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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.

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.

Development of electronic technology especially on flight controller-related aspect and engine technology which can generate more thrust with relatively small size encourage development of aircraft design to reach boundary. For a long time, fixed wing aircraft needs long runway to achieve its lift-off velocity hence the take-off phase is able to be continued by climbing and other flight phases. However, the tiltrotor configuration aircraft which is equipped with reliable flight controller can reduce the length of the runway by changing the thrust direction which is commonly fixed towards the body axis. In this research, the authors attempt to analyze the flight dynamics behavior in transition mode by discretizing the transition mode to several pseudo-cruise flight configurations. The analysis is started by calculating the pseudo-cruise flight condition for several flight configurations based on analytical calculation and then, compared to trimming and linearization results in numerical simulation platform. After knowing the pseudo-cruise flight configuration by trimming the aircraft, the dynamics stability of the aircraft is obtained for several tilting angles of engine. The aircraft model is limited to 3 DoF non-linear model to represent the behavior of the aircraft. The results show that analytical and numerical calculations have similar trend and tendency in pseudo-trim condition. Afterward, the selected aircraft is concluded as unstable system in all variations.

An avoidance algorithm is proposed in this paper which works by generating avoiding waypoints, within the original predefined waypoints, when the vehicle finds obstacles obstructing its flightpath. The approach developed here bases the search for avoidance path on the utilization of ellipsoid geometry as a restricted zone containing the obstacle. The restricted ellipsoid zone is established by considering the identified obstacle geometry information, and further the ellipsoid becomes the basis for computing the new waypoints for avoiding the obstacle. These avoiding waypoints determined by computing the contact points between the ellipsoid and planes the normal vector of which are corresponded to the vehicle velocity vector. The information about geometry and dimension of the ellipsoid are computed from the information about obstacle geometry, which is assumed to be available, either from mission database or predicted from UAV’s ground detection system. In the development process, the algorithm is constructed in MATLAB environment and then simulated and analyzed in some scenarioed cases representing possible situations when an UAV has to avoid obstacles during its flight. This algorithm is intended to be integrated into the guidance system of UAV.

This study aims to implement a threat assessment and decision making system to UCAV’s rigid body model guidance system and evaluate the performance of the assigned system which is established in the form of Neutral-AdvantageDisadvantage-Mutual Disadvantage (NADM) method. First, a transformation process is conducted in order to gain combat parameters from the currently known UCAV state variables which were obtained from a state-space modeled UCAV. The obtained combat parameters are then presented to the NADM threat assessment program as inputs. After the assessment process is carried out, a decision making algorithm based on a predictive control scheme determines the best set of command inputs for the UCAV in order to achieve the desired combat situation through executed maneuvers. Comparison between the results of simulations and a set of provided reference conditions is later conducted. Finally, discussions regarding these results are presented.

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 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.

This paper focuses on the simulation of Beyond Visual Range (BVR) Air Combat using calculation of Probability of Kill (PK) and Supremacy to analyze the effectivity of the tactics and strategies. The developed simulation program is expected to be used as a means of developing tactics, weapons systems evaluation and to support other air combat system applications. The simulation system can be utilized so that strategy for using weapons and better tactics for maneuver can be determined and formulated in the face of an aerial battle. The simulation illustrates how an aerial battle in a 2D plane took place. The battle occurs between two sides and can also cover case of many against many air battle. Fighters maneuver will be generated based on logic and defined tactic by providing input in the form of initial conditions. Formulation / functions, parameters and variables influencing the fighter effectiveness will be taken into account. These elements are modelled using a dynamic and open scenario approach. The simulation used a method called 'Missile Launch Envelope Model' for computing PK of missile along its way to the target. Simulation results for one versus one and many versus many cases show that the developed system can produce good and relatively realistic prediction of the outcome. Further modification of simulation program include fly out modeling to the missile.

The purpose of this study is to acquire the correlation between composite tailoring at symmetric balanced laminate and gust response on high aspect ratio wing. As a flexible structure, the high aspect ratio wing had a susceptibility from dynamic loading. Gust or turbulence load is one of dynamic loading which can have a bad impact on the wing structure. There are many incidents caused by turbulence load, hence to obtain structural response is important. Because gust load is a stiffness problem, one way to handle these effects is to make the airframe more rigid which is not favorable due to the weight penalty in the gross weight of the aircraft. Plies orientations can be an alternative to make structure stiffer without adding structural weight. From this study, plies-orientations at symmetrically laminated composite had been varied starting from 0 to 90 degrees with 5 degrees increment at the skin of the wing. Bending and torsional natural frequencies have changed where changed in play orientation as well as the vibration response from the gust load. The discrete gust cases defined in the EASA requirements are used in this study. The structural response from the gust problem had different characteristics between angle plies laminate, cross-plies laminate, and various plies orientation. As a result of this research that symmetrically balanced laminate composite with fiber direction ±15 or ±75 degrees had a good bending characteristic, nevertheless that direction have the lowest bending stiffness. Further research needs to optimize the plies orientation to design the high aspect ratio wing which has a good aeroelastic response and minimize weight.

