Flight Control System Design Research Articles

Algorithms for Validation of the Flight Data and Evaluation of the Nonlinearities of the Aerodynamic Coefficients of the Aircraft

He system identification methods make an important part of the flight test data analysis. Due to these methods accurate aircraft parameter estimations can be obtained, which meet the requirements of numerous applications, such as simulators and flight control systems design, aircraft modernization, etc. The aircraft parameter identification, when applied to the processing of the actual flight test data, faces the following principal problems: inevitable differences between a physical object and the adopted mathematical models; an essential incorrectness of the identification problem, since it belongs to the class of the inverted problems of dynamics. The presented techniques are intended to overcome those difficulties, using a system approach. That means that the identification problem is treated as an integral multistage system, designed on certain basic principles. These principles are the following: every stage includes a verification procedure; a decomposition of the general problem into minor subproblems; a systematic search for all the useful information in the relevant disciplines; an important role of a human operator. The article presents an algorithmic and methodological support for monitoring of the correctness of the airborne measurements and estimation of the aerodynamic nonlinearities for identification of the aerodynamic coefficients of the aircraft according to flights test. The paper also presents numerous examples of the practical applications of the considered methods and algorithms in the process of the parameters identification of the mathematical models of certain modern aircraft, using the flight test data. The authors believe that the choice of those examples would be helpful for understanding of the typical problems of the aircraft parameter identification.

Design of Flight Control System Using Gain Schedule Fractional PID Controller

The goal of this paper is to control the trajectory of the flight path of six degree of freedom flying body model using Fractional PID controller (FPID) and Gain Schedule Fractional PID controller. FPID and gain schedule FPID controllers gains with non linear flying body simulation are tuned by Simulink design optimization. Gain Schedule FPID controller is able to compensate for constraints that represent physical limits of actuators in pitch angle. The gain schedule FPID for the six degree of freedom flying body is designed in two phases. The first phase is boost phase where the thrust force is maximized. The second phase is sustain phase where the thrust force is minimized. The results of gain schedule FPID controller are compared with the results of FPID controller.

Design of Flight Control System Using Gain Schedule Fractional PID Controller

Design of Flight Control System Using Gain Schedule Fractional PID Controller

Effect of static shape deformation on aerodynamics and aerothermodynamics of hypersonic inflatable aerodynamic decelerator

The inflatable aerodynamic decelerator (IAD), which allows heavier and larger payloads and offers flexibility in landing site selection at higher altitudes, possesses potential superiority in next generation space transport system. However, due to the flexibilities of material and structure assembly, IAD inevitably experiences surface deformation during atmospheric entry, which in turn alters the flowfield around the vehicle and leads to the variations of aerodynamics and aerothermodynamics. In the current study, the effect of the static shape deformation on the hypersonic aerodynamics and aerothermodynamics of a stacked tori Hypersonic Inflatable Aerodynamic Decelerator (HIAD) is demonstrated and analyzed in detail by solving compressible Navier-Stokes equations with Menter's shear stress transport (SST) turbulence model. The deformed shape is obtained by structural modeling in the presence of maximum aerodynamic pressure during entry. The numerical results show that the undulating shape deformation makes significant difference to flow structure. In particular, the more curved outboard forebody surface results in local flow separations and reattachments in valleys, which consequently yields remarkable fluctuations of surface conditions with pressure rising in valleys yet dropping on crests while shear stress and heat flux falling in valleys yet rising on crests. Accordingly, compared with the initial (undeformed) shape, the corresponding differences of surface conditions get more striking outboard, with maximum augmentations of 379 pa, 2224 pa, and 19.0W/cm2, i.e., 9.8%, 305.9%, and 101.6% for the pressure, shear stress and heat flux respectively. Moreover, it is found that, with the increase of angle of attack, the aerodynamic characters and surface heating vary and the aeroheating disparities are evident between the deformed and initial shape. For the deformable HIAD model investigated in this study, the more intense surface conditions and changed flight aerodynamics are revealed, which is critical for the selection of structure material and design of flight control system.

バックステッピング法と最適サーボ系の組み合わせによるシングルロータヘリの高速旋回飛行制御系の設計とその実験的検証

Aiming at the realization of agile autonomous flight for the extension of the flying range of single-rotor helicopters in the limited flight time, the flight control system design and the control performance verification for the high-speed turn flight are conducted. To achieve the reference following control for large attitude angles and the rotor thrust, the control system is composed of the main-rotor thrust and body torque centralized controller based on the MIMO (Multiple-Input, Multiple-Output) helicopter model including the blade motion dynamics, rotor speed proportional-integral controller, the attitude controller designed via the backstepping method based on the quaternion attitude model, and the guidance controller based on translational motion model. Simulation results show that the hovering and 10-m/s straight cruise flight followed by over 45-degree bank turn is feasible by using the proposed control system. In addition, flight tests have been conducted using the experimental small unmanned electric helicopter equipped with the flight control computer, and the high speed turn agile autonomous flight was successful in the real environment.

