Flight Control Actuation System Research Articles

A Decentralized Voting and Monitoring Flight Control Actuation System for eVTOL Aircraft

The emergence of eVTOL (electrical Vertical Takeoff and Landing) aircraft necessitates the development of safe and efficient systems to meet stringent certification and operational requirements. The primary state-of-the-art technology for flight control actuation in eVTOL aircraft is electro-mechanical actuators (EMAs), which heavily rely on multiple redundancies of critical components to achieve fault tolerance. However, challenges persist in terms of insufficient reliability, immaturity, and a lack of a measurable evaluation method. This research addresses these issues by elucidating the design requirements for EMAs in eVTOL aircraft and proposing a systematic design and evaluation approach for EMA architecture. A key enhancement involves the incorporation of decentralized voting and monitoring (VoDeMo) mechanisms within the Electronic Control Units (ECUs) to improve the overall safety of the EMA. The paper introduces an innovative triple-dual redundant architecture for aircraft control effectors, comprising three dissimilar lanes of ECUs and two similar redundant parallel channels of power electronics and motors. The design is synergistically supported by a comprehensive evaluation that incorporates quantifiable Model-Based Safety Assessment (MBSA), utilizing both physical simulation and logical safety models. Hardware-In-the-Loop (HIL) tests are conducted on a constructed prototype to validate the proposed architecture.

Modeling a Hydraulically Powered Flight Control Actuation System

Many different types of aircraft designs have flight control systems (FCS) powered by hydraulic systems. With respect to the torques, moments, surface areas, and opposing forces to be acted upon, components introduce faults into the hydraulic system when these components are aging or degrading. The diagnostics of a hydraulically powered flight control actuation system (HPFCAS) rely on the faults produced within the subsystem components as well as the entire system’s mechanism itself. In this paper, a model for an HPFCAS is developed to analyze faults where the HPFCAS was approached as a system of systems (SOS). The identified faults were injected into the system. It is established that some of the faults from the different subsystems had similar characteristic effects and were propagated with attendant consequences. For instance, a measured decrease in the pressure value is observed because of the decrease in the pump speed. A similar characteristic is observed if there is leakage on the line or if there is a clogging valve. These form complex integrated responses in determining where the fault is coming from if only one component is analyzed since it involves components serving different subsystems. Results show that only models that can describe the real characteristics or attributes of the specific systems, due to their defined components, are sufficient for effective diagnostics. This is because the data obtained are more accurate at predicting the behavior of components.

A Review of Diagnostic Methods for Hydraulically Powered Flight Control Actuation Systems

Aircraft systems are designed to perform functions that will aid the various missions of the aircraft. Their performance, when subjected to an unfamiliar condition of operation, imposes stress on them. The system components experience degradation due to fault which ultimately results in failure. Maintenance and monitoring mechanisms are put in place to ensure these systems are readily available when required. Thus, the sensing of parameters assists in providing conditions under which healthy and faulty scenarios can be indicated. To obtain parameter values, sensor data is processed, and the results are displayed so that the presence of faults may be known. Some faults are intermittent and incipient in nature. These are not discovered easily and can only be known through a display of unusual system performance by error code indication. Therefore, the assessed faults are transmitted to a maintenance crew by error codes. The results may be fault found (FF), no fault found (NFF), or cannot display (CND). However, the main classification of the faults and their origins may not be known in the system. This continues throughout the life cycle of the system or equipment. This paper reviews the diagnostic methods used for the hydraulically powered flight control actuation system (HPFCAS) of an aircraft and its interaction with other aircraft systems. The complexities of the subsystem’s integration are discussed, and different subsystems are identified. Approaches used for the diagnostics of faults, such as model-based, statistical mapping and classification, the use of algorithms, as well as parity checks are reviewed. These are integrated vehicle health management (IVHM) tools for systems diagnostics. The review shows that when a system is made up of several subsystems on the aircraft with dissimilar functions, the probability of fault existing in the system increases, as the subsystems are interconnected for resource sharing, space, and weight savings. Additionally, this review demonstrates that data-driven approaches for the fault diagnostics of components are good. However, they require large amounts of data for feature extraction. For a system such as the HPFCAS, flight-management data or aircraft maintenance records hold information on performance, health monitoring, diagnostics, and time scales during operation. These are needed for analysis. Here, a knowledge of training algorithms is used to interpret different fault scenarios from the record. Thus, such specific data are not readily available for use in a data-driven approach, since manufacturers, producers, and the end users of the system components or equipment do not readily distribute these verifiable data. This makes it difficult to perform diagnostics using a data-driven approach. In conclusion, this paper exposes the areas of interest, which constitute opportunities and challenges in the diagnostics and health monitoring of flight-control actuation systems on aircraft.

