Primary Flight Control Research Articles

Stability analysis for a nonlinear model of a hydraulic servomechanism in a servoelastic framework

The effects of mounting structure stiffness on mechano-hydraulic servomechanisms actuating aircraft primary flight controls are modeled by a six-dimensional nonlinear system of ordinary differential equations. Stability analysis of equilibria reveals the presence of a critical case that is handled through the use of the Lyapunov–Malkin theorem. Stability charts are drawn using the Routh–Hurwitz criterion for the stability of a fifth-degree polynomial. Comparison with previous results shows how the stability of equilibria can be ensured exploiting the positive influence of structural feedback.

Real-time loading actuator control for an advanced aerospace test rig

The growing complexity of aircraft systems requires that extensive testing be performed on dedicated test rigs to grant the necessary flight clearance and to provide troubleshooting of the problems that may arise during prototype flight trials. A new test rig is described herein, that was recently built to provide comprehensive testing of the flight controls, landing gear, and hydraulic systems of the M-346 advanced trainer aircraft. All these subsystems are connected together on the rig as in the real aircraft to ensure that the entire system operates on the rig exactly as on the aircraft. To that purpose, a high-performance force control system was developed to provide accurate dynamic loading of the primary and secondary flight control actuators, and the structure onto which the flight control actuators are mounted provides a mechanical load path with the same stiffness as in the real aircraft. The present paper describes the main characteristics of the test rig and puts particular emphasis on the force control system that makes use of an adaptive control law to provide a very accurate control of the force acting upon the actuators during fast transients.

Worthy test programmes and developments of smart electromechanical actuators

Early aircraft flight control systems were totally manually operated, that is, the forcerequired to move flight control surfaces was generated by the pilot and transmitted bycables and rods. As aerodynamics and airframe technology developed and speeds increased,the forces required to move control surfaces increased, as did the number of surfaces. Inorder to provide the extra power required, hydraulic technology was introduced.To date, the common element in the development of flight control systems hasbeen, mainly, restricted to this type of technology. This is because of its provenreliability and the lack of alternative technologies. However, the technology to buildelectromechanically actuated primary flight control systems is now available.Motors developing the required power at the required frequencies are now possible(with the use of high energy permanent magnetic materials and compact highspeed electronic circuits). It is this particular development which may make theconcept of an ‘all electric aircraft’ realizable in the near future. The purposeof the all electric aircraft concept is the consolidation of all secondary powersystems into electric power. The elimination of hydraulic and pneumatic secondarypower systems will improve maintainability, flight readiness and use of energy.This paper will present the development of multi-lane smart electric actuators and offer aninsight into other subsequent fields of study. The key areas of study may be categorized asfollows.State of the art hydraulic actuators.Electromechanical actuator system test programmes.Development of electromechanical actuators.Modelling of electromechanical actuators.

New material extends seal life in aerospace applications

The newly engineered Turcon M30 non-abrasive material from Busak+Shamban has been shown to improve seal performance against tungsten carbide coated moving parts in primary flight controls and utility actuation systems.

제어법칙 개발을 위한 소프트웨어 전환장치 설계에 관한 연구

Relaxed Static Stability(RSS) concept has been applied to improve aerodynamic performance of modern version supersonic jet fighter aircraft. Therefore, the flight control systems are necessary to stabilizes the unstable aircraft and provides adequate handling qualities. The initial production flight control system are verified by flight test and it's always an elements of danger because of flight-critical nature of control law function and design error due to model base design method. These critical issues impact to flight safety, and it could be lead to a loss of aircraft and pilot's life. Therefore, development of an easily modifiable RFCS(Research Flight Control System) capable of reverting to a PFCS(Primary Flight Control System) of reliable control law must be developed to guarantee the flight safety. This paper addresses the concept of SSWM(Software Switching Mechanism) using the fader logic such as TFS(Transient Free Switch) based on T-50 flight control law. The result of the analysis based on non-real time simulation in-house software using SSWM reveals that the flight control system are switching between two computers without any problem.

