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The quadrotor type of UAV (QUAV), is arguably the most popular type for civilian purposes, since it does not mandate a complicated remote controlling skill to operate it. This popularity, however, has yet to be balanced with research in mathematical modelling, design, and analysis. This paper describes a mathematical modelling process of a QUAV, by exploiting the usage of MATLAB/Simulink software. The main aim is to provide a proper mathematical model to design and evaluate the control system to maintain QUAV stability. Along with the model, three basic automatic control systems were also designed, i.e., the forward-speed hold, sideward-speed hold, and altitude hold system. Several simulations were conducted afterward, to test both the model and the designed control system, which showed satisfying results in representing a QUAV system. QUAV dynamic characteristics can be observed from the simulation, including the logical commands to alter the rotors thrust, the rotor limitation in handling many commands at once, the vehicle dependency of automatic control systems, and the fact that a motion in one degree of freedom can easily affected others. The developed model evidently has a good potential as a base for QUAV analyses and control design.

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.

This paper presents the flight dynamics simulation and analysis of a tilt-rotor vertical takeoff and landing (VTOL) aircraft on transition phase, that is conversion from vertical or hover to horizontal or level flight and vice versa. The model of the aircraft is derived from simplified equations of motion comprising the forces and moments working on the aircraft in the airplane's longitudinal plane of motion. This study focuses on the problem of the airplane's dynamic response during conversion phase, which gives an understanding about the flight characteristics of the vehicle. The understanding about the flight dynamics characteristics is important for the control system design phase. Some simulation results are given to provide better visualization about the behaviour of the tilt-rotor. The simulation results show that both transition phases are quite stable, although an improved stability can give better manoeuver and attitude handling. Improvement on the simulation model is also required to provide more accurate and realistic dynamic response of the vehicle.

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.

This paper provides a model of rotor dynamics which is derived mathematically and develops it into a numerical simulation model. The mathematical model was derived by reviewing the kinematic phenomena of a rotating blade and connected them with corresponding forces and moments working on the blade. The mathematical model was then developed into a numerical simulation model created in MATLAB Simulink environment. The model is based on a rotor mounted on a rotary wing aircraft with tilt-rotor configuration, which uses two rotors at both ends of its wings, with four blades each. The results of the simulation are used to analyze the characteristics and dynamics of a rotor and how they affect the flight dynamics of an aircraft.

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 deals with a trajectory analysis of RKN200 multi-stage rocket system. The implementation of a multi-stage configuration means that a separation process has to be taken within the flight phase. The separation process is basically a transition process from a phase with booster as initial/launch propulsion to that with sustainer as the propulsion for the rest of flight time. RKN200 is a multi-stage rocket developed by LAPAN Indonesia for defence applications. In a multi-stage rocket development process, stage separation becomes a critical phase which should be considered carefully, since this phase will significantly affect the entire flight performance of the rocket. In this paper, the trajectory of RKN200 is analyzed in various cases with separation process included. The separation stage is modeled as a discrete process, i.e. by applying a sudden change on the rocket parameters values at the time of separation. In addition to that, impulsive force and moment which occur as a result of the separation ignition will also introduced into the mathematical model of the rocket dynamics, such that their influence to the rocket flight variables can be counted and computed . The modeling, simulation, and analysis of the flight trajectory are conducted using a simulation software already developed for rocket dynamic and performance analysis. Some simulation results are presented and analyzed to evaluate the RKN200 flight trajectory in some flight settings and conditions, and also to observe the effect of stage separation process on the rocket flight. Keywords: Rocket, Flight trajectory, Stage separation