Control Surface Position Research Articles

PWM Controlled Servomechanism

Precise control of aircraft flight control surfaces by the pilot is critical to flight safety. The joystick and control surface positions must be detected accurately and quickly, and the motion mechanism must be operated in such a way as to ensure synchronization between them. The analogue voltage value taken from the potentiometer, which is used as a position sensor in many small aircraft, can be used directly or by converting to digital. 
 In this applied study, the analogue position knowledge received from the potentiometer on the joystick was converted into a PWM signal with an oscillator circuit. The PWM position signal applied to the IR (InfraRed) LED in the transmitter unit is sent as infrared light. The PWM position knowledge detected by the photodiode in the receiver unit was passed through the LPF (Low Pass Filter) and the voltage value of the joystick potentiometer was obtained again. With the mechanism controlled according to the difference between the potentiometer values of the joystick and the control surface, the synchronization between the joystick and the control surface is ensured. In this study, a wireless and stable servomechanism connection is obtained by transmitting PWM position knowledge as IR.

Test rig dedicated for hardware used in wind tunnels

The paper present the results of work on measurement system dedicated to hardware used during wind tunnel tests, especially to servomechanisms. These devices could be applied to set specific position of control surface. Proposed system would ensure continuous monitoring of servo-rotor position and servo-motor temperature and would avoid uncontrolled change of control surface position. The application designed to monitor the operating status of the servomechanism was prepared in the LabVIEW software and was implemented on the myRIO platform. Developed test rig allow to register time histories of servo-rotor position and temperature during for different values of applied load. In the paper, test methodology were also presented. Experimental studies show that before wind tunnel tests, selected servomechanism should be tested in terms of maintaining the parameters declared by the manufacturer, especially during continuous operation. Developed measurement system can be used during wind tunnel tests, as well as only for servo-mechanism parameter testing.

Multiple-Input Multiple-Output Experimental Aeroelastic Control Using a Receptance-Based Method

This paper presents the first experimental study of multiple-input multiple-output (MIMO) active vibration suppression by pole placement using the receptance method. The research is based on a purpose-built modular flexible wing equipped with leading- and trailing-edge control surfaces and two displacement sensors for measuring its position. The MIMO controller has the advantage of being designed entirely on frequency response functions, which are measured between the control surface position (control inputs) and the structural displacements (outputs), and include the actuator dynamics. There is no requirement to evaluate or to know the structural , , and matrices or the aerodynamic loads; and the formulation eliminates the need for a state observer. The controller is first implemented numerically and then experimentally on the aeroelastic system. Both frequencies and damping are assigned (together or independently) for the first two vibration modes. The research includes a procedure for assessing the control effort required and demonstrates an effective means of increasing the flutter margin while the effort is minimized.

A longitudinal flight control law to accommodate sensor loss in the RECONFIGURE benchmark

The feedback gains in state-of-the-art flight control laws for commercial aircraft are scheduled as a function of values such as airspeed, mass, and centre of gravity (CoG). If measurements or estimates of these are lost due to multiple simultaneous sensor failures, the pilot must revert to an alternative control law, or, in the ultimate case, directly command control surface positions. This work develops a robust backup load-factor tracking control law, that does not depend on these parameters, based on application of theory from robust MPC and H2 optimal control. Firstly, the methods are applied with loss only of airdata, and subsequently also with loss of mass and CoG estimates. Local linear analysis indicates satisfactory performance over a wide range of operating points. To keep the aircraft within an acceptable operating region, an outer protection loop is implemented using an override approach, based on ground speed, a model of the trim angle of attack and variation of load factor with respect to angle of attack, and a priori bounds on the wind speed. Finally, the resulting control laws are demonstrated on the nonlinear RECONFIGURE benchmark, which is derived from an Airbus high fidelity, industrially-validated simulator.

Set-membership fault detection under noisy environment with application to the detection of abnormal aircraft control surface positions

The paper develops a set membership detection methodology which is applied to the detection of abnormal positions of aircraft control surfaces. Robust and early detection of such abnormal positions is an important issue for early system reconfiguration and overall optimisation of aircraft design. In order to improve fault sensitivity while ensuring a high level of robustness, the method combines a data-driven characterisation of noise and a model-driven approach based on interval prediction. The efficiency of the proposed methodology is illustrated through simulation results obtained based on data recorded in several flight scenarios of a highly representative aircraft benchmark.

