• International Journal of Aeronautical and Space Sciences
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• v.18 no.4
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• pp.719-728
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• 2017

This paper proposes a synthesis of new guidance law to generate an evasive maneuver against enemy's missile interception while considering its impact angle, acceleration, and field-of-view constraints. The first component of the synthesis is a new function of repulsive Artificial Potential Field to generate the evasive maneuver as a real-time dynamic obstacle avoidance. The terminal impact angle and terminal acceleration constraints compliance are based on Time-to-Go Polynomial Guidance as the second component. The last component is the Logarithmic Barrier Function to satisfy the field-of-view limitation constraint by compensating the excessive total acceleration command. These three components are synthesized into a new guidance law, which involves three design parameter gains. Parameter study and numerical simulations are delivered to demonstrate the performance of the proposed repulsive function and guidance law. Finally, the guidance law simulations effectively achieve the zero terminal miss distance, while satisfying an evasive maneuver against intercept missile, considering impact angle, acceleration, and field-of-view limitation constraints simultaneously.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.38 no.3
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• pp.264-273
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• 2010

In this work, autonomous formation flight tests of multiple UAVs are experimentally studied. After a guidance and control system for a UAV is designed and tested, PID formation controller for follower UAV is tested using longitudinal and lateral distance feedback. It is shown that more stable and efficient formation guidance system is obtained by using position and attitude of the leader aircraft, which is exploited to calculate virtual waypoint for follower. In order to improve transient response during turn, part of roll command of the leader is added to the guidance command. Finally, autonomous formation flight test results of 3 UAVs are shown by using the best guidance algorithm suggested.

• Journal of the Korea Institute of Military Science and Technology
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• v.16 no.6
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• pp.739-744
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• 2013

In this paper, we propose a method of generating pitch commands for consistent initial trajectories irrelevant to flight conditions in the initial boosting phase of a thrust-vector-controlled vehicle. After shape assumption of the pitch command profile, parameters of the profile are determined in real time in order for the summit height of the trajectory to be a desired value by deriving the summit height considering thrust performance, gravity, and other flight conditions. Computer simulation results demonstrate good performance of the proposed method.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.45 no.8
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• pp.662-670
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• 2017

This paper proposes an optimization-based procedure to determine the parameters of the Apollo guidance law for Korean lunar lander mission. A lunar landing mission is formulated as a trajectory optimization problem to minimize the fuel consumption and the reference trajectory for the lander is obtained by solving the problem in the pre-flight phase. Some parameters of the Apollo guidance, which are coefficients of the polynomial used to define the guidance command, are selected based on the reference trajectory obtained in the pre-flight phase. A case study for the landing guidance of Korean lunar lander mission using the proposed procedure is conducted to demonstrate its effectiveness.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.35 no.11
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• pp.1006-1012
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• 2007

In this paper, a suboptimal homing guidance law for the homing missiles with an impact angle constraint is presented. Unlike general LQ optimal control, the aided loop ensuring some degrees of freedom for the constraint is introduced. Then an optimal feedback loop in consideration of the aided loop is designed by using Schwartz inequality. The aided loop is synthesized with the optimal control to produce the guidance command. Furthermore, to investigate the characteristics of the guidance law we carry out the comparative studies with other guidance laws. The results of the various computer simulations show the good performance of the proposed law.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.32 no.7
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• pp.69-75
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• 2004

An avoidance/guidance problem of an aircraft against moving obstacle is considered in two dimensional space. The aircraft is modelled as a point mass flying with constant speed. The lateral acceleration is assumed the control input. Artificial potential functions are applied to the terminal point and moving obstacles in order that repulsive forces and an attractive force are produced by the obstacles and the terminal point respectively. A real time guidance/avoidance law is proposed by using the potential forces and relative velocity. The guidance law for a logarithm potential function results the well-known proportional navigation law. The avoidance control command is inverse proportional to the time-to-go to the obstacle and turns the aircraft toward the negative direction of the line-of-sight change. The performance of the proposed guidance/avoidance law is verified with simulations.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.49 no.12
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• pp.981-988
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• 2021

This paper presents a decentralized guidance algorithm for swarm flight of fixed-wing UAVs (Unmanned Aerial Vehicles). Considering swarm flight missions, we assume four representative swarm tasks: gathering, loitering, waypoint/path following, and individual task. Those tasks require several distinct maneuvers such as path following, flocking, and collision avoidance. In order to deal with the required maneuvers, this paper proposes an integrated guidance algorithm based on vector field, augmented Cucker-Smale model, and potential field methods. Integrated guidance command is synthesized with heuristic weights designed for each guidance method. The proposed algorithm is verified through flight tests using up to 19 small fixed-wing UAVs.

• Journal of Institute of Control, Robotics and Systems
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• v.8 no.7
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• pp.613-623
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• 2002

Ascent trajectory optimization and explicit guidance problems for a satellite launch vehicle with yaw maneuver in a 3-dimension are considered. The trajectory optimization problem with boundary conditions is formulated as a nonlinear programming problem by parameterizing the inertial pitch and yaw attitude control variables, and is solved by using the SQP algorithm. The flight constraints such as gravity-turn and range safety conditions are imposed. An explicit inertial guidance algorithm in the exoatmospheric phase is also presented. The guidance algorithm provides steering command and time-to-go value directly using the current states of the vehicle and the desired orbit insertion conditions. The liquid propelled Delta 2910 launch vehicle is used as a numerical model.

• Journal of the Korea Institute of Military Science and Technology
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• v.18 no.2
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• pp.181-188
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• 2015

This paper present, the result of the guidance and control law for a course correction munitions(CCM) with 2sets of canards positioned in the rotating nose section. The nonlinear simulation model of the CCM was developed based on 7DOF equation of motion. The ability of correcting position was verified by open-loop control input with nonlinear model. The guidance and control command was constructed by reference trajectory which can be obtained with no control. Finally, the performance of the guidance and control law was evaluated through Monte-carlo simulation. The CEP(Circular Error Probability) was obtained by considering the errors in muzzle velocity, aerodynamic coefficient, wind, elevation and azimuth angle and density.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.38 no.7
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• pp.664-672
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• 2010

This paper provides the target observability analysis of time-to-go polynomial guidance law with bearing-only measurement. In this study, a direct approach is used to analyze the target observability. Since the observability condition of a constant-velocity target is given by the function of LOS angle only, the target observability characteristic is determined by substituting the closed form solution of LOS angle to the observability condition directly. The analysis results show that the target observability is depended on the choice of guidance gain, initial intercept condition and guidance command shape. After that this mathematical analysis result is evaluated and demonstrated by number of simulation.