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In this study, an aerial tour guide drone was developed based on the existing Coanda‐drone. The research integrated a Voice Information Presentation System with ROS2 using the open‐source PX4 flight control system to update the drone's control module. This system can recognize and read QR codes on displayed items and convert the text information from QR codes into speech for playback. It also utilizes left and right laser sensors for obstacle avoidance. Furthermore, the study conducted both simulation experiments and actual flight tests to validate the feasibility of equipping the Coanda‐drone with the voice information presentation system and to assess the flight capabilities of the Coanda‐drone. This system offers a novel solution for guidance and obstacle avoidance, demonstrating the potential of drones as flexible, highly secure, and efficient mobile service robots in the future. © 2024 Institute of Electrical Engineers of Japan and Wiley Periodicals LLC. [ABSTRACT FROM AUTHOR]

Drones have witnessed extensive popularity among diverse smart applications, and visual Simultaneous Localization and Mapping (SLAM) technology is commonly used to estimate the six-degrees-of-freedom pose for drone flight control systems. However, traditional image-based SLAM cannot ensure the flight safety of drones, especially in challenging environments such as high-speed flight and high dynamic range scenarios. The event camera, a new vision sensor, holds the potential to enable drones to overcome these challenging scenarios if fused with the image-based SLAM. Unfortunately, the computational demands of event-image fusion SLAM have grown manifold compared with image-based SLAM. Existing research on visual SLAM acceleration cannot achieve real-time operation of event-image fusion SLAM on on-board computing platforms for drones. To fill this gap, we present TrinitySLAM, a high-accuracy, real-time, low-energy consumption event-image fusion SLAM acceleration framework utilizing Xilinx Zynq, an on-board heterogeneous computing platform. The key innovations of TrinitySLAM include a fine-grained computation allocation strategy, several novel hardware–software co-acceleration designs, and an efficient data exchange mechanism. We fully implement TrinitySLAM on the latest Zynq UltraScale+ platform and evaluate its performance on one custom-made drone dataset and four official datasets covering various scenarios. Comprehensive experiments show that TrinitySLAM improves the pose estimation accuracy by 28% with half end-to-end latency and 1.2× energy consumption reduction compared with the most comparable state-of-the-art heterogeneous computing platform acceleration baseline. [ABSTRACT FROM AUTHOR]

This paper investigates the coupling of successive convex optimization guidance with robust structured H ∞ control for the descent and precise landing of Reusable Launch Vehicles (RLVs). More particularly, this Guidance and Control (G&C) system is foreseen to be integrated into a nonlinear six-degree-of-freedom RLV controlled dynamics simulator which covers the aerodynamic and powered descent phase until vertical landing of a first-stage rocket equipped with a thrust vector control system and steerable planar fins. A cost function strategy analysis is performed to find out the most efficient one to be implemented in closed-loop with the robust control system and the vehicle flight mechanics involved. In addition, the controller synthesis via structured H ∞ is thoroughly described. The latter are built at different points of the descent trajectory using Proportional-Integral-Derivative (PID)-like structures with feedback on the attitude angles, rates, and lateral body velocities. The architecture is verified through linear analyses as well as nonlinear cases with the aforementioned simulator, and the G&C approach is validated by comparing the performance and robustness with a baseline system in nominal conditions as well as in the presence of perturbations. The overall results show that the proposed G&C system represents a relevant candidate for realistic descent flight and precise landing phase for reusable launchers. [ABSTRACT FROM AUTHOR]

In the modern aviation field, flight control systems' reliability and safety are paramount. This paper presents a triplex redundant flight control system based on the M1394B bus and designs several key algorithms to enhance system performance. Firstly, a triplex redundant flight control system with a redundant bus structure is constructed based on the characteristics of the M1394B bus. Secondly, a periodic synchronization algorithm with automatic adjustment capabilities is designed to achieve periodic synchronization among the Vehicle Management Computers. An improved voting algorithm based on a sliding window is proposed to enhance the decision-making accuracy and reliability of the control commands output by the flight control system. Additionally, a system reconstruction algorithm is designed to promptly identify and isolate faults, enabling the recovery and reallocation of system resources. Finally, experiments validate the effectiveness of the synchronization algorithm, voting algorithm, and system reconstruction algorithm. The results indicate that the system can effectively address practical engineering challenges and significantly improve reliability and stability. This research provides an essential theoretical foundation and practical reference for the design of future flight control systems for unmanned aerial vehicles and aircraft, holding significant relevance to application. [ABSTRACT FROM AUTHOR]

The objective of the currentwork was to develop an all-electric distributed propulsion and flight control (DPFC) architecture that will have no more than 10-9 catastrophic failures per flight hour (pfh). The DPFC architecture was broken into four system design teams: (i) electrical power and distribution system, (ii) drive and power system, (iii) thermal management system, and (iv) flight control system. System designs were updated and iterated upon, working with reliability and safety analysis teams to develop compliant designs. The firm designs were reflected in a preliminary system safety assessment (PSSA) for initial verification of compliance. Additionally, design and analysis excursions are presented in which aircraft attributes were modified to investigate sensitivities to propulsion type, number of rotors, and control schemes. Excursion results found that all aircraft evaluated likely have paths to comply with the stringent, probabilistic catastrophic failure criteria. However, stability and control models showed large power transients that must be addressed, and PSSA results show that future work is needed in single load path structures, high voltage power storage and distribution, and motor/rotor overspeed protection. [ABSTRACT FROM AUTHOR]

