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In vehicle control, control allocation is often used to abstract control variables from actuators, simplifying controller design and enhancing performance. Surveying available literature reveals that explicit solutions are restricted to strong assumptions on the actuators, or otherwise fail to exploit the capabilities of the actuator constellation. A remedy is to formulate hierarchical minimization problems that take into account the limits of the actuators at the expense of a longer computing time. In this paper, we compared the most common norms of the objective functions for linear or linearized plants, and show available numeric solver types. Such a comparison has not been found in the literature before and indicates that some combinations of linear and quadratic norms are not sufficiently researched. While the bulk of the review is restricted to control-affine plant models, some extensions to dynamic and nonlinear allocation problems are shown. For aerial vehicles, a trend toward linearized incremental control schemes is visible, which forms a compromise between real-time capabilities and the ability to resolve some nonlinearities common in these vehicles.

The hydrofoil vessel is among the few types of ships known to demonstrate negative added resistance in waves. This paper investigates the potential for maximizing this effect through the development of a neural network-based flight control system trained via reinforcement learning to actively extract wave energy. Furthermore, it examines the impact of such a control system on passenger comfort, and whether a feasible trade-off can be made between efficiency and comfort. The results suggest that significant energy savings can be achieved through the proposed methodology, achieving a power reduction of 61.8 % compared to that in calm water, for a simplified model of a hydrofoil vessel operating in 1.5 m regular head waves of 5 s wave period. If optimizing for passenger comfort, a reduction in vertical accelerations of 98% can be achieved, however, at the cost of achieving virtually no power reduction as compared to calm-water operation. By using a reward function with weighted focus on comfort and energy efficiency, it was demonstrated that the presented methodologies can produce a good trade-off between these two performance indicators. The presented cases differ only in terms of the flight control software, demonstrating the profound potential for altering hydrofoil vessel performance through intelligent control. • A method for DRL-based hydrofoil flight control. • Significant negative added resistance in waves, < −60%. • Cancellation of wave-induced motions with no resistance penalty. • Effective balancing of efficiency and comfort. • Discussions of the physics of negative added resistance of hydrofoil vessels. [ABSTRACT FROM AUTHOR]

The flight control system of coaxial compound helicopters (CCHs) is a complex multi-variable system characterized by redundant control effectors, strong nonlinear characteristics, and multiple flight modes. This paper presents a unified robust control framework tailored for CCH full-envelope flight, utilizing nonlinear dynamic inversion and extended state observer techniques. The framework is specifically designed to tackle issues arising from a multi-mode nature, parametric uncertainties, and external disturbances. Within the unified framework, the study addresses inherent complexities such as nonlinearities and cross-coupling effects in CCH by leveraging nonlinear dynamic inversion to formulate flight control laws for both inner and outer-loops. To mitigate model uncertainties and external disturbances, the extended state observer is employed to estimate lumped disturbance effectively. Control allocation for both inner and outer-loops is devised to adapt to changes in control effector authorities and flight modes. Furthermore, an exponential control allocation strategy is proposed for the outer-loop to ensure a smooth pitch angle during transition flight. The effectiveness of the proposed control laws and control allocation strategies in achieving precise attitude tracking, transition flight and mission task elements (MTEs), even in the presence of notable model uncertainties and external disturbances, is demonstrated through numerical simulations. • A unified robust nonlinear dynamic inversion control framework is presented for full envelope flight of the CCH. • An outer-loop control allocation strategy is proposed to address the issue of pitch angle command oscillation during transition flight. • A simplified nonlinear dynamic model of CCH is employed to mitigate the effects of model uncertainties during the design of inner-loop control law. [ABSTRACT FROM AUTHOR]

In this paper, an innovative dynamic observer is applied to complex nonlinear interconnected systems with matched and mismatched uncertainties. This dynamic observer can estimate system states, which cannot be achieved during the design process for nonlinear interconnected systems with uncertainties. The proposed method has great identification ability with small estimated errors for the states of nonlinear interconnected systems with matched and mismatched uncertainties. It should be pointed out that the considered uncertainties of nonlinear interconnected systems have general forms, which means that the proposed method can be effectively used in more generalised nonlinear interconnected systems. Finally, simulation results for the lateral flight control system are presented to ensure the effectiveness of this strategy. [ABSTRACT FROM AUTHOR]

A fault detection design is proposed for addressing the Oscillatory Failure Case (OFC) detection problem, introduced in the joint Airbus-Stellenbosch university aerospace industrial-benchmark competition called at the IFAC 2020 World Congress 1 1 https://www.ifac2020.org/program/competitions/aerospace-industrial-fault-detection.. The detection scheme is comprised of an output estimator, a wavelet decomposition and an energy-based denoising method, and the residual evaluation unit. The detection problem of wide frequency range OFCs is also addressed. According to the achieved simulation results, the proposed fault detection method is able to satisfy the competition prescriptions in the frequency range [1 10] Hz for those OFC's having an amplitude greater than 2.3 mm for OFCs at rod position sensor, or 1.4 mA for OFCs at servo input current, regardless of disturbances level, uncertainties and load factor control input. In other cases, faults are detected slightly after the prescribed detection limit, with some interesting exceptions. [Display omitted] • An oscillatory failure case detection scheme is presented to be used in the flight control systems. • A general detection solution using discrete wavelet decomposition is provided. • OFCs with amplitude beyond 2.3 mm at the rod sensor and 1.9 mA in the servo input are detectable. [ABSTRACT FROM AUTHOR]

