Development Of Flight Control System Research Articles

基于集成数字化的月地高速再入返回飞行器飞控支持系统

Circumlunar return and reentry spacecraft accurately landed within the predetermined area in SiZiWangQi, Inner Mongolia on Nov. 1, 2014, which marks the successful completion of the reentry test in the third phase of ChangE lunar exploring project, and indicates that the key technology of deep space reentry return, as the speed of the spacecraft is close to the second cosmic velocity, is fully mastered by China. It ensures the smooth implementation of ChangE-5 program and lays a solid foundation of the lunar exploring project. During the flight control of circumlunar return and reentry spacecraft, the integrated digital flight control support system (IDFCS) is played a key role in ether flight control strategy development, simulation & validation and design optimization. In this paper, the principle and architecture design of the IDFCS for Spacecraft is proposed, and the key points of the system, such as acceleration of processor simulation, virtual chip modeling, automatic test script, general user interface, data dynamic interpretation, are analyzed. Finally, the practical effect of the IDFCS for circumlunar return and reentry spacecraft is present. The actual in-orbit validation shows that some of the satellite flight control problems, such as the complexity of constraints, the difficulty of strategy determination, are solved by the present system, and rich experience in both theory and engineering is accumulated for the subsequent development of the satellite flight control system.

유도형 활공 탄약 비행제어시스템 개발 Part II : 유도 및 제어

유도형 활공 탄약 비행제어시스템 개발 Part II : 유도 및 제어

유도형 활공 탄약 비행제어시스템 개발 Part I : 운용 개념 및 항법

유도형 활공 탄약 비행제어시스템 개발 Part I : 운용 개념 및 항법

Advanced Flight Control System Failure States Airworthiness Requirements and Verification

For advanced airplane, the Safety of Flight tends to be dependent on complex flight control system. The result is a greatly increased emphasis on flight control system failure effects. FAR25 airworthiness standards are based on, and incorporate, the techniques of the fail-safe design concept. Critical FAR25 airworthiness clauses and Special Conditions about flight control system failures are analyzed. Airworthiness requirements and methods of compliance are studied. Key works in flight control system development are summarized. System safety assessments and failure state verification test are conducted for showing the compliance with airworthiness requirements.

Investigation of Small Scale Helicopter Servo-Actuator Performance Using Frequency Sampling Filters

This paper addresses the performance of electric servo-actuators as part of a helicopter control linkage system. The development of a helicopter flight control system initially requires a non-linear model capable of accurately predicting the aircraft dynamics. The first step of this non-linear model represents the servo-actuator system, which transmits the pilots control inputs to the main rotor system. A series of system identification experiments were performed on a cyclic system servo-actuator under different operating conditions, and the servo-actuator step responses were estimated using a Frequency Sampling Filter. A performance comparison is made for a zero load bench test and the aircraft load condition in order to determine the most accurate representation of the servo-actuator dynamics for helicopter flight applications.

Integrated flight control system development using CEASIOM

Flight control system design typically involves hardware and software development, requiring the definition of both continuous and discrete design parameters. The control surface topology is an example of a discrete optimization problem, whereas the tuning of control system gains is an example of a continuous one. Contemporary design practice classically involves decoupling these two problems, however, they are intrinsically linked; for example, the maximum performance attainable by the control system software is determined by the control system topology employed. The cascaded effect is that no definitive sequential roadmap of development can be established, resulting in the requirement for an iterative framework using dynamic programming or optimization methods. This paper presents a framework for aircraft design together with the Flight Control System Designer Toolkit, and the Computerised Environment for Aircraft Synthesis and Integrated Optimisation Methods. The software platform was developed for the European Framework 6 Program Simulating Stability and Control and integrates both the reliability-driven hardware analysis and the performance-driven software design. A case study is presented, based on the NASA Boeing 747–100 model, in order to demonstrate the potential impact of introducing control design early into the design process and to show the intrinsic nature of the hardware–software FCS problem.