In order to improve the safety of formation flight, a method using wingtip normal blowing is proposed to weaken the influence of the wake generated by the leading aircraft. Numerical simulation methods are employed to analyze the aerodynamic characteristics of the following aircraft affected by the wake of the leading aircraft during different formation positions. Additionally, numerical simulation is conducted on the wingtip normal blowing. The simulation results indicate that during formation flight, the following aircraft experiences significant disturbance torque when transiting laterally through the vortex area of the leading aircraft, which increases the difficulty of rudder manipulation and aircraft control, posing a serious threat to the safety of the following aircraft. After introducing wingtip normal blowing, the wingtip vortex of the following aircraft is weakened, and the lift-to-drag ratio increases with the augmentation of the blowing momentum coefficient. Subsequently, a formation flight control system with wingtip normal blowing is established in Simulink for simulation. The simulation results demonstrate that the proposed wingtip normal blowing mitigates the influence of the wake from the leading aircraft, reducing both the lateral position and velocity deviations of the following aircraft as it traverses the vortex area of the leading aircraft. Concurrently, the aileron deflection of the following aircraft is significantly decreased, thereby improving the safety of the following aircraft. [ABSTRACT FROM AUTHOR]

In this study, we investigated inspection robots that can safely, accurately, and quickly perform hammer sound inspections on tile exterior walls and present new wall hammering inspection systems using two-wheeled multi-copters. Our results yielded the following five advantages. First, the proposed multi-copter can use its wheels not only to move freely on the walls of buildings but also to overcome obstacles on the walls. Therefore, almost any tile wall surface can be inspected for hammering sound. Second, the multi-copter was equipped with both hammer-shaped rods to hit the tile exterior wall and microphones to collect the hammer sound of the tile, while moving quickly on the wall surface. Third, the wall hammering inspection systems can be used safely with a work assistance mechanism using a wire even in densely populated areas, such as urban areas, because the multi-copter has high wind resistance against crosswinds by both pushing against the wall and operating by the wire. Fourth, we presented some automatic control systems that make the multi-copter operation easy during hammering inspections and demonstrated its usefulness through experiments. Fifth, we proposed useful methods for detecting floating tiles based on the hammering results, conducted some outdoor flight experiments on building exterior walls, and demonstrated that floating tiles can be determined with high accuracy. [ABSTRACT FROM AUTHOR]

In Japan today, aging infrastructures such as dams and bridges are becoming an issue. We considered that the use of a hydrofoil multicopter would be an effective method for inspecting and surveying such facilities. When using a hydrofoil multicopter, it is possible to take off from the shore, land on the water, and dive without using a ship. The aircraft developed in this study has the capability to communicate with the land. The aircraft developed in this study is equipped with a communication buoy, and after landing on the water, a communication cable is extended from the buoy to enable underwater maneuvering via wired communication. We report on the development of a precise control system for flight and diving of this aircraft. [ABSTRACT FROM AUTHOR]

The design of a flying vehicle autopilot poses a significant challenge to the controller performance and the effectiveness of the flight control system. This challenge arises from uncertainties and extreme variations in parameters. To enhance the performance of the flight autopilot, the application of adaptive control techniques becomes crucial. Flight control systems extensively employ adaptive control methods due to their capacity to swiftly and automatically adapt to environmental changes and system dynamics, consequently achieving greater speed and accuracy. This paper aims to simulate scenarios where in the flight control system of an flying vehicle approach, with a specific focus on aerodynamic control in the roll, pitch, and yaw channels. The main purpose of the paper is to be able to provide an adaptive flight control system that can perform properly in the face of sharp and sudden changes in parameters. The paper compares the performance of the multiple model adaptive control method, particularly the switching and tuning technique, for the autopilot's roll, pitch, and yaw with other methods such as the reference model adaptive control used in previous research. The results demonstrate the superiority of the multiple adaptive controller utilizing switching and tuning techniques in handling parameter changes, exhibiting faster and more responsive behavior compared to traditional adaptive controllers. [ABSTRACT FROM AUTHOR]

Modeling a high-speed flying vehicle is imperative to ensure the success of vehicle development missions. Moreover, adherence to research protocols mandates a stepwise approach to testing the vehicle model, encompassing simulation trials using software-in-the-loop simulation (SILS), hardware-in-the-loop simulation (HILS), as well as diverse ground and environmental tests prior to flight testing. This study entailed a collaborative effort between MATLAB/Simulink and LabVIEW to seamlessly integrate the model developed in MATLAB/Simulink into LabVIEW for the implementation of model predictive control (MPC) strategy, aimed at trajectory correction (TC) missions for the vehicle. This MPC strategy was directly applied to the onboard flight control system (OBFCS) of the vehicle. Simulation results indicate the successful control of roll and pitch conditions by OBFCS in both SILS and HILS, ensuring the maintenance of flight conditions in accordance with predicted trajectories despite the presence of simulated disturbances. Notably, the simulation demonstrates the independence or absence of interference between each simultaneous MPC control for roll and pitch adjustments. [ABSTRACT FROM AUTHOR]

