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친환경 추진제를 이용한 200N급 엔진의 설계 및 성능에 관한 연구
이양석,전준수,황오식,고영성,김유,김선진,Lee, Yang-Suk,Jun, Jun-Su,Hwang, Oh-Sik,Ko, Young-Sung,Kim, Yoo,Kim, Sun-Jin 한국군사과학기술학회 2010 한국군사과학기술학회지 Vol.13 No.6
In the last decade, hydrogen peroxide has received renewed interest as a green propellant which is non-toxic, environmentally clean and relatively easy to handle. This study was performed to acquire the design technique and combustion performance of a 200N bi-propellant engine using hydrogen peroxide and kerosene. The engine which used a catalytic ignition method was designed and cold flow tests were carried out to investigate atomization characteristics. Combustion tests including a pulse mode operation were performed to investigate the combustion performance on various O/F ratios. The results showed that the combustion efficiency and the repeatability of the engine performance were enough to use as an essential database for the development of a high performance engine.
액체로켓엔진 연소기를 이용한 고고도 환경 모사용 디퓨저 시동특성 연구
이양석(Yang-Suk Lee),전준수(Jun-Su Jeon),고영성(Young-Sung Ko),김유(Yoo Kim),김선진(Sun-Jin Kim) 한국항공우주학회 2010 韓國航空宇宙學會誌 Vol.38 No.12
본 연구에서는 액체 로켓 엔진의 고온 연소 가스를 이용하여 축소형 고고도 환경 모사용 초음속 디퓨저 성능 실험을 수행하였다. 실험 장치는 연소실, 진공 챔버, 디퓨저로 구성되어 있다. 고고도 환경 모사 시험은 연소실 압력이 약 26, 29, 32barg 세 조건으로 수행하였고, 세가지 조건에서 모두 디퓨저는 성공적으로 시동되었으며 진공 챔버 압력이 약 140torr로 형성하였다. 이전의 상온 고압 가스를 이용한 디퓨저의 시동 특성과 비교하였을 때 시동 압력과 압력 분포 등의 시동 특성의 경향성은 유사하였으나, 고온 환경으로 인하여 진공 챔버에 형성되는 압력은 2배 정도 높은 것을 확인하였다. 본 연구 결과는 향후 실물형 고고도 환경 모사 시험 설비를 구축하는데 기초 자료로 활용될 수 있을 것으로 판단된다. Performance tests of a supersonic exhaust diffuser were conducted by using a liquid rocket engine for simulating high-altitude environment. The experimental setup consisted of a combustion chamber, a vacuum chamber and a diffuser. The combustion tests for simulating high-altitude environment were carried out at three cases by chamber pressure variation(26, 29, 32barg). The test results showed that the diffuser was started at all case and vacuum chamber pressures were approximately 140torr. The starting pressure using combustion gas was similar with that of cold gas, but the vacuum chamber pressure was relatively high because of high temperature in the vacuum chamber. The results of this test can be used as an essential database for the design of real-scale high-altitude simulation test facility in the future.
과산화수소-케로신 엔진을 이용한 지상 및 고고도 추력에 대한 실험적 연구
이양석(Yang-Suk Lee),김중일(Joong-Il Kim) 한국산학기술학회 2019 한국산학기술학회논문지 Vol.20 No.10
고농도 과산화수소와 케로신을 추진제로 하는 액체 로켓 엔진을 이용하여 수직형 연소 실험대에 고고도 모사용 디퓨저와 기 검증된 추력 측정 장치를 장착하여 지상 및 고고도 모사 연소 실험 설비를 구축하였으며, 고도에 따른 추력특성을 고찰하였다. 선행으로 고고도 모사용 디퓨저의 특성 및 시동압력을 검증하기 위하여 1:4.8 스케일로 축소한 디퓨저를 설계 및 제작하였다. 축소형 디퓨저는 질소 가스를 이용하여 cold flow test를 수행하여 성능 및 시동 특성을 확인하였으며, 그 결과 연소 실험용 디퓨저의 성능 안정성과 시동 특성을 확보하였다. 수직형 연소 실험대에 고고도 모사용 디퓨저와 추력 측정 장치를 장착하고, 시스템 저항에 대한 추력 보정식을 도출하였다. 추력 보정식은 실제 연소 실험전에 수행한 추력 단계별 실험과 진공 단계별 실험을 통하여 도출하였다. 작동 고도가 10km인 노즐을 설계, 제작하여 지상 연소 실험 및 고고도 모사 연소 실험을 수행하여 작동 고도 변화에 따른 추력 특성을 분석하였다. 추력 측정 장치에서 계측한 추력값을 이용하여 실제 추력을 각각의 보정식을 이용하여 계산하였다. Ground and high altitude simulated combustion experiments were conducted using a liquid rocket engine with hydrogen peroxide and kerosene as the propellant. A ground and high altitude simulated combustion test facility was constructed by installing a high altitude model diffuser and TMS (Thrust Measuring System) on a vertical combustion test bench. The thrust characteristics according to altitude were investigated using the combustion test equipment. The diffuser was designed on a 1:4.8 scale to verify the characteristics of the high diffusing diffuser and starting pressure. The cold flow tests were conducted using nitrogen gas, and the performance characteristics and starting characteristics of the scale down diffuser were verified. A diffuser and TMS were installed on the vertical combustion test bench, and the thrust correction equations for the system resistance were derived. The thrust correction equations were derived from the step test and vacuum step test before the actual hot firing test. Nozzles with an operating altitude of 10km were designed. Hot firing tests were conducted to analyze the thrust characteristics according to the operating altitude changes. The actual thrust was calculated using each correction equation with the thrust value measured by the TMS.
