Flight Control Computer Research Articles

Design, Analysis, and Experimental Investigation of a Cyclocopter with Two Rotors

A cyclocopter is a vertical takeoff and landing aircraft that has cycloidal blade systems that consist of several blades rotating about a horizontal axis. The cycloidal blade system allows changes in direction and magnitude of the generated thrust. This paper describes the design and experimental studies of a 110 kg cyclocopter having two cycloidal blade system rotors and one tail propeller rotor powered by a -class rotary engine. The tail rotor generates pitching motion and compensates for the torque generated by the front two rotors. The upward thrust direction of the tail rotor provides 10% of the lifting force. The design thrust of the cycloidal blade system rotor is estimated in analytical and numerical analyses. An accurate analytical aerodynamic model is developed. The analytically estimated results have a trend similar to that of computational fluid dynamics analysis. Both analysis results match well with the results of a ground test. The onboard flight control computer system has two embedded processors, and a proportional-integral-derivative algorithm is used as the flight control scheme. The gains of the proportional-integral-derivative algorithm are adjusted during tethered flight tests. Finally, the developed cyclocopter demonstrates hovering and maneuvering flight in a tethered test.

Design of a field programmable gate array based test launch and control system for hybrid vehicle

In order to satisfy the throttling and multiple restart characteristics of the hybrid rocket engine and guarantee the controllability and security of hybrid rocket, the test launch and control system for hybrid vehicle needs to achieve ground power control, serial communication with the flight control computer and the telemetry system, onboard thermal battery activation, battery voltage acquisition, power conversion, pressure monitoring, valve control, booster ignition, power-off and remote pressure relief in emergency. Considering the complexity of the test launch and control system for hybrid system and the advantages of Field Programmable Gate Array (FPGA), such as, a large number of I/O ports, ability to efficiently implement large-scale systems and programmability, a FPGA based test launch and control system for a type of hybrid vehicle was designed. Many conducted ground and flight experiments have proved that the system has superior performance in integration, universalization, miniaturization, automation and can perfectly meet the needs of the hybrid vehicle.

무인항공기 비행제어 HILS 시험환경 연구

A UAV(Unmanned Aerial Vehicle) flies along pre-programed navigation points(in-flight, take-off, or landing) automatically without pilot input. Even though UAVs fly differently from general piloted aircraft as the pilot controls the aircraft from a ground station through means of a data-link system. Occasionally, the data-link connection can be lost for any number of reasons, in which case, the FLCC(Flight control Computer) must automatically switch to autopilot to continue flying. Hence, the FLCC is a flight-critical component that must be throughly tested and validated. This paper discusses the development of a HILS(Hardware in the Loop Simulation) test environment designed to simulate real flight conditions to verify the FLCC satisfies flying quality requirements and maintains robustness despite any potential malfunctions or emergency situations.

A Data-Driven Approach to Detect Faults in the Airbus Flight Control System

This paper presents a data-driven strategy for the detection of failures impacting the flight control system. Early and robust detection of Oscillatory Failure Case (OFC) allows the aircraft structural design to be optimized, which in turn helps improve the aircraft environmental footprint thanks to weight saving. Compared to existing model-based techniques already used on in-service Airbus aircraft, this paper studies a novel signal processing approach based on distance and correlation. It is shown that a mixed similarity index between Euclidean distance and logarithmic invariant divergence gives promising detection results. This paper details the proposed approach by insisting on practical constraints due to implementation in embedded real-time systems such as the flight control computer. Preliminary results obtained from a Verification & Validation (V&V) on-going campaign are presented.

Practical design considerations for successful industrial application of model-based fault detection techniques to aircraft systems

This paper discusses some key factors which may arise for successful application of model-based Fault Detection (FD) techniques to aircraft systems. The paper reports on the results and the lessons learned during flight V&V (Validation & Verification) activities, implementation in the A380 Flight Control Computer (FCC) and A380 flight tests at Airbus (Toulouse, France). The paper does not focus on new theoretical materials, but rather on a number of practical design considerations to provide viable technological solutions and mechanization schemes. The selected case studies are taken from past and on-going research actions between Airbus and the University of Bordeaux (France). One of the presented solutions has received final certification on new generation Airbus A350 aircraft and is flying (first commercial flight: January 15, 2015).