Flight control system is typical safety critical system, and the reliability of flight control system plays an important role in ensuring the safe operation of aircraft. Traditional reliability analysis methods have a heavy reliance on the experience of analysts, which makes it easy for inconsistencies between reliability models and design models. The fault propagation behavior of the system is thoroughly described by the Architecture Analysis Design Language（AADL）and stochastic Petri nets（SPN）, and a method for model-based reliability analysis is proposed. The nominal model and error model of a lateral fly-by-wire flight control system are constructed using AADL. A method for extracting error propagation information from the AADL model is proposed, and the SPN model described the fault propagation behavior of the system is automatically generated by extracting the information of AADL model. Based on the SPN model, Monte Carlo simulation is used to evaluate the reliability of the lateral fly-by-wire flight control system, compared with the fault tree analysis method, the error is less than 0.018%, which can be neglected in practice. Through the method of this study, the reliability model is automatically generated by the AADL model, which ensures the consistency between the reliability model and the design model and avoids reliance on the experience of designers.

The technology of firefighting unmanned aerial vehicles (UAVs) used in fire scene reconnaissance, water disaster investigation, and rescue missions, has become a crucial asset in responding to fires and emergencies. The key to this technology is to swiftly and effectively improve the operational efficiency of UAVs. Based on the premise of rapid and effective rescue, the stability, tracking, and safety of hexacopter UAVs technology improved the design and optimization innovative design of the storage and transportation convenience structure of unmanned aerial vehicles were carried out. An innovative high-pressure ejection mechanism for fire bombs (delivery items) has been designed. So, a portable, efficient, and fast unmanned aerial vehicle structure was designed. The PX4 flight control system was used to conduct simulation experiments on the newly designed UAV, and it was found that the structural design was reasonable, and the UAV was operated smoothly and accurately during the simulation flight. This structural design and simulation analysis can provide new ideas for the design of new fire-fighting equipment. [ABSTRACT FROM AUTHOR]

Current weather modification rockets with a single operation mode, limited operating height, and fixed and uncontrollable operating time cannot achieve seeding in different layers, stages, and quantities for different cloud systems. Therefore, a temporally controllable weather modification rocket with spatial seeding capability is developed in this study. This new rocket features an electronic fuse-controlled intelligent ignition system, with eight channels of ignition tube outputs. Additionally, carrier wave communication technology is incorporated to set the seeding time for eight sets of ignition tubes. The temporally controllable rocket is capable of initiating seeding within 2–26 s and can conduct operations for layering, arbitrary altitude, and fractional-dose seeding within the altitude range of 500–5500 m (at a launch angle of 70°), which is calculated based on the flight trajectory of rockets. The minimum time interval of the rocket for seeding can be set to 0.1 s, and all 48 catalyst bullets loaded in a single rocket can be launched within 0.8 s. Thus, the rocket can achieve both concentrated and continuous seeding. Consequently, during weather modification operations, the parameters such as altitude, thickness, and operating temperature of target clouds can be obtained through detection, and they can be used to automatically calculate the suitable seeding time, seeding altitude, and seeding dose by the launch control system based on the flight trajectory of rockets to improve the accuracy and scientificity of cloud catalytic operations. Ground tests show that the reliabilities of the electric ignition tube output, new electronic fuse input and output, and electronic fuse output energy all meet the design specification. The temporally controllable spatial seeding rocket can achieve adjustable and controllable seeding times for catalytic bullets, meeting the safety and reliability requirements of rockets. [ABSTRACT FROM AUTHOR]

In order to ensure the flight safety of small unmanned aerial vehicles (UAVs), a deep neural network-fused mathematical modeling approach is put up for reliable flight control of small UAVs. First, engine torque, thrust eccentricity and initial stop angle are taken into full consideration. A six-degree-of-freedom nonlinear model is formulated for small UAVs, concerning both ground taxiing and air flight status. Then, the model was linearized using the principle of small disturbances. The linearized model expressions for both ground taxiing and air flight were provided. In addition, radial basis function neural networks are used for online approximation to address the nonlinearity and uncertainty caused by changes in aircraft aerodynamic parameters. At the same time, to compensate for the external disturbance and the approximation error of the neural network, the system robustness is improved by selecting reasonable design parameters. This helps the whole flight control system obtain better tracking control performance. At last, some simulation experiments are carried out to evaluate the performance of the proposed mathematical modeling framework. The simulation results show that the proposal has stronger convergence ability, smaller prediction error, and better performance. Thus, proper proactivity can be acknowledged. [ABSTRACT FROM AUTHOR]

This work focuses on the three-dimensional integrated guidance and control (IGC) problem for a flight vehicle with a body-aligned strapdown seeker. The strapdown seeker cannot provide the line-of-sight (LOS) angular rate information and causes difficulties in the controller design. Additionally, external disturbance and gain–loss actuator faults also lead to the loss of control performance. To solve these problems, an extended state observer (ESO) is firstly developed to estimate the LOS angular rate by applying the observed body-line-of-sight angles provided by the body-aligned strapdown seeker. Based on backstepping and dynamic surface control techniques, the fault-tolerant IGC is then designed to deal with the gain–loss actuator fault, and adaptive approaches are applied to improve the robustness of the system. Finally, the uniformly ultimately bounded stability of the flight control system is guaranteed via Lyapunov synthesis, and numerical simulations are conducted to verify the effectiveness of the system. [ABSTRACT FROM AUTHOR]