과산화수소/케로신을 이용한 다중 분사기 엔진 설계 및 수류 실험
이양석(Yang-Suk Lee),전준수(Jun-Su Jeon),고영성(Young-Sung Ko),김유(Yoo Kim),김선진(Sun-Jin Kim) 한국추진공학회 2012 한국추진공학회지 Vol.16 No.1
Multi-injector rocket engine using high-concentrated hydrogen peroxide and kerosene was designed and manufactured. Design requirements of a rocket engine were determined and main geometrical parameters of rocket engine were determined on the basis of fundament. Six coaxial swirl injectors were mounted on the multi-injector engine. Flow analysis in the hydrogen peroxide manifold was performed to minimize stagnation and recirculation zones. Finally, the optimized hydrogen peroxide manifold was manufactured and cold flow test was carried out to confirm mass flow rate per uni-element, spray pattern and atomization characteristics. The results of cold flow test showed that the mixing head design process was successful and enough to use as a essential database for the development of a full-scale engine.
정세용,이양석,Jung, Se-Yong,Lee, Yang-Suk 한국군사과학기술학회 2009 한국군사과학기술학회지 Vol.12 No.3
The experimental research was conducted to setup a performance prediction logic for the regenerative cooling system on a small scale liquid rocket engine using kerosene and LOX. Total heat flux of the combustion gas side was determined for the flow rate of the coolant, combustion pressure using the calorimeter thrust chamber. Based on the experimental investigation, a performance prediction scheme for the regenerative cooling system is setup in our own way. A performance prediction logic for the regenerative cooling system has been developed by the correction scheme of the combustion gas side. The key parameters determining the temperature limitation of the coolant are the mass flow rate of the coolant and the length of the combustion chamber and the nozzle. And the parameters to control the limitation of the usable wall temperature are the number of channels and wall thickness.
친환경 추진제인 과산화수소와 액체메탄의 활용 역사와 연구 동향
김선진(Sun-Jin Kim),이양석(Yang-Suk Lee),고영성(Young-Sung Ko) 한국추진공학회 2010 한국추진공학회지 Vol.14 No.4
Hydrogen peroxide(HP) and liquid methane have deserved renewed considerations as green propellants in recent years, because main design concerns in the development of the new generation propulsion system for spacecrafts are concentrated on low operation cost and environmental cleanness. Although HP has a long history of application to aerospace propulsion systems due to high density, mono-propellant characteristics and low toxicity, it had been replaced by hydrazine and liquid oxygen due to extreme performance requirement during the cold war. But HP has received a renewed interest due to its increased stability and many researches have been conducted to develop high performance LREs(Liquid Rocket Engines) using HP. Liquid methane has also received a new interest in rocket propulsion system for the future space exploration according to its possibility of ISRU(In-Situ Resource Utilization).
과산화수소/케로신을 사용하는 액체로켓엔진의 촉매 점화기 설계에 관한 연구
채병찬(Byoungchan Chae),이양석(Yang-Suk Lee),전준수(Jun-Su Jun),고영성(Young-Sung Ko) 한국추진공학회 2011 한국추진공학회지 Vol.15 No.6
An experimental study on design of a catalytic ignitor was performed to use an ignition source for a small bi-propellant liquid rocket engine which use hydrogen peroxide and kerosene as propellants. In the catalytic ignitor, hot gas of hydrogen peroxide which was decomposed by a catalyst induced autoignition of kerosene. Mass flow rate and O/F ratio for the ignitor were calculated by CEA code. A combustion chamber which had a quartz window and thermocouples was manufactured to determine whether the ignition is successful. Ignition performance was investigated according to exit area of fixed rings and mixture ratio. Results showed that reliable ignition performance was achieved at non-choking exit area of fixed ring and O/F ratio of 6~8.
소형 가스터빈 연소기 고공환경 점화 시험 설비 구축 및 검증 실험
김태완(Tae-Woan Kim),이양석(Yang-Suk Lee),김기우(Ki-Woo Kim),김보연(Bo-Yean Kim),고영성(Young-Sung Ko),김선진(Sunjin Kim),김형모(Hyung-Mo KIM),정용운(Yong-Wun Jung) 한국추진공학회 2010 한국추진공학회지 Vol.14 No.3
A small high altitude test facility has been developed to investigate ignition performance of a small gas-turbine combustor under high altitude conditions. Supersonic diffusers and a heat exchanger were used to perform a low pressure and a low temperature condition, respectively. Experimental results showed that the low pressure environment could be controlled by upstream pressure of primary nozzle flow and low temperature environment by mixture ratio of cooled air and ambient air. Ignition performance tests were performed to verify the performance of the facility under simulated high altitude conditions. Conclusively, it was proven that the test facility could be used for ignition performance test of a small gas-turbine combustor under high altitude condition of approximately 6,100m.