Search

This paper discusses an approach for determining the initial engagement point and steering missile during its terminal phase such that it can hit a ground target from predefined direction. The guidance algorithm generates maneuver reference, in terms of normal acceleration, that is fed to the flight control system. The algorithm is based on Proportional Navigation (PN) approach, with scheduled navigation gains. The navigation gains are scheduled as a function of Line of Sight (LOS) and desired flight path angles at engagement stage, hence it may be exploited for determining the initial engagement point. For a desired navigation gain value, the LOS is also related to the maximum value of normal acceleration reference, hence it can be used for limiting the maneuver that must be executed by the missile. The guidance system then is integrated with the flight maneuver controller. The flight control system regulates the normal acceleration, so that the missile will follow the desired trajectory and hit the target. The algorithm then is implemented and tested on a hypothetical cruise missile model. The simulation results show that the guidance approach can generate good maneuver reference, and that the performance of the normal acceleration tracking controller will significantly affects the accuracy of the system.

This paper presents a Lyapunov stability approach to design a controller for an altitude control system of a bicopter. The altitude control system enables the bicopter to fly at a desirable altitude. A representative nonlinear model of the bicopter required for designing the controller can be derived based on a first-principle modeling approach. Control law of the controller is derived by applying the Lyapunov stability theorem. The resulting controller gives rise to a global-asymptotically-stable closed-loop system when it is applied to the bicopter. Such a stability property is satisfied via the LaSalle's invariance principle. Performance of the bicopter equipped with the altitude control system is evaluated through numerical simulations. Simulation results show that the bicopter succeeds to achieve a designated altitude level with a relatively fast-and-less-overshoot transient response. For instance, the controlled bicopter is capable of reaching 10 meters from 6 meters of altitude within 1.35 seconds with only 3% overshoot. Thus, these results signify the merit and potential of applying the proposed control method to an actual bicopter.

This paper presents methodology to extract Low Order Equivalent System (LOES) model from non-linear model of flight control actuation system. Synthesis of flight control law in conceptual phase only need simple model of flight control actuation system but can represents real conditions that might happen in operation such as reduced supply pressure, broken piston seal, active-standby actuator condition, etc. Simple model of flight control actuation system is represented by LOES model which is a transfer function for one specific operation condition. This work is started by get actuation responses from non-linear model for sinusoidal input with variation in frequency. With variation of input frequency, actuation response gain and phase for each frequency can be obtained and used to reconstruct frequency response of nonlinear model. Frequency response of non-linear model is used to estimate frequency response of approximate linear model. By obtaining frequency response of approximate linear model, transfer function of approximate linear model can be derived. The methodology that is presented in this work might be applied to any modern aircraft from fighter aircraft, trainer aircraft, cargo & military transport aircraft and civilian transport aircraft. This work is performed for hypothetical aircraft that has two elevators, two ailerons and one rudder. All the control surfaces have active-standby actuation configuration to ensure availability of control. At the end of this paper, transfer function of all control surfaces of hypothetical regional aircraft are obtained for some considered real condition and will be used for flight control law synthesis. Keywords: Non-linear, Actuation, LOES.

Air superiority is a crucial thing when it comes to modern defense and conventional war because it enables one side to optimize its military operation and inhibit the enemy’s operation at the same time, whether it’s on the ground or at the sea. This enhance the development of multirole fighters which one of its properties is high maneuverability. On the other hand, this aerodynamically unstable design requires advanced flight control system which makes the development of flight control system a highly important matter. In this paper, the model of a fighter aircraft flight dynamics and the model of disturbance in terms of forces and moments caused by the missile launching are built and simulated on the open-loop system initially. From open-loop simulation, the aircraft becomes unstable after releasing right-wing missile. Furthermore, same processes are operated on two different closed-loop systems to obtain proper controller that can compensate the effect of missile releasing. Both closed-loop system are the aircraft equipped with Stability Augmentation System and aircraft equipped with Stability Augmentation System, Basic Autopilot System, and Flight Path Control System. The results show that these closed-loop systems successfully improve the response of the aircraft after launching a missile since the aircraft flight trajectory is still maintained in desired value.

A highly maneuverable aircraft, such as fighter aircraft, relies significantly on the performance of its control devices for executing aggressive and non-conventional maneuvers, and also for executing other routine tasks. Any defect or failure that may happen on the control device actuators can reduce the controllability of the aircraft, even can bring it to a catastrophic condition. A study on the effect of defected/failure control device has on the aircraft closed-loop response during some predefined maneuvers is described in this paper. It is intended to identify and analyze the aircraft dynamic behavior when one of its control surfaces is failed in particular flight phases. The failure is modeled as changes in the dynamic characteristics of the actuator, which is affected by the load generated on the control surface during maneuvers. The change of the actuator behavior affects the performance of the closed- loop system and may result in excessive control command/deflection in order to maintain control authority. The proposed approach may be exploited to examine the effect of control device failures has on the performance of fighter aircraft in executing a mission.

