than linear robust method, although its efficiency depends essentially on the NN topology as well as on the RL strategy.Research limitations/implicationsOnly numerical analysis was carried out; thus, the proposed solution is of theoretical value, and its application to the real suppression of aeroelastic vibrations requires further research.Practical implicationsThe work shows the NN/RL method has a great potential in improving suppression of highly nonlinear aeroelastic vibrations, opposed to the classical robust methods that probably reach their limits in this area.Originality/valueThe work raises the questions of controllability of the highly nonlinear aeroelastic systems by means of classical robust and NN/RL methods of control.

Purpose The separation of energy conversion and propulsor is a promising aspect of hybrid-electric propulsion systems, allowing for increased installation efficiencies and setting the basis for distributed propulsion concepts. University of Stuttgart’s Institute of Aircraft Design has a long experience with electrically powered aircraft, starting with Icaré 2, a solar-powered glider flying, since 1996. Icaré 2 recently has been converted to a three-engine motor glider with two battery-powered wing-tip propellers, in addition to the solar-powered main electric motor. This adds propulsion redundancy and will allow analyzing yaw control concepts with differential thrust and the propeller-vortex interaction at the wing-tip. To ensure airworthiness for this design modification, new ground vibration tests (GVTs) and flutter calculations are required. The purpose of this paper is to lay out the atypical approach to test execution due to peculiarities of the Icaré 2 design such as an asymmetrical aileron control system, the long wing span with low frequencies of the first mode and elevated wing tips bending under gravity and thus affecting the accuracy of the wing torsion frequency measurements. Design/methodology/approach A flutter analysis based on GVT results is performed for the aircraft in basic configuration and with wing tip propulsors in pusher or tractor configuration. Apart from the measured resonant modes, the aircraft rigid body modes and the control surface mechanism modes are taken into consideration. The flutter calculations are made by a high-speed, low-cost software named JG2 based on the strip theory in aerodynamics and the V-g method of flutter problem solution. Findings With the chosen atypical approach to GVT the impact of the suspension on the test results was shown to be minimal. Flutter analysis has proven that the critical flutter speed of Icaré 2 is sufficiently high in all configurations. Practical implications The atypical approach to GVT and subsequent flutter analysis have shown that the effects of wing-tip propulsors on aeroelasticity of the high aspect ratio configuration do not negatively affect flutter characteristics. This analysis can serve as a basis for an application for a permit to fly. Originality/value The presented methodology is valuable for the flutter assessment of aircraft configurations with atypical aeroelastic characteristics.

A closed-loop coupling model at the system-level is developed to analyze the effects of joint clearance on the dynamic responses of electromechanical aileron actuation system. The proposed model considers the coupling effects between the electromechanical actuator (EMA) control performance and dynamic characteristics of linkage mechanism with joint clearance. Besides, the experiments are conducted in a test rig, which verifies the effectiveness of the proposed closed-loop coupling model. The nonlinear contact force model and modified Coulomb friction model are adopted in the joint clearance of the linkage mechanism, and the influences of clearance size on the dynamic behaviors of electromechanical aileron actuation system are studied. The numerical and experimental results indicate that the novel closed-loop coupling model, considering the EMA control performance and dynamics of linkage mechanism with joint clearance at the same time, is an effective model to predict the dynamic characteristics of electromechanical aileron actuation with joint clearance, which provides a practical method to analyze the dynamic performance of electromechanical coupling multibody systems with joint clearance.

When the wing of Oblique Wing Aircraft (OWA) is skewed, the center of gravity, inertia and aerodynamic characteristics of the aircraft all significantly change, causing an undesirable flight dynamic response, affecting the flying qualities, and even endangering the flight safety. In this study, the dynamic response of an OWA in the wing skewing process is simulated, showing that the three-axis movements of the OWA are highly coupled and present nonlinear characteristics during the wing skewing. As the roll control efficiency of the aileron decreases due to the shortened control arm in an oblique configuration, the all-moving horizontal tail is used for additional roll and the control allocation is performed based on minimum control energy. Given the properties of pitch-roll-yaw coupling and control input and state coupling, and the difficulty of establishing an accurate aerodynamic model in the wing skewing process due to unsteady aerodynamic force, a multi-loop sliding mode controller is formulated by the time-scale separation method. The closed-loop simulation results show that the asymmetric aerodynamics can be balanced and that the velocity and altitude of the aircraft maintain stable, which means that a smooth transition is obtained during the OWA’s wing skewing. Keywords: Decoupling, Dynamic response, Oblique wing aircraft, Sliding mode control, Wing skewing process