Flight control surfaces guarantee a safe and precise control of the aircraft. As a result, hinge moments are generated. These moments need to be estimated in order to properly size the aircraft actuators. Control surfaces include the ailerons, rudder, elevator, flaps, slats, and spoilers, and they are moved by electric or hydraulic actuators. Actuator sizing is the key when comparing different flight control system architectures. This fact becomes even more important when developing more-electric aircraft. Hinge moments need to be estimated so that the actuators can be properly sized and their effects on the overall aircraft design are measured. Hinge moments are difficult to estimate on the early stages of the design process due to the large number of required input. Detailed information about the airfoil, wing surfaces, control surfaces, and actuators is needed but yet not known on early design phases. The objective of this paper is to propose a new methodology for flight control system sizing, including mass and power estimation. A surrogate model for the hinge moment estimation is also proposed and used. The main advantage of this new methodology is that all the components and actuators can be properly sized instead of just having overall system results. The whole system can now be sized more in detail during the preliminary design process, which allows to have a more reliable estimation and to perform systems installation analysis. Results show a reliable system mass estimation similar to the results obtained with other known methods and also providing the weight for each component individually. [ABSTRACT FROM AUTHOR]

Helicopter dronization is expanding, for example, the VSR700 project. This leads to the integration of electromechanical actuators (EMAs) into the primary flight control system (PFCS). The PFCS is in charge of controlling the helicopter flight over its four axes (roll, pitch, yaw, and vertical). It controls the blade pitch thanks to mechanical kinematics and actuators. For more than 60 years, the actuators have been conventionally using the hydraulic technology. The EMA technology introduction involves the reconsideration of the design practices. Indeed, an EMA is multidisciplinary. Each of its components introduces new design drivers and new inherent technological imperfections (friction, inertia, and losses). This paper presents a methodology to specify and pre-design critical EMAs. The description will be focused on two components: the electrical motor and the housing. This includes a data-driven specification, scaling laws for motor losses estimation, and surrogate modeling for the housing vibratory sizing. The tools are finally applied to two study cases. The first case considers two potential redundant topologies of actuation. The housing sizing shows that one prevails on the other. The second case considers the actuators of helicopter rotors. The electrical motor sizing highlights the importance of designing two separate actuators. [ABSTRACT FROM AUTHOR]

A rice vortex generated by the downwash airflow of the agricultural UAV has a great connection to pesticide droplet deposition. In this study, an unmanned aircraft system (UAS) was designed to obtain rice vortex parameters including position coordinates, area, and relative distance to the UAV. The system uses the flight control system mounted Beidou RTK positioning system and Manifold onboard computer to localize the spraying UAV and trains a YOLOv3-tiny model to recognize motion images of dynamic rice vortices in real-time, achieving a mean accuracy of 98.15% with 0.9326 IOU. Moreover, to optimize the system, a data time axis alignment method based on the robot operating system (ROS) was proposed to alleviate the transmission delay and image processing time-consumption, which increased the accuracy of the vortex parameter by 3.48%. The rice vortex area acquired by the UAS was highly relevant to UAV flight parameters and the relative distance was just correlated to the UAV height. The instantaneous status of the rice vortex helps to lay a data foundation for adjusting the flight parameters of the plant protection UAV. [ABSTRACT FROM AUTHOR]

Copyright of Systems Engineering & Electronics is the property of Journal of Systems Engineering & Electronics Editorial Department and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

This survey deals with the problem of the group motion of spacecraft, which is rapidly developing and relevant for many applications, in terms of developing the onboard control algorithms to ensure the fulfillment of a given mission. The paper provides a comprehensive overview of spacecraft formation flight control. The bibliography is divided into three main sections: the multiple-input–multiple-output approach, in which the formation is treated as a single entity with multiple inputs and multiple outputs; the leader–follower formation, in which individual spacecraft controllers are linked hierarchically; and a virtual structure formation, in which spacecraft are treated as rigid bodies embedded in a common virtual rigid body. This survey expands a 2004 survey and updates it with recent results. [ABSTRACT FROM AUTHOR]