Sensing Data-Based Degradation Estimation of Electromechanical Actuator Under Dynamic Operating Conditions

In flight control actuation systems, electromechanical actuators (EMAs) as energy-efficient components play critical roles in next-generation aircraft. Reliable degradation estimation models based on sensing data from sensors installed in EMA can help to improve its availability and safety, which need to be studied. However, the dynamic operating conditions, which will lead to the mismatch issues of the sensing data-based degradation estimation models, reduce the overall accuracy of the degradation estimation. To address this issue, an improved adaptive Wiener process model for EMA degradation estimation based on sensing data with dynamic operating conditions is proposed. First, a health indicator (HI) is extracted from sensing data through physical analysis or extracted using data processing techniques. Then, a novel model for EMA degradation estimation with the improved strong tracking filter and the Wiener process model (i.e., ISTF-WPM) is formulated based on the obtained HI. By introducing a saturation function to the improved strong tracking filter (STF), the performance of sensing data-based degradation estimation with dynamic operating conditions can be improved. The proposed model presents a novel routine for EMA degradation estimation based on sensing data with dynamic operating conditions. In the experiments of this study, the effectiveness of the ISTF-WPM is validated using simulation data as well as practical data collected by sensors of EMA. Experimental results indicate that EMA sensing data-based degradation estimation with the proposed model has a high overall accuracy.

Flight control actuation system modeling and engineering application

The flight control actuation system drives the rudder deflection and is an important part of aircraft handling system. This paper combines the development process of engineering flight simulator, takes a civil aircraft multi-function spoiler actuation system as an example, builds a simulation model of flight control multi-function spoiler actuation system based on MATLAB/Simulink modeling tool, and completes the model real-time code conversion, software installation and platform testing. The test results show that the simulation model can better reflect the function and characteristics of the spoiler actuation system, which has important engineering application value for the development of engineering simulator.

Research on energy flow charachteristics between elector-hydrostatic actuartor and power supply network

The characteristics of energy feedback, high power mutation and constant power load of electro-hydrostatic actuator have impact on power supply network of aircraft. Based on the theoretical analysis, this paper establishes the layered model of electro-hydrostatic actuator and power supply network with energy characteristics for the first time, and determines the parameter setting method of key devices in the model. The motor speed and electromagnetic torque direction and energy transfer rule were obtained by simulation when electro-hydrostatic actuator was motoring or braking. Four Quadrant Operation of electro-hydrostatic actuator's impact on the power supply network is quantitative analyzed. In the case of no brake resistance, the energy feedback from electro-hydrostatic actuator makes the bus voltage of the power supply network of the aircraft far beyond the safe voltage range. Finally, the comprehensive effect of multiple electro-hydrostatic actuators on power supply network is analyzed. When multiple electro-hydrostatic actuators have the same operating instructions and the same loads, their impact on the power supply network will be superimposed. When multiple electro-hydrostatic actuators have opposite operating instructions and the same loads, their impact on the power supply network will be counterbalanced. The command of flight control actuation system and real-time aerodynamic load are the key factors affecting the impact of multiple electro-hydrostatic actuators on the power supply network. The results of this paper have guiding significance for restraining the impact of electric actuator energy consumption on the power supply network of aircraft.