This paper presents a concomitant observer-based multi-level fault-tolerant control (FTC) for near-space vehicles (NSVs) with a new type dissimilar redundant actuation system (NT-DRAS). When NSV flight control system faults occur in NT-DRAS and attitude-corresponding sensors, the NSV hybrid output states, including the concomitant observer usable states and the real system states, are applied to solve the FTC gain by using the linear quadratic regulator (LQR) technique. Furthermore, since NT-DRAS is used in NSVs, a multi-level (actuation system level and flight control level) FTC strategy integrating NT-DRAS channel switching and flight control LQR is proposed for complex and worsening fault cases. The most important finding is that though the proposed strategy is applicable for worsening fault cases in NSVs, systematic and accurate criteria for the process being performed are necessary and can improve the FTC efficiency with minimal FTC resources. Additionally, such criteria can improve the NSV's responsiveness to comprehensive faults, provided that the real-time performance of the fault detection and diagnosis (FDD) scheme can be further optimized. The concomitant observer convergence and the multi-level FTC strategy have been verified by numerical simulations based on the Matlab/Simulink platform. [ABSTRACT FROM AUTHOR]

Model-based systems engineering (MBSE) has been widely used in complex system design. A multi-level progressive architecture design process based on MBSE is proposed by constructing the relationship between function, behavior and structure. Subsequently, guided by the demand for altitude control, an example modeling of the civil aircraft flight control system is conducted. The results indicate the multi-level progressive architecture design of civil aircraft flight control system architecture based on MBSE leverages the advantages of reusable digital models, ensuring a close integration of requirements with functional, logical, and physical architecture. Meanwhile, the method improves the traceability of system design and provides an architectural reference for subsequent domain level model design. [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.)

The article discusses the importance of research in the field of flight dynamics, control, and simulation for advancing aeronautical sciences. It emphasizes the need for a deep understanding of flight dynamics to enhance aircraft performance, safety, and efficiency. The integration of advanced control strategies and the use of simulation play a crucial role in optimizing aircraft behavior and reducing costs. The article also provides an overview of 14 papers published in a special issue of the Aerospace journal, covering various topics such as hybrid electric propulsion, aerodynamic parameter identification, and autonomous flare control. The research presented in the papers contributes to the development of new technologies and innovations in aviation. [Extracted from the article]

Copyright of Journal of Engineering (17264073) is the property of Republic of Iraq Ministry of Higher Education & Scientific Research (MOHESR) 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.)

The AquaFly project is a research collaboration between the Flight Systems and Control (FSC) Laboratory at the University of Toronto Institute for Aerospace Studies (UTIAS) and Queen’s University Biology Department in Canada to offer a fast, reliable, and efficient method of sampling water bodies by using autonomous Unmanned Aerial Systems (UAS). To this end, two UAS are custom-designed and flown for autonomous water sampling, incorporating an octocopter configuration and a ratchet and pawl mechanism for retracting and releasing payloads. This design offers easy-to-use and safe deployment and retraction without continuous motor engagement while eliminating load swinging and the risk of instability by positioning the sampler close to the drone’s center of mass. Furthermore, the system enables autonomous water sampling in remote or inaccessible areas with pre-programmed flight paths and sampling locations, ensuring consistent data collection. The sampling mechanism allows controlled retrieval and release of payloads and further facilitates multi-drone operations to support a variety of sampling missions. This innovative UAS has the potential to revolutionize water sampling and enhance efficiency and safety in environmental monitoring, research, and resource management. Flight tests over several water bodies demonstrate successful simultaneous water sampling missions with two UAS. To the best of the authors’ knowledge, this is the first work to include multiple UAS, each equipped with a payload deployment and retraction mechanism, for autonomous water sampling. [ABSTRACT FROM AUTHOR]

With the rapid development of electronic information technology, the performance of embedded microcontroller is constantly improved, and the software of UAV flight control system is also developing in the direction of multi-function and multi-task. Traditional flight control software is usually written in C language, resulting in its scalability and modular design is limited. In view of the relatively consistent core content of the flight control system software, object-oriented thought and component technology are introduced to carry out modular design and implementation of the core part, improve the scalability and flexibility of the system, and facilitate the rapid addition of new functions. Based on the analysis of the shortcomings of the existing flight control software design, this paper proposes to use C++ language to componentize the driver code, design each submodule in the form of classes, and build the overall software framework according to the dependence relationship of the software level. Through component technology, the design and implementation of the driving component are completed, and the design principle of the driving component is discussed, and the implementation process of the component is explained in detail through the concrete code, and the derivation relationship between the components is analyzed. The subscription-release communication mechanism based on component relationship is designed, the communication interface between modules is standardized, and the independence of each module of flight control software is improved. Based on the driver component, the design and implementation of each task thread are completed, the data interaction and synchronization relationship between the task threads are discussed in detail, and the implementation details and execution process of the task thread are analyzed in combination with specific components. The relationship between the driver component and the task is sorted out, the integration of flight control software is completed, and the expansibility of the driver component and the independence of the task thread are analyzed. Through functional tests, including mission execution efficiency, cycle delay, and the effectiveness of the inertial navigation module, and test flights on the quadrotor platform, the results show that the optimized flight control software has significantly improved in performance and stability. The research and optimization in this paper not only enhance the scalability and modularization design of UAV flight control system, but also provide a valuable reference for the development of flight control software in the future. [ABSTRACT FROM AUTHOR]

