Primary Flight Control Research Articles

Performance and Reliability Evaluation of Innovative High-Lift Devices for Aircraft Using Electromechanical Actuators

In the last decades, electromechanical actuators started to be introduced in transport aircraft for primary and secondary flight control surfaces. Some innovative architectures have been proposed in the literature to use these actuators for high-lift devices (flaps and slats). The state-of-the-art architecture is built with a central mechanical shaft powered by a power distribution unit connected to ballscrew actuators that actuate the flap and slat surfaces. New innovative concepts have the potential to improve the state-of-the-art architectures. However, there is a lack of quantitative results for such innovative architectures. A new methodology is proposed to preliminarily estimate performance and reliability aspects of conventional and innovative architectures. This allows quantitative comparisons to finally be obtained. The methodology is applied to a new architecture that uses electromechanical actuators for flaps and slats, providing results in terms of performance and reliability and comparing them to the current state-of-the-art high-lift devices. Results show that the new architecture is lighter than the reference one and can be more reliable. This is achieved thanks to the removal of the mechanical links among components, which allows each control surface to be deployed independently from the others. This highly increases the operational reliability of the system. Two cases are analyzed, with and without actuator jamming. This provides more realistic results since this failure mode is currently the main reason why electromechanical actuators are not being used for more applications. The innovative architecture outperforms the conventional one in the case where the electromechanical actuators are not affected by the jamming failure mode.

Nonlinear dynamics in the primary flight control system model

Nonlinear dynamics in the primary flight control system model

Safety Design of Uncontained Engine Rotor Failure for Civil Aircraft Primary Flight Control System

The safety design of uncontained engine rotor failure for civil aircraft primary flight control systems is more difficult than other airborne systems. This paper presents a safety design method that establishes a two-stage collision model inspection mechanism and introduces an automatic analysis algorithm in the process of comprehensive impact analysis of the cascade system. The method applies to the detailed design stage of civil aircraft systems. It can reduce the difficulty of the safety analysis process and improve the accuracy of the analysis results. The verification of a civil aircraft primary flight control system design shows that after using the safety design method, the time cost of the analysis process is significantly reduced, and the analysis results are more accurate, thereby reducing the difficulty of design changes.

DESIGN AND DEVELOPMENT OF SURFACE PRIMARY FLIGHT CONTROL TOOLS FOR CESSNA 172 AS LEARNING MEDIA IN POLYTECHNIC AVIATION ACADEMIC SURABAYA

The purpose of designing this final project is for teaching aids and modeling the workings of Primary Flight Control in the form of simulations that make it easier for cadets and in learning about Primary Flight Control, especially in Primary Flight Control props which demonstrate Primary Flight Control and can be visualized with 3 axis axis movement and portable. With the method of visualizing the 3 axis axis movement in real terms, this makes cadets understand more about the movement of the aircraft from the movement of the axis directly in the movement by Radio Control. This research uses several methods, namely data collection, problem identification. This props uses wood, polyfoam, wire. The technique of making tools to assemble the aircraft is to use gluing techniques and for the stand using nailing techniques and adding glue to strengthen the structure of the props stand. The design which has the aim as a learning media support tool for cadets at the Surabaya Aviation Polytechnic campus, especially in basic Primary Flight Control learning. The results of making optimal props are then made in the form of a prototype as an analysis material for comparison between the simulation of design work tools carried out in the Radio Control-based Primary Flight Control design process. The system of the trainer uses Radio control as the main controller and tools supported by other components such as Servo, Esc, Battery, and which can control Primary Flight Control.

In-Flight Monitoring of a Primary Flight Control Electromechanical Actuator⋆

We aim at designing a monitoring algorithm able to perform early detection of degradation that may lead to jamming in the electromechanical actuators (EMAs) used for primary flight control. The challenge is to construct a detection algorithm performing an online monitoring using in-flight data. As the EMAs only move during specific phases of flight, the data subsets containing relevant information for the identification of at least some friction parameters must be determined. This is achieved by using tools from sensitivity analysis. The resulting data analysis is used to govern the parameter estimation part of a Dual Extended Kalman Filter (DEKF) in such a way that all informative data are exploited at best for friction estimation. The method is validated on healthy and faulty synthetic data resulting from a detailed EMA simulator.