© 2021, The Author(s), under exclusive licence to Springer-Verlag London Ltd. part of Springer Nature.In this paper, a mini unmanned aerial vehicle (UAV) is planned to be used in applications such as spraying pesticide and weed control in agricultural areas. According to literature review, proportional + integral + derivative (PID) structure is used to control many of these UAVs. This controller is insufficient against uncertain weather conditions and disturbance effects. In this study, many different control techniques are evaluated to select the controller structure that can respond to these uncertainties. The structure having the best result was chosen as the UAV controller. Ultrastick-25e mini UAV model is used to control the roll and yaw angle lateral dynamics. State-space presentation of the UAV longitudinal and lateral dynamics is explained, and it is just obtained for the lateral dynamics to control the attitude of the UAV under 60 km/h flight velocity condition. According to the aileron and rudder inputs, lateral dynamics simulations have successfully done by using five different controller methods such as classical PID, artificial neuro-fuzzy inference system (ANFIS), fuzzy logic controller, combined ANFIS-PID, and PD-Fuzzy-PI controllers. Moreover, three different input signals are assumed to evaluate the system response. Additionally, transient response and the time performance parameters such as overshoots, peak, rise and settling times, and steady-state error have analyzed for the designed different controllers. The simulated results for the five different controller designs showed that combined PD-fuzzy-PI and ANFIS-PID controllers have more acceptable performance than other controllers at the steady level flight condition. It is aimed that the simulation findings obtained in this study will contribute to experimental studies.

This paper relates to airfoil selection and wing design for an unmanned aerial vehicle. The proposed wing design is analyzed numerically by FEM using Ansys. Changes in aerodynamic characteristics are illustrated through examples of different wing airfoil geometries. Quantitative aerodynamic parameters are determined, and obtained analysis results are compared with data from NACA airfoil databases. Geometries of individual structural members of the wing are specified, including ribs, spars, spar stiffeners (at fuselage attach points) and skin shape. The shape and fastening of high-lift devices such as ailerons and flaps are proposed. Also, different variants of spars and rib holes are tested and compared using structural simulations.

This paper presents a multidisciplinary and multiobjective optimal design of an aircraft wing with two, three, and four control surfaces. The study aims to compare the performance of the wing in terms of aerodynamic loads rejection, stability robustness, and energy consumption. An LQR (Linear Quadratic Regulator) is designed for each control surface. The geometrical parameters of the control surfaces such as the span-wise and chord lengths, and the diagonal elements of the LQR weighting matrices are optimally adjusted by the NSGA-II (Non-dominated Sorting Genetic Algorithm). The algorithm returns a set of solutions called the Pareto set and its function evaluation forms another set known as the Pareto front. The solution set holds optimal geometrical and control decision variables that produce various degrees of optimal trade-offs among the design goals. To facilitate the comparison between the three optimization problems, a post-processing algorithm that operates on the Pareto front is utilized. Then, the knee points and portions of the Pareto fronts are compared. The optimal solutions show that there are conflicting relationships between the design objectives. The disturbance rejection of the wing with the two ailerons is the least effective however control energy consumption is the smallest as compared to the other configurations. The wing with the three ailerons at 18 different design options has the best relative stability. At the knee point, a wing having four control surfaces can offer the best disturbance rejection but at the expense of the control energy. With these considerations, a wing with three surfaces can be the best compromise among the other configurations.

High aspect ratio aircraft have a significantly reduced induced drag, but have only limited installation space for control surfaces near the wingtip. This paper describes a multidisciplinary design methodology for a morphing aileron that is based on pressure-actuated cellular structures (PACS). The focus of this work is on the transient dynamic system behavior of the multi-functional aileron. Decisive design aspects are the actuation speed, the resistance against external loads, and constraints preparing for a future wind tunnel test. The structural stiffness under varying aerodynamic loads is examined while using a reduced-order truss model and a high-fidelity finite element analysis. The simulations of the internal flow investigate the transient pressurization process that limits the dynamic actuator response. The authors present a reduced-order model based on the Pseudo Bond Graph methodology enabling time-efficient flow simulation and compare the results to computational fluid dynamic simulations. The findings of this work demonstrate high structural resistance against external forces and the feasibility of high actuation speeds over the entire operating envelope. Future research will incorporate the fluid–structure interaction and the assessment of load alleviation capability.