Concepts for position and load control for hybrid actuation in primary flight controls

Moving towards the More Electric Aircraft, a hybrid actuator configuration, in which an electromechanical actuator (EMA) and an electrohydraulic servoactuator (EHSA) operate on the same control surface, provides an opportunity to introduce electromechanical actuators into primary flight controls. Besides the operation in active/passive or active/active mode, an “active/no-load” mode is promising. In this mode the EMA is controlled such that it actively follows the movement of the control surface without carrying external air loads, thereby reducing power dissipation compared to active/active mode and failure transients compared to active/passive mode. However, force fighting will occur if both actuators are actively controlled. In this paper, control concepts for a hybrid configuration, extending the original actuator control loops, are presented, enabling active/active as well as active/no-load operation. Nonlinear as well as linear models for an EMA, an EHSA, and a control surface structure are derived from technical data for an airworthy EHSA and combined to a model of the hybrid configuration. These models are used for matching of actuator dynamics and simulation of the developed control laws. For active/active mode, maximum force fighting between the actuators is reduced from about 500% to 7% of the stall load. For active/no-load mode, a force control loop is added to the EMA control, causing the EMA to follow movements of the control surface such that the external loads on the EMA are zero in steady-state. Force fighting is reduced to 30% of the stall load.

Swashplateless Helicopter Rotor System with Trailing-Edge Flaps for Flight and Vibration Controls

The objective of this study is to demonstrate the concept of active trailing-edge flaps as primary flight control and vibration reduction devices for a typical full-scale helicopter. A comprehensive rotorcraft analysis based on UMARC was developed to analyze the swashplateless rotor. A parametric study of various key design variables involved in the trailing-edge flap design was carried out. An optimal design of a trailing-edge flap system that provides effective control authority within the complete range of advance ratios as well as minimum actuation requirements was achieved. Trailing-edge flaps demonstrated the capability of performing both primary flight control and active vibration control functions. At a high forward speed (advance ratio of 0.32), the 4/rev vertical force and roll and pitch moments at hub are successfully eliminated (by 90%), and the 4/rev in-plane hub forces are reduced by more than 40%. The half peak-to-peak value of the trailing-edge flap deflection for primary flight control is 7.1 deg, and an additional 4.7 deg is required for active vibration control.

Flight Control System Design for Propulsion-Controlled Aircraft

The loss of an aircraft's primary flight controls can lead to a fatal accident. However, if the engine thrust is available, controllability and safety can be retained. This article describes flight control using engine thrust only when an aircraft has lost all primary flight controls. This is a kind of flight control reconfiguration. For safe return, the aircraft must first descend to a landing area, decelerate to a landing speed, and then be capable of precise flight control for approach and landing. For these purposes, two kinds of controllers are required: a controller for descent and deceleration and a controller for approach and landing. The former controller is designed for longitudinal motion using a model-following control method, based on a linear quadratic regulator. The latter is designed by an H∞ state-feedback control method for both longitudinal and lateral-directional motions. Computer simulation is conducted using linear models of the Boeing 747. The results indicate that flight path control, including approach and landing, is possible using thrust only; however, speed control proves more difficult. However, if the horizontal stabilizer is available, the airspeed can be reduced to a safe landing speed.

German contribution to the X-38 CRV demonstrator in the field of guidance, navigation and control (GNC)

In the frame of the ESA/NASA cooperation on the X-38 project, different GNC-related contributions have been made by German industry and universities. First, the primary flight control software for the autonomous guidance and control of the X-38 parafoil descent and landing phase has been developed, integrated and successfully flown during the aerial drop test campaign conducted by NASA. In addition, a fault-tolerant computer similar to the one used onboard the ISS has been delivered to JSC. Together with an alternate re-entry GNC software using onboard flight path optimization for the guidance task and dynamic inversion methods for attitude control, this computer shall be flown as a flight experiment onboard the V201 space flight test vehicle. Finally, the German project team provided a real-time X-38 vehicle simulator, which was supposed to be used as an independent validation tool for the X-38 re-entry simulation and onboard software. This paper will focus on the European parafoil guidance and control software across the different phases of the X-38 mission. Flight test results from the X-38 aerial drop test campaigns will be presented and discussed. In addition, the flight experiment of the fault tolerant computer will be described briefly.

Design criteria for conceptual sizing of primary flight controls

AbstractThe European Supersonic Commercial Transport’s control surface configuration is based largely on Concorde’s and has been scaled to provide comparable un-augmented stability and manoeuvre performance. Hence, optimising the surface size could provide significant performance benefits in terms of reduced drag, noise, structural and actuator power requirements. Adequate control power will be required to meet current civil aviation regulations whose primary aim is to ensure the aircraft can be flown safely during both normal and emergency operation. Additional design criteria, combined with the optimum longitudinal control laws, are required to ensure desirable handling qualities with minimum pilot workload. Two critical low-speed flight conditions, normal and emergency, together with associated aircraft configurations for control surface sizing have been identified. The rudder must provide sufficient control power to achieve positive heading changes subsequent to a double asymmetric engine failure during normal operation. The fin should be sized to satisfy Dutch roll stability criteria with the un-augmented aircraft in its emergency configuration. The dual functionality of the elevons require that they are sized using both pitch and roll performance and handling quality criteria. The bank angle capture requirement provides the most critical elevon design case, the satisfaction of which also ensures adequate pitch control power. Validation using ‘pilot-in-the-loop’ simulation will be required whilst more explicit control surface size optimisation would require the definition of limiting airspeeds and operating conditions applicable to the European Supersonic Commercial Transport. Additional studies of control power requirements during transonic and supersonic cruise may also be required.