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 discusses the modeling, simulation, and analysis of air combat in a within visual range (WVR) condition. The battle is occurred between two opposing aircraft in a 3-dimensional space. The air combat simulation is done using software MATLAB®. The analysis only covers within visual range air combat where the maneuver is much more aggressive than the beyond visual range (BVR) air combat. Both aircraft will be evaluated to determine the supremacy level based on their relative position. The relative position is used to determine the threat condition for both aircraft, whether they are in neutral, advantage, disadvantage, or mutual disadvantage (NADM) condition. Each possible state then will be quantified to determine the dominant condition. The pilot decision for taking particular maneuver is determined based on this threat assessment. Another method to assess the effectiveness of the fighter's maneuver is based on a parameter Probability of Kill (PK). This method will be used as a comparison to the developed NADM method.

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 discusses the development of a numerical simulation platform that can be used for representing the principal works of aircraft systems. The platform consists of some parts each of which is intended to replicate the operation of a system implemented in a modern aircraft, such as hydraulic line, electrical system, landing gear system, control system, etc. The platform is intended to be a tool for modeling and analyzing the operation and performance of certain systems configuration. To some extend the platform can be viewed as a virtual Iron Bird System. Iron Bird is a term representing a platform for simulating the works of aircraft systems using real components, which is very important for aircraft system development. At this stage, the numerical platform will only involve some sub-systems, namely main hydraulic line, control surface actuation, landing gear, and control system. These sub-systems are chosen to be the basis for further development where other sub-systems will be added and integrated to the platform.Keywords—Aircraft System, Numerical Simulation, System Analysis

This paper will talk about AIM 120 AMRAAM missile guidance for pursuing a moving target. The missile guidance system itself consists of missile dynamics, control system, seeker, and guidance methods. For general purpose, the missile dynamics approach will use non-linear equation of motions. The control surface that will be discussed follow the rule BTT (Bank-to-Turn) and control system that will be used is PID control system that widely used for control design. Lastly, the guidance method that will be studied here is proportional navigation and constant bearing course approach. The simulation will be conducted using MATLAB Simulink. The Simulink model consist of target dynamics, and guidance system. From the result of simulation, it will be shown that the missile can pursue its target quite well. Hence, the simulation system can be used well for preliminary design purpose.Keywords—Homing System, Control System, Proportional Navigation, Missile Dynamics, Seeker, Bank-to-Turn, Constant Bearing Course.

This paper presents the development of a 3D obstacle avoidance algorithm for Unmanned Aerial Vehicle (UAV) system. The avoidance algorithm works by generating avoiding waypoints, within the original predefined waypoints, when the vehicle faces obstacles obstructing its flightpath. The proposed approach bases its algorithm on the utilization of ellipsoid geometry for defining a restricted zone containing the obstacle. Once the restricted ellipsoid zone has already been established, the algorithm computes the locations of new waypoints on the edge of the ellipsoid. These new waypoints represent the contact points between the ellipsoid and planes the normal vector of which are determined form the vehicle velocity vector. The geometry and dimension of the ellipsoid are determined using the known information about the size of obstacle which is assumed to be available, either from mission database or predicted from UAV's ground detection system. At this stage, the algorithm computation is implemented in MATLAB environment and then evaluated for a couple of cases representing some situations when an UAV has to avoid obstacles during its flight to a predefined destination. This algorithm then can be integrated into a waypoint following system and/or guidance system of UAV.

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.

In this paper, modelling and simulation of RKX200 rocket flight dynamics are presented. The dynamics model is constructed by first-principle method based on forces and moments acting the rocket rigid body. In the development of the model, the aerodynamics coefficients are determined by using missile DATCOM software. A six Degree of Freedom (DoF) nonlinear model, as well as linear decoupled longitudinal and lateral-directional models are constructed for representing the flight dynamics of the rocket. The resulting models are numerically modelled and simulated using Simulink™ while its motion visualization is constructed using FlightGear™. Further, the dynamics model and the visualization application then are integrated and tested in Hardware In The Loop Simulation (HILS) environment.

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.

Developing an autonomous flight control system for a fixed-wing unmanned aerial vehicle (UAV) requires the mathematical representation of the system dynamics which can be obtained by applying analytical and/or empirical methods. This paper presents a method for modeling the flight dynamics of the fixed-wing UAV by analytically developing the model structure. The aerodynamic parameters/coefficients of the UAV, to be inserted into the dynamic model, are computed using DATCOM software, by considering various flight condition determined by the angles of attack, altitudes and velocities. In addition to that, an empirical model is also formed based on flight test data, which has been already processed using Kalman filter, to validate the analytical model. By comparing the simulation result of these two systems representations, it can be seen that the both approaches produce dynamic models that have similar characteristics.

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

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