• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.04a
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• pp.281-285
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• 2007

The flow characteristics in the thruster should be analyzed considering its geometry and the pressure ratio to estimate its performance and etc. This paper suggests the computational result of an axisymmetric real nozzle for the altitude control of a satellite to find out the application limit that the assumption of continuum mechanics holds. The steady non-reacted compressible flow field in the unstructured grid system is computed and analyzed with varying the environmental pressure (or the degree of vacuum) under the fixed pressure ratio in a real thruster of which the area ratio of exit to throat is 56. The assumption of the continuum mechanics is not approved when the environmental pressure is reduced less than $10^{-3}$ atm.

• The Journal of Korea Institute of Information, Electronics, and Communication Technology
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• v.17 no.5
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• pp.313-322
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• 2024

To ensure the stable flight of aerospace electric propulsion systems, it is necessary to measure the supplied flow rate and control it to an appropriate level. However, conventional flow sensors are costly and face limitations in space environments, making heat-based flow estimation a promising alternative. In this study, the Taguchi method, one of the experimental design techniques, was applied to perform thermal analysis simulations using Ansys under various variables and conditions. The Taguchi method was used to set heat supply and the positioning of inlet and outlet temperature sensors as key variables, and the optimal distance conditions were derived. Thermal analysis was conducted through Ansys to analyze the flow estimation results under each experimental condition. Therefore, this study demonstrates the practicality of the heat-based flow estimation method for fuel management systems in electric propulsion systems, presenting a new approach for the efficient and economical operation of electric propulsion. Additionally, this research contributes to the development of fuel management technologies that can be effectively utilized in the constrained environment of space.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 1998.10a
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• pp.9-9
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• 1998

다목적 실용위성의 궤도전이 및 위성체 자세제어를 위한 추진시스템의 설계요소에는 구조적 안전성, 우주환경에서의 열제어를 위한 회로 및 구성하드웨어 설계, 연료계통 맥압강하를 위한 장치설계 및 추력기 배기가스 영향을 고려한 형상설계 등이 있으며, 설계검증을 위해 부분해석이 수행된다. 또한 발사환경과 우주 궤도환경에서의 추진시스템 성능평가를 위한 연제어계 기능시험, 압력인증시험, 청정도시험 및 내부/외부 누설시험이 수행된다. 본 논문에서는 추진시스템 설계 및 조립공정에 대해 기술하였고, 시험분석을 통해 시스템의 설계 및 조립공정상의 신뢰성을 검증 분석하였다.

• 제어로봇시스템학회:학술대회논문집
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• 2000.10a
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• pp.3-3
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• 2000

Limit cycle analysis of attitude control system using gas jet thrusters is performed. Schmitt-Trigger and PD control laws are applied and solenoid valve time delay is considered. Phase plane method is used for calculation of characteristics of limit cycle. Important characteristics of resultant limit cycle such as frequency, amplitude, maximum rate, and duty ratio could be expressed analytically by proposed method.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.11a
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• pp.7-10
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• 2006

In order to inject satellites into a target orbit, launch vehicles should have a precise attitude and control system capable of controlling three axises of pitch, yaw and roll. For launch vehicles, there are two types of attitude control system currently in popular use; the first one is a cold gas method, and the other is a liquid propulsion system using a single and dual property propellant. The purpose of this paper is to analyze the characteristics of thrust control system using said propellant, thereby providing for a rationale for its application to the upper stages of launch vehicles, in terms of the simplicity of the system, economics of structure and operation.

• Journal of the Korean Society of Propulsion Engineers
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• v.17 no.4
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• pp.1-9
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• 2013

An active anti-ship missile decoy system was designed conceptually to analyze propulsion system requirements and feasibility to use a liquid bi-propellant rocket engine. Overall mass, size, and shape were assumed referring to specifications of Nulka which was developed by US and Australia in 1990s. The propulsion system was assumed to be a 1,000 N-class $H_2O_2$/kerosene rocket engine with a pressurized feed system. A three-degree-of-freedom optimal trajectory was calculated based on the assumptions, and mass budget was designed from the calculation results. It was found that the requirements for the propulsion system is that it shall be operated more than 100 sec; it shall be re-ignitable; it shall have a throttle capability of a range from 35% to 100% when the maximum thrust at sea level is 1,000 N.

