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Huh, Jeongmoo,Seo, Daeban,Kwon, Sejin Elsevier 2017 Sensors and actuators. A Physical Vol.263 No.-
<P><B>Abstract</B></P> <P>This paper reports a feasibility study of regenerative micro-cooling channels in a liquid microthruster composed of thermally fragile materials. Glass, which is among the most thermally insulating materials, has been used as microthruster fabrication material to suppress excessive heat loss in micro scale thruster. However, the frangibility of glass has remained a challenge to be solved. To thermally manage the fragile structure, the use of regenerative micro-cooling channels in a microthruster is suggested in this work, and the feasibility was tested through design, fabrication and experimental performance of a glass microthruster with microchannels. Nine photosensitive glass layers were wet etched and integrated to fabricate the microthruster. Before integration of the layers, a fabricated Pt/Al<SUB>2</SUB>O<SUB>3</SUB> catalyst was inserted into the chamber of the microthruster for propellant decomposition. Hydrogen peroxide (90wt%) was used as a monopropellant and served as the working fluid for regenerative cooling. A liquid microthruster with micro-cooling channels was successfully fabricated with a photosensitive glass MEMS process. Experimental performance tests were conducted while measuring the microthruster chamber pressure, chamber temperature, and surface temperatures. The test results showed normal operation of the microthruster, which had an estimated thrust of approximately 48 mN and temperature efficiency of approximately 41%. The decreasing surface temperatures of the microthruster during thruster operation successfully validated the cooling effect of the micro-cooling channels and demonstrated their practicality for the regenerative cooling of liquid microthrusters.</P> <P><B>Highlights</B></P> <P> <UL> <LI> The use of micro-cooling channels to cool a liquid propellant microthruster is suggested. </LI> <LI> A liquid propellant microthruster with micro-cooling channels was fabricated with a photosensitive glass MEMS process. </LI> <LI> The feasibility of using micro-cooling channels was successfully validated through the design, fabrication and testing of a liquid propellant microthruster. </LI> </UL> </P>
Development of a University-Based Simplified H₂O₂/PE Hybrid Sounding Rocket at KAIST
Jeongmoo Huh,Byeonguk Ahn,Youngil Kim,Hyunki Song,Hosung Yoon,Sejin Kwon 한국항공우주학회 2017 International Journal of Aeronautical and Space Sc Vol.18 No.3
This paper reports development process of a university-based sounding rocket using simplified hybrid rocket propulsion system for low-altitude flight application. A hybrid propulsion system was tried to be designed with as few components as possible for more economical, simpler and safer propulsion system, which is essential for the small scale sounding rocket operation as a CanSat carrier. Using blow-down feeding system and catalytic ignition as combustion starter, 250 N class hybrid rocket system was composed of three components: a composite tank, valves, and a thruster. With a composite tank filled with both hydrogen peroxide(H₂O₂) as an oxidizer and nitrogen gas(N₂) as a pressurant, the feeding pressure was operated in blowdown mode during thruster operation. The MnO₂/Al₂O₃ catalyst was fabricated for propellant decomposition, and ground test of propulsion system showed the almost theoretical temperature of decomposed H₂O₂ at the catalyst reactor, indicating sufficient catalyst efficiency for propellant decomposition. Auto-ignition of the high density polyethylene(HDPE) fuel grain successfully occurred by the decomposed H₂O₂ product without additional installation of any ignition devices. Performance test result was well matched with numerical internal ballistics conducted prior to the experimental propulsion system ground test. A sounding rocket using the developed hybrid rocket was designed, fabricated, flight simulated and launch tested. Six degree-of-freedom trajectory estimation code was developed and the comparison result between expected and experimental trajectory validated the accuracy of the developed trajectory estimation code. The fabricated sounding rocket was successfully launched showing the effectiveness of the simplified hybrid rocket propulsion system.