L1 adaptive control for safety-critical systems

This article presents the development of L1 adaptive-control theory and its application to safety critical flight control system (FCS) development. Several architectures of the theory and benchmark examples are analyzed. The key feature of L1 adaptive-control architectures is the decoupling of estimation and control, which enables the use of arbitrarily fast estimation rates without sacrificing robustness. Rohrs's example and the two-cart system are used as benchmark problems for illustration. NASA's flight tests on subscale commercial jet verify the theoretical claims in a set of safety-critical test flights.

Quadrotor Autonomous Flight Control System Design

An autonomous quadrotor is an aerial helicopter with four horizontal rotors designed in a square configuration capable of locating lost or jeopardized victims, gathering military intelligence, and surveillance. Controlling of unmanned quadrotor is an interesting problems for control researchers. The literatures as assume a linear model and control it about operating point using the RPM , thrust, and torque In this paper, the proposed controllers are utilized on a full non linear model starting from voltages input able to determine its own attitude through an onboard sensor modeling to have a desired trajectory. The major goal of this ongoing research is the development and implementation of an autonomous flight control system for a quadrotor helicopter. Controlling a Vertical Taking-Off and Landing (VTOL) flying vehicle is basically dealing with highly unstable dynamics and strong axes coupling. The simulation results showed a satisfactorily performance of the proposed controllers adopted on the quadrotor in face off the presence of coupling and nonlinearties.Vertical flight, hover, landing and horizontal flight are among the considered flight features in a predetermined trajectory.

항공기의 수평속도에 대한 비행영역 보호 시스템 설계

Recently developed aircrafts use Fly-By-Wire(FBW) or Fly-By-Light(FBL) system. These systems have some merits; they can perform very complicated missions, they can expand the flight region and improve the reliability of the aircrafts. With the development of flight control systems that use FBW technique, flight envelope protection concept is introduced to guarantee reliability of the aircraft and improve the efficiency of mission achievement. In this study, flight envelope protection system is designed using a dynamic trim algorithm, a peak response estimation, and a gain scheduling technique. The performance of these methods are compared by performing numerical simulation.

Development and Conversion Flight Test of a Small Tiltrotor Unmanned Aerial Vehicle

T HE tiltrotor configuration was selected as the unmanned aerial vehicle (UAV) platform for the Smart UAV (SUAV) development program, which is one of the 21st Frontier research and development programs supported by the Korean government [1]. Some of the subsystems of the SUAV, such as the rotor system, drive system, andflight control system, were especially challenging for the Korea Aerospace Research Institute to develop, due to lack of previous development experience. To reduce the risk of failure during the development, the ironbird test of the rotor and drive system and flight simulation was conducted. Although the flight simulation mitigated the risk in the flight control system development, a downscaled model was constructed in an effort to ensure the safety in the flight test of the full-scale vehicle. The small-scale platform is shown in Fig. 1. It is expected from the flight test of the small scale model to enhance the understanding of the features of the actual tiltrotor vehicle. The size of the downscaledmodelwas determined to be 1=2:5 (2m overall length) of the full-scale SUAV (5 m overall length) so as to have large enough space for the simple flight control computer, the navigation system, and the other components such as the engine and actuators. The aerodynamic performance of the 1=2:5-scale tiltrotor was analyzed using the in-house performance code. The calculated performance data were used for scheduling the control devices, such as the collective pitch of the rotors and the flaperon deflection, in the flight control logic. In this Note, sizing and performance analysis of the downscaled tiltrotor are presented in which simple codes based on blade element and momentum theory were used. The conversion corridor of the tiltrotor was predicted and the nacelle tilt angle and air speed were compared with flight-test results. After a series of progressive flight tests, conversion flight from helicopter to fixed-wing mode was accomplished. This verified that the stability and control augmentation algorithmwork properly in the flight control software. This small tiltrotor flight test is expected to reduce risks in flight test of the 5-m-span full-scale tiltrotor UAV called Smart UAV. II. Sizing of the Tiltrotor