Flight testing in a realistic three-dimensional virtual environment is increasingly being considered a safe and cost-effective way of evaluating aircraft models and their control systems. The paper starts by reviewing and comparing the most popular personal computer-based flight simulators that have been successfully interfaced to date with the MathWorks software. This co-simulation approach allows combining the strengths of Matlab toolboxes for functions including navigation, control, and sensor modeling with the advanced simulation and scene rendering capabilities of dedicated flight simulation software. This approach can then be used to validate aircraft models, control algorithms, flight handling chatacteristics, or perform model identification from flight data. There is, however, a lack of sufficiently detailed step-by-step flight co-simulation tutorials, and there have also been few attempts to evaluate more than one flight co-simulation approach at a time. We, therefore, demonstrate our own step-by-step co-simulation implementations using Simulink with three different flight simulators: Xplane, FlightGear, and Alphalink's virtual flight test environment (VFTE). All three co-simulations employ a real-time user datagram protocol (UDP) for data communication, and each approach has advantages depending on the aircraft type. In the case of a Cessna-172 general aviation aircraft, a Simulink co-simulation with Xplane demonstrates successful virtual flight tests with accurate simultaneous tracking of altitude and speed reference changes while maintaining roll stability under arbitrary wind conditions that present challenges in the single propeller Cessna. For a medium endurance Rascal-110 unmanned aerial vehicle (UAV), Simulink is interfaced with FlightGear and with QGroundControl using the MAVlink protocol, which allows to accurately follow the lateral UAV path on a map, and this setup is used to evaluate the validity of Matlab-based six degrees of freedom UAV models. For a smaller ZOHD Nano Talon miniature aerial vehicle (MAV), Simulink is interfaced with the VFTE, which was specifically designed for this MAV, and with QGroundControl for the testing of advanced H-infinity observer-based autopilots using a software-in-the-loop (SIL) simulation to achieve robust low altitude flight under windy conditions. This is then finally extended to hardware-in-the-loop (HIL) implementation on the Nano Talon MAV using a controller area network (CAN) databus and a Pixhawk-4 mini autopilot with simulated sensor models. [ABSTRACT FROM AUTHOR]

The Flight Control System (FCS) is considered as the brain of an aerial vehicle. It is a mechanism through which pilot's commands are transferred to the actuators of the aircraft control surfaces. In order to ensure safety and increase reliability of aerial vehicles, development of fault tolerant FCSs has been the focus of research community for past few decades. Fault tolerant ability enables an aircraft to maintain satisfactory performance even in the state of a fault. Fault Tolerant Control Systems (FTCS) are categorized as passive and active control systems. Passive FTCS are designed to mitigate the effects of certain known faults. These faults can be related to sensor failure, actuator failure, or system component failure. On the other hand, active FTCS contain a controller reconfiguration mechanism, whereby, they can adjust the controller input online to mitigate the effects of the faults. In this way, they can accommodate complicated and versatile faults as compared to their passive counterparts. This paper presents a review of significant research during last decade in active fault tolerant control with applications to FCSs. A review of state-of-the-art works in this domain has also been presented. Upon review, these state-of-the-art research interests have been categorized into respective categories. Furthermore, research works have been cataloged based on their technology readiness levels. Based on these reviews, future research directions have also been highlighted. [ABSTRACT FROM AUTHOR]

Copyright of Journal of Computer Engineering & Applications is the property of Beijing Journal of Computer Engineering & Applications Journal Co Ltd. and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

In this paper, a robust flight control system is proposed for an autonomous quadrotor to quickly and accurately achieve the targeted trajectories. A novel supertwisting nonsingular terminal sliding mode control (STNTSMC) has been developed to ensure that the tracking errors vanish in a short finite-time. A nonlinear disturbance observer (DO) is incorporated into the control system to estimate the unknown external perturbations and to strengthen the system's robustness. The Lyapunov theory is used to verify the closed-loop stability of the synthesized controller. Moreover, processor-in-the-loop (PIL) implementations are performed to validate the efficacy of the suggested method. The merit of the proposed DO-STNTSMC is evaluated under multiple flight scenarios. The obtained results demonstrate that the proposed controller has a highly reduced tracking error and strong robustness against random parameter changes and external disturbances, compared to conventional nonsingular terminal sliding mode control. Finally, experimental tests are conducted to validate the performance of the suggested method. [ABSTRACT FROM AUTHOR]

This paper proposes an actuator fault diagnosis method for a class of discrete-time linear systems subject to parametric uncertainties, unknown but bounded disturbance and noise. The actuator faults of the systems are taken as unknown inputs, and the proposed unknown input observer has a new structure that provides more robust and more design degrees of freedom. A zonotope-based method is used to estimate the threshold of residual generated by the unknown input observer. Based on the proposed method, a fault detection and isolation strategy using a set of unknown input set-membership observers is presented. The effectiveness of the proposed method is verified by numerical simulation of a flight control system. [ABSTRACT FROM AUTHOR]

