• Journal of the Korean Society for Industrial and Applied Mathematics
• /
• v.19 no.3
• /
• pp.305-325
• /
• 2015

In this paper, missile homing guidance laws to control the impact angle and time are proposed based on the polynomial function. To derive the guidance commands, we first assume that the acceleration command profile can be represented as a polynomial function with unknown coefficients. After that, the unknown coefficients are determined to achieve the given terminal constrains. Using the determined coefficients, we can finally obtain the state feedback guidance command. The suggested approach to design the guidance laws is simple and provides the more generalized optimal solutions of the impact angle and time control guidance.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.39 no.1
• /
• pp.16-24
• /
• 2011

In this paper, we propose a new guidance law for target observability enhancement, which can control both terminal impact angle and acceleration. The proposed guidance law is simple form, combined conventional time-to-go polynomial guidance and a additional bias term which consists of relative position and proportional gain. The guidance law provides oscillatory flight trajectory and it maintains the conventional time-to-go polynomial guidance performance. To investigate the characteristics of the guidance law, we derive the closed-form solution, and various simulations are performed for proving the validity of the proposed guidance.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.38 no.8
• /
• pp.774-780
• /
• 2010

This paper deals with the choice of guidance gain for the time-to-go polynomial (POLY) guidance law when the acceleration limit is existed. POLY is derived based on the assumption that guidance commands are formed by a time-to-go polynomial function. The main characteristic of POLY is that any positive values can be used for its guidance gain. For this reason, it is ambiguous to choose a proper guidance gain. To relieve this difficulty, we firstly derive the closed-form solution of acceleration command and figure out the relationship between the maximum acceleration and guidance gain. From this analysis, we provide a guideline for choosing a guidance gain which satisfies the desired acceleration limit. Finally, the proposed method is demonstrated by simulation study.

• International Journal of Aeronautical and Space Sciences
• /
• v.8 no.1
• /
• pp.46-53
• /
• 2007

In this paper, a guidance law for vision-based automatic landing of unmanned aerial vehicles (UAVs) is proposed. Automatic landing is a challenging but crucial capability for UAVs to achieve a fully autonomous flight. In an autonomous landing maneuver of UAVs, the decision of where to landing and the generation of guidance command to achieve a successful landing are very significant problem. This paper is focused on the design of guidance law applicable to automatic landing problem of fixed-wing UAV and rotary-wing UAV, simultaneously. The proposed guidance law generates acceleration command as a control input which derived from a specified time-to-go ($t_go$) polynomial function. The coefficient of $t_go$-polynomial function are determined to satisfy some terminal constraints. Nonlinear simulation results using a fixed-wing and rotary-wing UAV models are presented.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.38 no.7
• /
• pp.664-672
• /
• 2010

This paper provides the target observability analysis of time-to-go polynomial guidance law with bearing-only measurement. In this study, a direct approach is used to analyze the target observability. Since the observability condition of a constant-velocity target is given by the function of LOS angle only, the target observability characteristic is determined by substituting the closed form solution of LOS angle to the observability condition directly. The analysis results show that the target observability is depended on the choice of guidance gain, initial intercept condition and guidance command shape. After that this mathematical analysis result is evaluated and demonstrated by number of simulation.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.46 no.10
• /
• pp.814-822
• /
• 2018

This paper proposes a guidance law for missiles with control time constraint. Because the proposed guidance law is based on a time-to-go polynomial, it has a generalized form. Also, acceleration of the proposed law converges to zero at the end of the control time, which reduces the sensitivity to the time-to-go estimation error and can increase the flight stability when the separation of the missile appears. A prediction method of the time-to-go is proposed for implementing the proposed law, and the possibility of application to the midcourse and terminal guidance phases is dealt with for exoatmospheric interception. The characteristics and performance of the proposed law are analyzed throughout various simulations.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.45 no.8
• /
• pp.662-670
• /
• 2017

This paper proposes an optimization-based procedure to determine the parameters of the Apollo guidance law for Korean lunar lander mission. A lunar landing mission is formulated as a trajectory optimization problem to minimize the fuel consumption and the reference trajectory for the lander is obtained by solving the problem in the pre-flight phase. Some parameters of the Apollo guidance, which are coefficients of the polynomial used to define the guidance command, are selected based on the reference trajectory obtained in the pre-flight phase. A case study for the landing guidance of Korean lunar lander mission using the proposed procedure is conducted to demonstrate its effectiveness.

• Journal of the Korean Society for Industrial and Applied Mathematics
• /
• v.19 no.3
• /
• pp.217-234
• /
• 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
• /
• 1989.10a
• /
• pp.59-64
• /
• 1989

In this paper, the launch trajectory of the Japan scientific satellite M-3H-3 from launch to orbit injection is investigated. For the terminal conditions at a guidance target point, a guidance and control system is used. An open-loop and a closed-loop guidance schemes are used simultaneously. For the closed-loop guidance scheme, the velocity polynomial algorithm represented by the velocity difference between the target point and present velocity is used. A PD control system is used for activating gimbal type engines. The simulation result shows that all the terminal position and velocity conditions are satisfied and the trajectory for the M-3H-3 scientific satellite is reasonable.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.47 no.6
• /
• pp.429-437
• /
• 2019

This paper deals with a ZEM (Zero-Effort-Miss) based guidance law for the interception of moving targets and characteristics of the guidance law according to time-to-go estimation methods. To derive the ZEM vector feedback guidance command, we introduce a polynomial function with unknown coefficient, and then we determine the coefficient to satisfy initial and terminal constraints. Since the directions of the guidance command and ZEM vectors are adjusted by the time-to-go, general time-to-go estimation methods are proposed, which can generate the vertical and horizontal guidance commands with respect to an arbitrary reference frame. By performing various numerical simulations, the performance and characteristics of the proposed methods are investigated.