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Adaptive flight control system for flight safety improvement in reentry and other high-velocity vehicles

ACTA ASTRONAUTICA 2023年 204卷 900-911页

作者： Efremov, A. V. Efremov, E. V. Tiaglik, M. S. Irgaleev, I. Kh Shcherbakov, A. I. Mbikayi, Z. Moscow Inst Aviat Technol Moscow Russia Tech Univ Munich Munich Germany

An adaptive controller and a nonlinear rate limiter are presented in order to improve the flight safety of manually controlled reentry vehicles and supersonic transport (SST) aircraft. The adaptive controller is based on the principle of dynamic inversion. The calculation of the controller parameters corresponding to the aerodynamic coefficients of the vehicle was carried out by a proposed online identification procedure. The effectiveness of the adaptive controller was tested through a set of experiments performed using a ground-based simulator. The obtained results were used to estimate the probability of an accident due to pilot error for the basic Space Shuttle flight control system and for its improved version with the adaptive controller. It was found that using the adaptive controller causes high required rates of elevon deflection. Because of the actuator's rate limiting, this is a source of pilot-induced oscillations. To counteract these oscillations in the pilot-vehicle closed loop system, an adaptive nonlinear rate limiter is introduced. The effectiveness of using the adaptive prefilter for improving the flight safety and flying qualities of the Space Shuttle reentry vehicle and an SST was demonstrated in experiments involving a ground-based simulator.

关键词： flight safety Adaptive controller Nonlinear prefilter flight control system Re-entry vehicle dynamics Supersonic transport Nonlinear inverse dynamics Pilot-induced oscillation Pilot-vehicle system

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Interval Type-2 Fuzzy Cognitive Map-Based flight control system for Quadcopters

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INTERNATIONAL JOURNAL OF FUZZY systemS 2020年第8期22卷 2504-2520页

作者： Amirkhani, Abdollah Shirzadeh, Masoud Kumbasar, Tufan Iran Univ Sci & Technol Sch Automot Engn Tehran *** Iran Amirkabir Univ Technol Dept Elect Engn Tehran 158754413 Iran Istanbul Tech Univ Fac Elect & Elect Engn Dept Control & Automat Engn Istanbul Turkey

In this paper, we propose a novel Interval Type-2 (IT2) Fuzzy Cognitive Map (FCM)-based flight control system to solve the altitude, attitude and position control problems of quadcopters. The proposed IT2-FCM encompasses all concepts related to drone for a satisfactory path-tracking and stabilizing control performance. The degree of mutual influences of the concepts is designed with opinions of three experts that take account the dynamics of drone and rules governing proportional integral derivative (PID) controllers. To model the inter-uncertainty of the experts' opinions, IT2 fuzzy logic systems are utilized as they are powerful tools to model high level of uncertainties. Thus, the proposed IT2-FCM has a qualitative representation as it merges the advantages of IT2 fuzzy logic systems and FCMs. We present comparative simulations results in presence of uncertainties where the superiority of the proposed IT2-FCM-based flight control system is shown in comparison with its type-1 fuzzy counterpart.

关键词： Interval type-2 fuzzy sets Quadcopters Fuzzy Cognitive Map flight control system

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New methodology for flight control system sizing and hinge moment estimation

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PROCEEDINGS OF THE INSTITUTION OF MECHANICAL ENGINEERS PART G-JOURNAL OF AEROSPACE ENGINEERING 2022年第12期236卷 2375-2390页

作者： de la Hoz, Carlos Cabaleiro Fioriti, Marco Politecn Torino Mech & Aerosp Engn Dept Cso Duca Abruzzi 24 I-10100 Turin Italy

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.

关键词： flight control system hinge moment more electric aircraft

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Design of sliding mode flight control system for a flexible aircraft

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PROCEEDINGS OF THE INSTITUTION OF MECHANICAL ENGINEERS PART G-JOURNAL OF AEROSPACE ENGINEERING 2022年第9期236卷 1851-1860页

作者： Mohamed, Majeed Gopakumar, Madhavan CSIR Natl Aerosp Labs Bangalore 560017 Karnataka India Coll Engn Trivandrum Thiruvananthapuram Kerala India

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.