Assigning development assurance levels for items/functions and conducting corresponding development assurance activities can minimize the possibility of errors in the development process. The development assurance levels of the items/functions consisting of the system are taken as decision variables, the assignment principle of development assurance levels and the probability requirement of the top failure conditions are taken as constraints, and the minimization of the development cost is taken as the optimizing objective, the model of the development assurance level assignment is established. Taking the vector composed of the development assurance levels of all items/functions as the individual chromosome, a method of solving the model is proposed based on the a genetic algorithm and particle swarm optimization (GA-PSO) hybrid algorithm. Finally, an application instance is given based on a hypothetical air-borne system and a certain fly-by-wire flight control system. The results show that the proposed method can reduce the dependence on designers' experiences or skills effectively, and has higher accuracy and efficiency compared with other algorithms. [ABSTRACT FROM AUTHOR]

Competing failures are time domain contention situations between the propagated failures (PFs) that originate from dependent components and the failure isolation caused by the trigger component. The methods based on combinatorial analysis commonly used in the analysis of competing failures require a complicated formula derivation and model reduction process. This paper proposes an integrated model based on generalized stochastic Petri nets (GSPNs) for analyzing the competing failures in the system, and further considers the effect of common cause failures (CCFs). The proposed modeling method inherits the advantages of GSPNs and provides a simplified method to compute the reliability of systems, which affect by competing failures and CCFs. Finally, the proposed method is applied in the flight control system (FCS) and demonstrated by the efficient decomposition and aggregation (EDA) method and combinatorial analysis method. [ABSTRACT FROM AUTHOR]

Because the Polish Air Force currently uses F-16 and MiG-29 aircraft, the aim of the study was to conduct a comparison of the susceptibility to destruction in combat of these two aircraft. The first part of the work concerned the analysis of individual critical components, such as: general characteristics of the airframe structure, aircraft engine, flight control system, fuel system, aircraft weapons, radar system. The index of susceptibility to destruction in combat was defined considering the listed critical components and the following types of enemy weapon: aircraft gun, air-to-air missile, anti-aircraft gun, surface-to-air missile. The analysis proved that the aircraft have similar susceptibility to damage in combat, a slight advantage of the F-16 aircraft in this respect was determined. The presented scheme can be used to analyze other aircraft. Proposals of aircraft modifications, directions of further actions, possibilities of using the described method were presented. The method can be used to making decisions by governments regarding the purchase aircraft for their fleets and identify aircraft critical components with high susceptibility to destruction in order to introduce appropriate modifications by military aircraft manufacturers. [ABSTRACT FROM AUTHOR]

The objective of this investigation is threefold. First, to assess the flight dynamics of an electric vertical take-off and landing (eVTOL) concept aircraft with a propeller-driven rotor. Second, to develop an automatic flight control system (AFCS) for this concept aircraft. Third, to verify the potential safety benefits of the concept aircraft by analyzing the autorotation performance following a total loss of power. The paper begins with an overview of the design of the aircraft and description of the simulation model, including a detailed discussion on the inflow model of the propellers that drive the main rotor. Next, the flight dynamics are assessed at hover and in forward flight. An AFCS based on dynamic inversion is developed to provide stability and desired response characteristics about the roll, pitch, yaw, and heave axes for speeds ranging from hover to 80 kt. Additionally, an RPM governor is implemented to hold the main rotor angular speed constant at its nominal value. Finally, simulations that make use of the AFCS are performed to analyze the autorotation performance following total loss of power. [ABSTRACT FROM AUTHOR]

The conventional aircraft layout will contain the elevator, aileron, rudder three control surfaces, and thus generate three control moments to adjust the aircraft trajectory and flight attitude. However, the aircraft with multi-control plane flying wing layout needs to choose redundant new control plane configuration, which is quite different from the conventional aircraft layout in both control method and operation mode. Since there are many problems in the aircraft motion mathematical model during the ownership period, which will directly affect the control effect of the flight control system, in order to reduce dynamic error and improve the robustness and accuracy of the flight control system, this paper proposes an adaptive neural network algorithm. Combined with the current operating status of tailless flying wing aircraft and the research status of aircraft control system analysis, this paper mainly discusses the design differences between the adaptive control system and the original control system, based on the empirical case, to clarify the compensation effect of the adaptive control system. [ABSTRACT FROM AUTHOR]