A longitudinal flight control law based on robust MPC and H2 methods to accommodate sensor loss in the RECONFIGURE benchmark

The feedback gains in state-of-the-art flight control laws for commercial aircraft are scheduled as a function of values such as airspeed, mass, and centre of gravity (CoG). If estimates of these are lost due to multiple simultaneous sensor failures, the pilot must revert to an alternative control law, or, in the ultimate case, directly command control surface positions. This work develops a robust backup load-factor tracking control law, that does not depend on these parameters, based on application of theory from robust MPC and H2 control. Firstly, the methods are applied with loss only of airdata, and subsequently also with loss of mass and CoG estimates. Local linear analysis indicates satisfactory performance over a wide range of operating points. Finally, the resulting control laws are demonstrated on the nonlinear RECONFIGURE benchmark, which is derived from an Airbus high fidelity, industrially-validated simulator.

Data fusion and primary image processing for aircraft identification

The primary scope of this study lays on system technique solutions of collecting data required for the identification of an aircraft’s nonlinear dynamic model. It is assumed that the aircraft has no inbuilt navigational system, nor any sensors mounted on its control surfaces. The control column and pedals manipulated by the pilot can only visually be observed. For the time of data logging, an external sensory system (GPS, IMU) and a camera system were deployed on the airplane supporting the collection of flight data. The paper presents the data acquisition solutions required for aircraft’s nonlinear model identification, with an emphasis on the determination of the control surface positions as the system’s input signals using image processing. During flight, the control column and pedal positions manipulated by the pilot are recorded using a video camera and with post processing, data is converted to control surface (rudder, elevator, aileron) positions. The 3D positions of the pilot’s control column are determined from 2D pixel values. The input signals are then calculated using this information and the control surface characteristics. The input signals and state variables determined with a state estimator are regarded as input signals for the identification of an aircraft’s nonlinear model.

Fault Diagnosis and Fault-Tolerant Control Strategy for the Aerosonde UAV

A fault detection and diagnosis (FDD) and a fault-tolerant control (FTC) system for an unmanned aerial vehicle (UAV) subject to control surface failures are presented. This FDD/FTC technique is designed considering the following constraints: the control surface positions are not measured and some actuator faults are not isolable. Moreover, the aircraft has an unstable spiral mode and offers few actuator redundancies. Thus, to compensate for actuator faults, the healthy controls may move close to their saturation values and the aircraft may become uncontrollable; this is critical due to its open-loop unstability. A nonlinear aircraft model designed for FTC researches has been proposed. It describes the aerodynamic effects produced by each control surface. The diagnosis system is designed with a bank of unknown input decoupled functional observers (UIDFO) which is able to estimate unknown inputs. It is coupled with an active diagnosis method in order to isolate the faulty control. Once the fault is diagnosed, an FTC based on state feedback controllers aims at sizing the stability domain with respect to the flight envelope and actuator saturations while setting the dynamics of the closed-loop system. The complete system was demonstrated in simulation with a nonlinear model of the aircraft.

Dynamic gain-scheduled control of the ICE 101-TV

AbstractThis paper shows the theoretical development and application of dynamic gain scheduled control – a novel method for the control of nonlinear systems – to an aircraft model. The idea behind this method is to schedule the control law gains with a fast varying state variable rather than with a slow varying state or an input parameter. This is advantageous as it is then possible to schedule the gains with a state variable that is dominant in the mode that we are most interested in controlling. The use of this type of gain scheduling is shown to improve the transient response of the aircraft model when stepping between trim conditions and to overcome some of the problems associated with conventional gain scheduled controllers (such as control surface position limit saturation). ‘Hidden coupling terms’ that introduce unwanted dynamics when scheduling gains with a fast state (rather than the input design parameter) are eliminated directly by applying a transformation to the classical parameter-scheduled gain distributions which are calculated using eigenstructure assignment. A second order longitudinal model and a 5th order longitudinal/lateral model of the ICE 101-TV tailless delta-wing aircraft configuration are used to demonstrate the design process.

Alternate table look-up routine for real-time digital flight simulation

HE increase in complexity of flight simulators in the last decade has placed an increasingly larger computational burden on the simulation host computer. In the past, this has been dealt with by using separate processors for different tasks. This is not always financially feasible, especially in a university setting, where an entire simulation might be limited to one or two processors. While a perturbation model, sometimes used in a partial task trainer, can be regarded as a linear system, the total force and moment model of a real-time man-in-the-loop simulation requires a more complex mathematical representation of the vehicle. Coefficients and derivatives are often functions of several variables, including control surface position, angle of attack, Mach number, and several other parameters. These functions seldom can be represented in closed form by a reasonably small number of equations. Often a table is created that consists of an array of known dependent variable values, tabulated at certain known values of one or more independent variables. Table 1 is an example of such a table. On each pass through the simulation program, a search is initiated through the appropriate tables, and an interpolation between known values is performed in each of the independent variable directions. This has been shown to be one of the most computationally intense aspects of the simulation process.1 In order to retain model complexity and still achieve an acceptable frame rate, table look-up algorithms that reduce the computational load are highly desirable. Two widely used search procedures are outlined by Rolfe and Staples.2 The first, Method la, performs a top-down search between successive array values and exits once the proper interval is located. The second procedure, Method Ib, also performs a top-down search, but only if the position in the table has changed significantly. This is a method that has been used in industry.3 The search procedure is employed for each independent variable in a multidimensional table, $nd the process is repeated for each simulation pass. Once the proper location in the array is reached, an approximation of the dependent variable must be calculated. This dependent variable is considered to be a function/of n independent variables and is approximated, as F, by carrying out successive linear interpolations in each of the n directions as outlined in Refs. 2 and 3. Alternate algorithms presented in this paper reduce the table search time by performing a top-down search during only the initialization pass. On all subsequent passes, the algorithm returns to the interval where the value will most likely reside. This is accomplished by tracking the direction and the rate of

