• Journal of the Korea Institute of Military Science and Technology
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• v.11 no.3
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• pp.5-12
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• 2008

This paper presents the features of an impact angle constrained open-loop optimal trajectory which is given by a function of initial conditions and optimal guidance gains. Using missile motion described by linearized kinematic equations and a proper form of performance index, an inverse optimal problem is suggested to investigate the gains related to the performance index. The flight trajectory and time-to-go can be shaped in terms of the optimal guidance gains. The results are evaluated by 3-DOF simulation.

• Journal of Mechanical Science and Technology
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• v.16 no.6
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• pp.770-775
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• 2002

In this paper, a game optimal receding horizon guidance law (GRHG) is proposed, which does not use information of the time-to-go and target maneuvers. It is shown that by adjusting design parameters appropriately, the proposed GRHG is identical to the existing receding horizon guidance law (RHG), which can intercept the target by keeping the relative vertical separation less than the given value, within which the warhead of the missile is detonated, after the appropriately selected time in the presence of arbitrary target maneuvers and initial relative vertical separation rates between the target and missile. Through a simulation study, the performance of the GRHG is illustrated and compared with that of the existing optimal guidance law (OGL).

• Journal of Institute of Control, Robotics and Systems
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• v.7 no.2
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• pp.173-182
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• 2001

Ascent trajectory optimization and optimal explicit guidance problems for a satellite launch vehicle in a 2-dimensional pitch plane are studied. The trajectory optimization problem with boundary conditions is formulated as a nonlinear programming problem by parameterizing the pitch attitude control variable, and is solved by using the SQP algorithm. The flight constraints such as gravity-turn are imposed. An optimal explicit 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. To verify the optimality and accuracy of the algorithm simulations are performed.

• Journal of the Korean Society for Industrial and Applied Mathematics
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• v.19 no.3
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• pp.289-303
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• 2015

In this paper, an optimal guidance law with terminal angle constraint considering the seeker's lock-on condition, in which the target is located within the field-of-view (FOV) and detection range limits at the end of the midcourse phase, is proposed. The optimal solution is obtained by solving an optimal control problem minimizing the energy cost function weighted by a power of range-to-go subject to the terminal constraints, which can shape the guidance commands and the missile trajectories adjusting guidance gains of the weighting function. The proposed guidance law can be applied to both of the midcourse and terminal phases by setting the desired relative range and look angle to the final interception conditions. The performance of the proposed guidance law is analyzed through nonlinear simulations for various engagement conditions.

• 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.

• 제어로봇시스템학회:학술대회논문집
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• 2001.10a
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• pp.125.5-125
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• 2001

Until now, many guidance laws have been developed. They mainly used the classical tail-pursuit guidance method based on geometric angle information, the proportional navigation method based on the line of sight(LOS) rate, and the optimal guidance law based on optimal control theorem. In the augmented guidance law, target acceleration information and autopilot characteristics are added the guidance command. In this study, new guidance laws considering geometric angle are proposed. Two guidance laws are developed for the midcourse guidance law, and a guidance law is developed for the terminal guidance respectively. The proposed guidance laws utilize the LOS rate and the geometric angle information simultaneously. In the midcourse guidance, the guidance command is ...

• International Journal of Aeronautical and Space Sciences
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• v.10 no.1
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• pp.46-58
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• 2009

This paper deals with design procedure of online guidance and control law for future missiles that requires agile maneuverability. For the purpose, the missile with high powered side thruster is proposed. The guidance and control law for such missiles is discussed from a point of view of optimal control theory in this paper. Minimum time problem is solved for the approximated system. It is derived that bang- bang control is optimal input from the necessary conditions of optimal solution. Feedback guidance without iterative calculation is useful for actual systems. In this paper. multiple design point method is applied to design feedback gains and feed forward inputs of the guidance and control law. The numerical results show that the proposed guidance and control law has a high -performance for wide-ranging boundary conditions.

• Journal of Institute of Control, Robotics and Systems
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• v.10 no.11
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• pp.1065-1070
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• 2004

Optimal collision avoidance algorithm for unmanned aerial vehicles based on proportional navigation guidance law is investigated this paper. Although proportional navigation guidance law is widely used in missile guidance problems, it can be used in collision avoidance problem by guiding the relative velocity vector to collision avoidance vector. The optimal navigation coefficient can be obtained if an obstacle if an obstacle moves at constant velocity vector. The stability of the proposed algorithm is also investigated. The stability can be obtained by choosing a proper navigation coefficient.

• 제어로봇시스템학회:학술대회논문집
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• 1992.10a
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• pp.578-583
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• 1992

In this paper, we show that various navigation constant values in PNG law can result in as optimal gains when we introduce proper time-varying weighting functions into the cost function of an optimal control problem. we then apply this idea to the guidance problem where we are required to achieve a given impact angle as well as the zero miss distance. As a result, we obtain a set of optimal guidance laws each of which could be related to a navigation constant in PNG. Some basic properties of these guidance laws are also presented.

• 제어로봇시스템학회:학술대회논문집
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• 1992.10a
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• pp.601-606
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• 1992

In this paper, we consider an optimal guidance problem with both the terminal impact angle and control constraints in addition to the usual zero miss distance constraint. We first present the optimal solution of the problem for the missile of an arbitrary order, and show that it is a linear combination of a step response and a ramp response of the missile. Therefore the usual practice of using the control obtained by saturating the optimal solution for the case of unlimited control may result in a large terminal miss. A method called the initial command saturation is suggested to reduce this terminal miss, where the control in the initial phase of guidance is forced to be saturated until a certain condition for a guidance variable is met.