The impact of operating voltage on Unmanned Aerial Vehicle (UAV) inspection and control is mainly manifested as electromagnetic interference, where the electric field mainly affects the distribution of space charges and ions, exhibiting adsorption effects on UAVs, and the magnetic field interferes with airborne magnetometers, disrupting the navigation system of UAVs. Under power frequency conditions, the electromagnetic field of alternating current exhibits alternating characteristics, and it only polarizes near the wire to form space charges or ion currents, with little effect at further distances. However, the variation in the magnetic field in one cycle is zero (positively correlated with alternating current), so its impact on UAVs is not particularly significant. Under direct current conditions, a constant current is introduced into the wire, and the electric field polarized around the wire generates a constant property of charge or ion current, resulting in a relatively larger electric field strength and a wider range of influence. At the same time, the constant current generates a constant magnetic field, which is applied to the airborne magnetometer, equivalent to adding a constant interference source, thus having a significant impact on the inspection and control of UAVs. This article uses ANSYS software to conduct electromagnetic field simulation analysis on DC transmission lines of different voltage levels. [ABSTRACT FROM AUTHOR]

The strong airflow beneath a rotary drone generates a wind vortex within the rice canopy; precise control of the wind vortex distance and wind vortex area can improve pesticide utilization efficiency. This paper calculates the flight parameter curve based on the wind vortex target from the wind vortex target parameter control model of the four-rotor plant protection drone, designs a flight control system using a Cube Orange flight controller and a Jetson AGX Xavier onboard computer, and implements flight parameter control using both PID control and fuzzy control algorithms. Experimental results indicate that when using PID control and fuzzy control, the average deviation values of UAV flight altitude and speed are 0.08 m, 0.08 m/s, 0.06 m, and 0.08 m/s, respectively. When using PID control, the average distance and area errors of the target downwind and upwind are 0.17 m and 0.37 m2 and 0.20 m and 0.46 m2, respectively. The corresponding values for fuzzy control are 0.12 m, 0.38 m2, 0.09 m, and 0.31 m2. In the twelve voyage experiments, the target parameter variance using fuzzy control was relatively smaller for eight voyages compared to PID control, which had a smaller variance for four voyages. On the whole, the effect of fuzzy control is superior. The wind vortex control method proposed in this paper can effectively enable precise pesticide spraying by drones. This has significant implications for reducing agricultural production costs and safeguarding the natural environment. [ABSTRACT FROM AUTHOR]

A coaxial helicopter with a maximum take-off weight of 600 kg was converted to an unmanned aerial vehicle. A minimally invasive robotic actuator system was developed, which can be retrofitted onto the copilot seat of the rotorcraft in a short period of time to enable automatic flight. The automatic flight control robot includes electromechanical actuators, which are connected to the cockpit inceptors and control the helicopter. Most of the sensors and avionic components were integrated into the modular robotic system for faster integration into the rotorcraft. The mechanical design of the control system, the development of the robot control software, and the control system architecture are described in this paper. Furthermore, the multi-body simulation of the robotic system and the estimation of the linear low-order actuator models from hover-frame flight test data are discussed. The developed technologies in this study are not specific to a coaxial helicopter and can be applied to the conversion of any crewed flight vehicle with mechanical controls to unmanned or fly-by-wire. This agile development of a full-size flying test-bed can accelerate the testing of advanced flight control laws, as well as advanced air mobility-related functions. [ABSTRACT FROM AUTHOR]

Novel vertical take-off and landing advanced air mobility aircraft which are overactuated and transition between vertical and forward flight modes pose unique challenges for the design of safe and robust full-envelope fly-by-wire flight control systems. This paper presents a methodology for designing and optimizing a control system architecture for a tilt-wing urban air mobility concept. It features the total energy control system algorithm, which is extended to be applicable to hover and transitioning flight and explicit model following inner loops. Control system parameters are optimized using a genetic algorithm optimization scheme, subject to constraints on closed-loop dynamic stability and control response characteristics. Control system performance is demonstrated by simulation of lateral, longitudinal, and directional maneuvering for hover, transition, and forward flight conditions, followed by simulation of departure and arrival transitions while tracking a prescribed flightpath and speed profile in calm and turbulent air. The results demonstrate the effectiveness of the proposed control system architecture and the demonstrated optimization methodology. [ABSTRACT FROM AUTHOR]

This paper investigates a problem of autonomous aerial refueling point parallel rendezvous guidance and flight control subject to exogenous wind field disturbances. The tanker and uncrewed receiver aircraft approach the common refueling rendezvous point according to their respective flight plans, ultimately maintaining a fixed altitude and the same speed. According to the relative position of refueling target point and uncrewed receiver, a new continuous terminal sliding-mode guidance law combined with the prescribed performance control theory is designed under the end angle and speed constraints. Meanwhile, a finite-time observer is proposed to estimate the parameter uncertainties and wake vortex disturbances during refueling rendezvous. Based on the estimated information, a backstepping double-integral sliding-mode controller is developed for the uncrewed receiver flight control system. In the process of the guidance and flight control system, the integral variable is introduced to improve the steady-state performance. The designed guidance and control system can enable the uncrewed receiver to achieve the refueling rendezvous task within finite time and improve the efficiency of uncrewed aerial refueling. Comparative numerical simulations are provided to demonstrate the effectiveness of the proposed guidance and control scheme. [ABSTRACT FROM AUTHOR]