This paper discusses the development of a terrain following system for aerial vehicle based on ground elevation data obtained by an online detection system. The developed approach utilizes a set of elevation data obtained sequentially to generate vertical trajectory references that includes a safe clearance with respect to the ground. This reference then is fed to a flight path control system such that the aircraft can change its trajectory to follow the terrain contour. The trajectory reference generating algorithms have some parameters that can be tuned so that it can produce a reference path that will not require high g-load maneuvers. The flight path control system tracks the reference by controlling the altitude change or the flight path angle of the aircraft. The developed approach is then implemented and simulated numerically on a hypothetical air vehicle model. The results show that the approach can steer the vehicle to follow a terrain contour model, while maintain relatively moderate maneuvers.

Developing a nonlinear adaptive control system for a fixed-wing unmanned aerial vehicle (UAV) requires a mathematical representation of the system dynamics analytically as a set of differential equations in the form of a strict-feedback systems. This paper presents a method for modeling a nonlinear flight dynamics of the fixed-wing UAV of BPPT Wulung in any conditions of the flight altitude and airspeed for the first step into designing a nonlinear adaptive controller. The model was formed into 10-DOF differential equations in the form of strict-feedback systems which separates the terms of elevator, aileron, rudder and throttle from the model. The model simulation results show the behavior of the flight dynamics of the Wulung UAV and also prove the compliance with the actual flight test results.

This paper presents the result and analysis of a tilt-rotor VTOL aircraft's longitudinal flight dynamics in transition phase. This phase bridges the airplane and helicopter mode in a tilt-rotor, and is an area that is crucial to be understood in designing the control system of a tilt-rotor. The aircraft dynamics are calculated using simplified equations of motion., and considers forces and moments generated from the fuselage, wing, rotors, and both the vertical and horizontal tail. A simple control system using pole placement and PI-like controllers are used to stabilize the aircraft. The resulting simulation shows that a fully stable transition is achieved, although its dynamic behavior can be improved.

An obstacle avoidance system is being proposed. The system will combine available flight controller with a proposed avoidance method as a proof of concept. Quadcopter as a UAV is integrated with the system which consist of several modes in order to do avoidance. As the previous study, obstacle will be determined using ultrasonic sensor and servo. As result, the quadcopter will move according to its mode and successfully avoid obstacle.

The main role of Rocket Navigation System is to deliver an attitude command to Flight Control System hence the rocket can be directed to a particular path or a specific target. This command is produced from Trajectory Planning algorithm. The Navigation System also performs attitude estimation by using the Extended Kalman Filter. The inputs of Navigation System are obtained from GPS, IMU, Pitot tube, Altitude meter and actuator encoder which have been modeled in SIMULINK. From the test results, the Trajectory Planning algorithm is successful enough to give an attitude command and the rocket can reach the target. Also, the Extended Kalman Filter has successfully estimated pitch angle, u-velocity, and w-velocity. Moreover, the Extended Kalman Filter has successfully filtered out the pitch rate noise from the simulated IMU sensor.

This paper presents the design and implementation of Three-loop autopilot that functions as an attitude control system of a guided rocket. Three-loop autopilot is designed to manage the guided rocket flight path on a longitudinal plane trajectory by controlling the rocket attitude angle. Three-loop autopilot consists of three successive closed-loop systems which manipulate short period, phugoid oscillation, and other oscillation mode to have adequate damping. Control system is designed on a certain linear condition. Classical stability margin such as gain margin and phase margin is needed to handle error between the assumed model and real system. The system is also equipped with anti-windup to overcome integral windup when the actuator gets saturated. Testing of control system is done by using hardware in the loop simulation (HILS). Control system test shows that the results are good enough. With 0° initial value of pitch and angle of attack, pitch angle can be maintained at a value of 5.5° with a 0.6 seconds rise time, 0.8 peak time, 12% overshoot and zero steady state error. In addition, roll and yaw angle values can be maintained at a value of 0°. On changed command testing, the control system shows good command tracking for small angles, but the system performance decreases when angle value large enough. This is due to the saturation in the actuators and change on system parameters.

This paper presents the development of a guidance system for a small Unmanned Aerial Vehicle (UAV) system which combines a waypoints following and an obstacle avoidance systems so that the UAV has a capability to operate in an environment whose ground condition is not completely known. The waypoints following system works by adopting the VOR-Hold approach, in which a correction command will be produced to reduce the angle difference between the desired path and the line connecting the actual UAV position and the subsequent destination point. An avoidance algorithm is developed and integrated with the path following system. In case of the UAV faces an obstacle lying on its flight path, then the avoidance system will generate a set of new waypoints for correcting the flight path, so that the UAV can avoid the obstacle and then returns to the previous flight path. The proposed avoidance approach bases its algorithm on the utilization of ellipsoid geometry for defining a restricted zone containing the obstacle, which is assumed to be already identified by the detection system. When the restricted ellipsoid zone has already been established, the algorithm then computes the locations of new waypoints on the edge of the ellipsoid. The algorithm then is simulated and evaluated in some cases representing situations when an UAV has to avoid obstacles during its flight to a predefined destination.Copyright © 2011 by ASME