D-SEND#2の制御系設計

This paper is concerned with flight controller design of D-SEND#2. D-SEND is a project to demonstrate a low sonic boom aerodynamic design concept. In the #2 part of this project, an unpowered test vehicle is lifted to an altitude of 30km by a balloon from which it then separates. After the separation, the vehicle's on-board flight control computer selects a target Boom Measurement System (BMS) according to the separation point. The vehicle then autonomously flies to the selected BMS and establishes prescribed sonic boom measurement flight conditions. The design of the GNC system for D-SEND#2 project is exceptionally challenging since 1) there is no nominal trajectory, 2) there is only limited control available to meet various requirements, 3) the flight envelope is quite wide, and 4) the development time is short. Dynamic inversion method and time-scale separation technique are applied to deal with these problems, and a systematic controller design procedure is presented. The controller is designed with this procedure and evaluated through linear analysis and Monte-Carlo simulations which shows it has sufficient performance and safety margines. This controller design method is confirmed to be a practical method for UAVs.

Implementation of real-time moving horizon estimation for robust air data sensor fault diagnosis in the RECONFIGURE benchmark

This paper presents robust fault diagnosis and estimation for the calibrated airspeed and angle-of-attack sensor faults in the RECONFIGURE benchmark. We adopt a low-order longitudinal model augmented with wind dynamics. In order to enhance sensitivity to faults in the presence of winds, we propose a constrained residual generator by formulating a constrained moving horizon estimation problem and exploiting the bounds of winds. The moving horizon estimation problem requires solving a nonlinear program in real time, which is challenging for flight control computers. This challenge is addressed by adopting an efficient structure-exploiting algorithm within a real-time iteration scheme. Specific approximations and simplifications are performed to enable the implementation of the algorithm using the Airbus graphical symbol library for industrial validation and verification. The simulation tests on the RECONFIGURE benchmark over different flight points and maneuvers show the efficacy of the proposed approach.

Safety risk assessment of navigation and flight control computer system of unmanned aerial vehicle

The article deals with the problem solution of reducing of safety risk level of the navigation and flight control computer system (NAFCCS) of unmanned aerial vehicle (UAV) owing to well-grounded reliability increasing of the NAFCCS critical components. The problem solution lies in development of methods (technique) of calculation of NAFCCS safety risk level, namely probabilities of catastrophic failure occurrence, without Fault Tree Evaluation. The technique allows solving the problems of reducing of safety risk level of the NAFCCS on the systems engineering design stage. The offered technique includes new mathematical models of NAFCCS subsystems with detail presentation of the of critical fault state.

Modeling, Control, and Experimental Validation of a High-Speed Supercavitating Vehicle

Underwater vehicles that travel inside a bubble or supercavity offer possibilities for high-speed and energy-efficient transportation of cargo and personnel. Validation and testing of mathematical models and control systems for these vehicles is a challenge due to the cost and complexity of experimental facilities and testing procedures. A cost-efficient and low-complexity approach to the experimental validation of mathematical models and control systems for a supercavitating test vehicle is presented in this paper. The proposed method enables the testing of control algorithms subject to steady and unsteady flows in a high-speed water tunnel. The method combines a real-time simulation of the vehicle dynamics, force measurements from an experimental scale vehicle, and flight control computer to reproduce the vehicle motion subject to realistic flow conditions and hardware constraints as actuator saturation and time delay. The model of the vehicle dynamics, used for the validation infrastructure and control design, is derived using experimental data. A controller designed to track pitch angle reference commands was tested on the experimental platform. The test cases validated the operation of the vehicle and controller subject to steady and unsteady flows.