Real-Time Flight Trajectory Optimization and its Verification in Flight

T HIS Note presents a real-time flight trajectory optimization method. The flight trajectory that minimizes the fuel consumption or flight time of an aircraft can be solved with optimization. Because the optimization is time-consuming, the optimal flight trajectory needs to be obtained before flight. This, however, cannot cope with unexpected situations in flight. Thus, the real-time optimization, which optimizes the flight trajectory with the transition of the flight state, is important. The final goal of our study is to establish the real-time trajectory generation applicable to emergency landing approaches. Though the present flight management system provides the optimal flight path for a flight plan in normal operations, it cannot operate in emergency situations. This Note presents a fundamental real-time trajectory optimization algorithm and its flight validation. Many studies about real-time flight-path generation have produced a flight path by connecting the trim conditions in prestored databases [1,2]. In these studies, computer-assisted simulations were performed, but there were few actual flight experiments. Additionally, it is difficult to generate nonstational trajectories and the wind effects were not considered. In Japan, the Society of Japanese Aerospace Companies has promoted research on the development of a fault-tolerant flight control system. In a proceeding study, Suzuki et al. proposed a flight trajectory search method composed of a realtime A* algorithm and a random tabu search method [3]. They also conducted flight experiments. This random-based method, however, lacks convergent stability, so it often does not produce the appropriate trajectory. We have taken over this study, and in this Note we apply a direct collocation method including some schemes to solve the preceding problems. Actual flight experiments are also conducted to prove the effectiveness of the method. The experimental aircraft is MuPAL, the Multi-Purpose Aviation Laboratory, developed by JAXA (Japan Aerospace Exploration Agency) [4]. It has a high-precision Global Positioning System/inertial navigation system and a three-axis airspeed sensor. The system allows the estimation of the wind velocity and direction during flight. MuPALalso has a tunnel-in-the-sky display, which was developed by JAXA [5]. The experiments were conducted under manual tracking control using the display. In the future, the trajectory generated by the real-time optimization method will be tracked by an automaticflight control systembeing developed in the research on the fault-tolerant flight control system.

Identification of the Transonic Aerodynamic Model for a Re-entry Vehicle

The development of flight control systems for aerospace vehicles requires the availability of reliable aerodynamic models. Pre-flight models, obtained by means of wind tunnel tests and computational fluid dynamics analyses, are usually to be refined using test flight data in order to reduce the level of uncertainty on aerodynamic coefficients. In this paper we present a methodology for the estimation of the lateral-directional aerodynamic model of a re-entry vehicle in subsonic, transonic and supersonic regimes. The identification is formulated as a nonlinear filtering problem and solved through a multi-step approach using the Unscented Kalman Filter. The exploitation of all the available a priori information for the stochastic characterization of the filter models and sensors noises and a rigorous management of all the uncertainties involved in the system identification process allow to obtain reliable figures of estimation accuracy, that are of paramount importance for the design of guidance, navigation and control (GNC) system. The methodology is applied to the simulated data of the CIRA Dropped Transonic Flight Test 2 mission, that will be performed by the end of 2009, with the objective of refining the lateral-directional aerodynamic model.

비행제어시스템 설계 및 검증 절차

Relaxed static stability(RSS) concept has been applied to improve aerodynamic performance of modern version supersonic jet fighter aircraft. Therefore, flight control systems are necessary to stabilize an unstable aircraft, and provides adequate handling qualities and achieve performance enhancements. Standard FCSDVP (Flight Control System Design and Verification Process) is provided to reduce development period of the flight control system. In addition, if this process is employed in developing flight control system, it reduces the trial and error for development and verification of flight control system. This paper addresses the flight control system design and verification process for the RSS aircraft utilizing design goal based on military specifications, linear and nonlinear system design and verification based on universal software, handling quality test based on HILS(Hardware In-the-Loop Simulator) environment, and ground and flight test results to verify aircraft dynamic flight responses.

Practical near hover flight control of a ducted fan (SLADe)

The development of a near hover flight control system for an electrically powered, coaxial, counter-rotating ducted fan aircraft (SLADe—surface launched aerial decoy) is presented. The control and estimation algorithms are designed to operate with low-cost sensors, be computationally efficient, robust and easily adaptable from flight test data. Successive loop closure is employed in the controller design with feedback signals from linear decoupled estimators. Simulation results verify the controller functionality and show that the decoupled estimators perform well near hover when compared to a coupled nonlinear 16 state Kalman filter. Flight test results are presented that verify the practical success of the autopilot.