关键词： Flexible aircraft flight control system flight dynamics sliding mode control state estimator

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Integrated rotor-flight control system optimization with aeroelastic and handling qualities constraints

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JOURNAL OF GUIDANCE control AND DYNAMICS 1997年第2期20卷 217-224页

作者： Sahasrabudhe, V Celi, R Tits, AL UNIV MARYLAND DEPT ELECT ENGNSYST RES INSTCOLLEGE PKMD 20742

The results are presented of a study concerning the integrated optimization of rotor and flight control systems with simultaneous aeroelastic and handling qualities constraints. A representative subset of the ADS-33C handling qualities specification is reformulated into inequality constraint form, A simplified design model is used, consisting of three rotor design variables and five control system design variables for each Eight condition. It is not possible to satisfy the entire set of aeroelastic and handling qualities constraints using just the three rotor design parameters of this study, If the design is optimized for just one flight condition, optimizing simultaneously the rotor and the Eight control system leads to substantially better designs than by sequentially optimizing the rotor first and the flight control system next, The designs obtained using the two optimization strategies are quite different, confirming that rotor and flight control systems are strongly coupled.

关键词： Handling Qualities flight control system control Optimization Yaw Airframes Helicopters Dutch Roll control system Design Pilot Rotor Dynamics

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A Modified NSGA-II for Solving control Allocation Optimization Problem in Lateral flight control system for Large Aircraft

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IEEE ACCESS 2019年 7卷 17696-17704页

作者： Bian, Qi Nener, Brett Wang, Xinmin Northwestern Polytech Univ Sch Automat Xian 710129 Shaanxi Peoples R China Univ Western Australia Dept Elect Elect & Comp Engn Perth WA 6009 Australia

Coordinated aileron and rudder control is crucial to the lateral control stability augmentation of an aircraft. In this paper, a modified non-dominated sorting genetic algorithm II is proposed to not only optimize the control allocation between the aileron and rudder channels on different flying quality levels but also explore the relationships between the optimum solutions and the state variables of the aircraft. In doing so, a digital, nets-based stratification method is used to initialize the search chromosomes more evenly. To improve the search efficiency of the algorithm, crowding-distance-based interpolation and elimination strategies are developed to approach the optimum Pareto frontier as close as possible. Moreover, a dynamic depth search method is proposed to balance between the global and local explorations. Finally, the control allocation relationships between the aileron and rudder channels on different flying quality levels are illustrated. The comparative simulations on a six-degree-of-freedom Boeing 747 model are carried out to verify the feasibility of the proposed algorithm.

关键词： MNSGA-II control allocation optimization flight control system

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Integrated flight Path Planning system and flight control system for Unmanned Helicopters

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SENSORS 2011年第8期11卷 7502-7529页

作者： Jan, Shau-Shiun Lin, Yu-Hsiang Natl Cheng Kung Univ Dept Aeronaut & Astronaut Tainan 70101 Taiwan

This paper focuses on the design of an integrated navigation and guidance system for unmanned helicopters. The integrated navigation system comprises two systems: the flight Path Planning system (FPPS) and the flight control system (FCS). The FPPS finds the shortest flight path by the A-Star (A*) algorithm in an adaptive manner for different flight conditions, and the FPPS can add a forbidden zone to stop the unmanned helicopter from crossing over into dangerous areas. In this paper, the FPPS computation time is reduced by the multi-resolution scheme, and the flight path quality is improved by the path smoothing methods. Meanwhile, the FCS includes the fuzzy inference systems (FISs) based on the fuzzy logic. By using expert knowledge and experience to train the FIS, the controller can operate the unmanned helicopter without dynamic models. The integrated system of the FPPS and the FCS is aimed at providing navigation and guidance to the mission destination and it is implemented by coupling the flight simulation software, X-Plane, and the computing software, MATLAB. Simulations are performed and shown in real time three-dimensional animations. Finally, the integrated system is demonstrated to work successfully in controlling the unmanned helicopter to operate in various terrains of a digital elevation model (DEM).