Effect of engine failure on aerodynamic characteristics of a light transport aircraft model

In-flight engine failure has always been a hazardous situation in case of a single-engine aircraft associated with losing altitude and making a landing within the available gliding range. In the event of a multi-engine aircraft, this situation improves markedly due to the potential flight continuation. The necessary conditions to continue a flight are available excess thrust, satisfactory aerodynamic performance, and a fixed-wing control. If the half of engine thrust is lost, which adversely affects a rate of climb, the most critical flight modes are the take-off and go-around procedures performed at low altitudes using the engine asymmetric thrust. Ensuring flight safety in these modes requires extensive experimental studies in wind tunnels to simulate the basic flight envelope with an inoperative engine. The aim of studies is the effect of a critical engine failure on the aerodynamiccharacteristics of an aircraft model as well as ensuring the required efficiency of the flight controls to dampen the yaw and roll moments that arise during an asymmetric thrust flight. The complexity of solving the problem is determined by the necessity of recovering from the substantial yaw, roll moments and by a significant decrease in the wing lifting efficiency along with an increase in a drag force which limits a climb rate and aircraft control. This article presents an analysis of the effect of critical engine failure on the aerodynamic characteristics of a light twin-engine transport aircraft model in the wing take-off and landing configurations. The aircraft aerodynamic configuration is made according to the classic pattern with the high-mounted tapered wing and deck-type empennage. The high lift devices comprise a double-slot hinged flap with a fixed deflector. The aircraft is equipped with a loading ramp with a relatively short flat rotatable part of the lower fuselage surface. Experimental studies of the longitudinal and lateral characteristics of the model with installed simulators of a power plant were carried out in TsAGI low-speed wind tunnel T-102. The analysis of engine failure effect on the model aerodynamics was executed in changing the load factor within the range B = 0.3…2. The capabilities to dampen the yaw and roll moments, using the primary flight controls (rudder and ailerons), were determined.

Preliminary Sizing of the Electrical Motor and Housing of Electromechanical Actuators Applied on the Primary Flight Control System of Unmanned Helicopters

Helicopter dronization is expanding, for example, the VSR700 project. This leads to the integration of electromechanical actuators (EMAs) into the primary flight control system (PFCS). The PFCS is in charge of controlling the helicopter flight over its four axes (roll, pitch, yaw, and vertical). It controls the blade pitch thanks to mechanical kinematics and actuators. For more than 60 years, the actuators have been conventionally using the hydraulic technology. The EMA technology introduction involves the reconsideration of the design practices. Indeed, an EMA is multidisciplinary. Each of its components introduces new design drivers and new inherent technological imperfections (friction, inertia, and losses). This paper presents a methodology to specify and pre-design critical EMAs. The description will be focused on two components: the electrical motor and the housing. This includes a data-driven specification, scaling laws for motor losses estimation, and surrogate modeling for the housing vibratory sizing. The tools are finally applied to two study cases. The first case considers two potential redundant topologies of actuation. The housing sizing shows that one prevails on the other. The second case considers the actuators of helicopter rotors. The electrical motor sizing highlights the importance of designing two separate actuators.

An Improved Fault Identification Method for Electromechanical Actuators

Adoption of electromechanical actuation systems in aerospace is increasing, and so reliable diagnostic and prognostics schemes are required to ensure safe operations, especially in key, safety-critical systems such as primary flight controls. Furthermore, the use of prognostics methods can increase the system availability during the life cycle and thus reduce costs if implemented in a predictive maintenance framework. In this work, an improvement of an already presented algorithm will be introduced, whose scope is to predict the actual degradation state of a motor in an electromechanical actuator, also providing a temperature estimation. This objective is achieved by using a properly processed back-electromotive force signal and a simple feed-forward neural network. Good prediction of the motor health status is achieved with a small degree of inaccuracy.