Survey on Flight Control Technology for Large-Scale Helicopter

A literature review of flight control technology is presented for large-scale helicopter. Challenges of large-scale helicopter flight control system (FCS) design are illustrated. Following this, various flight control methodologies are described with respect to their engineering implementation and theoretical developments, whose advantages and disadvantages are also analyzed. Then, the challenging research issues on flight control technology are identified, and future directions are highlighted.

Design of the Nonlinear Structure Adaptive Model Inversion Flight Control System

[abstFig src='/00280006/17.jpg' width='300' text='Design of the nonlinear SAMI flight control system' ] The advanced aircraft has those technological characteristics: stealth, supersonic cruise, super maneuverability, multi-target attack, multi-role, high load, long range cruise, integrated avionics, short takeoffs and vertical landings. These characteristics need multi-subsystem integration in the advanced aircraft. In order to meet the need of advance aircraft, the nonlinear control system is necessary. In this paper, the nonlinear structure adaptive model inversion system was first used for the nonlinear control problem of super-maneuver aircraft. The sufficient and necessary condition of the control law was analyzed. Finally the results of the height-angle maneuver simulation show that the designed control system has good performance.

Research on the Unmanned Aerial Vehicle Adaptive Control System based on Fuzzy Control and Chaos Mechanics

In this paper, we conduct research on the unmanned aerial vehicle adaptive control system based on fuzzy control and chaos mechanics. Four rotor aircraft is a kind of nonlinear systems with underactuated, strong coupling characteristic. Although in existing research, through the design of the control algorithm effectively inhibits both for flight control effect, but not fundamentally eliminate the effect of aircraft. Dynamic model of unmanned helicopter flight control system design is very approximate, need to gradually improve the modeling accuracy, so as to get the exact autonomous flight control, so you need to practice constantly required to modeling in the flight information, so the unmanned helicopter flight control system to have the ability to retrieve information modeling. This paper proposes the new idea on the issues that will be meaningful.

Tuning of controller for an aircraft flight control system based on particle swarm optimization

Purpose This study aims to present a method for the conceptual design and simulation of an aircraft flight control system. Design/methodology/approach The design methodology is based on particle swarm optimization (PSO). PSO can be used to improve the performance of conventional controllers. The aim of the present study is threefold. First, it attempts to detect and isolate faults in an aircraft model. Second, it is to design a proportional (P) controller, a proportional derivative (PD) controller, a proportional-integral (PI) controller and a fuzzy controller for an aircraft model. Third, it is to design a PD controller for an aircraft using a PSO algorithm. Findings Conventional controllers, an intelligent controller and a PD controller-based PSO were investigated for flight control. It was seen that the P controller, the PI controller and the PD controller-based PSO caused overshoot. These overshoots were 18.5, 87.7 and 2.6 per cent, respectively. Overshoot was not seen using the PD controller or fuzzy controller. Steady state errors were almost zero for all controllers. The PD controller had the best settling time. The fuzzy controller was second best. The PD controller-based PSO was the third best, but the result was close to the others. Originality/value This study shows the implementation of the present algorithm for a specified space mission and also for study regarding variation of performance parameters. This study shows fault detection and isolation procedures and also controller gain choice for a flight control system. A comparison between conventional controllers and PD-based PSO controllers is presented. In this study, sensor fault detection and isolation are carried out, and, also, root locus, time domain analysis and Routh–Hurwitz methods are used to find the conventional controller gains which differ from other studies. A fuzzy controller is created by the trial and error method. Integral of squared time multiplied by squared error is used as a performance function type in PSO.