Huh, Jeongmoo,Ahn, Byeonguk,Kim, Youngil,Song, Hyunki,Yoon, Hosung,Kwon, Sejin The Korean Society for Aeronautical and Space Scie 2017 International Journal of Aeronautical and Space Sc Vol.18 No.3
This paper reports development process of a university-based sounding rocket using simplified hybrid rocket propulsion system for low-altitude flight application. A hybrid propulsion system was tried to be designed with as few components as possible for more economical, simpler and safer propulsion system, which is essential for the small scale sounding rocket operation as a CanSat carrier. Using blow-down feeding system and catalytic ignition as combustion starter, 250 N class hybrid rocket system was composed of three components: a composite tank, valves, and a thruster. With a composite tank filled with both hydrogen peroxide($H_2O_2$) as an oxidizer and nitrogen gas($N_2$) as a pressurant, the feeding pressure was operated in blowdown mode during thruster operation. The $MnO_2/Al_2O_3$ catalyst was fabricated for propellant decomposition, and ground test of propulsion system showed the almost theoretical temperature of decomposed $H_2O_2$ at the catalyst reactor, indicating sufficient catalyst efficiency for propellant decomposition. Auto-ignition of the high density polyethylene(HDPE) fuel grain successfully occurred by the decomposed $H_2O_2$ product without additional installation of any ignition devices. Performance test result was well matched with numerical internal ballistics conducted prior to the experimental propulsion system ground test. A sounding rocket using the developed hybrid rocket was designed, fabricated, flight simulated and launch tested. Six degree-of-freedom trajectory estimation code was developed and the comparison result between expected and experimental trajectory validated the accuracy of the developed trajectory estimation code. The fabricated sounding rocket was successfully launched showing the effectiveness of the simplified hybrid rocket propulsion system.
허정무(Jeongmoo Huh),김영일(Youngil Kim),안병욱(Byeonguk Ahn),정우석(Woosuk Jung),김현탁(Hyuntak Kim),최석민(Sukmin Choi),송현기(Hyunki Song),임재민(Jaemin Lim),유기정(Kijeong Yu),김종학(Jonghak Kim),윤호성(Hosung Yoon),권세진(Sejin Kwo 한국추진공학회 2016 한국추진공학회 학술대회논문집 Vol.2016 No.12
250N 급 하이브리드 로켓 추진기관을 이용한 소형 과학로켓 개발과 비행시험이 진행되었다. 시스템 간편성을 확보하기 위해 추가적인 가압기체 탱크나 압력조절기 없이 블로우다운 가압 방식을 사용하였다. 촉매 점화 방식의 하이브리드 추진기관을 구성하여 내탄도 예측과 지상테스트가 진행되었으며 성공적인 자연점화와 내탄도 예측과 유사한 추진성능을 파악하였고 점화신뢰성을 확보하였다. 비행시험에 앞서 비행 시뮬레이션이 진행되었고, 비행시험결과 로켓이 성공적으로 비행하여 예측된 값과 유사한 최고점 부근 고도 95m, 비행시간 10.5 초의 결과를 보여주어, 구성된 추진기관의 시스템 간편성과 높은 점화신뢰성, 그리고 소형 과학로켓의 추진기관으로써 유효성을 성공적으로 입증하였다. Small scale sounding rocket was developed and flight tested using 250N class hydrogen peroxide/polyethylene hybrid rocket propulsion system. Pressure-fed system was used for system simplicity, which was desired for small scale sounding rocket. Internal ballistics and ground test were conducted for catalyst ignited hybrid rocket stand-alone system, and the results were well matched with successful auto-ignition and reliability. Experimental flight test of the sounding rocket showed successful flight with 95 m maximum altitude and 10.5 sec flight time, which was suitable for the flight simulation. These results show effectiveness of blow-down feeding and catalyst ignition hybrid rocket propulsion system for small scale sounding rocket application.