Performance analysis of electrical flight control actuation system in a commercial transport aircraft

PurposeIn more than 100 years of aviation, significant progress has been made in flight control systems. The aircrafts that have entered service for the past ten years tend towards power-by-wire flight control with electrical actuators. The purpose of this study is to analyse the effects of electrical actuation on power consumption, weight and fuel consumption on a commercial transport aircraft.Design/methodology/approachThe Airbus A321-200 aircraft was chosen as a case study for analysing the effects of electrical actuation on the flight control actuation system (FCAS) architecture, and Pacelab SysArc software was used for design, modelling and analysis. As alternatives to the existing system, hybrid and all-electric models are built to a set of design guidelines with certain limitations.FindingsCompared to the existing FCAS architecture model, 80 kg weight savings in the hybrid FCAS architecture model and 171 kg weight savings in the all-electric FCAS architecture model were observed. In terms of fuel consumption, it has been observed that there is 0.25% fuel savings in the hybrid FCAS architecture model, and 0.48% fuel savings in the all-electric FCAS architecture model compared to the existing FCAS architecture model at 3200 NM.Practical implicationsIn line with the data obtained from this study, it is predicted that electrical actuation is more preferable in aircraft, considering its positive effects on weight and fuel consumption.Originality/valueIn this study, three different models were created: the existing FCAS architecture of a commercial transport aircraft, the hybrid FCAS architecture and the all-electric FCAS architecture. Hybrid and all-electric models are built according to a set of design guidelines, with certain limitations. Then, similar flight missions consisting of the same flight conditions are defined to analyse the effects of power consumption, weight, and fuel consumption comparatively.

Diagnostics of actuation system by Hadamard product of integrated motor current residuals applied to electro-mechanical actuators

The paper presents diagnostics methodology that can identify the event of occurrence of fault in the actuator or the linkage system of the flight control actuation system driven by Linear Electromechanical Actuators (LEMA). The standard data analysis like motor current signature analysis (MCSA) is good at identifying the incipient faults within the elements of the actuators in situations where-in the actuators are driving control surfaces. But in back driven cases, where-in LEMA is driven back by control surfaces, the faults outside the LEMAs are difficult to be detected due to higher mechanical advantages of transmission elements like roller screws, gear train and linkage arms scaling down their effects before reaching the motor. One such event occurred in a ground test, wherein the jet vanes were sheared when back driven by excessive gas dynamic forces. Neither the motor current nor the LEMA position feedback data has any clue of the instance of occurrence of such shearing. The case study is discussed in detail and diagnostics solution for such failures is proposed. A new methodology to pin point the event of occurrence is arrived at based on ground static test data of four independent channels. The same is reassured for its applicability using lab experiments on three samples mimicking the failure. The method's applicability is also extended for extracting events in actual flight, by comparing the flight telemetry data with the mimicked lab level (dry runs) data. The methodology uses the analysis of LEMA motor current data to arrive at the vital diagnostic information. The current data of LEMA directly cannot be interpreted due to non-stationary nature arising from variable speed and its pulsating form because of the pulse width modulation (PWM) switching, threshold voltages and closed loop dynamics of the servo. Hence the motor current is integrated using cumulative trapezoidal method. This integrated data is spline curve fitted to arrive at residuals vector. The Hadamard product is used on the residuals vector to amplify the information and suppress the noise. Further, normalizing is done to compare data across tests and samples. With this, necessary diagnostic information was extracted from static test data. The method is extended for extracting diagnostics information from actual flight using comparison analysis of, the test data in actual environment with mimicked lab level dry runs. It is also verified for applicability in faults directly driven by actuators in lab level experiments on three samples.