안정성증강 작동기를 이용한 Fly-By-Wire 헬리콥터 제어법칙 설계에 대한 연구

The previous limited authority system capable of implementing attitude command response type and translational command response type by operating SAS actuator has the problem of early saturation of SAS actuator since SAS actuator should compensate the mechanical linkage displacement caused by control sick movement. In this paper, a limited authority system where flight control computer receives the command from the control stick which is not connected to the mechanical linkage is described. In this system the compensation by the SAS actuator is not necessary and SAS actuator saturate later. SAS actuator saturation problem can be further relaxed by using the trim actuator. This new limited authority system is applied to BO-105 model, simulation is performed for the doublet input and pirouette maneuver is also simulated and analyzed.

English

System redundancy in aircraft is very important as it ensures faulty components would not jeopardize the safety of crew and passengers. High priority system such as flight control computers, instruments system and controller for control surface mechanism have double or triple redundancy to maintain the aircraft safety. While full scale aircraft has multiple redundancy system installed, small unmanned aerial vehicle on the other hand do not have these system installed due to cost and weight. Propulsion, either fuel or electric powered and control surface movement are prone to damage due to wear and tear of mechanical movement. Severe damage could occur if the UAV lost control of its control surface. Elevator system fault would result in the most severe as it controls the vertical movement of the UAV. This research presents the initial development of a fault recovery system by changing the pitching movement control by the elevator to stabilator control. This paper in particular would discuss the approach to construct the sensor system for detecting and identifying the fault of control surface of small UAV during the flight. Servo fault detection would be used as the basis of providing suitable recovery scheme. Fast detection and recovery with light weight system are the main parameters that are considered during evaluation and development stage. The outputs of these sensors in normal and abnormal condition of elevator have been determined for the development of fault detection and identification scheme.

파티션 컴퓨팅 기반의 무인기 고장 감내 관성 항법 시스템

When new inertial navigation systems for an unmanned aerial vehicles are being developed and tested, construction of a fault-tolerant system is required because of various types of hazards caused by S/W and H/W faults. In this paper, a new fault-tolerant flight system that can be deployed into one or more FCCs (Flight Control Computers) is introduced, based on a partition scheme wherein each OFP (Operational Flight Program) partition uses an independent CPU and memory slot. The new fault-tolerant navigation system utilizes one or two FCCs, and executes a primary navigation OFP under development and a stable shadow OFP partition on each node. The fault-tolerant navigation system based on a single FCC can be used for UAVs with small payloads. For larger UAVs, an additional FCC with two OFP partitions can be used to provide both H/W and S/W fault-tolerance. The developed fault-tolerant navigation system significantly removes various hazards in testing new navigation S/Ws for UAVs.

Guidance of Air Vehicles: A Sliding Mode Approach

This paper presents a novel nonlinear guidance scheme for ground track control of aerial vehicles. The proposed guidance logic is derived using the sliding mode control technique, and is particularly suited for unmanned aerial vehicle (UAV) applications. The main objective of the guidance algorithm is to control the lateral track error of the vehicle during flight, and to keep it as small as possible. This is achieved by banking the vehicle, that is, by executing roll maneuvers. The guidance scheme must perform well both for small and large lateral track errors, without saturating the roll angle of the vehicle, which serves as the control input for the guidance algorithm. The limitations of a linear sliding surface for lateral guidance are indicated; a nonlinear sliding surface is thereafter proposed which overcomes these limitations, and also meets the criterion of a good helmsman. Stability of the nonlinear surface is proved using Lyapunov theory; control boundedness is also proved to ensure that the controls are not saturated even for large track errors. The proposed guidance law is implemented on the flight control computer of a scaled YAK-54 UAV and flight results for different scenarios (consisting of both small and large errors) are presented and discussed. The flight test results confirm the effectiveness and robustness of the proposed guidance scheme.