허정무(Jeongmoo Huh),안병욱(Byeonguk Ahn),김영일(Youngil Kim),송현기(Hyunki Song),고수정(Sujeong Ko),윤호성(Hosung Yoon),권세진(Sejin Kwon) 한국추진공학회 2016 한국추진공학회 학술대회논문집 Vol.2016 No.5
250 N 급 하이브리드 로켓을 추진기관으로 이용하는 소형 과학로켓의 시스템 구성과 발사 준비가 진행되었다. 산화제와 연료로는 각각 90wt% 과산화수소와 폴리에틸렌이 사용되었다. 추진기관은 추력기, 공급관, 산화제탱크, 밸브류로 구성되었으며, 케이싱은 가볍고 투명하여 내부가 보이는 폴리카보네이트로 제작되어 이후 알루미늄과 스테인리스 스틸 프레임으로 추진기관에 체결되었다. 과학로켓 발사를 위한 발사대를 제작하여 비행 테스트 준비를 완료하였다. 제작된 과학로켓은 궤적 시뮬레이션 결과 최대고도 약 700 m의 운용 범위를 보여주었다. Small scale sounding rocket system was configured using 250N class hybrid rocket and prepared for flight testing. 90wt% hydrogen peroxide was used as an oxidizer and high density polyethylene was used as a fuel grain. Propulsion system was composed with thruster, feeding line, oxidizer tank, and valves. Casing was manufactured using polycarbonate and assembled with propulsion system using metal frame. Launch pad was also manufactured and flight testing was successfully prepared. Fabricated sounding rocket was expected to reach approximately 700m altitude based on trajectory estimation.
소형 사운딩 로켓 적용을 위한 H2O2/PE 하이브리드 로켓 시스템
허정무(Jeongmoo Huh),정상우(Sangwoo Jung),김영일(Youngil Kim),안병욱(Byeonguk Ahn),최석민(Sukmin Choi),이재완(Jaewan Lee),송현기(Hyunki Song),김종학(Jonghak Kim),윤호성(Hosung Yoon),권세진(Sejin Kwon) 한국추진공학회 2015 한국추진공학회 학술대회논문집 Vol.2015 No.11
250 N 급 과산화수소/폴리에틸렌 하이브리드 로켓을 추진시스템으로 이용하는 캔 위성 발사를 위한 소형 과학로켓 설계가 진행되었다. 내탄도 계산을 통한 추진기관 성능예측, 추진기관 제작 및 지상성능 시험이 수행되었으며, 과학로켓 운용 고도 계산을 위한 궤적 시뮬레이션 코드가 작성되었고 공개 코드계산결과와 비교를 통하여 정확성이 검증되었다. 추진 기관의 지상 시험결과는 촉매점화방식으로 촉매반응기 외에 추가적인 점화시스템이 요구되지 않는 과산화수소/폴리에틸렌 하이브리드 로켓 시스템의 높은 점화 신뢰성과 간소화된 시스템 구성이 가능함을 보여주었으며, 설계된 하이브리드 소형 과학로켓은 궤적 시뮬레이션 결과 최대 고도 700m의 운용 가능성을 보여주었다. Small scale sounding rocket as CanSat carrier was conceptually designed using 250 N class H2O2/PE hybrid rocket. Propulsion system was manufactured and ground tested. Internal ballistics was calculated for performance estimation of the propulsion system. The performance test of the propulsion system was successfully conducted showing ignition reliability and system simplicity, using catalyst ignition hybrid rocket with blow-down feeding system. For flight simulation of the designed sounding rocket, trajectory code was developed and validated. The trajectory simulation was conducted with expected altitude of 700 m for the designed sounding rocket.
Measuring the Reaction Rate of Hypergolic Propellants with a Microelectromechanical Systems Reactor
Kang, Hongjae,Huh, Jeongmoo,Kwon, Sejin American Institute of Aeronautics and Astronautics 2017 Journal of spacecraft and rockets Vol.54 No.2
<P>A novel experimental method to evaluate the ignition characteristics of hypergolic propellants was investigated. Three hypergolic liquid fuels with different ignition characteristics were prepared, and their reaction rates with hydrogen peroxide were measured in a microreactor fabricated using a microelectromechanical systems lithography process. To prevent evaporation and boiling by the reaction heat, a low-concentration of hydrogen peroxide of 34.5wt % was used as a reference oxidizer. The reaction rates of the hypergolic fuels were obtained from temperature measurements in the microreactor and correlated with the ignition delays that were separately measured by drop tests with high-test peroxide. The reaction rate measurements were fully compatible with the drop test results and proved to be useful as an alternative method for the selection of hypergolic propellants by ignition characteristics.</P>