• Journal of the Korean Society of Propulsion Engineers
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• v.24 no.6
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• pp.126-142
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• 2020

This study aims to provide basic design data for a pintle injector and its combustion device through case study on the research and development of combustion devices to which a liquid-liquid pintle injector was applied. From data analysis, it was possible to provide the initial dimension of the combustion chamber and pintle injector based on the engine thrust, and the geometric characteristics of the high-efficiency injector. In addition, the pintle tip heat damage prevention mechanism and materials, face-shutoff pintle injector implementation method, and central propellant selection criteria were summarized. Theses results will be used as basic data for the design criteria of an initial pintle injector combustion device.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2005.11a
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• pp.196-200
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• 2005

A theoretical model for the calculation of chemical equilibrium composition of propellant combustion product is briefly presented for the performance analysis of monopropellant hydrazine rocket engine. Analysis result is compared to that of test and evaluation of 1-lbf class thruster and is scrutinized primarily from the view point of ammonia dissociation fraction. Chemical equilibrium composition and average molecular weight is additionally depicted according to the variation of propellant inlet pressures and the varying nozzle area ratio. The theoretical analysis is tried as a way of derivation of design parameters for mid- and large-thrust class of monopropellant rocket engines.

• The Bulletin of The Korean Astronomical Society
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• v.37 no.2
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• pp.163.2-163.2
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• 2012

다목적실용위성 3호는 2012년 5월 발사되어, 위성 기능점검을 위한 시험을 성공적으로 완료하였다. 위성이 발사체로부터 분리된 이후 임무궤도(고도 685km, 승교점 지방시 13시 30분을 갖는 태양동기궤도)를 획득하기 위해서는 궤도조정이 필요하다. 본 논문에서는 다목적실용위성 3호의 초기운영 기간 동안 수행한 총 10번의 궤도조정 계획 및 결과에 대해 기술하였다. 궤도조정 1 단계에서는 궤도조정 절차 및 기능을 점검하기 위해 6번의 시험 궤도조정을 순차적으로 수행하였고 이후 2 단계에서는 임무궤도 진입을 위해 4번의 궤도조정을 실시하였다. 궤도조정을 위해서는 원하는 추력분사 방향을 맞추기 위해 롤 방향 또는 피치 방향의 자세제어가 필요한데, 추력기를 사용하여 자세를 기동하는 모드(Del-V Mode)와 휠을 사용하여 자세를 기동하는 모드(Fine Del-V Mode)로 구분된다. 시험 궤도조정에서는 우선적으로 두 가지 모드에 대한 모드전환 시험을 실시하여 위성체 및 지상국 운영절차에 대한 이상 유무를 점검하였고, 이후 추력기 분사량을 10초로 설정하여 예측 대비 실제 궤도변경 결과값을 확인하였다. 시험 궤도조정의 결과를 토대로 본 궤도조정에서는 임무궤도를 획득하기 위한 경사각 조정 및 고도 조정을 수행하였다. 경사각 조정 시에는 승교점 지방시의 변화량을 줄이고, 이후 자연 교란력에 의한 궤도변화를 고려하여 목표궤도를 계획하였다. 또한, 고도 조정 단계에서는 연료 사용량 및 이심률 변화를 최소화 할 수 있도록 전형적인 호만 궤도천이 방식을 적용하였다. 궤도조정 결과 당초 목표한 값을 정확하게 달성하였고, 궤도조정 이후 궤도변화도 장기간 동안 임무궤도 범위를 유지함을 확인할 수 있었다.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.11a
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• pp.105-108
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• 2008

Nowadays, as there is so much interest in environment, hydrogen peroxide attracts attention as an eco-propellant. Hydrogen peroxide is widely used for mono-propellant of thruster, and oxidizer of bi-propellant rocket. Especially, it is used as mono-propellant of the thruster for attitude control of satellite and military weapons. So, the need of long time storage of hydrogen peroxide appears and storage test is required. In this paper, necessity of storage test of hydrogen peroxide and some conditions and methods are introduced. In addition, the results of storage tests under some condition are compared and analyzed.