Actuator controller based on fuzzy sliding mode control of tilt rotor unmanned aerial vehicle

The Smart Unmanned Aerial Vehicle (SUAV) is the tilt rotor type aircraft, which has been developed by the Korea Aerospace Research Institute (KARI). The actuation system of the SUAV is composed of actuators, controllers, and mechanical linkages for the control elements, i.e. nacelle tilt, flaperon, elevator, and rotor. It should be able to control the vehicle under a variety of aerodynamic loads while the control performance is fulfilled in terms of accuracy, response time, and stability. In order to satisfy the target performance, the system employs the sliding mode control that guarantees stability under varying loads. In addition, the system is enhanced with the fuzzy control logic that enables intelligent selection of the control switch. Based on this fuzzy sliding mode control (FSMC) scheme, the actuator controller of the SUAV receives the control commands from the Digital Flight Control Computer (DFCC) and transmits them to actuators under the feedback control logic. This paper presents the application of the FSMC to the SUAV and provides the description of the actuation system architectures, modeling and analysis for the actuation system, implementation of the control logic into the actuator controller, and verification of the actuator control system through ground and flight tests of the SUAV.

A Design and Implementation of Bootloader Based on MPC8280

The core unit of the flight control system being the central part of the Unmanned Aerial Vehicle (UAV) system is the flight control computer playing a vital role in the stability and security of the entire system, which has fast and reliable requirements for system startup. Bootloader is the first section of software code that runs after powering on, mainly responsible for initializing hardware devices. By loading an application or operating system kernel, the bootloader completes the system startup. Based on MPC8280 processor’s hardware platform of the central processing unit of distributed flight control computer, this paper designs and implements a non-operating system boot scheme. Under the boot scheme, an optimal boot scheme aiming at increasing efficiency of software development for VxWorks operating system is provide.

Lateral track control of UAVs using the sliding mode approach: from design to flight testing

This paper develops sliding-mode-based nonlinear logic for guidance of unmanned aerial vehicles (UAVs) for curved and straight path following. UAV trajectories generally consist of straight path segments, curved arcs, circular loiters and other manoeuvres; tight ground track control is desired throughout the trajectory. This is achieved by controlling the lateral (cross-track) deviation of the vehicle in flight. The main objective of the guidance algorithm is to keep the lateral track error of the vehicle as small as possible while performing graceful and stable manoeuvres despite the presence of uncertainties and disturbing winds. Lateral track control is usually achieved by banking the vehicle, that is, by executing roll manoeuvres. The scheme must perform well without saturating the roll angle of the vehicle, which serves as the control input for the guidance algorithm. The algorithm proposed here is shown to perform well for both straight and circular path tracking while ensuring control boundedness, and hence no saturation. Crosswinds are a major source of disturbance for the guidance problem. This is incorporated into the design formulation and guidance gains are selected to provide the desired performance despite the presence of disturbing winds. A sliding-mode-based scheme is developed which includes a feedforward component related to the rate of change of the desired path heading. Stability of the algorithm is proved using an appropriate Lyapunov function. The algorithm is implemented in the flight control computer of a scaled YAK-54 research aircraft; flight test results are presented and compared with those from other guidance algorithms. Flight results demonstrate the effectiveness and performance of the proposed guidance scheme. The algorithm considers guidance in the 2D lateral plane only and minimizes deviations from the desired ground track of the vehicle.

Development and application of sliding mode LPV fault reconstruction schemes for the ADDSAFE Benchmark

This paper describes the development and the evaluation of a robust sliding mode observer fault detection scheme applied to an aircraft benchmark problem as part of the ADDSAFE project. The ADDSAFE benchmark problem which is considered in this paper is the yaw rate sensor fault scenario. A robust sliding mode sensor fault reconstruction scheme based on an LPV model is presented, where the fault reconstruction signal is obtained from the so-called equivalent output error injection signal associated with the observer. The development process includes implementing the design using AIRBUS׳s the so-called SAO library which allows the automatic generation of flight certifiable code which can be implemented on the actual flight control computer. The proposed scheme has been subjected to various tests and evaluations on the Functional Engineering Simulator conducted by the industrial partners associated with the ADDSAFE project. These were designed to cover a wide range of the flight envelope, specific challenging manoeuvres and realistic fault types. The detection and isolation logic together with a statistical assessment of the FDD schemes are also presented. Simulation results from various levels of FDD developments (from tuning, testing and industrial evaluation) show consistently good results and fast detection times.