Optimum HVOF surface finish recommendation for high-performance dynamic seal applications

Testing at the product test laboratory in Kulpsville, Pennsylvania has helped Greene, Tweed, in conjunction with several partners in the aerospace industry, to develop a recommendation for the optimum surface finish for high-velocity oxygen fuel (HVOF) coatings for use in high-performance dynamic seal applications. Testing included the evaluation of surface finish parameters that define the surface profile topography, and the influence these parameters have on the performance of dynamic seal components in critical aerospace applications such as landing gear and primary flight controls.

Enhanced eigenstructure assignment for aeroelastic control application

A sophisticated robust controller design method for regulation of weakly damped aeroelastic structures using the primary flight controls (PFC) is presented. The approach is based on eigenvalue eigenvector assignment, and is capable of taking into account potentially competing objective specifications in either time or frequency domains, as well as parametric constraints. The infinite set of flight conditions and fuel load distributions lead us to represent the underlying aeroservoelastic plant as a linear multivariable multi-model system, for which a robust output feedback controller is sought in line with technical feasibility, parametric and dynamic uncertainty, architectural and sensor layout constraints. While the method of eigenvalue eigenvector assignment is well known in the fields of flight control design and modal decoupling, here the primary objective is to apply it to constrained simultaneous robust eigenvalue stabilisation. Since the intrinsic controller calculation is implemented as an multiobjective optimisation problem, aeroservoelastic performance specifications can be specified in the problem formulation. The suitability of the resulting low-order static output controller is demonstrated via real-time simulation of a real PFC hardware-in-the loop actuation system, both for normal and degraded configurations.

Enhanced eigenstructure assignment for aeroelastic control application

A sophisticated robust controller design method for regulation of weakly damped aeroelastic structures using the primary flight controls (PFC) is presented. The approach is based on eigenvalue eigenvector assignment, and is capable of taking into account potentially competing objective specifications in either time or frequency domains, as well as parametric constraints. The infinite set of flight conditions and fuel load distributions lead us to represent the underlying aeroservoelastic plant as a linear multivariable multi-model system, for which a robust output feedback controller is sought in line with technical feasibility, parametric and dynamic uncertainty, architectural and sensor layout constraints. While the method of eigenvalue eigenvector assignment is well known in the fields of flight control design and modal decoupling, here the primary objective is to apply it to constrained simultaneous robust eigenvalue stabilisation. Since the intrinsic controller calculation is implemented as an multiobjective optimisation problem, aeroservoelastic performance specifications can be specified in the problem formulation. The suitability of the resulting low-order static output controller is demonstrated via real-time simulation of a real PFC hardware-in-the loop actuation system, both for normal and degraded configurations.

Neuro‐fuzzy synthesis of flight control electrohydraulic servo

Presents a switching type neuro‐fuzzy control synthesis. The control algorithm supposes as a component part a neurocontrol designed to optimize a performance index. Whenever the neurocontrol saturates or a certain performance parameter of the system decreases, the scheme of control switches to a feasible and reliable fuzzy logic control. Describes the procedure of return to the optimizing neurocontrol which is essential. This methodology of control synthesis ensures antisaturating, antichattering and robustness properties of the controlling system, as illustrated by numerical simulation in the case of a primary flight controls electrohydraulic servo actuator

Robust Synthesis With Antiwindup Compensation for Electrohydraulic Servomechanism

The robust control of a general servomechanism problem is applied in the case of electrohydraulic servo actuating primary flight controls. The servocompensator structure is proved to be close to the one designed for step signals. The stabilizing compensator is ensured by the way of a linear quadratic optimal procedure. An antiwindup compensation is added to deal with the adverse effects caused by control saturation. The proof is made involving the form of a feedback compensation matrix. Numerical simulations are presented to validate the proposed procedure.