In this paper, a novel design approach for low-speed recovery of a high-performance fighter aircraft is presented. It is shown that the phugoid mode has an important bearing on the problem of low-speed departure. Based on the analysis of the phugoid mode trajectories, a novel low-speed protection algorithm is presented in this paper. The proposed low speed recovery is achieved in three phases. The first phase consists of detecting the incipient departure followed in the second phase by the application of suitable recovery controls and finally the third phase ends with the transfer of controls to the pilot. The design of the first and the third phase consist of choosing the correct trigger conditions which ensures safe recovery of the aircraft in all conditions. The proposed Automatic low speed recovery is triggered when the aircraft trajectory crosses a fixed boundary in the region spanned by the dynamic pressure and its rate of decrease. It is observed that this boundary is approximately a straight line, implying that it is equivalent to a forward prediction in time to indicate when the aircraft will reach the lowest controllable airspeed. This Automatic Low Speed Recovery with Forward Prediction (ALSR-FP) algorithm is found to be simpler than other existing design methods and effective in preventing low speed departure for a variety of pilot inputs that result in the aircraft losing airspeed leading to stall. In the second phase control inputs are chosen to align the velocity vector to the direction of local gravity. The recovery phase is considered complete after the aircraft reaches the dynamic pressure which is approximately 10 % higher than the minimum dynamic pressure for control. Performance of the ALSR-FP is demonstrated using the high-performance fighter aircraft ADMIRE model which has a delta wing configuration, canards and multiple redundant controls. It is also shown that the proposed algorithm can be easily implemented on board for any other fighter and civil aircraft. [ABSTRACT FROM AUTHOR]

Adoption of electromechanical actuation systems in aerospace is increasing, and so reliable diagnostic and prognostics schemes are required to ensure safe operations, especially in key, safety-critical systems such as primary flight controls. Furthermore, the use of prognostics methods can increase the system availability during the life cycle and thus reduce costs if implemented in a predictive maintenance framework. In this work, an improvement of an already presented algorithm will be introduced, whose scope is to predict the actual degradation state of a motor in an electromechanical actuator, also providing a temperature estimation. This objective is achieved by using a properly processed back-electromotive force signal and a simple feed-forward neural network. Good prediction of the motor health status is achieved with a small degree of inaccuracy. [ABSTRACT FROM AUTHOR]

Copyright of Control Theory & Applications / Kongzhi Lilun Yu Yinyong is the property of Editorial Department of Control Theory & Applications and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

Copyright of Radio Electronics, Computer Science, Control is the property of Zaporizhzhia National Technical University and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

Currently, quite accurate measurements of atmospheric gusts are carried out by airport systems only in the vicinity of the runways. There is a still open issue of availability of information about real wind gusts at cruising altitudes and during approach at a considerable distance from the airfield. Standard on-board systems of a jet transport airplane provide some information which is desirable to have knowledge of how flight parameters reflect real gust parameters and their impact on the aircraft dynamics. The paper proposes an algorithm for headwind gust magnitude estimation in relation to aircraft response. The analysed estimation algorithms assume the use of data available from the existing on-board systems only without the employment of any extra sensors or ground and satellite systems. In this way, many problems caused by different structures, configurations, and ways of installation of additional sensors and structural changes are rejected. The algorithms use the classical method for estimation of wind parameters as well as a linear longitudinal model of aircraft dynamics, taking into account the influence of wind gusts. Data fusion was realised with the use of three filtration methods. Results were evaluated to select the most accurate method of the estimation. Test data were obtained from advanced flight simulation. The experimental scenario considered a flight of a passenger twin-engine jet airplane through a layer of programmed gusts. The results of the flight simulations allowed us to determine the accuracy of the proposed gust estimation algorithms in reference to the ideal wind-speed data analysis obtained directly from the simulation environment (with the accuracy of the simulation process). The use of the proposed gust estimation algorithms may provide more accurate signal for integrated on-board systems, especially for wind shear detection and sped-up response time of flight control systems, protecting aircrafts against the adverse impact of encountered wind shear or gusts, e.g., auto-thrust or auto-throttle systems. The dedicated algorithm presented in the paper may increase the safety level of take-off and approach phases in gusty conditions and also during significant changes in wind speed at cruising altitudes in the case of crossing the area of jet stream occurrence. [ABSTRACT FROM AUTHOR]

The evolution of large transport aircraft is characterized by longer fuselages and larger wingspans, while efforts to decrease the structural weight reduce the structural stiffness. Both effects lead to more flexible aircraft structures with significant aeroelastic coupling between flight mechanics and structural dynamics, especially at high speed, high altitude cruise. The lesser frequency separation between rigid body and flexible modes of flexible aircraft results in a stronger interaction between the flight control system and its structural modes, with higher flexibility effects on aircraft dynamics. Therefore, the design of a flight control law based on the assumption that the aircraft dynamics are rigid is no longer valid for the flexible aircraft. This paper focuses on the design of a flight control system for flexible aircraft described in terms of a rigid body mode and four flexible body modes and whose parameters are assumed to be varying. In this paper, a conditional integral based sliding mode control (SMC) is used for robust tracking control of the pitch angle of the flexible aircraft. The performance of the proposed nonlinear flight control system has been shown through the numerical simulations of the flexible aircraft. Good transient and steady-state performance of a control system are also ensured without suffering from the drawback of control chattering in SMC. [ABSTRACT FROM AUTHOR]

