• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2002.04a
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• pp.16-17
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• 2002

재생냉각은 엔진 경량화 및 높은 추력을 발생시킬 수 있으며 엔진이 장시간 작동할 경우에도 추력의 변화가 일어나지 않는 우수성으로 인해, 액체로켓엔진에서 보편적으로 사용되고 있는 냉각방식이며, 고성능 액체로켓엔진 개발에 있어서 핵심기술이다. 일반적으로 재생냉각 방식은 연소기 내벽에 형성된 냉각유로에 연료 또는 산화제를 흘려보내 고온고압의 연소실내에 온도 경계층을 생성시키면서 벽면온도를 적정온도 이내로 유지하는 것이 목적이며, 또한 냉각유로에서의 압력강하가 추진제 공급 시스템의 공급 압력의 한계값을 넘어서지 않도록 하며, 냉각후의 연료 또는 산화제의 열역학적 상태가 엔진 작동 조건에 적합하도록 제어하여야 한다.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.488-491
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• 2011

The thermo-structural analysis of the solid-fuel propulsion Nozzle is studied to estimate the thermo-structural safety of the metal nozzle. The thermal load is determined to be significantly large, Because the metal nozzle in a short combustion time is directly exposed to high pressure and temperature of combustion gas. Through a analysis result, the influence of a thermal load is estimated and henceforward a design data of thruster is used.

• Journal of Aerospace System Engineering
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• v.9 no.2
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• pp.13-18
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• 2015

Micro planar type liquid propellant thruster was fabricated by MEMS manufacturing process for micro/nano satellites applications. 90 wt.% hydrogen peroxide was used as propellant and for propellant decomposition, Pt/Al2O3 was used as catalyst. Micro thruster structure was made by 5 photosensitive glasses patterned with thruster component profiles. Objective thrust was 50 mN and required hydrogen peroxide mass flow was 2.1 ml/min, which was supplied by syringe pump and teflon tube in experimental test. Performance test said that average steady thrust was approximately 30 mN, around 60% of objective thrust, and transient time was about 5 sec. It is estimated that extended response time was due to high thermal energy loss of micro scale thruster and low enthalpy input by propellant mass flow.

• 유체기계공업학회:학술대회논문집
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• 2006.08a
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• pp.45-46
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• 2006

Usual LEO satellite for earth observation use a blowdown hydrazine monopropellant propulsion system for attitude hold and orbit maintenance. For precision control, thruster valve has very short closing time, but this can cause water hammering and pressure surge. Since water hammering and pressure surge can cause damage of propulsion system and ununiform thrust, Thruster valve closing is one of the special concern during satellite propulsion system design. In this paper, an analysis for propellant feed system is conducted using the method of characteristics. The results represent water hammer effect is negligible even at the worst case and pressure surge can be decreased effectively with a trim orifice.

• Aerospace Engineering and Technology
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• v.2 no.1
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• pp.124-132
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• 2003

The function and performance test of the developed sub-system should be conducted before integrating into the total system. For the successful achievement of the flight test, the integration tests of sub-systems must be performed to eliminate and prevent abnormal conditions, which may happen due to ground problems, Electro-Magnetic Interference noises, and so on. This paper describes the certification test procedure and results for the roll axis attitude control system of Korea Sounding Rocket-III.

• 한국전산유체공학회:학술대회논문집
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• 2006.10a
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• pp.101-105
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• 2006

The computational fluid dynamic analysis has been conducted for the thermo-chemical flow field in an arcjet thruster with mono-propellant Hydrazine (N2H4) as a working fluid. The Reynolds Averaged Navier-Stokes (RANS) equations are modified to analyze compressible flows with the thermal radiation and electric field. The Maxwell equation, which is loosely coupled with the fluid dynamic equations through the Ohm heating and Lorentz forces, is adopted to analyze the electric field induced by the electric arc. The chemical reactions of Hydrazine were assumed to be infinitely fast due to the high temperature field inside the arcjet thruster. The chemical and the thermal radiation models for the nitrogen-hydrogen mixture and optically thick media respectively, were incorporated with the fluid dynamic equations. The results show that performance indices of the arcjet thruster with 1kW arc heating are improved by amount of 180% in thrust and 200% in specific impulse more than frozen flow. In addition to thermo-physical process inside the arcjet thruster is understood from the flow field results.

• Proceedings of the KSME Conference
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• 2003.11a
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• pp.1578-1583
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• 2003

Time-Optimal solutions for attitude control with reaction wheels as well as with thrusters are studied. The suggested varying-time-sharing ratio thrusting is found to reduce the maneuvering time enormously. The hybrid control such as sequential hybrid and simultaneous hybrid with reaction wheels and thrusters are considered. The results show that simultaneous hybrid method reduces the maneuver time very much. Spacecraft model is KOrea Multi-Purpose SATellite(KOMPSAT)-II, which is being developed by KARI in KOREA as an agile maneuvering satellite.

• Proceedings of the KIEE Conference
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• 2002.11d
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• pp.30-32
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• 2002

선형동기전동기는 고추력, 고정밀 운전을 요하는 시스템에 많이 사용되고 있는데, 추력 증가에 따른 추력 리플의 증가는 고정밀 위치제어에 어려움을 주기도 한다. 철심형 선형동기전동기에서는 자석 배열과 철심 치형상의 상호 관계에 의해 발생하는 디텐트력의 영향이 매우 크므로 이를 저감하기 위한 철심 구조의 설계가 필요하다. 본 논문에서는 실제 제작성을 고려하여 치와 철심부의 챔퍼링, 길이 변경, 자석 배열의 스큐 적용 등을 통해 선형동기전동기의 디텐트력을 저감하기위한 설계법을 제시하고 그 효과를 시험기의 제작 및 실험을 통해 증명하였다.

• Journal of the Korean Society for Aviation and Aeronautics
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• v.15 no.3
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• pp.70-76
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• 2007

The benefits for using reduced takeoff thrust are many, ranging from lower maintenance and operating costs to improved engine and dispatch reliabilities. Some pilots, however, are apprehensive about using reduced thrust. They are particularly reluctant to use the maximum permissible level of reduced thrust. Two common arguments are (1)If reduced thrust is used, then the airplane will not be able to clear the obstacles if an engine fails during takeoff, and (2)If the maximum allowable assumed temperature is used, then there will be no stopping margin left if the takeoff is aborted. There is the notion that using reduced thrust sacrifices safety. The intent of this discussion is to: (1)Show that reduced thrust performance meets all regulatory requirements (2)Show that the Assumed Temperature method includes inherent extra performance margins (3)Show how to maximize performance margins while maximizing thrust reduction.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.04a
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• pp.49-52
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• 2007

Vacuum facility is required for high altitude space environment test to develop small thruster. We, at Chungnam National University, developed vacuum test facility up to $10^{-5}$ torr to simulate $100{\sim}120km$ altitude environment. Prior to operation, we predict vacuum pump performances and present preliminary calculation and experiments.