• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.44 no.4
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• pp.324-332
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• 2016

In this paper, the terminal homing loop with a IIR-type strapdown seeker and a roll rate gyroscope is proposed. Basically, the proposed homing loop is based on the look-angle-control guidance. Since the range of the seeker is strictly limited, the missile is delivered to a point to lock the target on the seeker via non-guided flight during the midcourse guidance. The non-standard firing table is developed to compensate the wind and the target movement. To secure the delay margin is very important to prevent the instability of the homing loop when the time delay of the seeker is included. To validate the proposed homing loop, the 6-DOF nonlinear simulation is performed, and the Monte-Carlo simulation is also done for checking the robustness for the various kinds of uncertainty.

• Journal of the Korean Society of Propulsion Engineers
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• v.19 no.3
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• pp.54-64
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• 2015

This paper deals with an acceleration controller design for a kill vehicle equipped with a divert and attitude control system (DACS). In the proposed method, the attitude control system (ACS) is used to produce the thrust command to nullify angle-of-attack. For the angle-of-attack control, a nonlinear angle-of-attack controller is proposed based on the feedback linearization methodology. Since the flight path angle is identical to the attitude angle under the condition of zero angle-of-attack, the divert control system (DCS) can directly produce the lateral acceleration which is demanded from the guidance loop. In the proposed method, we can minimize the aerodynamic uncertainty due to the propulsive force. Additionally, we can simplify the operation logic of DCS and ACS. In this paper, nonlinear simulations are performed to show the performance of the proposed method.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.50 no.6
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• pp.393-400
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• 2022

The Divert and Attitude Control System(DACS) used in high-altitude engagements is expensive and complex. In this paper, we design a high-altitude terminal guidance and control loop of guided-missile equipped with a Thrust Vector Control(TVC) that is less expensive and simpler than DACS. The proposed system utilizes a quaternion feedback control technique to track the thrust attitude command converted from the acceleration command of true proportional navigation guidance. The performance analysis of the proposed terminal guidance and control loop is conducted through engagement simulations against ballistic targets at a high altitude.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.39 no.6
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• pp.505-516
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• 2011

The concept of the IGC(Integrated Guidance and Control) has been introduced to overcome the performance limit of the SGC(Separated Guidance and Control) loop. A new type of IGC with impact angle constraint has been proposed in this paper. Angle of attack, pitch angle rate, pitch angle and line of sight angle are considered as state variables. A controllability analysis and equilibrium point analysis have been carried out to investigate the control characteristic of the prposed IGC. The LQR(Linear Quadratic Regulator) has been adopted for the control law and detailed explanations about the adoption has been provided. The performance comparison between the IGC and the SGC has been carried out. The result of numerical simulations shows that the IGC guarantees better guidance performance than the SGC when the agile maneuver is needed for a specific guidance geometry.

• Journal of the Korea Institute of Military Science and Technology
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• v.21 no.5
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• pp.623-629
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• 2018

The acceleration and the angular velocity that include natural frequencies of a missile detected by Inertial Measurement Unit(IMU) are transmitted to the control loop of a missile. The control loop command that is calculated using above signals can cause the resonance of the missile while it flies. Hence it is common to adapt the filter and the control loop for attenuating or eliminating the undesired signals such as natural frequencies. This paper introduces the new test technique using a floating system for performance evaluation of the designed filter and the control loop prior to a flight test. The proposed scheme can check out the degradation property of vibration in the filter and the control loop, while the conventional hardware-in-the-loop simulation(HILS) scheme cannot.

• Proceedings of the KIPE Conference
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• 2017.07a
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• pp.194-195
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• 2017

본 논문은 독립형 마이크로그리드에서 BESS-디젤발전기 연계운전 시, 동기 인덕턴스와 LCL 필터의 상호간섭을 고려한 CVCF 제어기를 설계하고 해석하였다. CVCF 제어의 외부루프는 비례-공진(PR)을 이용한 기본파 전압제어이고, 내부루프는 제어 안정도 향상을 위한 커패시터 전류제어로 구성된 이중 루프 구조이다. 인버터 $1.5T_s$ 시지연을 고려하여 모델링을 설계하고, 이산시간 영역에서의 근궤적 기법을 통해 CVCF 제어의 안정도를 해석하였다. 50kW BESS-디젤발전기로 구성된 시뮬레이션을 통해 안정도 해석과 동일한 결과를 도출함으로써 설계된 CVCF 제어기를 검증하였다.

• Proceedings of the KIPE Conference
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• 2015.07a
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• pp.327-328
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• 2015

본 논문은 불평형부하 조건에서 전력품질 향상을 위한 독립형 인버터의 다중루프 제어기를 설계하고 실험적으로 검증하였다. CVCF(Constant Voltage Constant Frequency) 제어를 수행하는 독립형 인버터에 정지좌표계 상에서 다중루프 특성을 갖는 전압-전류 제어기를 설계하여 불평형 부하에 안정적인 전력공급을 수행하도록 구현하였다. 전압 제어기는 기본파 및 선택적 고조파 제어특성을 갖는 PR(Proportional-Resonant, 비례-공진)제어기를, 전류제어기는 P제어기를 적용하였다. 50kW 독립형 인버터 프로토타입을 구현하였으며, 불평형부하 조건에서 기존 동기좌표계 PI 제어와 비교할 때, 선간전압 THD는 4.3%에서 0.9%, 불평형률은 4.3%에서 0.35%로 개선됨으로써 제안된 기법의 타당성을 검증하였다.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.34 no.7
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• pp.35-44
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• 2006

When control surface failure occurs, it is conventional to correct a current control or to transform to other control. In this paper, instead of adopting a conventional way, a reconfiguration method which compensate the failure with alternative control surface deflection, depending on the level of failure, by using neural network and PCH(Pseudo-Control Hedging). The Conroller is designed of inner-loop(SCAS : Stability Command Augmentation System) with DMI(Dynamic Model Inversion) and outer-loop with Y axis acceleration feedback for a coordinate turn. Additionally, double PCH method was adopted to prevent actuator saturation and input command was generated to compensate for failure. At the end, The feasibility of the method is validated with randomly selected failure scenarios.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.30 no.6
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• pp.101-106
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• 2002

In a bank-to-turn(BTT) missile, if a 2-gimbaled seeker was stabilized using a 2-axis rate gyro mounted along its primary axis, the change of line of sight(LOS) measured by the seeker would be induced by rolling effects due to bank-to-turn(BTT) steering as well as an actual change. It is observed that the body-coupled effects in a homing loop of BTT missile are mainly concerned with the spurious target maneuver and the coupling due to the rate gyro misalignments. In this paper, we formulate a simple linear BTT homing loop model with seeker model including each body-coupling. With the model, we analyze the effects of the couplings on the homing loop stability, and propose a direct linear compensator for the coupling to recover the stability.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.39 no.4
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• pp.332-340
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• 2011

A missile with a strapdown seeker should properly estimate line-of-sight(LOS) rate using its attitude information and the look angle of the seeker because LOS rate information in an inertial coordinate system, which is used for a proportional navigation(PN) homing guidance, can not be obtained directly. However, an unnecessary feedback loop(Parasite Loop) is formed in the guidance and control loop, and it may cause the guidance performance degradation or even the unstability of the system(Parasite Effect). This paper presents estimation methods for the LOS rate information and effective ways to minimize the parasite effect using Routh-Hurwitz stability criterion. Various numerical simulations are also included to verify the proposed methods.