H ∞ ∕ H − LPV solutions for fault detection of aircraft actuator faults: Bridging the gap between theory and practice

SummaryThe work presented in this paper is undertaken within the European FP7 funded Advanced Fault Diagnosis for Sustainable Flight Guidance and Control (ADDSAFE) project. The goal is to propose new fault detection and fault diagnosis techniques that could significantly help developing environmentally‐friendlier aircraft. In this paper, LPV model‐based fault detection schemes are proposed and compared for robust and early detection of faults in aircraft control surfaces servo‐loop. The proposed methodologies are based on a slight modification of the H ∞ ∕ H_ LPV optimization techniques proposed in [1] and [2] for systems modelled in, first polytopic manner, second linear fractional representation fashion. Both the methods aim at designing fault detection filters with enhanced fault transmission H − gain and large h ∞ nuisance attenuation. A complete Monte Carlo campaign from a high representative simulator, provided by Airbus as a part of the ADDSAFE project, demonstrates the potential of the proposed techniques. It is shown that both the proposed fault detection schemes can be embedded within the structure of in‐service monitoring systems as a part of the Flight Control Computer software. Copyright © 2014 John Wiley & Sons, Ltd.

스마트무인기 비행운용프로그램 개발

스마트무인기의 비행제어컴퓨터는 탑재장비와의 입/출력 신호처리, 비행제어법칙 연산 그리고 고장진단 및 이중화 기능 등이 구현된 비행운용프로그램이 탑재되어 있다. 비행운용프로그램은 PowerPC 755 프로세서와 VxWorks 5.5 실시간운영체제 환경에서 개발되었다. 비행운용프로그램은 메모리 참조 모듈, 탑재장비 입/출력 신호처리 모듈 그리고 비행제어법칙 모듈로 구성되었고 각각의 모듈은 계층 구조로 설계되었다. 메모리 참조 모듈과 신호처리 모듈은 벤치 테스트를 통해 검증되었고, 비행제어법칙 모듈은 시뮬레이션 시험, 지상통합시험, 안전줄 시험 그리고 비행시험을 통해 검증되었다. 본 논문에서는 비행운용프로그램의 개발환경, 소프트웨어 구조 그리고 검증 및 관리방법에 대해서 기술하였다. The operational flight program(OFP) which has the functions of I/O processing with avionics, flight control logic calculation, fault diagnosis and redundancy mode is embedded in the flight control computer of Smart UAV. The OFP was developed in the environment of PowerPC 755 processor and VxWorks 5.5 real-time operating system. The OFP consists of memory access module, device I/O signal processing module and flight control logic module, and each module was designed to hierarchical structure. Memory access and signal processing modules were verified from bench test, and flight control logic module was verified from hardware-in-the-loop simulation(HILS) test, ground integration test, tethered test and flight test. This paper describes development environment, software structure, verification and management method of the OFP.

Lateral Control Implementation for an Unmanned Aerial Vehicle

This paper presents practical aspects of guidance and control design for UAV and its flight test results. The paper focuses on the lateral-directional control and guidance aspects. An introduction to the mission and guidance problem is given first. Waypoints for straight and turning flight paths are defined. Computation of various flight path parameters is discussed, including formulae for real-time calculation of down-range (distance travelled along the desired track), cross-track deviation, and heading error of the vehicle; these are then used in the lateral guidance algorithm. The same section also describes how to make various mission-related decisions online during flight, such as when to start turning and when a waypoint is achieved. The lateral guidance law is then presented, followed by the design of a robust multivariable H∞ controller for roll control and stability augmentation. The controller uses the ailerons and rudder for control of roll angle and stabilization of yaw rate of the vehicle. The reference roll angle is generated by the nonlinear guidance law. The sensors available on-board the vehicle do not measure yaw rate; hence, a practical method of its estimation is proposed. The entire guidance and control scheme is implemented on the flight control computer of the actual aerial vehicle and taken to flight. Flight test results for different mission profiles are presented and discussed.