Development of flight control system for 2D differential geometric guidance and control problem

PurposeThe paper aims to provide further study on the development and analysis of flight control system for two‐dimensional (2D) differential geometric (DG) guidance and control system based on the application of a set‐point weighting proportional‐integral‐derivative (PID) controller.Design/methodology/approachThe commanded angle‐of‐attack is developed in the time domain using the classical differential geometry theory. Then, a set‐point weighting PID controller is introduced to develop a flight control system so as to form the 2D DG guidance and control system, and the gains of the PID controller are determined by the Ziegler‐Nichols method as well as the Routh‐Hurwitz stability criterion. Finally, the classical frequency method is utilized to study the relative stability and robustness of the designed flight control system.FindingsThe results demonstrate that the designed controller yields a fast responding and stable system which is robust to the high frequency parameters variation. Moreover, the DG guidance law is viable and effective in a realistic missile defense engagement.Originality/valueThis paper provides a novel approach on the development of DG guidance and control system associated with its stability analysis.

THE DEVELOPMENT OF FLIGHT CONTROL SYSTEM FOR THE CLOSE RANGE SURVEILLANCE UAV

This paper presents the development of flight control system for the close range surveillance UAV. The developed flight control system consists of hardware and software including flight control and guidance algorithms. It includes essential functions of UAV operations such as autopilot, point navigation and preprogram mode. The results of flight and HILS tests show that the developed system possesses performances suitable for the purpose of close range surveillance.

DEVELOPMENT OF AUTOMATIC FLIGHT CONTROL SYSTEMS FOR WING IN SURFACE EFFECT CRAFT

This paper will discuss the development of an automatic flight control system of Wing in Surface Effect Craft (WiSE). Since the vehicle must fly at very low altitude to gain an improved lift to drag ratio, the altitude-holding system is designed to hold the desired optimal altitude. The design for the altitude holding system is done by using the classical control law of gain scheduling. The gains of three different altitudes are determined by analysing the root locus diagram. Some simulation results are presented and evaluated in this paper.

Interactive Design and Simulation Platform for Flight Vehicle Systems Development

The need for integrated systems development arises from complexity of a system. This paper presents an interactive design and simulation platform for flight vehicle systems development. Its “connect-and-play” capability and adaptability enable “on-line” interaction between design and simulation during the integrated development. As a case study, the implementation of the proposed platform and an aircraft flight control system development example are demonstrated on an experimental test bed including a real-time systems simulator and a flight training device.

Full-Envelope Robust Control of a Shrouded-Fan Unmanned Vehicle

The development is described of a rate-command system for the control of a novel unmanned vehicle, the baseline model of which is highly nonlinear and presents fast and unstable open-loop modes. Structured singular-value design methodology is used to achieve the desired command response characteristics under specified uncertainties taking into consideration typical problems of small-size helicopters and ducted-fan vehicles such as rate-limited servos and significant time delays. Two robust linear controllers are designed for the low- and high-speed subsets of the operating envelope, and full-envelope flight control is achieved by switching between controllers as the threshold airspeed is traversed. Following an analysis of scaling effects, controller performance is evaluated against rotorcraft handling-qualities specifications. Flight control system development is assessed by piloted, hardware-in-the-loop simulation in the full range of operating conditions, with the controller implemented in the flight computer and pilot commands transmitted to the vehicle via radio link. Simulation testing also shows that the control system has good turbulence pust rejection performance and is robust to significant variations of c.g. position.

Decentralized Flight Control Integration and Performance Evaluation

Decentralized flight control shows great advantage in implementation, and is well fitted into a systematic flight control systems development and engineering process. However, the preservation of locally achieved design specifications under the integrated system environment is largely unknown. This work proposes two control integration approaches to address this challenge. It describes both open-loop and closed-loop integration strategies within an integrated control framework. Furthermore, the selection of candidate subsystems is also taken into account during integration. Using an integrated longitudinal pitch attitude and speed control example, the performance of both strategies is evaluated. Numerical simulation results are included for further demonstration.