This paper presents an overview on the state of the art in copter drones and their components. It starts by providing an introduction to unmanned aerial vehicles in general, describing their main types, and then shifts its focus mostly to multirotor drones as the most attractive for individual and research use. This paper analyzes various multirotor drone types, their construction, typical areas of implementation, and technology used underneath their construction. Finally, it looks at current challenges and future directions in drone system development, emerging technologies, and future research topics in the area. This paper concludes by highlighting some key challenges that need to be addressed before widespread adoption of drone technologies in everyday life can occur. By summarizing an up-to-date survey on the state of the art in copter drone technology, this paper will provide valuable insights into where this field is heading in terms of progress and innovation. [ABSTRACT FROM AUTHOR]

This work addresses the problem of planning and tracking collision-free trajectories with a laboratory helicopter restrained to three degrees of freedom (DOF) using model predictive control (MPC) with mixed-integer programming (MIP) encoding. We propose a multirate structure that divides the system dynamics into slower and faster modes. The MPC-MIP is then employed to directly control only the slower portion, allowing it to operate in a closed loop in spite of its relatively time-demanding optimizations. Experimental results show that the proposed control structure allows the computation of globally optimal trajectories and controls with the required frequency, resulting in the successful maneuvering of the helicopter toward a target in an obstacle-filled environment. [ABSTRACT FROM AUTHOR]

In this paper, a fixed wing, tail fin controlled gliding munition has been designed. A guidance and control system is developed for the designed model and the glide towards the target is tested in the XPLANE 11 flight simulator environment. The simulation was carried out in two stages, both in the software loop and in the hardware loop. After the munition system is released from the aircraft, it measures the angles that will enable gliding to the target coordinates with the guidance system and starts to control fin movements. It compares the instantaneous position data received from GPS with the target position data. With the LOS method, the pitch angle is found by comparing the heading angle required to reach the target and the instantaneous altitude information with the target altitude information. The angles found are compared with the information from the IMU sensor and the errors are processed in the PID controller. The output of the PID controller is converted as a PWM signal to the tail fins in accordance with the munition dynamics. The system was transferred to XPLANE 11 and scenarios with different initial conditions were tested. [ABSTRACT FROM AUTHOR]

There have been numerous studies on the control of quadcopters. These studies mainly aim to control the flight behavior of quadcopters. To achieve this, researchers have been developing new tools and testing new methods. One of the developed tools is the 3-DOF Hover system, which enables researchers to analyze the flight behaviors of quadcopters, such as roll, pitch, and yaw, even in a physically limited area or only in a computer environment. The control method applied in the control of the 3-DOF Hover system has been determined by the manufacturer as Linear-Quadratic Regulator (LQR). LQR has control parameters that are complex to calculate. This complex calculation process creates an optimization problem. Beyond controlling the 3-DOF Hover system using LQR, this study focuses on calculating the complex control parameters of LQR using optimization algorithms when controlling a dynamic system with LQR. This study includes well-known algorithms such as Genetic Algorithm (GA), Particle Swarm Optimization (PSO), and Simulated Annealing (SA), as well as an innovative approach known Gray Wolf Optimization (GWO). These algorithms were selected due to their proven effectiveness in various studies. Based on the results obtained from these algorithms, a hybrid algorithm incorporating SA and GWO is proposed. The aim of this hybrid algorithm is to combine the advantages of different methods and achieve a more effective and efficient optimization process. The mentioned hybrid algorithm, obtained by combining SA and GWO, is named hSA-GWO. This hSA-GWO is compared with traditional algorithms, and the comparison results show that the proposed hybrid algorithm can be used as an alternative and competitive method for controlling the flight behaviors of quadcopters. [ABSTRACT FROM AUTHOR]

The paper studies the control problem of an uncertain nonlinear system with tracking error constraint, unknown nonlinear functions and unknown control input gains. An asymmetric barrier Lyapunov function composed with predefined time prescribed performance function is constructed to ensure tracking error enters into the predefined asymmetric constraint in a given time. Then radial basis function neural network is adopted to approximate unknown functions. To reduce computation load, minimal-learning parameter technique is applied. Meanwhile, adaptive method is used to solve actuator faults and the unknown control input gains. Moreover, an adaptive neural control strategy is designed in the framework of backstepping method. An adaptive fixed time filter is developed for avoiding the "explosion of complexity" problem, where the convergence speed of the filter error is improved compared with fixed time filter. It is proved that all signals of the closed-loop system are bounded and tracking error is kept in its constraint boundary. At the end, compared numerical simulations and application simulation of a hypersonic vehicle are demonstrated to verify the efficiency of the designed control scheme. • A composed performance function is designed to ensure the tracking error enters into the constraint in a given time. • An adaptive fixed time filter is developed to solve the 'explosion of complexity' problem. • Based on minimal-learning parameter method, neural networks estimate unknown functions, which reduce computation load. • Compared simulation and application simulation of a hypersonic vehicle are given to test the efficiency of the controller. [ABSTRACT FROM AUTHOR]