AN H∞ DESIGN METHOD FOR THE PITCH ATTITUDE HOLD AUTOPILOT OF A FLYING UAV

The paper presents an H∞ loop shaping method for the design of the automatic flight control system of an unmanned air vehicle's (UAV's) longitudinal dynamics. The design objectives include robust stabilization with respect to modeling uncertainties, pitch angle tracking of an ideal model and reduced sensitivity with respect to low frequency measurement errors. The design is illustrate by a case study for a flying wing UAV. The applications based on Unmanned Air Vehicle's (UAV's) received much attention over the last decades. Among them, the flying wing configuration of the UAVs was intensively analyzed due to their advantages concerning the reduced drag and the fuel consumption. A disadvantage of this configuration with respect to the classical one is its instability which fact requires an automatic control system able to stabilize the aircraft and to provide acceptable maneuverability qualities. The design problem of such a system has been previously considered in the control literature (see e.g. (1), (2), (3)). The proposed design methods include both classical techniques based for instance on the well- known PID (Proportional Integral Derivative) control laws, optimal approaches based on the linear quadratic problem, the H∞ norm minimization and nonlinear methods including nonlinear inversion, etc ((4), (5), (6), (7), (8)). The methods mentioned above have advantages and drawbacks; the main difficulty in choosing the control design method is to ensure a trade-off between the complexity of the automatic flight control system and its performances. The actual applications require a wide spectrum of performances including not only the stabilization of the aircraft but also robustness with respect to modeling uncertainties and flying conditions changes, time response performances and reduced sensitivity to disturbances and measurement errors. The aim of this paper is to present a design methodology for the Pitch Attitude Hold (PAH) autopilot of a flying wing configuration. The design procedure is illustrated and investigated for the Hirrus UAV designed and manufactured by a private Romanian company in collaboration with academics from Faculty of Aerospace Engineering of University Politehnica of Bucharest. The design approach is based on a modified version of the so-called two degrees of freedom (2 DOF) H∞ loop-shaping able to accomplish simultaneously several objectives as robustness stability, model tracking and disturbances attenuation requirements. The original 2 DOF H∞ method presented in (9) have been previously used for the autopilot design of the aircraft longitudinal and lateral SCIENTIFIC RESEARCH AND EDUCATION IN THE AIR FORCE-AFASES 2016

An H<sub>∞</sub> Design Approach for the Heading Hold Autopilot of a Flying Wing UAV

The paper presents a modified loop shaping approach for the design of the automatic flight control system of an unmanned air vehicle’s lateral-directional dynamics. The design objectives include robust stabilization with respect to modeling uncertainties, heading angle tracking of an ideal model, coordinated turn and reduced sensitivity with respect to low frequency measurement errors. Based on the resulting full order controller one determined a reduced order controller. The numerical results show that both the full order and the reduced order autopilots accomplish the design objectives.

Design of a tolerant flight control system in response to multiple actuator control signal faults induced by cosmic rays

Due to continued miniaturization, semiconductor-based components used in high-performance digital microelectronics are becoming increasingly sensitive to cosmic rays and solar particle events. In the context of high-altitude flight control systems based on fly-by-wire techniques, this may produce sensor noise or affect actuator control signals. Although the consequences so far have been simply reductions in aircraft performance, catastrophic scenarios may be envisioned. In this article, we propose a novel architecture for a fault-tolerant flight control system able to detect and compensate for cosmic ray-induced multiple-bit upsets that affect actuator control signals in modern fly-by-wire avionics systems while assuming that the actuator itself remains healthy. A fault detection and diagnosis procedure was designed using a geometric approach combined with an extended multiple-model adaptive estimation technique. This procedure is able to process multiple faulty actuator-control signals and identify their parameters. The parameters thus obtained are then used with a reconfigurable sliding-mode control to compensate for such errors by mobilizing the remaining actuators’ healthy control signals. Lyapunov stability theory is used to analyze the closed-loop system stability. Simulation results using Matlab /Simulink showed the effectiveness of the proposed approach in the case of a system challenged with double faults.

Impacts of safety on the design of light remotely-piloted helicopter flight control systems

This paper deals with the architecture definition and the safety assessment of flight control systems for light remotely-piloted helicopters for civil applications. The methods and tools to be used for these activities are standardised for conventional piloted aircraft, while they are currently a matter of discussion in case of light remotely-piloted systems flying into unsegregated airspaces. Certification concerns are particularly problematic for aerial systems weighing from 20 to 150kgf, since the airworthiness permission is granted by national authorities. The lack of specific requirements actually requires to analyse both the existing standards for military applications and the certification guidelines for civil systems, up to derive the adequate safety objectives. In this work, after a survey on applicable certification documents for the safety objectives definition, the most relevant functional failures of a light remotely-piloted helicopter are identified and analysed via Functional Hazard Assessment. Different architectures are then compared by means of Fault-Tree Analysis, highlighting the contributions to the safety level of the main elements of the flight control system (control computers, servoactuators, antenna) and providing basic guidelines on the required redundancy level.