Technology selection for holistic analysis of hybrid-electric commuter aircraft

Electric powertrains have different characteristics than conventional powertrains with combustion engines and require unconventional aircraft designs to evolve their full potential. Therefore, this paper describes a method to identify potential aircraft designs with electrified powertrains. Promising technology options in the fields of powertrain architecture, aerodynamic interactions, onboard systems and operating strategies were collected by the project partners of the LuFo project GNOSIS. The effect of the technology options on a commuter aircraft was evaluated in terms of global emissions (hbox {CO}_{2}), local emissions (text {NO}_{text {X}} and noise) and operating costs. The evaluation considers an entry into service in 2025 and 2050 and is based on the reference aircraft Beechcraft 1900D. Literature review and simplified calculations enabled the evaluation of the aerodynamic interactions, systems and operating strategies. A preliminary aircraft design tool assessed the different powertrain architectures by introducing the two parameters ’power hybridization’ and ’power split’. Afterwards, compatible technology options were compiled into technology baskets and ranked using the shortest euclidean distance to the ideal solution and the farthest euclidean distance to the worst solution (Technique for Order Preference by Similarity to Ideal Solution (TOPSIS) method). An analysis of the CS 23 regulations leads to a high-wing design and excluded the partial turbo-electric powertrain architecture with the gas turbine in the aircraft tail. For 2025, a partial turbo-electric powertrain with two additional electric driven wingtip propellers was selected. A serial hybrid powertrain, which uses a gas turbine or fuel cell in combination with a battery, powers distributed electric propulsors at the wing leading edge in 2050. In both scenarios, the aircraft design includes an electric environmental control system, an electric driven landing gear and electro-hydraulic actuators for the primary flight control and landing gear.

PFCS Four-redundant Sidestick Randomly Dynamically Grouping Vote Technology

For civil aircraft, the reliability of control device sensors in flight control system is very important for aircraft control and safe flight. Redundancy signal voting and monitoring technology can improve the availability and integrity of sensor signals in fly by wire primary flight control system of civil aircraft. In this paper, a simplified four-redundant signal voting and monitoring method is proposed which randomly and dynamically divides the four redundant signals into two groups for comparison. The safety analysis and simulation results reveal that this simplified method meets the safety requirements and has advantages in robustness and availability.

Research and modeling of force fighting equalisation for aircraft rudder’s triple active actuation system

AbstractThis paper presents a new approach to force fighting equalisation in a redundant active-active-active rudder actuation system that is used for the primary flight control system of a turboprop regional aircraft. The related coupled problem of force fighting scenario, and the hydraulic architecture of electronic-hydrostatic actuator (EHA) are analysed, the mathematical model of the EHA system is built. The virtual test bench is designed to evaluate the performance of the force fighting equalisation strategy. The proposed methodology is tested on an iron bird test rig. The physical experiment shows that the fighting force is minimised under all flight conditions, meets the low cost requirement and can be a very reliable system. The proposed methodology can be applied to other types of aircraft’ flight actuation systems.

Structural Health Monitoring of Electromechanical Actuators in Aviation—Challenges Ahead and Case Study

Abstract Electrical actuation concept fits perfectly with the worldwide strategy of more electric aircraft to reduce carbon footprint. However, the integration of linear electro-mechanical actuators is promising yet challenging in safety critical systems due to the jamming of the driven load. That fault is a critical mechanical transmission failure giving rise to concern in many applications such as primary flight controls or landing gears extension and steering. This article critically reviews electric actuation solutions currently available for aerospace application, the limits for their upcoming deployment and the different solutions to achieve an on-condition maintenance to reduce any safety risk during lifetime. A particular attention is given to the jamming and possible strategies to avoid any hazard induced by this failure. The most promising approach relies on the establishment of a predictive maintenance by monitoring continuously the actuator to timely warn any structural alteration, which is prone to induce jamming. Finally, a case study is presented, showing the ability of engine current monitoring to warn any alteration in the load maneuvering. The current spectrum analysis is found to be sensitive to both mechanical fault and control failure or malfunctioning. On top of that, no addition sensors are needed, making this technique affordable and cost-effective.

Unmanned helicopter flight control actuator specification through mission profile analysis

Helicopter dronization is expanding, as for example with the VSR700 project, and leads to the design and the integration of electromechanical actuators (EMA) into the primary flight control system (PFCS). The PFCS is in charge of controlling the helicopter flight over its 4 axis (roll, pitch, yaw, vertical). It controls the blade pitch through dedicated mechanical kinematics and actuators. The hydraulic technology has been conventionally used in actuators for more than 60 years. On the other hand, the introduction of the EMA technology requires the reconsideration of design practices right at development start. Indeed, the establishment and synthesis of the specification need to deal with – new design drivers (high performance points, wear, fatigue) and - new inherent technological imperfections (friction, inertia and reduction ratio). To address these topics, this paper draws a list of the main EMA design drivers to focus on along with a brief description of the main EMA components. Then, it proposes indicators evaluated over a complete mission profile in time coming from measurement on a given applicative helicopter flight. These indicators are chosen and elaborated to provide an image of the design drivers responsible for rapid and gradual degradations of the actuator components. Also, they give an idea of the importance taken by the actuator imperfections into the global performance. Furthermore, we explain how mission profiles are processed depending of data sources. Finally, through a comparison with a standard aircraft mission profile, we emphasize the specificity of the helicopter application use case.