A passive cooling design for an aircraft electromechanical actuator by using heat pipes

The more electric aircraft using the fly-by-wire technology is a developing trend for future aircrafts, in which the electromechanical actuator (EMA) is the key technique to realize the power-by-wire in the flight control actuation systems. The EMAs often suffer high heat generation during a short time, causing severe thermal problem, however, energy consumption and space for cooling systems are very limited on aircrafts. In this paper, a passive cooling design is proposed by transferring heat from the major sources of the EMA directly to fuel tanks by using heat pipes. The heat generation characteristics and the cooling performance of the EMA are experimentally investigated. The results show that the heat generation rates increase with the velocity and output force of the EMA, and efficient cooling for the EMA is performed by such a passive cooling device which decreases the temperature rise of the motor by 17.8% to 58.6% under different operation conditions. It is also found that the performance of this cooling device is majorly influenced by the coolant temperature and environmental temperature, while the impact of the coolant flow rate is not obvious. This work lays a foundation for the cooling design of the electromechanical actuators in an aircraft.

Experimental research of dynamic characteristics of electro-backup hydraulic actuator at the load test-rig of flight control systems

The trend in aircraft systems development to overall electrification includes electrification of flight control actuation system. Generally, there’re two main technologies of «electrical» actuation, which use the power of on-board electrical system to move the aerodynamic surfaces. These technologies are electro-hydrostatic (EHA) and electromechanical (EMA) actuation systems. This paper describes the results of the dynamic testing of electro-backup hydraulic actuator (EBHA), which is one of the possible solutions for actuation systems of the next-generation flying vehicles.

Remaining Useful Life Prognostics for the Electrohydraulic Servo Actuator Using Hellinger Distance-Based Particle Filter

The electrohydraulic servo actuator (EHSA) has been widely used in the flight control actuation system of civil aircraft. Accurate remaining useful life (RUL) prediction for EHSA is of great significance in inhibiting the unexpected failures of aircrafts, improving the scheduling ability, and reducing the maintenance costs, which has become highlight for airline operators in recent years. The particle filtering (PF) is a commonly accepted technique in RUL prediction due to its superiority in coping with nonlinear and non-Gaussian behaviors. However, the irrational selection of importance density can induce the problem of particle degeneracy and restrict the estimation performance, resulting in the decrease of RUL prediction accuracy. In this paper, an importance density choice scheme grounded on the Minimum Hellinger Distance principle, namely, MHD-PF, is proposed to fight against the degeneracy problem at the early phase. In addition, the strict mathematical derivation process of the proposed scheme is proffered. Then, the MHD-PF is used to track the degradation state and predict the RUL of EHSA. Finally, the specific performance of the MHD-PF method is performed by a series of actual experiments. The effectiveness of the proposed RUL prognostics approach is also experimentally verified.

An intelligent design method for actuation system architecture optimization for more electrical aircraft

The design of flight control actuation system is facing major challenge due to the development of more electrical aircraft. The task is to find the combinations of power sources, actuators and computers, which becomes more complex because of the new power sources and actuator types of more electrical aircraft. It is impossible to determine optimal architecture by traditional trial-and-error method within acceptable time. Therefore, the need for new methodology for actuation system architecture design emerges. This study proposes an intelligent design method which has steps of design space exploration of actuation system architectures by constraint satisfaction problem (CSP) method, safety assessment process to exclude unsafety solution, multi-objectives optimization to get Pareto optimal front and comprehensive decision for final architecture via analytic hierarchy process. And the design method is implemented in python and a software platform is developed. Furthermore, within the paper a case study for A350 flight control actuation system is presented to testify the application of this methodology. Compared to the traditional hydraulic architecture, the optimal architecture is more competitive in weight, power and cost. At the same time, the optimal architecture is found in less than 30 minutes among 1075 candidates, which greatly reduces the design cycle. This method deals with the problem in the design of flight control actuation system.

