• 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.

• International Journal of Aeronautical and Space Sciences
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• v.18 no.4
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• pp.709-718
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• 2017

This paper proposes the guidance laws for an agile turn of air-to-air missiles during the initial boost phase. Optimal solution for the agile turn is obtained based on the optimal control theory with a simplified missile dynamic model. Angle-of-attack command generating methods for completion of agile turn are then proposed from the optimal solution. Collision triangle condition for non-maneuvering target is reviewed and implemented for update of terminal condition for the agile turn. The performance of the proposed method is compared with an existing homing guidance law and the minimum-time optimal solution through simulations under various initial engagement scenarios. Simulation results verify that transition to homing phase after boost phase with the proposed method is more effective than direct usage of the homing guidance law.

• Journal of the Korea Institute of Military Science and Technology
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• v.1 no.1
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• pp.211-218
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• 1998

As a part of trajectory modulation to increase system survivability and terminal effectiveness, impact angle control is required in the terminal phase of tactical missile systems. The missile systems are not allowed to have high altitude to reduce probability of detection by sensors of missile defense systems. In this paper, an analytic form of a time-optimal control law is suggested in the case of constrained missile maneuverability and impact angle under the assumption of a zero-lag autopilot. The control law is obtained by establishing optimal missile-target engagement geometry in the vertical plane. Extension of the law for missiles with autopilot response lags requiring a numerical solution is studied by introducing an iterative algorithm for optimal switching time determination of which the initial switching instants are obtained from the analytic solution. Also suggested is a closed-form impact angle control law derived by an energy-optimal approach. The performances of the proposed guidance laws are evaluated by a series of computer runs.

• Journal of the Korean Society for Industrial and Applied Mathematics
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• v.19 no.3
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• pp.271-287
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• 2015

In this article, a more generalized form of the impact time and angle control guidance law is proposed based on the linear quadratic optimal control methodology. For the purpose on controlling an additional constraint such as the impact time, we introduce an additional state variable that is defined to be the jerk (acceleration rate). Additionally, in order to provide an additional degree of freedom in choosing the guidance gains, the performance index that minimizes the control energy weighted by an arbitrary order of time-to-go is considered in this work. First, the generalized form of the impact angle control guidance law with an additional term which is used for the impact time control is derived. And then, we also determine the additional term in order to achieve the desired impact time. Through numbers of numerical simulations, we investigate the superiority of the proposed guidance law compared to previous guidance laws. In addition, a salvo attack scenario with multiple missile systems is also demonstrated.

• Journal of the Korean Society for Industrial and Applied Mathematics
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• v.19 no.3
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• pp.365-386
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• 2015

In this article, the linear quadratic (LQ) optimal guidance laws with arbitrary weighting functions are introduced. The optimal guidance problems in conjunction with the control effort weighed by arbitrary functions are formulated, and then the general solutions of these problems are determined. Based on these investigations, we can know a lot of previous optimal guidance laws belong to the proposed results. Additionally, the proposed results are compared with other results from the generalization standpoint. The potential importance on the proposed results is that a lot of useful new guidance laws providing their outstanding performance compared with existing works can be designed by choosing weighting functions properly. Accordingly, a new optimal guidance law is derived based on the proposed results as an illustrative example.

• 제어로봇시스템학회:학술대회논문집
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• 2001.10a
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• pp.156.3-156
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• 2001

This paper handles synthesis and investigation of a neural network guidance law based on pursuit-evasion games. This work considers two-dimensional pursuit-evasion games solved by using the gradient method. The procedure of developing a guidance law from the game-optimal solutions is deeply examined, and important features of the neural network guidance are investigated. The proposed neural network guidance law takes the range, range rate, line-of-sight rate, and heading error as its input variables. By reconstructing the trajectory, the accuracy of the neural network approximation is verified. Afterwards, robustness of the neural network guidance to the autopilot lag, which results from its feedback structure, is investigated ...

• International Journal of Aeronautical and Space Sciences
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• v.13 no.2
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• pp.154-169
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• 2012

This paper presents a comprehensive review of spacecraft guidance algorithms for asteroid intercept and rendezvous missions. Classical proportional navigation (PN) guidance is reviewed first, followed by pulsed PN guidance, augmented PN guidance, predictive feedback guidance, Lambert guidance, and other guidance laws based on orbit perturbation theory. Optimal feedback guidance laws satisfying various terminal constraints are also discussed. Finally, the zero-effort-velocity (ZEV) error, analogous to the well-known zero-effort-miss (ZEM) distance, is introduced, leading to a generalized ZEM/ZEV guidance law. These various feedback guidance laws can be easily applied to real asteroid intercept and rendezvous missions. However, differing mission requirements and spacecraft capabilities will require continued research on terminal-phase guidance laws.

• Proceedings of the KIEE Conference
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• 1995.11a
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• pp.222-224
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• 1995

This paper deals with 3-dimensional missile guidance law. This presents the general optimal solution of the state equation which includes the target maneuvering as the Gauss-Markov processing. The main results ore about the transformation between the Cartesian coordinates on which both the guidance law and the filter are bused and the polar coordinates system in real missile guidance and measurement information. And the extended Kalman filter and adjustment of the estimated target acceleration by triangular functions is proposed solution to this transformation problem. It is shown that this proposed transformation is valid in real 3-dimensional guidance problem by the computer simulation.

• The Transactions of The Korean Institute of Electrical Engineers
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• v.68 no.1
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• pp.172-181
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• 2019

This paper deals with the problem of generating the energy optimal trajectory which is intended to enhance the target tracking performance of a passive homing missile. Noticing that the essence of passive target tracking is the range estimation problem, the target information gathered by passive measurements can be readily analyzed by introducing the range estimator designed in line-of-sight(LOS) frame. Moreover, for the linear filter structure of the suggested range estimator, the cost function associated with the target information is clearly expressed as a function of the line-of-sight rate. Based on this idea, the optimal missile trajectory maximizing the target information is obtained by solving the saddle point problem for an indefinite quadratic cost which consists of the target information and the energy. It is shown that, different from the previous heuristic approaches, the guidance command producing the optimal passive homing trajectory is produced by the modified proportional navigation guidance law whose navigation constant is determined by the weighting coefficient for target information cost.

• Proceedings of the IEEK Conference
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• 2006.06a
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• pp.853-854
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• 2006

Due to a recently developed approach to autopilot controller design for highly-maneuvering BTT (bank-to-turn) missiles, we now can derive explicitly the final-phase optimal guidance (OG) law considering the autopilot dynamics through direct use of the well-known linear optimal control theory. The proposed OG law can decrease the miss distance (MD) remarkably with small acceleration and roll rate profile at the time of interception.