• 제어로봇시스템학회:학술대회논문집
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• 2005.06a
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• pp.463-467
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• 2005

In this paper, the guidance law applicable to formation flight of UAV in three-dimensional space is proposed. The concept of miss distance, which is commonly used in the missile guidance laws, and Lyapunov stability theorem are effectively combined to obtain the guidance commands of the wingmen. The propose guidance law is easily integrated into the existing flight control system because the guidance commands are given in terms of velocity, flight path angle and heading angle to form the prescribed formation. In this guidance law, communication is required between the leader and the wingmen to achieve autonomous formation. The wingmen are only required the current position and velocity information of the leader vehicle. The performance of the proposed guidance law is evaluated using the complete nonlinear 6-DOF aircraft system. This system is integrated with nonlinear aerodynamic and engine characteristics, actuator servo limitations for control surfaces, various stability and control augmentation system, and autopilots. From the nonlinear simulation results, the new guidance law for formation flight shows that the vehicles involved in formation flight are perfectly formed the prescribed formation satisfying the several constraints such as final velocity, flight path angle, and heading angle.

• International Journal of Aeronautical and Space Sciences
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• v.7 no.1
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• pp.129-136
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• 2006

A new guidance synthesis for anti-ship missiles to control impact angle and impact time is proposed in this paper. The flight vehicle is assumed as a 1st order lag system to consider more practical system. The proposed guidance synthesis enhances the survivability of anti-ship missiles because multiple anti-ship missiles with the proposed synthesis can hit the target simultaneously. The control input to satisfy constraints of zero miss distance and impact angle, and the feedforward bias control input to control impact time constitute the guidance law. The former is from trajectory shaping guidance, the latter is from neural network. And particle swarm optimization method is introduced to furnish reference input and output for learning in neural network. The performance of the proposed synthesis in the accuracy of impact time and angle is validated by numerical examples.

• 제어로봇시스템학회:학술대회논문집
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• 1996.10b
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• pp.482-485
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• 1996

We design a homing guidance law based on the proportional navigation for the fast-rolling, single-axis control missiles and analyse the misdistance of the designed guidance system. The guidance law includes a compensation scheme which compensates for the phase-shift between the commanded and achieved acceleration which is peculiar to the fast rolling airframe with single-axis control. In the error analysis of the guidance system, we calculate the misdistance with respect to the target maneuver on the 3-dimensional space via direct simulations. Also, we conduct adjoint simulation on the 2-dimensional plane in case that phase-shift is perfectly compensated. Finally we approximate the linear time-varying dynamics of the missile with autopilot to a linear time-invarient system, and as a result we can find the misdistance as a closed-form.

• 제어로봇시스템학회:학술대회논문집
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• 1996.10b
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• pp.245-248
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• 1996

This paper suggests a new concept for missile guidance control, called linearized common curvature guidance law that enhances the probability to kill a target. The proposed guidance system is composed of two switching modes; one for the midcourse guidance and the other for the terminal guidance, which is switched by a specified critical value (.epsilon.). And the system and the commands are formulated and its simulations are provided in comparison with the conventional commanded line of sight guidance algorithm. Miss distance and angle of attack are denoted as performance of parameters. This new concept, common curvature guidance algorithm, revises the navigation guidance and accompanies, various considerations.

• 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.

• International Journal of Aeronautical and Space Sciences
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• v.13 no.3
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• pp.369-376
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• 2012

A nonlinear robust guidance law is designed for missiles against a maneuvering target by incorporating sliding-mode and optimal control theories based on the state dependent Riccati equation (SDRE) to achieve robustness against target accelerations. The guidance law is derived based on three-dimensional nonlinear engagement kinematics and its robustness against disturbances is proved by the second method of Lyapunov. A new switching surface is considered in the sliding-mode control design. The proposed guidance law requires the maximum value of the target maneuver, and therefore opposed to the conventional augmented proportional navigation guidance (APNG) law, complete information about the target maneuver is not necessary, and hence it is simple to implement in practical applications. Considering different types of target maneuvers, several scenario simulations are performed. Simulation results confirm that the proposed guidance law has much better robustness, faster convergence, and smaller final time and control effort in comparison to the sliding-mode guidance (SMG) and APNG laws.

• 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.

• 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.

• International Journal of Aeronautical and Space Sciences
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• v.15 no.3
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• pp.281-292
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• 2014

We propose a novel cooperative guidance law design method based on the finite time disturbance observer (FTDO) for multiple near space interceptors (NSIs) with impact time control. Initially, we construct a cooperative guidance model with head pursuit, and employ the FTDO to estimate the system disturbance caused by target maneuvering. We subsequently separate the cooperative guidance process into two stages, and develop the normal acceleration command based on the super-twisting algorithm (STA) and disturbance estimated value, to ensure the convergence of the relative distance. Then, we also design the acceleration command along the line-of-sight (LOS), based on the nonsingular fast terminal sliding mode (NFTSM) control, to ensure that all the NSIs simultaneously hit the target. Furthermore, we prove the stability of the closed-loop guidance system, based on the Lyapunov theory. Finally, our simulation results of a three-to-one interception scenario show that the proposed cooperative guidance scheme makes all the NSIs hit the target at the same time.

• 제어로봇시스템학회:학술대회논문집
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• 2003.10a
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• pp.1822-1825
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• 2003

A practical range estimator based on the robust Kalman filter is proposed to solve the range estimation problem for surface to air missile(SAM) homing guidance. Apart from the previous works based on the extended Kalman filter(EKF) with bearing only measurement, the proposed scheme makes use of line-of-sight(LOS) rate to ensure the fast convergency at long-range. In this reason, the robust Kalman filter is considered to deal with LOS rate measurement error. The recursive linear structure of proposed filter is easy to implement and make it possible to reduce computational burdens. Moreover, it shows good estimation performance without specific guidance law such as oscillation proportional navigation guidance(OPNG).