Simulating actuator energy consumption for trajectory optimisation

This work aims to construct a high-speed simulation tool which is used to quantify the dynamic actuator power consumption of an aircraft in flight, for use within trajectory optimisation packages. The purpose is to evaluate the energy penalties of the flight control actuation system as an aircraft manoeuvre along any arbitrary trajectory. The advantage is that the approximations include major transient properties which previous steady state techniques could not capture. The output can be used to provide feedback to a trajectory optimisation process to help it compute the aircraft level optimality of any given flight path. The tool features a six degree of freedom dynamic model of an aircraft which is combined with low frequency functional electro-mechanical actuator models in order to estimate the major transient power demands. The actuator models interact with the aircraft using an aerodynamic load estimator which generates load forces on the actuators that vary as a function of flight condition and control surface demands. A total energy control system is applied for longitudinal control and a total heading control system is implemented to manage the lateral motion. The outer loop is closed using a simple waypoint following guidance system with turn anticipation and variable turn radius control. To test the model, a simple trajectory analysis is undertaken which quantifies a heading change executed with four different turn rates. The tool shows that the actuation system requires 12.8 times more electrical energy when performing a 90° turn with a radius of 400 m compared to 1000 m. A second test is performed to verify the model’s ability to track a longer trajectory under windy conditions.

Prognostic and Health Management System for Fly-by-wire Electro-hydraulic Servo Actuators for Detection and Tracking of Actuator Faults

Maintenance of flight control actuation systems is currently performed on a scheduled basis, however air fleet operators and component manufacturers are willing to move from scheduled maintenance to Condition Based Maintenance (CBM) in order to reduce maintenance costs and improve aircraft dispatchability. Prognostics and Health Management (PHM) systems are a critical part of CBM and are perceived as a breakthrough technology to effectively respond to an urgent and critical need to improve the readiness, availability, reliability, safety and maintainability of aerospace vehicles. This paper presents the results of an ongoing research activity focused on the development of a PHM system for fly-by-wire Electro-Hydraulic Servo Actuators (EHSA) without adding new sensors. The PHM system is being developed with the objective of detecting the most common faults, according to a failure mode effects and criticality analysis (FMECA). The paper describes in particular the tools used for detection and tracking of internal leakage faults of the hydraulic actuator, which is one of the most common faults of hydraulic servo-actuators in service, and for predicting its remaining useful life (RUL). The research work has been supported by the development of a nonlinear model for a reference EHSA, that has been implemented using physical equations and system parameters, taking into account environmental condition and disturbances. The model was validated through tests runs on a flight control actuator of a civil aircraft. Simulations are performed in nominal conditions and with progressive injection of degradation to verify the PHM algorithm. The performances of the PHM algorithms are evaluated by means of proper metrics.

Health and Usage Monitoring System (HUMS) Strategy to enhance the Maintainability & Flight Safety in a Flight Control Electromechanical Actuator (EMA)

Concerning flight control actuation systems, the current state of the art is the fly-by-wire control of servo actuators hydraulically operated. In the last decades several research programs have been dedicated to analyze the replacement of these equipments with the implementation of Electro-Hydrostatic Actuators (EHA) or Electro-Mechanical Actuators (EMA), pursuing the challenge to get More Electrical Aircraft (MEA).This paper focuses on the HUMS strategy proposed to provide diagnosis and prognosis of the EMA as one way to enhance Flight Safety and Maintainability for its use in Primary Flight Controls.Considerations related to Flight Safety are analyzed, specially linked to the main problem associated to the EMA, the critical effects derived from the mechanical jamming of some parts of the EMA, and the advantages of the HUMS implementation to anticipate and mitigate these consequences.