Validation of Aero-Load Estimator for Convair 880 Aircraft Flight Simulator Benchmark with Electro-Hydrostatic Actuators

Simulating an aircraft model using of high fidelity models of subsystems for its primary and secondary flight control actuators requires measuring or estimating aero-load data acting on flight control surfaces. One solution would be to incorporate the data recorded from flight tests, which is a time-consuming and costly process. This paper proposes another solution based on the validation of an aero-loads estimator or on the hinge moments predictor for fully electrical aircraft simulator benchmark. This estimator is based on an aerodynamic coefficient calculation methodology, inspired by Roskam’s method that uses the geometrical data of the wing and control surfaces airfoils. The hinge moment values are found from two-dimensional lookup tables where the deflections of the control surfaces, aircraft altitude, and aircraft angles of attack are the input vectors of the tables; and the resulting hinge moment coefficients are the output vectors. The resulting hinge moment coefficients of the Convair 880 primary flight control surfaces are compared to those of its recorded flight test data; the results from the new software solution were found to be very accurate. Hinge moment lookup tables are integrated in the Convair 880 high fidelity flight simulation benchmark using mathematical models of energy-efficient Electro-Hydrostatic Actuators (EHA). Autopilot controls are designed for the roll, pitch, attitude and yaw damper motions using Proportional Integral (PI) controller scheduled for different flight conditions. Several different aircraft simulation scenarios are evaluated to demonstrate the efficacy and accuracy of the predicted hinge moment results.

A Case Study on the Detection and Prognosis of Internal Leakages in Electro-Hydraulic Flight Control Actuators

Electro-hydraulic servo-actuators (EHSAs) are currently considered the state-of-the art solution for the control of the primary flight control systems of civil and military aircraft. Combining the expected service life of a commercial aircraft with the fact that electro-hydraulic technology is employed in the vast majority of currently in-service aircraft and is planned to be used on future platforms as well, the development of an effective Prognostic and Health Management (PHM) system could provide significant advantages to fleet operators and aircraft maintenance, such as the reduction of unplanned flight disruptions and increased availability of the aircraft. The occurrence of excessive internal leakage within the EHSAs is one of the most common causes of return from the field of flight control actuators, making this failure mode a priority in the definition of any dedicated PHM routine. This paper presents a case study on the design of a prognostic system for this degradation mode, in the context of a wider effort toward the definition of a prognostic framework suitable to work on in-flight data. The study is performed by means of a high-fidelity simulation model supported by experimental activities. Results of both the simulation and the experimental work are used to select a suitable feature, then implemented within the prognostic framework based on particle filtering. The algorithm is at first theoretically discussed, and then tested against several degradation patterns. Performances are evaluated through state-of-the-art metrics, showing promising results and providing the basis towards future applications on real in-flight data.

Collaborative Robotics: Enhance Maintenance Procedures on Primary Flight Control Servo-Actuators

Electro-Hydraulic Servo-Actuators (EHSAs) are mainly used to command primary flight control surfaces in military and commercial aircraft. Since these devices are crucial for vehicle stability and maneuverability, a correct assessment of their health status is mandatory. Within this framework, a joint research project (HyDiag), held by Politecnico di Torino and Lufthansa Technik AG (LHT), aims to provide a more efficient and reliable procedure to determine the operating conditions of the EHSA. A smart and automatic sequence, able to extract several health features of the Unit Under Test (UUT), has been developed and integrated. The present paper discusses the implementation of a collaborative robot, equipped with a vision system and customized tools, for both health features extraction, and maintenance tasks on unserviceable servo-actuators. The main challenges related to the automation of such complex tasks in a real working environment are highlighted, togetherwith the advantages brought by the proposed approach. The paper also presents the first results of an ongoing experimental campaign. Specifically, it reports the enhancements of the maintenance procedures using collaborative robotics and possible future developments.