Control Law Synthesis for Flutter Suppression Using Linear Quadratic Gaussian Theory

This paper describes the application of linear quadratic Gaussian (LQG) methodology to the design of active control systems for suppression of aerodynamic flutter. A wind tunnel model of a research wing with associated sensors and actuators comprises the system to be controlled. Results of a synthesis methodology that provide small values of rms response and robust stability are presented. Results of control surface and sensor position optimization are also presented. Both frequency response matching and modal residualization are used to obtain practical flutter controllers. Scalars b c E( ) h (x,y,t) J km /_ q s t u V w x,y

Poststa/Spin Characteristics of an LK-1QA Sailplane

The post-stall and spin characteristics of a Laister-Kaufmann LK-10A sailplane were explored. The sailplane was instrumented to measure angular rates, linear accelerations, angle of attack, sideslip angle, airspeed, altitude, and control surface positions. The spin characteristics of the sailplane were investigated at 1%, 36%, 49%, and 99% of the allowable center of gravity range. Developed spins could not be attained for the 1% and 36% center of gravity positions. Developed spins were attained at the 49% eg position when the elevator deflection limit was increased from 15° to 25°. The LK-10A sailplane exhibited two mildly oscillatory spin modes when flown at the 99% aft eg position. The average angles of attack for these two spin modes were approximately 30° and 55°, respectively. Maximum instantaneous angles of attack of 70° were attained.

Linear optimal theory applied to active structural bending control.

To improve ride qualities and structural fatigue life of high-performance aircraft, control of structural flexing is desirable. Linear optimal control via the root-square locus was employed to design a simple, effective bending-control system for four XB-70 coupled longitudinal bending modes. The weighted sum of the mean-square differential angular acceleration and the control surface position was minimized for random wind-gust inputs. Approximation of frequency response characteristics of high-order optimal compensations led to a suboptimal mechanization (fixed-parameter, third-order transfer function with a programmed gain) which produced nearly optimal performance for three diverse low-altitude flight cases. The rms differential angular acceleration was reduced by a factor of 5 for a control expenditure of approximately 13°/sec rms elevon rate per fps of rms wind-gust velocity. Changed vehicle parameters for a fourth flight case require different sensor locations for effective control. Extensive design data show correlation of performance with plant gain and performance-index weighting factors, indicating a potential for predesign prediction of results. Comparable performance is provided by simple, uncompensated feedbacks for some cases, but stability margins are lower than for the near-optimal systems.

Some Jaguar Systems

THE basic autostabiliser system requirements are that the aircraft control surface position be controlled by signals derived from the aircraft's angular velocities. The gearing between these two quantities is required to be inversely proportional to a pressure derived signal, which may be cither dynamic pressure (at less than Mach 0·95) or static pressure (at greater than Mach 0·95), and limits are imposed on the range of this gain scheduling.

The describing functions for a constrained-range integration process with bang-bang input and dead zone

Recently, considerable attention has been devoted to the analysis of high-order systems containing severe nonlinearities separated by linear functions, particularly those using bang-bang or on-off controllers. These have proven to be very satisfactory in the attitude control systems of manned or unmanned spacecraft and satellites. This paper develops describing functions for a particularly complicated multiple nonlinearity: a tri-stable (bang-bang with dead zone) characteristic, followed by a linear integrator with a constrained range of integration. It should be noted that constraining range of integration is not equivalent to simple limiting of an integrator's output. This system of nonlinearities has not previously been treated in the literature, although it is founds for example, in satellite attitude controllers where on-off torques give rise to constrained momentum wheel angular velocities or in guided missile hydraulic actuators with bang-bang power spool and control surface position limits. This frequency-variant nonlinearity has three distinct modes of operation and, therefore, is quite different from the usual single nonlinearity considered for describing function application. The describing function and boundary equation for each mode are derived in the paper, and numerical examples are given. The analytical results were found to agree well with results from an analog computer simulation. These describing functions may be used to size power actuators ands therefores should be very useful for preliminary systems design.