Commercial aircraft have well-designed and optimized systems, the result of a huge experience in the field, due to the large fleet of aircraft in operation. For light, utility, or sports aircraft, with a multitude of shapes, tasks, and construction types, there are different solutions that seek to best meet the requirements of the designed aircraft. In this sense, for a sport plane, an increased maneuverability is desired, and the system that controls flaps and wing must be properly designed. A new flap mechanism command solution is proposed and justified in the paper, for use in sports and recreational aviation, in order to achieve angles of braking greater than 40°, take-off and landing in a shorter time and over a shorter distance, as well as the gliding of the aircraft in critical flight conditions or when fuel economy is needed. A finite element model is used to verify the optimized command system for the flap and wing and to check if the strength structure of the aircraft is properly designed. The main result consists of the new design command system for flaps and wings and in verifying, by calculation, the acceptability of the new mechanism proposed from the point of view of the strength of the materials.

Since 2019, a key element of simulator-based training of airline pilots is stall training. A major and still largely open research question is which level of model fidelity is required for effective training. As part of an effort to answer this question, a model of the quasi-steady stall dynamics of a Cessna Citation II aircraft is identified from flight test data that was specifically collected for this research at an altitude of 5,500 m. To ensure any reductions in elevator and aileron effectiveness during stall were also explicitly measured, the test pilots used additional quasi-random flight test inputs. The considered stall model structure is based on Kirchoff's theory of flow separation. During identification, the nonlinear and linear parameters of the model are estimated in separate, recursively executed, steps. This separation enables the application of a semi-objective model structure selection method using multivariate orthogonal functions for the aerodynamic coefficients included in the model. This approach shows that stall-related effects should be included in the model equations for lift, drag, and pitch moment. Overall, it is found that the model parameters were consistently estimated from the flight test data and that the model accurately describes the aircraft's stall dynamics in the considered flight condition. The developed methodology is concluded to be well-suited for the direct identification of stall models from flight test data.

This paper presents an active control scheme aimed at the mitigation of the buffeting phenomenon on the wing of the high-performance aircraft. Buffeting is a common occurrence in flight where the aircraft undergo sharp turn rates and maneuvers. The high energy flow strikes the wing and the vertical tail of the aircraft and consequently, it excites the aeroelastic modes. The phenomenon of buffeting is modeled for the Finite Element (FE) model of Aluminum Fighter Aircraft (AFA). Buffet loads acquired through flight tests are mapped onto the wing of AFA model through Pade approximation. An accurate description of aeroelastic behavior is acquired in terms of Linear Time Invariant (LTI) state space system. In-house built software MASSA is used to formulate the Multiple Input Multiple Output (MIMO) system, where the buffet loads serve as the exogenous inputs to the aeroelastic system of the AFA model. The active control scheme based on Static Output Feedback (SOF) is used to actively control the ailerons. Two aileron deflection strategies optimized by the heuristic algorithm are proposed and compared in the presented research: firstly, the results are presented for the actuator which produces the deflection directly proportional to the excited disturbance and termed here as unconstrained movement, then an optimized strategy for aileron deflection is presented in which deflection is energy efficient with a compromise is made on the attenuation of the excited modes.

In the structural design of aircraft wings, aeroelastic tailoring is used to control the aeroelastic deformation to improve the aerostructural performance by making use of directional stiffness. Recently, tow-steered composites, where the fibre angles continuously vary within each ply, have been proven to have the potential to further expand the advantages of aeroelastic tailoring. This work extends TU Delft aeroelastic tailoring framework PROTEUS by introducing a lay-up retrieval step, so that it can be used for the conceptual design of tow-steered composite wing structures. In the extended framework, aeroelastic tailoring and lay-up retrieval are sequentially and iteratively performed to take static and dynamic loads, manufacturing and cruise shape constraints into consideration. The first step is carried out using PROTEUS, in which the lamination parameters and thickness of the wing sections are optimised under manoeuvre and gust load conditions. Further, for ensuring optimal aircraft performance in cruise flight conditions, the jig twist distribution is allowed to be optimised to maintain a desired prescribed cruise shape. In the second step, the stacking sequence, including minimum steering radius constraint, is retrieved. Since the lamination parameters cannot be matched exactly during the retrieval step, the constraints are checked, and tightened to take the performance loss during retrieval into account. The first step is repeated until all constraints are satisfied after fibre angle retrieval. To demonstrate the usefulness of the proposed optimisation framework, it is applied to the design of the NASA Common Research Model (CRM) wing, of which the objective is minimizing wing mass subjected to aerostructural design constraints, such as aeroelastic stability, aileron effectiveness, material strength and buckling load.