A methodology for vehicle and mission level comparison of More Electric Aircraft subsystem solutions: Application to the flight control actuation system

As part of the More Electric Initiative, there is a significant interest in designing energy-optimized More Electric Aircraft, where electric power meets all non-propulsive power requirements. To achieve this goal, the aircraft subsystems must be analyzed much earlier than in the traditional design process. This means that the designer must be able to compare competing subsystem solutions with only limited knowledge regarding aircraft geometry and other design characteristics. The methodology presented in this work allows such tradeoffs to be performed and is driven by subsystem requirements definition, component modeling and simulation, identification of critical or constraining flight conditions, and evaluation of competing architectures at the vehicle and mission level. The methodology is applied to the flight control actuation system, where electric control surface actuators are likely to replace conventional centralized hydraulics in future More Electric Aircraft. While the potential benefits of electric actuation are generally accepted, there is considerable debate regarding the most suitable electric actuator – electrohydrostatic or electromechanical. These two actuator types form the basis of the competing solutions analyzed in this work, which focuses on a small narrowbody aircraft such as the Boeing 737-800. The competing architectures are compared at both the vehicle and mission levels, using as metrics subsystem weight and mission fuel burn, respectively. As shown in this work, the use of this methodology aids the decision-making process by allowing the designer to rapidly evaluate the significance of any performance advantage between the competing solutions.

Detection of faults in flight control actuators with hard saturation and friction non-linearity

Expeditious diagnosis and prognosis in mission critical flight control actuation system is essential to prevent loss of aerospace vehicles on impending failures. This paper presents the application of the additive form of the Unscented Kalman Filter (UKF-A) estimator for detecting faults in actuators with significant saturation and friction nonlinearity. An electromechanical flight actuator based on a DC torque motor is considered. The complete dynamics of the actuator including mounting structure stiffness, load dynamics, compensators, power amplifier with current loop and saturation nonlinearity, along with friction nonlinearity is considered. Capability of the method for detecting faults like changes in motor coil resistance and inductance, torque constant and viscous friction coefficient in the actuator under realistic input conditions is demonstrated. The method presented is attractive for real time implementation as only a single estimator is used for fault diagnosis. Robustness of the scheme to sensor noise and external disturbances is also presented.

A fault tolerant single sided matrix converter for flight control actuation systems

We describe a single sided matrix converter (SSMC) designed for safety critical applications like flight control actuation systems. Dynamic simulations of multi-phase SSMC using Matlab Simulink are carried out to evaluate the fault tolerance capabilities. Investigation into different numbers of phases and power converter topologies under single phase open circuit, single switch open circuit, and single switch short circuit has been executed. The simulation results confirm 5-phase SSMC design as a compromise between fault tolerance and converter size/volume. A 5-phase SSMC prototype was built. Experimental results verify the effectiveness of our design.

Challenges for Health Monitoring of Electromechanical Flight Control Actuation Systems

Flight control systems of civil aircraft have undergone huge developments in the last decades. The current more/all electric aircraft concepts lead to the replacement of the hydraulic actuators in the primary flight control systems by electromechanical systems. Integrating electromechanical systems in safety critical applications implies three main challenges: (a) the detection of all fault cases which could lead to a safety critical state, (b) finding measurement parameters capable to detect faults, and (c) the development of algorithms to detect faults under all flight conditions. Putting the scope on the health monitoring of the mechanical components of a direct drive actuator, a new technology based on piezoresistive thin film sensors (TFS) is presented and its potential shown by using defective ball bearings as an example.

Architecture Optimization of More Electric Aircraft Actuation System

The optional types of power source and actuator in the aircraft are more and more diverse due to fast development in more electric technology, which makes the combinations of different power sources and actuators become extremely complex in the architecture optimization process of airborne actuation system. The traditional “trial and error” method cannot satisfy the design demands. In this paper, firstly, the composition of more electric aircraft (MEA) flight control actuation system (FCAS) is introduced, and the possible architecture quantity is calculated. Secondly, the evaluation criteria of FCAS architecture with respect to safe reliability, weight and efficiency are proposed, and the evaluation criteria values are calculated in the case that each control surface adopts the same actuator configuration. Finally, the optimization results of MEA FCAS architecture are obtained by applying genetic algorithm (GA). Compared to the traditional actuation system architecture, which only adopts servo valve controlled hydraulic actuators, the weight of the optimized more electric actuation system architecture can be reduced by 6%, and the efficiency can be improved by 30% based on the safe reliability requirements.