In the paper, a new cascade control system for an autonomous flight of an unmanned aerial vehicle (UAV) based on Proportional–Integral–Derivative (PID) and finite-time State-Dependent Riccati Equation (SDRE) control is proposed. The PID and SDRE controllers are used in a hybrid control system for precise control and stabilization, which is necessary to increase the drone's flight stability and maneuver precision. The hybrid PID-SDRE control system proposed for the quadrotor model is quasi-optimal, since the suboptimal control algorithm for the UAV stabilization is used. The combination of the advantages of PID and SDRE control gives a significant improvement in the quality of control while maintaining the simplicity of the control system. Furthermore, the use of the suboptimal control technique provides the UAV attitude tracking in finite time. These remarks are drawn from a series of simulation tests conducted for the drone model. [ABSTRACT FROM AUTHOR]

Development of a society in last decades is characterized by automation of person activity in various spheres. In spite of achievements in the field of computer science and information technology, in most cases control systems by aircraft, railway vehicle, nuclear reactor can be considered as automated control systems. From the point of view of control theory, an automated control system implies the presence of human-operator in control loop. Statistics of anthropogenic and aviation accidents in most cases indicates a significant influence of the human factor on the cause of their initiation. This allows us to make a significant general conclusion that the presence of a human operator in the control loop reduces the reliability of the system or, in other words, the low reliability of the ergatic component as an integral part of the control system. The analysis of the principles of aircraft control system development, from Mechanical flight control system, Conventional mechanical, Fly-by-Wire, and up to Fly-by-Wireless, shows that there is no information about the cybernetic ability of a human operator (pilot) in the control loop. The presented study proposes a conceptual approach to the possibility of integrating the quantitative markers of a human operator cyber ability into the control loop of the system to improve its reliability. [ABSTRACT FROM AUTHOR]

• Demonstration of use of electrical converters ANPC in aircrafts applications offering huge technical advantages compared to the classical converters. • Theoretical, simulation and experimental comparative study of performances in terms of reliability for three types of converters. • Use of military handbook stress-part taking into account real environmental (aeronautics) parameters. • Comparison of results with manufacturers' parameters and reliability computations. In more electrical aircraft, most flight control wing surface actuators are either electro-hydrostatic or electro-mechanical, whether in manual or autopilot mode. Their reliability is essential for flight safety; failure of even one actuator can have catastrophic consequences. According to ARP4761 guidelines for the development and certification of critical aircraft systems, the probability of failure of flight-control actuator components must be less than 10-9 per hour of flight. In this paper, we compare the probability of failure per hour (PFH) of conventional two-level power converters and three-level NPC and ANPC converter topologies designed in our laboratory. In accordance with ARP procedures, PFH was calculated using an approach that includes qualitative analysis based on failure mode and effect analysis (FMEA) methodology as well as quantitative analysis based on truth tables. ANPC converter met all the requirements of the functional hazard assessment in all cases of defects. Finally, we discuss factors that determine the feasibility of implementing power converters in flight control systems. [ABSTRACT FROM AUTHOR]

The use of electromechanical actuators (EMAs) for aeronautical applications promises substantial benefits regarding efficiency and operability. To advance the design of power electronics and secondary power supply, there is a need for the ability to swiftly study the effects of aircraft mission and operational aspects on the actuator energy consumption. Pursuant to this, the aim of the work presented in this paper is twofold: (i) to build a generic mission-level flight control surface EMA power consumption simulation framework and (ii) to apply this framework to a case study involving a small all-electric aircraft, in which selected factors that impact energy consumption are investigated. The core of the framework comprises physics-based EMA power estimators, linked with a six-degree-of-freedom flight dynamics and control simulation module. The case study results show that the actuator power consumption correlates positively with the proportional gains in the flight control system but is inversely proportional to the trajectory radius and linearly dependent on turbulence intensity. The developed framework could aid in the selection of the actuator, as well as in the optimisation of airborne electronics and secondary power supply. [ABSTRACT FROM AUTHOR]

The paper presents the concept of a component of an aircraft's automatic flight control system, controlling the airplane when in longitudinal motion (i.e., pitch angle, sink rate, airspeed channels) during automatic landing, from a final approach until a touchdown. It is composed of two key parts: a vision system and an automatic landing system. The first part exploits dedicated image-processing algorithms to identify the number of red and white PAPI lights appearing on an onboard video camera. Its output data—information about an aircraft's position on a vertical profile of a landing trajectory—is used as one of the crucial inputs to the automatic landing system (the second part), which uses them to control the landing. The control algorithms implemented by the automatic landing system are based on the fuzzy logic expert system and were developed to imitate the pilot's control actions during landing an aircraft. These two parts were teamed together as a component of a laboratory rig, first as pure software algorithms only, then as real hardware modules with downloaded algorithms. In two test campaigns (software in the loop and hardware in the loop) they controlled an aircraft model in a simulation environment. Selected results, presenting both control efficiency and flight precision, are given in the final section of the paper. [ABSTRACT FROM AUTHOR]