A micro-aerial vehicle (MAV) has been developed to achieve high versatility through biological inspiration from bird’s wings. One way of achieving optimized flight performance during different flight regimes is by incorporating a sweeping mechanism in the vehicle. Sweeping mechanisms allow a change in the aspect ratio of the wing as well as the overall span of the wing. The main challenge involves creating the mechanism to control the outboard section of the wing for a robust sweeping authority. A tendon-actuator mechanism with an elastic spring recovery system has been developed for this cause. In addition, macro-fiber composites (MFCs) were used as the control effectors for this vehicle both as ailerons and elevator. The placement of the MFCs was optimized for the required aileron effectiveness and overall aircraft performance. The analytical models were compared with the experimental results obtained through DIC, a full-field deformation measurement technique, for the entire wing.

Aiming at the problem of sudden threat avoidance in the process of aircraft formation cooperative penetration, formation collision avoidance strategies based on improved artificial potential field method are proposed. The effective range of the artificial repulsion potential field and the collision avoidance airspace of the sudden threat source are established. To improve the artificial potential field, the formation velocity correction coefficient and the threat source entry angle correction coefficient are introduced. In addition, the improved artificial repulsion potential energy function is defined. Based on the improved artificial potential filed in three-dimensional space, the formation collision avoidance controller on the horizontal and altitude channel are proposed respectively. Flight control commands such as rudder angle, aileron rudder angle, and elevator yaw angle are given. The simulation results confirm the security and efficiency of the formation collision avoidance strategy.

Modern aircraft structures uses more lightweight materials such as composites for their design. This makes the aeroelastic study an extremely important aspect of aircraft design. By having more light weight control surfaces, the control effectiveness study becomes vital in today’s scenario. Control effectiveness is the ability of a control surface such as an aileron or a rudder to produce aerodynamic forces and moments to change the airplane orientation and manoeuvre it along the intended flight path. This paper presents the static aeroelastic analysis of a typical reusable launch vehicle focusing on control effectiveness of elevon. A Reusable Launch Vehicle (RLV) is the space analogy of an aircraft. Ideally it takes off vertically on the back of a dispensable rocket and then glides back down like an aircraft. MSC PATRAN and MSC NASTRAN were the software’s used for Finite Element Modeling and Analysis. The main aim of the analysis is to compute the control effectiveness of launch vehicle along its trajectory to determine whether it is efficient to control the vehicle. The objectives of this work are, to study the control surface effectiveness of elevon by carrying out static aeroelastic analysis using NASTRAN inbuilt aerodynamics, for a typical Reusable Launch Vehicle (RLV).

In the present work, a reduced-order modeling (ROM) framework based on a recurrent neuro-fuzzy model (NFM) that is serial connected with a multilayer perceptron (MLP) neural network is applied for the computation of transonic aileron buzz. The training data set for the specified ROM is obtained by performing forced-motion unsteady Reynolds-averaged Navier Stokes (URANS) simulations. Further, a Monte Carlo-based training procedure is applied in order to estimate statistical errors. In order to demonstrate the method&rsquo, s fidelity, a two-dimensional aeroelastic model based on the NACA651213 airfoil is investigated at different flow conditions, while the aileron deflection and the hinge moment are considered in particular. The aileron is integrated in the wing section without a gap and is modeled as rigid. The dynamic equations of the rigid aileron rotation are coupled with the URANS-based flow model. For ROM training purposes, the aileron is excited via a forced motion and the respective aerodynamic and aeroelastic response is computed using a computational fluid dynamics (CFD) solver. A comparison with the high-fidelity reference CFD solutions shows that the essential characteristics of the nonlinear buzz phenomenon are captured by the selected ROM method.