• 제어로봇시스템학회:학술대회논문집
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• 1992.10a
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• pp.469-473
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• 1992

The conventional closed-loop guidance commands are generated from a simplified point mass model for real time operations. In real situations, the generated guidance commands are applied to the original rigid body. This can cause attitude instability of the vehicle. In this paper, in order to solve the attitude instability problem in the guidance system sense, the influence of the guidance commands on a launch vehicle attitude is derived quantitatively. The checking method of the attitude stability conditions that uses Liapunov theorem is proposed, and the attitude stabilizing method is also proposed. The attitude stability is accomplished by subtracting the influence of the guidance commands that destabilize the vehicle attitude. The closed-loop guidance commands generated from the simplified point mass model may destabilize the vehicle attitude, which is verified through simulations. In this case, the vehicle attitude can be always stabilized with the proposed attitude stabilizing method without additive fuel consumption.

• Journal of Institute of Control, Robotics and Systems
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• v.15 no.1
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• pp.15-22
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• 2009

The guidance controller of the conventional aircraft consists of inner-loop (autopilot) and outer-loop (guidance). If the guidance controller can be designed as an integrated guidance and control (IGC), the various advantages exist. The integrated guidance and control formulation can compensate for the effect of autopilot lag. An integrated approach also helps avoid the iterative procedure involved in tuning the guidance and autopilot subsystems, if designed separately. Integrated design is also less susceptible to saturation and stability problems. This paper presents an approach to IGC design for the unpowered air vehicle with the only flaperon using a combination of adaptive output feedback inversion and backstepping techniques. Adaptive neural networks are trained online with available measurements to compensate for unmodeled nonlinearities in the design process.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.36 no.7
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• pp.658-665
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• 2008

In this study a practical guidance law for missile guidance loop is established by considering the missile dynamics. Assuming the attitude controlled missile, several optimal guidance laws according to missile dynamical characteristics are derived by applying the optimal control theory and an effective guidance law in practical point is presented.

• 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 Institute of Control, Robotics and Systems
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• v.19 no.3
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• pp.275-280
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• 2013

Conventional proportional navigation guidance law is not adequate for missiles with a strapdown seeker, because the strapdown seeker cannot measure line-of-sight rate directly. This paper suggests a guidance loop design method, in which the look angle, measured by the strapdown seeker directly, is controlled to deliver a missile to a target. Basically, the look angle control loop is regarded as an attitude control loop. By using the proposed method, it is possible to shape the midcourse trajectory by choosing the reference look angle properly. The look angle control loop can robustly maintain target lock-on against disturbances because the target is always captured in the field of view of the seeker. The performance of the proposed method is verified via 6-DOF simulations of a true short range tactical missile model.

• 제어로봇시스템학회:학술대회논문집
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• 1991.10a
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• pp.268-273
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• 1991

Missile and target closing velocity is used in the proportional navigation(PN) missile guidance loop. But it is difficult to estimate the closing velocity when passive seeker is used and only the Line-of-Sight(LOS) rate is available in the guidance loop. In this study, new closing velocity estimation method is developed. This method uses LOS rate measurement only and uses some characteristics of PN guidance law. The Lyapunov method is used to analyze the stability of the developed estimation method.

• 제어로봇시스템학회:학술대회논문집
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• 1992.10a
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• pp.567-571
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• 1992

A guidance loop of the missile system which uses a command to line of sight(CLOS) guidance law is designed based on LQG theory. In the environment of the severe tracking system noise, the system requires small aerodynamic control fin travel and small miss distance simultaneously. Results from a sample airframe shows good performance against a randomly maneuvering target.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.36 no.9
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• pp.874-883
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• 2008

This study considers powered explicit guidance, one of the closed-loop guidance laws for satellite launch vehicles. The guidance algorithm employed here does not include the iterative procedure of the original algorithm. Also, the single-target algorithm to treat the general time-varying thrust profiles is described. The computer simulations for the 6-DOF launch vehicle model are performed to investigate the orbit injection accuracy of the guidance algorithm in the nominal/off-nominal flight conditions.

• 제어로봇시스템학회:학술대회논문집
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• 1995.10a
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• pp.283-286
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• 1995

Due to finite bandwidth of missile dynamics, guidance commands in PN guidance tend to diverage as the missile approaches to the target. In this paper, a new method based on the short-time stability theorem is introduced to extend the stability region.

• Journal of Advanced Navigation Technology
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• v.2 no.2
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• pp.84-99
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• 1998

As a part of closed-loop guidance law studies for anti-ballistic missile defense, a mid-course guidance law is proposed to engage the target with the predetermined attitude for increased terminal effectiveness. The proposed guidance law is based on the predicted target position calculated from a simplified solution of target motion and the estimates of an extended Kalman filter utilizing noisy nonlinear radar measurements. Extension of the proposed mid-course guidance to 3 dimensional engagements are also studied. Performance of the proposed mid-course guidance law together with a terminal guidance law in the form of conventional proportional navigation guidance is evaluated by a series of simulation studies.