The paper concentrates on the problem of fault-tolerant control of UAV against actuator faults from the perspective of flight control system architecture. Using backstepping control method and inverse optimization theory, the design of backstepping optimal control law was constructed. Based on the fault monitoring mechanism of vehicle management computer in the distributed flight control and control allocation system, a fault-tolerant control design method was established in the case of multiple failure modes of the actuators, which compensates the influence caused by the failures. Finally, the effectiveness of the proposed strategy was verified by numerical simulation. [ABSTRACT FROM AUTHOR]

This study is aimed in order to study the data protection software of civil aviation management flight information system based on FPE algorithm. The author established the data warehouse software of the air traffic control production information statistics system. First, using the Jquery browser application framework + service-oriented Web Service architecture technology, a B/S/S multilayer data application based entirely on browser applications is established. Then, with the support of the ATC production information statistical data warehouse software and the system application framework, the system realizes a series of production data processing related to production information management, such as data collection, data editing, data review, and business data statistics. Finally, combined with the attributes of business data, we analyze and compare all kinds of production information from multiple dimensions such as time and region; it also provides users with data analysis results of curve, pie, column, and other performance situations, display mode, and report output tool. SQL statement can be implemented to add data to the database, modify the existing data in the database, delete the existing data in the database, and can be queried in a variety of ways to the data in the database, and VB provides a human-computer interaction interface for various operations. [ABSTRACT FROM AUTHOR]

This paper describes the design and flight test of the control system of a novel octocopter with new topology to improve aerodynamic performance and efficiency. The article analyzes the advantages of the new configuration and analyzes the feasibility of the control method which can ensure fault-tolerant control when one rotor of the aircraft stops in theory. The feasibility verification is carried out through a prototype flight test. The power and hardware platform of the octocopter UAV was set up according to the requirements of structural and control system design. Flight tests were carried out multiple times, and the UAV's redundant actuation was able to successfully stabilize the vehicle, even after a single rotor stopped functioning. Under normal flight conditions, smooth flight and effective control could be ensured for the UAV, while under the condition of a single rotor having stopped functioning, the control method was able to effectively utilize the other seven rotors to provide a proper lift force and to control the aircraft to perform basic motions such as pitch, roll, and yaw. [ABSTRACT FROM AUTHOR]

Model predictive control (MPC) is receiving increasing attention in space applications, as a key technology for enhancing autonomy of the flight control system. Sum-of-norms formulations are specifically suited to this context, because they allow to optimize meaningful performance figures and to promote control sparsity. This brief presents a sum-of-norms MPC scheme for linear periodically time-varying systems. Closed-loop stability is proven by suitably defining periodic sequences of terminal weights and terminal sets. The proposed solution is applied to a rendezvous case study involving periodic dynamics due to geopotential effects and solar eclipses. [ABSTRACT FROM AUTHOR]

One of the main concerns associated with diagnosis, prognosis, and health management (DPHM) of engineering systems is the accuracy of estimates that are derived from Bayesian tracking methods. Estimating the exiting degradation based on stochastic models and evaluating the remaining useful life (RUL) of the system is inherently associated with variances that characterize the inaccuracy of estimation techniques. Furthermore, there are scenarios where a single measurement does not necessarily generate sufficient information regarding the system states, leading one to require multiple readings (and, hence, multidimensional analysis) to deduce diagnostic and/or prognostic decisions. This article introduces a novel approach for solving complex nonlinear multivariable Bayesian models that are utilized for estimation and prediction problems that would be, otherwise, challenging or impractical to solve through available methods, such as particle filters (PFs). Theoretical derivation and strategies that are developed in this article are verified through numerical case study simulations for electrohydraulic servo valves (EHSVs) that constitute a core component of many hydraulic actuators, such as multifunctional spoilers (MFSs), which are widely utilized in aircraft flight control systems. Our developed results are compared with those that are derived through PF in order to illustrate and demonstrate the advantages, benefits, and improvements that are accomplished by applying our proposed methodologies. [ABSTRACT FROM AUTHOR]

This study investigates the relationship of pilot inceptor workload (PIW) with the workload buildup flight test technique (WBFTT) with various conditions of pilot workload, for objective aircraft handling qualities (HQ) evaluation. HQ evaluation of a piloted aircraft remains an elusive area of flight testing, due to the existing subjective evaluation technique for the pilot workload. The pilot control inceptor being an important interface for a pilot with the aircraft, it is the best tool to estimate pilot workload. The statistical model of this research involved 3 independent variables, namely aircraft flying qualities, secondary task, and boundary effect of WBFTT, each differently affecting the pilot workload. Two studies were undertaken on a fixed-base, variable-stability HQ research flight simulator with military test pilots. In-flight study of pilot inceptor movement was undertaken in an advanced jet trainer aircraft during high pilot-gain air-to-ground target tracking tasks. The results of simulator studies validated the relationships of PIW and WBFTT with statistical significance from a wide set of data, with variations in pilot workload in terms of flying qualities and secondary task. The in-flight studies validated the effects of high pilot gain and proximity to ground (boundary), on the pilot inceptor movements, in training combat maneuvers. Studies manifested PIW as a simple and direct measure to estimate pilot workload and WBFTT as an effective technique for HQ stress testing with high pilot gain. [ABSTRACT FROM AUTHOR]