Results of a Feasibility Study of a Prognostic System for Electro-Hydraulic Flight Control Actuators

Electro-Hydraulic Servo-Actuators (EHSA) are currently the most used actuation technology for primary flight control systems of civil and military aircrafts. Although some alternatives have emerged in the last decade, such as electromechanical or electro-hydrostatic solutions, electrohydraulic systems are still considered the most effective technology in flight-critical application of new commercial aircrafts. Moreover, the vast majority of aircraft currently in service are equipped with this technology. Considering the number of actuators typically employed in a primary flight control system and the expected service life of a commercial aircraft, the development of an effective PHM system could provide significant benefits to fleet operators and aircraft maintenance. This paper presents the results of a feasibility study of such a system for electro-hydraulic actuators used in fly-by-wire primary flight control systems, considering the actuator of a wide body commercial aircraft as use case. Aim of the research is the implementation of a PHM system without the addition of dedicated sensors, solution which would allow for the application of the proposed prognostic solution on both new and existing platforms. This paper describes the methodology and the results of the feasibility study through simulation and experimental activities, which shows how the novel PHM technologies proposed for a PHM system for the EHSAs of primary flight control actuators can allow the migration from scheduled to condition-based maintenance.

Electrical Machines for the More Electric Aircraft: Partial Discharges Investigation

Modern electrical machines employed in transportation applications are required to provide high performance in terms of power (and torque) density. At the same time, being these applications safety-critical, a significant level of reliability and/or fault tolerance is expected. Among all the factors which can compromise motors reliability, partial discharges (PDs) inception is one of the most crucial, in particular for low voltage, random wound machines. This article presents an extensive experimental investigation on PDs in electrical machines for aerospace applications. Measurements are carried out using both sinusoidal and fast-rising pulses in a representative aerospace environment, emulating the typical ambient conditions encountered throughout a commercial aircraft mission, and beyond (i.e., down to 30 mbar). As a main result of the investigation, it is proved that electrical machines employed for actuating primary flight control surfaces feature a higher risk of PDs inception. Therefore, their insulation system demands an extremely careful design.

Windings Fault Detection and Prognosis in Electro-Mechanical Flight Control Actuators Operating in Active-Active Configuration

One of the most significant research trends in the last decades of the aeronautic industry is the effort to move towards the design and the production of “more electric aircraft”. Within this framework, the application of the electrical technology to flight control systems has seen a progressive, although slow, increase: starting with the introduction of fly-by-wire and proceeding with the partial replacement of the traditional hydraulic/electro-hydraulic actuators with purely electro-mechanical ones. This evolution allowed to obtain more flexible solutions, reduced installation issues and enhanced aircraft control capability. Electro-Mechanical Actuators (EMAs) are however far from being a mature technology and still suffer from several safety issues, which can be partially limited by increasing the complexity of their design and hence their production costs. The development of a robust Prognostics and Health Management (PHM) system could provide a way to prevent the occurrence of a critical failure without resorting to complex device design. This paper deals with the first part of the study of a comprehensive PHM system for EMAs employed as primary flight control actuators; the peculiarities of the application are presented and discussed, while a novel approach, based on short pre-flight/post-flight health monitoring tests, is proposed. Turn-to-turn short in the electric motor windings is identified as the most common electrical degradation and a particle filtering framework for anomaly detection and prognosis featuring a self-tuning non-linear model is proposed. Features, anomaly detection and a prognostic algorithm are hence evaluated through state-of-the art performance metrics and their results discussed.

Proposal of a simplified Coulomb friction numerical model for the preliminary design of electrohydraulic servomechanisms

Electrohydraulic servomechanisms (EHAs) are particularly interesting for aviation application, in fact tanks to the high power to weight ration are widely diffused in medium to large cargo and passengers planes or fighters. This work is focused on the proposal of a new dry friction numerical algorithm, based upon Coulomb’s approach, which can be integrated into simulation algorithms obtained by degrading the systems dynamic models (e.g. an over-damped second-order system reducible to a simpler first-order one). This approach, if correctly applied, significantly reduces the computational burden, without significant losses in simulation accuracy. The authors evaluated the approach proposed by a numerical test bench simulating the behaviour of an electrohydraulic linear actuator commonly used in primary flight controls.