关键词： unmanned helicopter flight path planning system flight control system digital elevation model

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Sequential loop closure in design of a robust rotorcraft flight control system

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JOURNAL OF GUIDANCE control AND DYNAMICS 1997年第6期20卷 1235-1240页

作者： Gorder, PJ Hess, RA UNIV CALIF DAVIS DEPT MECH & AEROSP ENGNDAVISCA 95616

Rotorcraft flight control systems present many significant design challenges. First, large variations in the response characteristics of rotorcraft result from the wide range of airspeeds of typical operation (hover to over 100 kn). Second, the assumption of vehicle rigidity often employed in the design of fixed-wing flight control systems is rarely justified in rotorcraft where rotor degrees of freedom can have a significant impact on the system performance and stability. A methodology is proposed for the design of robust rotorcraft flight control systems utilizing quantitative feedback theory (QFT). QFT is a technique that accounts for variability in the dynamic response of the controlled element in the design of robust control systems. It was developed to address a multiple-input/single-output (MISO) design problem, but the extension of this technique to address multiple-input/multiple-output (MIMO) systems was employed to address the Eight control system (FCS) design of a UH-60 Black Hawk Helicopter. This was accomplished by constructing a set of MISO systems mathematically equivalent to the MIMO system. QFT was applied to each member of the set independently, Inherent conservatism in this design technique leads to limitations in its utility. A second approach utilizing the sequential closure of control loops proved an improved method for designing of robust MIMO FCSs. An analysis of the two FCS design methodologies is presented.

关键词： Rotorcrafts Multiple Input Multiple Output flight control system Closed Loop Transfer Function Dynamic Response UH 60 Black Hawk Helicopter Airspeed Yaw Frequency Domain Pilot

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Decoupling the flight control system of a supersonic vehicle

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AEROSPACE SCIENCE AND TECHNOLOGY 2007年第7-8期11卷 553-562页

作者： Tsay, Tain-Sou Natl Formosa Univ Dept Aeronaut Engn Huwei 63208 Yunlin Taiwan

In this paper, a nonlinear decoupling technique is proposed for analyzes and designs of a serious aerodynamic coupled missile flight control system. Decoupling is obtained from multiplications of two measurable accelerations with three controllable output commands of the autopilot. They give exact signs and magnitudes proportional to aerodynamic decoupling. No matrix inversion of the system dynamic for decoupling is needed. From frequency responses of the rolling channel, it will be seen that the decoupling and robustness for copping with aerodynamic couplings can be obtained simultaneously. Furthermore, the decoupling characteristics are kept almost unchanged for large system parameter variations and uncertainties. The 5-DOF simulations give same conclusions. (c) 2007 Elsevier Masson SAS. All rights reserved.

关键词： aerodynamic decoupling flight control system nonlinear multivariable system

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APPLICATION OF FEEDBACK LINEARIZATION METHOD IN A DIGITAL RESTRUCTURABLE flight control-system

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JOURNAL OF GUIDANCE control AND DYNAMICS 1993年第1期16卷 111-117页

作者： OCHI, Y National Defense Academy Yokosuka Kanagawa 239 Japan

The proposed control system is applied to aircraft failures occurring with the control effectors or airframe. The failures are identified as parameter changes in the six-degree-of-freedom nonlinear equations of motion by the recursive least-squares algorithm. The control parameters are updated using the latest estimated system parameters. In order to allow one to use a digital computer in implemetations, a discrete-time servo controller is designed for the actuators of the control surfaces and engine. The reference inputs are given by the continuous-time control law that is derived without considering actuator or engine dynamics. The frequency-dependent optimal regulator is employed in the design of the servo controller to eliminate input fluctuation. Performance of the control system is demonstrated through computer simulation using the nonlinear model of an aircraft that has half its right wing broken off.

关键词： flight control system Feedback Linearization Actuators control Surfaces Single Input Single Output system Mathematical Models Aircraft Configurations Six Degree of Freedom Digital Computers Longitudinal control

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