On one approach of observer-based flight control system design

On one approach of observer-based flight control system design

Observer-based flight control system design under linear matrix inequalities approach

Observer-based flight control system design under linear matrix inequalities approach

A structured [formula omitted]-based optimization approach for integrated plant and self-scheduled flight control system design

This paper presents a new procedure for the integrated plant and flight control system design in the presence of parametric uncertainties. The proposed approach is based on a design procedure consisting in casting gain-scheduling and robustness requirements into the framework of structured H∞ design. As both controller architecture and gain-scheduling structure are defined a priori, the scheduled gains can be tuned by the newly available MATLAB-based tool systune to minimize H∞ constraints related to performance requirements over both operating and uncertain domains. It is shown that this technique allows for optimization and adjustment of plant's physical parameters while satisfying closed-loop performance requirements. The proposed procedure is applied to the F-16 Fighting Falcon flight control system design under mass and center of gravity uncertainties by adjusting the location of an additional accelerometer to optimize the closed-loop performance while minimizing the horizontal tail area. Numerical studies are carried out to evaluate the effectiveness of the proposed approach.

Condition-based flight control for helicopters: An extension to condition-based maintenance

A new approach to flight control system design is presented which takes into account the operational usage (condition) of the helicopter. The flight control system gains are scheduled as a function of the condition of the helicopter, provided by the health and usage monitoring system. As a result, the utilization rate of the helicopter can be increased and maintenance costs can be reduced. The approach is demonstrated for vertical axis control of the Bell 412 helicopter operating at low speed. The control system is designed to provide a height rate response type combined with torque limit protection. For low to high levels of aggressiveness, level 1 handling qualities are achieved in piloted simulation of a bob-up maneuver compared to level 2 for the original aircraft without a vertical axis control law. The control system successfully prevents the torque from exceeding the limit. Two conditions are defined, based on which the control system gains are scheduled. The first condition is the number of transmission limit exceedances and the second is the number of rapid torque excursions. Four different control modes are possible due to these conditions. For implementation of this specific control law on a real-life aircraft, emergency situations must be considered in which the pilot needs to exceed the torque limit of the aircraft intentionally to ensure safety. This can be achieved through a tactile cueing system.

Design of Flight Control System Based on Adaptive Backstepping Method for a Space Transportation System

Design of Flight Control System Based on Adaptive Backstepping Method for a Space Transportation System

Input-Shaping and Model-Following Control of a Helicopter Carrying a Suspended Load

When a heavy load is suspended from a helicopter, excessive load swing degrades helicopter control and can result in obstacle collisions. This work examines the benefits of combining input-shaping and model-following control to improve performance when carrying a suspended load. Model-following control architectures are used in modern helicopter flight control systems to make the helicopter respond like a prescribed model. On its own, model-following control is ineffective when carrying a suspended load because excessive load swing degrades tracking of the prescribed model dynamics and thus control of the helicopter. Therefore, reducing load swing improves tracking of the prescribed model and increases safety and productivity. Input shaping is a control technique that has been used to control vibration of flexible machines, such as robots, satellites, and cranes. By combining input shaping with model-following control, helicopter payload swing is reduced and tracking of the prescribed model is improved. The design of an attitude-command flight control system that combines input-shaping and model-following control is illustrated using dynamic models of a Sikorsky S-61 helicopter. Simulation results are shown for lateral and longitudinal repositioning movements. These results show that applying input shaping to simulated pilot commands greatly improves helicopter performance when carrying a suspended load.

Adaptive Flight-Control Design Using Neural-Network-Aided Optimal Nonlinear Dynamic Inversion

A neural-network-aided nonlinear dynamic inversion-based hybrid technique of model reference adaptive control flight-control system design is presented in this paper. Here, the gains of the nonlinear dynamic inversion-based flight-control system are dynamically selected in such a manner that the resulting controller mimics a single network, adaptive control, optimal nonlinear controller for state regulation. Traditional model reference adaptive control methods use a linearized reference model, and the presented control design method employs a nonlinear reference model to compute the nonlinear dynamic inversion gains. This innovation of designing the gain elements after synthesizing the single network adaptive controller maintains the advantages that an optimal controller offers, yet it retains a simple closed-form control expression in state feedback form, which can easily be modified for tracking problems without demanding any a priori knowledge of the reference signals. The strength of the technique is demonstrated by considering the longitudinal motion of a nonlinear aircraft system. An extended single network adaptive control/nonlinear dynamic inversion adaptive control design architecture is also presented, which adapts online to three failure conditions, namely, a thrust failure, an elevator failure, and an inaccuracy in the estimation of . Simulation results demonstrate that the presented adaptive flight controller generates a near-optimal response when compared to a traditional nonlinear dynamic inversion controller.