• 제어로봇시스템학회:학술대회논문집
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• 1993.10a
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• pp.450-455
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• 1993

Guidance commands for attitude controlled missiles inevitably take the form of attitude angle commands. On the other hand, many guidance laws compute the accelerations required to achieve their goals. Therefore some integrators must be in use for the attitude controlled missiles to implement the guidance laws. Naturally, the use of the integrator raises the problem of choosing a proper initial value. In this paper, we compute the integrator initial value which minimizes the terminal miss and show that if the total flight time of the mission is long enough, the "optimal" initial value becomes some multiple of the initial heading error or of the given impact angle to the target. We demonstrate the validity of the analysis by showing some linear and nonlinear simulation results.n results.

• Journal of the Korea Institute of Military Science and Technology
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• v.11 no.4
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• pp.29-35
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• 2008

In this paper, a new guidance law to reduce sensitivity to the initial heading errors is proposed. In order for shaping the input weights over the flight, we introduce the distribution functions expressed in terms of time-to-go and its inverse term. By applying the optimal control theory with the synthesized weights, the homing guidance law is derived. Also the characteristics of the proposed law are examined. Various computer simulations show the good performance of the proposed guidance.

• The Transactions of the Korean Institute of Electrical Engineers A
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• v.48 no.7
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• pp.881-888
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• 1999

Target pointing guidance steers a vehicle to point at a target point at a given range Rs. In this paper, vehicle's motions relative to the target point are modeled by differential equations. Then a target pointing guidance law is derived using optimal control theories. In addition, it is shown that the proposed guidance law can achieve the goal of target pointing guidance whatever initial headings are.

• 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.

• Journal of the Korean Society for Precision Engineering
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• v.15 no.6
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• pp.97-106
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• 1998

In this Paper, a fuel minimizing closed loop explicit inertial guidance algorithm for orbit injection of a rocket is developed. In the formulation, the fuel burning rate and magnitude of thrust are assumed constant. The motion of rocket is assumed to be subject to the average inverse-square gravity, but negligible effects from atmosphere. The optimum thrust angle to obtain a given velocity vector in the shortest time with minimizing fuel consumption is first determined, and then the additive thrust angle for targeting the final position vector is determined by using Pontryagin's maximum principle. To establish real time processing, many algorithms of onboard guidance software are simplified. The explicit guidance algorithm is simulated on the 2nd-stage flight of the N-1 rocket developed in Japan. The results show that the explicit guidance algorithm works well in the presence of the maximum $\pm$10% initial velocity and altitude errors, and exhibits better performance than the open-loop program guidance. The effects of the guidance cycle time are also examined.

• International Journal of Control, Automation, and Systems
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• v.5 no.4
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• pp.397-409
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• 2007

In this paper, we propose a nonlinear guidance law for missiles against maneuvering targets. First, we derive the equations of motion described in the line-of-sight reference frame and then we define the equilibrium subspace of the nonlinear system to guarantee target interception within a finite time. Using Sontag's formula, we derive a nonlinear guidance law that always delivers the state to the equilibrium subspace. If the speed of the missile is greater than that of the target, the proposed law has global capturability in that, under any initial launch conditions, the missile can intercept the maneuvering target. The proposed law also minimizes the integral cost of the control energy and the weighted square of the state. The performance of the proposed law is compared with the augmented proportional navigation guidance law by means of numerical simulations of various initial conditions and target maneuvers.

• 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.

• International Journal of Aeronautical and Space Sciences
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• v.11 no.2
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• pp.110-117
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• 2010

We propose guidance laws based on a pursuit-evasion game. The game solutions are obtained from a pursuit-evasion game solver developed by the authors. We introduce a direct method to solve planar pursuit-evasion games with control variable constraints in which the game solution is sought by iteration of the update and correction steps. The initial value of the game solution is used for guidance of the evader and the pursuer, and then the pursuit-evasion game is solved again at the next time step. In this respect, the proposed guidance laws are similar to the approach of model predictive control. The proposed guidance method is compared to proportional navigation guidance for a pursuit-evasion scenario in which the evader always tries to maximize the capture time. The capture sets of the two guidance methods are demonstrated.

• 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).

• 제어로봇시스템학회:학술대회논문집
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• 1991.10a
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• pp.419-424
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• 1991

In this paper, a fuel minimizing closed loop explicit inertial guidance algorithm for the orbit injection of a rocket is developed. In this formulation, the fuel burning rate and magnitude of thrust are assumed constant, and the motion of a rocket is assumed to be subject to the average inverse-square gravity, but with negligible atmospheric effects. The optimum thrust angle for obtaining the given velocity vector in the shortest time with minimizing fuel consumption is first determined, and then the additive thrust angle for targeting the final position vectors is determined by using Pontryagin's Maximum Principle. To establish the real time processing, many algorithms of the onboard guidance software are simplified. Simulations for the explicit guidance algorithm, for the 2nd-stage flight of the N-1 rocket, are carried out. The results show that the guidance algorithm works well in the presence of the maximum .+-.10 % initial velocity and altitude error. The effects of the guidance cycle time is also examined.