Purpose: This study aims to discuss the simultaneous longitudinal and lateral flight control of the octorotor, a rotary wing unmanned aerial vehicle (UAV), for the first time under the effect of morphing and to improve autonomous flight performance. Design/methodology/approach: This study aims to design and control the octorotor flight control system with stochastic optimal tuning under morphnig effect. For this purpose, models of different arm lengths of the octorotor were drawn in the Solidworks program. The morphing was carried out by simultaneously lengthening or shortening the arm lengths of the octorotor. The morphing rate was estimated by using simultaneous perturbation stochastic approximation (SPSA). The stochastic gradient descent algorithm, which is frequently used in machine learning, was used to estimate the changing moments of inertia with the change of arm lengths. The proportional integral derivative (PID) controller has been preferred as an octorotor control algorithm because of its simplicity of structure. The PID gains required to control both longitudinal and lateral flight were also estimated with SPSA. Findings: With SPSA, three longitudinal flight PID gains, three lateral flight PID gains and one morphing ratio were estimated. PID gains remained within the limits set for SPSA, giving satisfactory results. In addition, the cost index created was 93% successful. The gradient descent algorithm used for the moment of inertia estimation achieved the optimum result in 1,570 iterations. However, in the simulations made with the obtained data, longitudinal and lateral flight was successfully carried out. Originality/value: Octorotor longitudinal and lateral flight control was performed quickly and effectively with the proposed method. In addition, the desired parameters were obtained with the optimization methods used, and the longitudinal and lateral flight of the octorotor was successfully carried out in the desired trajectory. [ABSTRACT FROM AUTHOR]

Purpose: This study aims to simultaneously and stochastically maximize autonomous flight performance of a variable wing incidence angle having an unmanned aerial vehicle (UAV) and its flight control system (FCS) design. Design/methodology/approach: A small UAV is produced in Iskenderun Technical University Drone Laboratory. Its wing incidence angle is able to change before UAV flight. FCS parameters and wing incidence angle are simultaneously and stochastically designed to maximize autonomous flight performance using an optimization method named simultaneous perturbation stochastic approximation. Obtained results are also benefitted during UAV flight simulations. Findings: Applying simultaneous and stochastic design approach for a UAV having passively morphing wing incidence angle and its flight control system, autonomous flight performance is maximized. Research limitations/implications: Permission of the Directorate General of Civil Aviation in Turkish Republic is necessary for real-time flights. Practical implications: Simultaneous stochastic variable wing incidence angle having UAV and its flight control system design approach is so useful for maximizing UAV autonomous flight performance. Social implications: Simultaneous stochastic variable wing incidence angle having UAV and its flight control system design methodology succeeds confidence, excellent autonomous performance index and practical service interests of UAV users. Originality/value: Creating an innovative method to recover autonomous flight performance of a UAV and generating an innovative procedure carrying out simultaneous stochastic variable wing incidence angle having UAV and its flight control system design idea. [ABSTRACT FROM AUTHOR]

This paper proposes a novel, robust flight control system using adaptive nonlinear sliding mode control for a quadrotor UAV in the presence of parametric uncertainties and aerodynamical disturbances. The proposed control system is based on a proportional derivative sliding surface combined with a non-singular fast terminal sliding mode control to enhance the tracking accuracy and reduce the chattering influence. Moreover, an adaptive mechanism is proposed to approximate the unknown upper limit of external disturbances/uncertainties. The Lyapunov criteria is used to prove the closed-loop stability and calculate the adaptive mechanism. The proposed adaptive PD-NFTSMC (APD-NFTSMC) can cope with the negative influence of modeling uncertainties, external disturbances, and measurement noise, allows null error in the steady state, and solves the chattering problem. Moreover, intensive simulation experiments under various external conditions are carried out to highlight the sovereignty of the developed control approach. Finally, comparisons with some well-known control techniques are performed to show the usefulness, smoothness, and robustness of the proposed APD-NFTSMC strategy. [ABSTRACT FROM AUTHOR]

This paper presents an active fault-tolerant flight control system (FCS) design for the dual-system/hybrid VTOL UAV configuration that maximally exploits its inherent over-actuation abilities to provide higher levels of fault-tolerance and reliability compared to that of the other VTOL UAVs configurations up to this point. The significance of this study is to convert the main drawback of having the vertical rotors as dead weights during the cruise to a key advantage where all the UAV's controls are kept active during all the flight phases raising the UAV's fault-tolerance and maneuverability abilities. The FCS proposed elaborates on a new architecture for the hierarchical automatic control loops using the dynamic inversion control to design the trajectory tracking virtual commands and employing optimal dynamic control allocation to optimally resolve the controls redundancy. Regarding fault-tolerance, the presented FCS handles simultaneous failures and jamming of all the control surfaces during flight where the control commands to the healthy controls (vertical rotors) were automatically adapted such that the trimming and trajectory tracking were successfully maintained till the mission end. Additionally, this paper addresses for the first time the employment of the controls redundancy to enhance the VTOL UAV's maneuverability in fault-free cases, where we present a study of reducing the minimum turn radius of the UAV when the vertical rotors are employed with the control surfaces during the maneuver. A reduction of 38% in the coordinated steady-level minimum turn radius turn at 35 m/sec flight speed was achieved. [ABSTRACT FROM AUTHOR]