The aim of this research was to understand risk, predictors of risk, and for pilots the relationship between personality as determined by the risk prediction scales and flight performance. Risk-taking is a topic that draws attention because of its link with the causation of injury. In the literature, there are a large number of papers linking personality traits to self-reported risk-taking, in addition to theoretical commentaries. In contrast, there are very few empirical studies that examine the link between personality traits, such as risk-taking and actual risk-taking behavior. One hundred university students with no flying experience (59 females) and 17 pilots (five females) completed a battery of personality and risk scales. Pilots were also tasked to complete a simulated flight involving a low-level flying task. The results revealed that pilots had a higher propensity for risk than the general population. However, this risk was tempered by the personality factor of self-control. Evidence of this juxtaposed position was present during the simulated flight; pilots' propensity for risky behavior was inversely related to actual risk-taking behavior. These results indicate a dichotomy in risk-takers, impetuous and calculative risk-takers. These findings have important implications for pilot selection and training. [ABSTRACT FROM AUTHOR]

The main advantage of the incremental dynamics formulation is its property of replacing state-dependent aspects of the system dynamics by measurements, solely requiring modeling the input-dependent dynamics. However, incremental controllers assume that state derivative measurements are available. The lack of accurate sensors and communication delay may appear as limitations for the use of incremental techniques. This article applies a second-order differentiator (SOD) for estimating state derivatives and attenuate measurement noise, and an input scaling gain (ISG) for attenuating the effects of high-frequency noise and delay in the feedback loop. A complete analysis of the combined solution is carried out using an incremental nonlinear dynamic inversion (INDI) strategy. First, a simple linear single-input single-output model is used to evaluate the influence of design parameters in the closed-loop response and stability. Then, the combined SOD and ISG approach is illustrated in a more complex case application, where INDI is used to control the lateral motion of an autonomous airship, results of which corroborate the ISG as an asset to increase the maximum delay allowed in the feedback loop. [ABSTRACT FROM AUTHOR]

The difficulty of obtaining accurate dynamical models of hypersonic flight has been generally recognized, and modeling inaccuracy can severely deteriorate the performance of the flight control systems. To address this issue, an adaptive control approach using two neural networks (NNs) is proposed with the aim of achieving precise and robust control for hypersonic flight when unknown dynamics are involved. Different from the existing adaptive control methods, two NNs are developed in this paper to learn the forward and inverse dynamics of hypersonic flight with guaranteed convergence. Particularly, this study focuses on the following three contributions. First, an iterative model learning algorithm is proposed to train the first NN to approximate the unmodeled system dynamics and achieve accurate observations of flight responses and unknown dynamics. Second, an iterative controller learning algorithm is proposed to guide the second NN to learn the control inputs from prior flight data and improve the dynamic performance of the adaptive controller. Third, an adaptive NN-based controller for trajectory tracking is developed combining the above two NNs. Simulations are provided to substantiate the effectiveness of the proposed techniques and demonstrate the excellent adaptability and robustness of the controller. • An iterative model learning algorithm is proposed to learn the unmodeled system dynamics. • An iterative controller learning algorithm is proposed to learn the control inputs from prior data. • A dual neural networks architecture is proposed to achieve intelligent adaptive control. [ABSTRACT FROM AUTHOR]

This article introduces an original model-based design methodology addressing a high-performance aircraft design challenge: conflicting performance requirements. The case study of the Global 7500 elevator actuation system also provides in-depth insight into the complex design process of today's fly-by-wire flight control systems. The methodology presented here redefines the aircraft manufacturer's involvement in the design process of the systems, implementing analysis and iteration capabilities early in system development. To this end, it introduces a novel modeling approach for analyzing loaded rate requirements by simulating closed-loop performance with a generic nonlinear second-order state filter, including the main performance limitations without requiring a preliminary design definition. In this way, it provides means to mature the system requirements and addresses requirement conflicts upfront. Then, a simulation-based preliminary sizing and performance assessment validates the candidate design concept. It also secures the preliminary design phase by implementing advanced design uncertainties and involving interfacing systems and disciplines early in the process. The redefined methodology identified directly that the problem's root cause was a conflict between stability and control and flutter protection requirements. It also indicated that the first sizing driver is the response time required under a specific failure case. These findings lead to an optimal elevator actuator design compliant with matured performance requirements. Thus, the methodology resolved a design challenge blocking the Global 7500 aircraft development and prevented redesign occurrences later during the detailed design phase. In this way, it directly contributed to the successful development of the Global 7500 and its optimal operational performance. This methodology applies to future aircraft design challenges, and the technical insight provides valuable lessons learned for high-performance T-tail business jets. [ABSTRACT FROM AUTHOR]