• Journal of Institute of Control, Robotics and Systems
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• v.14 no.1
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• pp.88-94
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• 2008

In this paper, a design method of a missile guidance command is presented considering the dynamics of missile control systems. The design of a new guidance command is based on the well-known PNG(propotional navigation guidance) laws. The missile control system dynamics cause the time-delays of the PN guidance command and degrade the performance of original guidance laws which are designed under the assumption of the ideal missile control systems. Using a backstepping method, these time-delay effects can be compensated. In order to implement the guidance command developed by the backstepping procedure, it is required to measure or calculate the successive time-derivatives of the original guidance command, PNG and other kinematic variables such as the relative distance. Instead of directly using the measurements of these variables and their successive derivatives, a simple disturbance observer technique is employed to estimate a guidance command described by them. Using Lyapunov method, the performance of a newly developed guidance command is analyzed against a target maneuvering with a bounded and time-varying acceleration.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.36 no.7
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• pp.658-665
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• 2008

In this study a practical guidance law for missile guidance loop is established by considering the missile dynamics. Assuming the attitude controlled missile, several optimal guidance laws according to missile dynamical characteristics are derived by applying the optimal control theory and an effective guidance law in practical point is presented.

• Journal of the Korea Institute of Military Science and Technology
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• v.19 no.5
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• pp.606-612
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• 2016

A Captive Flight Test(CFT) is used to verify the performance of missile component such as seeker and guiding algorithm of missile. Recently most of surface to air missile adapts multi-mode guidance method which include command guidance and active/passive guidance. A CFT system for missile system adapting multi-mode guidance method consists of pod equipment, command transmitting system and measuring system. In this paper, we proposed CFT system and testing method for missile which adapt multi-mode guidance, and system integration procedure by using distributed missile system integration method and procedure. The proposed CFT system and system integration method was applied to CFT of surface to air missile, and brought successful result.

• Journal of Institute of Control, Robotics and Systems
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• v.13 no.3
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• pp.273-278
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• 2007

In this paper, a performance comparison between CLOS(Command to Line-of-Sight) guidance law and PN(Proportional Navigation) guidance law is made, based on a short range surface-to-air missile simulation program called KNUCLOS. This simulation program has a full nonlinear aerodynamic missile model, a tracker model for missile and target, and target model. According to the simulation results, the PN guidance law has a better performance than CLOS guidance law under various target speed.

• Journal of the Korean Society for Industrial and Applied Mathematics
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• v.19 no.3
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• pp.217-234
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• 2015

In this tutorial the theoretical background of LQ optimal guidance is reviewed, starting from calculus of variations. LQ optimal control is then introduced and applied to missile guidance to obtain the basic form of LQ optimal guidance laws. Extension of LQ optimal guidance methodology for handling weighted cost function, dynamic lag associated with the missile dynamics and the autopilot, constrained impact angle, and constrained impact time is also described with a brief discussion on the asymptotic properties of the optimal guidance laws. Furthermore, an introduction to polynomial guidance and generalized impactangle-control guidance, which are closed related with LQ optimal guidance, is provided to demonstrate the current status of missile guidance techniques.

• 제어로봇시스템학회:학술대회논문집
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• 1994.10a
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• pp.647-652
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• 1994

In the previous paper, we presented a new guidance law for a missile during boost phase. Thus, this paper deals with the guidance law for a missile after the thrust cutoff against an accelerating and turning target. It is essentially based on the concept of proportional navigation. Some simulation studies were performed using a three dimensional mathematical model of an air-to-air missile and the effectiveness of the guidance law presented was shown.

• International Journal of Control, Automation, and Systems
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• v.6 no.3
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• pp.460-466
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• 2008

This paper deals with the stability analysis of a missile guidance loop employing an integrated proportional navigation guidance law. The missile guidance loop is formulated as a closed-loop control system consisting of a linear time-invariant feed-forward block and a time-varying feedback gain. Based on the circle criterion, we have defined the concept of absolute stability margins and obtained the gain and phase margins for the system assuming 1 st order missile/autopilot dynamics. The correlation between the absolute stability margins and the margins derived from the frozen system analysis is also discussed.

• Proceedings of the KIEE Conference
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• 2005.10b
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• pp.632-634
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• 2005

In this paper, The Frenet-Serret formula of classical geometric curve theory with the concept of a missile pointing velocity vector are used to analyze and design a missile guidance law. The capture capability of this guidance law is qualitatively studied by comparing the rotations of the velocity vectors of missile and target relative to the line of sight vector. when fuzzy Table look-up theory applied in target-missile distance & angle displacement, this research. It's performance is better then classical research.

• International Journal of Aeronautical and Space Sciences
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• v.9 no.2
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• pp.87-97
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• 2008

Modem anti-ship missiles employ complex and sophisticated guidance laws to hit the target and enhance their survivability by executing additional maneuvers. However, such maneuvers may cause the target to move out of the missile seeker's Field-Of-View (FOV). Maintaining seeker lock-on during an engagement is a critical factor for missile guidance. In this paper, a guidance law switching logic that maintains seeker lock-on and a simple guidance law that keeps the target look angle of the seeker constant is proposed. The proposed method can be used for the terminal homing phase, and can be switched from any kind of guidance laws if a proper switching condition is satisfied. The minimum and maximum flight time calculation method in consideration of the missile maneuver limit and the FOV of the seeker is also provided.

• International Journal of Aeronautical and Space Sciences
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• v.4 no.2
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• pp.26-36
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• 2003

The objective of this study is to enhance neural-network guidance to consider the impact condition. The optimal impact condition in this study is defined as an head-on attack. Missile impact-angle error, which is a measure of the degree to which the missile is not steering for a head-on attack, can also have an influence on the final miss distance. Therefore midcourse guidance is used to navigate the missile, reducing the deviation angle from head on, given some constraints on the missile g performance. A coordinate transformation is introduced to simplify the three-dimensional guidance law and, consequently, to reduce training data. Computer simulation results show that the neural-network guidance law with the coordinate transformation reduces impact-angle errors effectively.