• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.41 no.6
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• pp.473-480
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• 2013

Researches for various lunar landing technologies are in progress for the lunar exploration program planned for early 2020s in Korea. This paper shows the integrated simulation for safe lunar landing guidance/control system in powered descent phase. Generally, the lunar lander uses on/off(bang-bang) controller to control the RCS jet thrusters instead of proportional controller. In this paper, the on/off controller using phase-plane switching function, and thruster selection algorithm to control sixteen thrusters are applied. Also additional guidance commands are calculated by a proposed fuzzy logic guidance algorithm. The simulation results show that lunar lander can follow a reference trajectory which is generated by optimization method, then land on the surface safely.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.32 no.3
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• pp.88-94
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• 2004

The KSR-III rocket is a liquid propellant sounding rocket and thrust vector control actuators and cold gas thrusters are used to control pitch and yaw, roll attitude respectively during thrusting phase. In this paper, the structure of designed attitude controller and gain scheduling, results of stability analysis for KSR-III rocket are presented. The attitude controller is implemented with flight software in the domestically developed INS and successfully performed its function in the flight test. The flight data are coincident with simulation results.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.34 no.1
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• pp.95-102
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• 2006

The PPT being currently developed for the flight model represents a significant leap in techniques and technology compared to the previous flight ones. The electrical energy to be charged in the pulsed plasma thruster (PPT) is a very important aspect to provide an uniform impulse bit ,, and a specific impulse ,, for satellite attitude control. In this paper, we propose a nonlinear control technique and a stability analysis based on the Lyapunov function for the pulsed plasma thruster. Specifically, the proposed control law guarantees to charge and discharge the electrical energy generated from the power processing unit (PPU) within the specified time.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.43 no.8
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• pp.722-730
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• 2015

Drive mechanism of multi-axis pintle thrusters for DCS(Divert Control System) was designed to meet the needs of minimizing the number of driving motors. In this study, preliminary model was designed in order to implement appropriate pressure control and thrust distribution. Based on the preliminary model study, the drive mechanism for DCS multi-axis pintle thrusters using piston was designed and evaluated by using AMESim software. Results show that three driving motors are enough to actuate four pintle thrusters.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.32 no.9
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• pp.72-79
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• 2004

A method for thruster fault diagnosis for launch vehicle upper stage was developed. In order to protect the launch vehicle against the occurrence of faults, it is necessary to detect and identify the fault, as well as to reconfigure the controller of the vehicle. Considering the upper stage launch vehicle using reaction control system, an analytical method was adopted in order to detect the fault occurred in thruster. The fault detection scheme can be applied to the system regardless of the form of thruster fault occurred - leakage or lock-out. Results from processor-in-the-loop simulation are provided to demonstrate the validity of this fault detection and isolation scheme for the upper stage launch vehicle.

• Journal of the Korean Society of Propulsion Engineers
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• v.17 no.2
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• pp.63-70
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• 2013

To analyze flow characteristics of the side jet thruster with 4 shutters and rectangular nozzles, a 3-D simulation has been implemented. Numerical calculations for two rotation angles of the shutter, have been conducted. Internal recirculation in a chamber and asymmetric flow structure in a nozzle were observed. In addition, the more shutter rotated, the more asymmetries of flow increased, and this phenomena resulted in thrust bias. The degrees of thrust bias and thrust performance with the rotation angles of the shutter were predicted and compared with theoretical thrust.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 1998.04a
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• pp.30-30
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• 1998

다목적 실용위성의 자세 제어용으로 장착되는 하이드라진 단일 추진제 추력기의 연소성능 특성을 살펴보았다. 일차적으로, 연소기내의 화학 평형 계산을 통하여 하이드라진 연소 생성물인 암모니아의 분해율과 초기 엔탈피 수준에 따른 추력기의 성능 특성을 살펴보았다. 다음 순서로, 연소기 내의 비평형 화학 반응 계산을 통하여 연소진행 시간에 화학 조성 및 성능 특성을 살펴 볼 수 있었으며, 최종적으로, 점성 및 비점성, 동결 및 비평형 화학 반응 해석에 따른 성능 특성 변화를 살펴보았다. 본 연구 결과로부터 각 작동 변수에 따른 추격기 성능 특성의 변화를 이해할 수 있었으며, 이는 단일 추진제 추력기 연소실의 구성, 설계 및 추력기의 운용 조건 설정에 중요한 자료로 이용될 수 있을 것으로 기대된다.

• Proceedings of the KIPE Conference
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• 2002.07a
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• pp.133-137
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• 2002

리니어 모터 중에서 위치 정밀도가 우수하고 개루프 제어가 가능한 리니어 펄스모터(LPM : Linear Pulse Motor)에 대해 추력 성능 및 위치제어 정밀도 향상을 위해서 자기등가회로 모델링법에 의해 추력을 주로 계산하고 있다. 퍼미언스(Permeance) 모델법, 수치해석법(FEM), 속도 기전력법 등을 이용하여 자속밀도(B)를 구하는데 있어 해석적인 추력편차가 존재한다. 정확한 추력을 구하기 위해서는 이들 각각의 방법을 이용한 로렌츠 포스법(Lorentz Force Method), 자기수반 에너지법, Maxwell 응력법 등으로 추력을 계산하여 비교함으로써 모터용량, 파라미터 등의 각종 성능 비교시 중요 요소가 될 수 있다. 이에 본 연구는 위의 방법으로 LPM의 추력을 수식화하여 값을 얻고, 실제 추력값을 측정하여 얻어진 값과 비교 분석하고자한다. 이러므로 보다 정확한 LPM 추력 계산 방법을 얻을 수 있을 것으로 기대된다.

• Journal of the Korean Society of Propulsion Engineers
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• v.10 no.4
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• pp.77-84
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• 2006

With the development of micro/nano spacecraft, concepts of micro propulsion are introduced for orbit transfer and drag compensation as well as attitude control. Micro solid propellant thruster has been attention as one of possible solution for micro thruster. In this paper, micro solid propellant thruster is introduced and research on basic components of a micro solid propellant thruster is reported. Micro Pt igniter was fabricated through negative patterning and quantitative effect of geometry was estimated. The characteristic of HTPB/AP solid propellant was investigated to measure the homing velocity. A combustion chamber was fabricated by means of anisotropic etching of photosensitive glass. Finally, micro solid propellant thrusters having various geometries were fabricated and tested.

• Journal of the Korean Society of Propulsion Engineers
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• v.11 no.1
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• pp.18-26
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• 2007

A development of monopropellant thruster for microsatellite that uses concentrated hydrogen peroxide is described. Catalyst, the most important component in the thruster, was prepared and coated on a monolith honeycomb. Performance evaluation of thruster was peformed by considering the efficiency of characteristic velocity and ignition delay. As a result, 96.0% of $C^*$ efficiency was obtained at designed propellant flowrate and steady state operating condition.