Flight control law parameters should be designed to provide a sufficient stability margin for closed-loop aircraft while ensuring command tracking accuracy. The singular perturbation margin (SPM) and generalized gain margin (GGM), which are generalizations of the classical phase margin (PM) and gain margin (GM), respectively, from a linear time-invariant system to a nonlinear time-varying system, can be used to quantitatively characterize the maximum singular perturbation and regular perturbation allowed to maintain system stability. In this paper, the sliding mode flight control structure and the design parameters of the sliding mode control law are first introduced for an oblique wing aircraft (OWA), the SPM-gauge and GGM-gauge are added to this closed-loop aircraft model, and the analytical expressions of the SPM and GGM are derived with respect to the control law parameters. Second, the stability margin design requirements of closed-loop aircraft in flight control system design specifications are converted into limitations on the SPM and GGM to determine the value range of the flight control law parameters. Then, with the goal of reducing the sum of the approaching time and sliding time, the parameter value combination is selected within the control law parameter range that meets the stability margin requirements, thus forming a flight control law design method for OWA during the wing skewing process. Finally, the designed control law parameters are applied to a sample OWA, and the stability margin of closed-loop aircraft during the wing skewing process is verified. [ABSTRACT FROM AUTHOR]

The development of complex equipment has the feature of "main manufacturer-supplier", with a large amount of underlying data and a small amount of top-level data for reliability growth. This paper firstly introduces the basic model of reliability growth. For small samples and uncertain of complex equipment, Bayesian methods and grey system theory have also been investigated. It has been found that existing research cannot describe this distributed collaborative network and the inherent mechanism of reliability growth, and the information data of reliability growth had not been fully utilized. This paper proposes a new model that combines Bayesian and GERT networks. This model is suitable for evaluating the reliability growth process in collaborative development networks, and can make full use of supplier information, historical information, and experimental information to comprehensively evaluate the reliability growth. Through case analysis of the flight control system, it is proven that the new model performs well in comprehensively utilizing various types of information. [ABSTRACT FROM AUTHOR]

Purpose: This research study aims to minimize autonomous flight cost and maximize autonomous flight performance of a slung load carrying rotary wing mini unmanned aerial vehicle (i.e. UAV) by stochastically optimizing autonomous flight control system (AFCS) parameters. For minimizing autonomous flight cost and maximizing autonomous flight performance, a stochastic design approach is benefitted over certain parameters (i.e. gains of longitudinal PID controller of a hierarchical autopilot system) meanwhile lower and upper constraints exist on these design parameters. Design/methodology/approach: A rotary wing mini UAV is produced in drone Laboratory of Iskenderun Technical University. This rotary wing UAV has three blades main rotor, fuselage, landing gear and tail rotor. It is also able to carry slung loads. AFCS variables (i.e. gains of longitudinal PID controller of hierarchical autopilot system) are stochastically optimized to minimize autonomous flight cost capturing rise time, settling time and overshoot during longitudinal flight and to maximize autonomous flight performance. Found outcomes are applied during composing rotary wing mini UAV autonomous flight simulations. Findings: By using stochastic optimization of AFCS for rotary wing mini UAVs carrying slung loads over previously mentioned gains longitudinal PID controller when there are lower and upper constraints on these variables, a high autonomous performance having rotary wing mini UAV is obtained. Research limitations/implications: Approval of Directorate General of Civil Aviation in Republic of Türkiye is essential for real-time rotary wing mini UAV autonomous flights. Practical implications: Stochastic optimization of AFCS for rotary wing mini UAVs carrying slung loads is properly valuable for recovering autonomous flight performance cost of any rotary wing mini UAV. Originality/value: Establishing a novel procedure for improving autonomous flight performance cost of a rotary wing mini UAV carrying slung loads and introducing a new process performing stochastic optimization of AFCS for rotary wing mini UAVs carrying slung loads meanwhile there exists upper and lower bounds on design variables. [ABSTRACT FROM AUTHOR]

Forest fires can develop rapidly and may cause a wide range of hazards. Therefore, aerial firefighting, which has the ability to respond and reach fire fields quickly, is of great significance to the emergency response to and subsequent extinguishing of forest fires. The burning of forest fires generates a lot of heat and smoke, which changes the air flow environment and vision over the region and brings challenges to aerial firefighting. In the present work, aerial forest firefighting simulation was divided into the forest fire spread model, the air flow model and the aircraft flight dynamic and automatic control model. Each model was constructed based on a physical method. An integrated framework was designed to realize the interaction among fire fields, airfields, and aircraft, and is verified. The proposed framework can be used for the emergency response decision of aerial forest fire fighting and subsequent fire-fighting mission planning. [ABSTRACT FROM AUTHOR]

The flight control system is the integrated command center and core component of the UAV system. Aiming at the problems of poor accuracy and low efficiency in the field measurement of flight control system parameters, combined with the configuration of the flight attitude sensor-vertical gyroscope, hardware-in-the-loop simulation technology, and integrated automatic test methods are adopted. The hardware-in-the-loop simulation test platform for the UAV flight control system composed of a low-cost, high-precision miniaturized UAV attitude calibration platform and a control piston deflection angle test device is designed. The main controller replaces the ground master control station of the UAV and directly sends remote control commands to the aircraft. By changing the attitude of the aircraft, and measuring the voltage of the control signal, the feedback signal, and the rudder angle, the comprehensive performance test of the parameters of the flight control system is realized. The performance test of the platform shows that the platform can meet the test requirements of angular position, system voltage, and control piston deflection angle. [ABSTRACT FROM AUTHOR]

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