Magnetostrictively actuated control flaps for vibration reduction in helicopter rotors—design considerations for implementation

Vibration reduction using blade root pitch control introduces a significant power penalty and may adversely affect the airworthiness of the flight control system. Comparable levels of vibration reduction can be achieved using considerably less power through an actively controlled trailing edge flap mounted on the blade. Such a device would have no effect on helicopter airworthiness since it is controlled by a loop separate from the primary flight control system which utilizes the swashplate. Control flap actuation using the magnetostrictive material Terfenol-D is studied in this paper by designing a minimum weight actuator, subject to a set of actuation and stress constraints. An important ingredient in this design process is the precise characterization of the constitutive relations for the magnetostrictive material which provides valuable guidelines on the practical implementation of magnetostrictive actuators. It is shown that the magnetostrictively actuated flap is capable of producing vibration reduction in high speed forward flight in excess of 90%.

9.2 Propulsion Control of Aircraft: A Case Study in Systems Engineering

ABSTRACTOn July 19, 1989 United Airlines Flight 232 was cruising at approximately 37,000 feet between Denver and Minneapolis when it experienced a catastrophic failure of the number two (tail‐mounted) engine. At the moment of failure the engine ejected shrapnel through the tail section, destroying all three on‐board hydraulic systems. The complete loss of hydraulics left the pilots of the McDonnell Douglas DC‐10 aircraft without the ability to move the flight controls. In contrast to what most aviation experts would have expected as an outcome, the airplane was crash‐landed at Sioux City, Iowa with 185 of the 296 passengers and crew surviving. The flight crew was able to maintain rudimentary control of the aircraft with only the use of thrust from the two remaining wing‐mounted engines by independently moving their respective power levers. It was not until the final touchdown that the crew's command of the aircraft was lost due to the imprecise nature of manual propulsion control (Propulsion Controlled Aircraft [PCA], 1999).Following the accident, the National Transportation Safety Board (NTSB) recommended that Propulsion Control of Aircraft (PCA) be evaluated as a “pilot friendly” backup system using engine thrust as the only actuating mechanism in the case of primary flight control failure. This paper examines the use of systems engineering to effectively decompose the problem, identify the solution set, and integrate the solution with the potential for improved safety and reduced life cycle costs. An overview of aircraft flight dynamics and primary control systems is provided in order to understand the basic system properties. Analysis includes physical systems as well as software systems and their interface. Finally, the use of simulation to better understand system behavior and mitigate risk while arriving at a system‐based solution is examined (Blanchard, 1998). Control the aircraft in flight after a primary flight control failure and accomplish that with an interface which is intuitive to the pilot, that is the problem. The solution could only have been reached through systems engineering.

An integrated framework for development of real-time distributed control software based on can-bus

Embedded systems are computer-based systems which must respond to external stimuli within time scales determined by the external environment. Such systems are required to achieve ever more demanding behavioural. performance and safety requirements. Embedded systems are found in applications such as primary flight control, gas-turbine engine control and railway traffic managementOften complex embedded systems are distributed, typically to achieve demanding performance or dependability requirements, and must operate within hard real-time constraints. Considerable effort is required to select optimal design solutions and ensure adherence to specified requirements. In embedded systems safety, reliability and response-times are considered as strict constraints on the system design. Achievement of these constraints requires meticulous analysis, typically through the use of specialist modelling and simulation techniques. Integration of knowledge and understanding obtained from disparate detailed models remains a considerable challenge.Ideally, a system which has both continuous dynamics and event-driven parts. would be modelled in one environment, using the most appropriate techniques for each part, and allowing the analysis of the whole system to be carried out simultaneously.The paper presents an integrated approach in order to translate automatically the information manipulated during the different phases of the design: specification. analysis, design and implementation.

From robust control to antiwindup compensation of electrohydraulic servo actuators

A method of designing the controller to solve the robust servomechanism problem is applied in the case of an electrohydraulic servo actuating primary flight control. This method is based on the well‐known solution consisting of two separate devices: a servocompensator, in fact an internal model of the exogenous dynamics, including the reference commands and disturbance signals; and a stabilizing compensator. The proof is made involving the servocompensator structure which is close to the one designed for step signals. The stabilizing compensator is assured by way of a linear quadratic optimal procedure. An antiwindup compensation is added to deal with the adverse effects caused by actuator saturation.

Simulation of a prototype electrically powered integrated actuator for civil aircraft

Developments in the design and proposed operation of large civil aircraft have resulted in aircraft manufacturers and equipment suppliers developing new system concepts, one of which is the all or more electric aircraft. In the all or more electric aircraft the distribution of power for flight actuation will be through the electrical system, as opposed to the currently used bulk hydraulic system. In order to implement power-by-wire, high-performance electrically powered actuators will be required. The paper discusses the design details, and the simulation of an electrohydrostatic actuator suitable for use in primary flight control systems of a civil aircraft. The paper presents experimental and simulation results, and identifies the parameters that will critically affect the performance of an actuator.