기상조절용 하이브리드 로켓의 실험 설계 및 활용연구

차주완,김부요,노용훈,고아름,김선희,박동오,박지만,구해정,장기호,이용희,김수종,Joo Wan Cha,Bu-Yo Kim,Miloslav Belorid,Yonghun Ro,A-Reum Ko,Sun Hee Kim,Dong-Ho Park,Ji Man Park,Hae Jung Koo,Ki-Ho Chang,Hong Hee Lee,Soojong Kim 한국기상학회 2024 대기 Vol.34 No.2

The National Institute of Meteorological Sciences in Korea has developed the Weather Modification Hybrid Rocket (WMHR), an advanced system that offers enhanced stability and cost-effectiveness over conventional solid-fuel rockets. Designed for precise operation, the WMHR enables accurate control over the ejection altitude of pyrotechnics by modulating the quantity of oxidizer, facilitating specific cloud seeding at various atmospheric layers. Furthermore, the rate of descent for pyrotechnic devices can be adjusted by modifying parachute sizes, allowing for controlled dispersion time and concentration of seeding agents. The rocket's configuration also supports adjustments in the pyrotechnic device's capacity, permitting tailored seeding agent deployment. This innovation reflects significant technical progression and collaborations with local manufacturers, in addition to efforts to secure testing sites and address hybrid rocket production challenges. Notable outcomes of this project include the creation of a national framework for weather modification technology utilizing hybrid rockets, enhanced cloud seeding methods, and the potential for broader meteorological application of hybrid rockets beyond precipitation augmentation. An illustrative case study confirmed the WMHR's operational effectiveness, although the impact on cloud seeding was limited by unfavorable weather conditions. This experience has provided valuable insights and affirmed the system's potential for varied uses, such as weather modification and deploying high-altitude meteorological sensors. Nevertheless, the expansion of civilian weather rocket experiments in Korea faces challenges due to inadequate infrastructure and regulatory limitations, underscoring the urgent need for advancements in these areas.

PE/LN<SUB>2</SUB>O 소형 하이브리드 시험발사 및 비행궤도 분석

이민호(Minho Lee),김재욱(Kim Jaewook),신준호(Sin Junho),엄용경(Um Yongkyung),오유진(Oh Yujin),이선재(Lee Sunjae),정영규(Jung Youngkyu),조재윤(Jo Jaeyun),최영록(Choi Youngrok),이정표(Jung-pyo Lee),김진곤(Jinkon Kim) 한국추진공학회 2010 한국추진공학회 학술대회논문집 Vol.2010 No.11

본 연구는 소형 하이브리드 로켓의 제작 및 발사를 통해 하이브리드 발사체의 기초 발사기술을 확보하는 것이다. 연료로는 HDPE, 산화제는 LN2O를 적용하였으며, 알루미늄 외형재질의 설계 총 무게 12.5 kg, 외경 114 mm, 전장 1.8 m의 소형 하이브리드 로켓을 설계하였다. 로켓의 목표고도는 500 m로 설정하였고, 목표 추력 50 kgf와 연소시간 2.5 초의 연료 그레인 및 인젝터를 설계 및 제작 하였다. 발사 후 실시간 압력 및 속도 등의 데이터를 수집하기 위한 데이터 획득장치와 로켓의 안정적인 회수를 위한 스프링-모터를 이용한 사출장치를 제작·탑재 하였다. 로켓의 발사는 성공적으로 수행되었으나, 로켓 중량의 증가 및 추력의 부족으로 설계 최고 고도에는 로켓이 미치지 못했고, 로켓의 비행 궤도를 분석하였다. The purpose of this study is to develop basic technology of hybrid rocket vehicle by constructed and launched. This small hybrid rocket using HDPE/LN2O and Aluminium steel for its body (weight : 12.5 kg, diameter : 114 mm, height : 1.8 m) was designed. The fuel grain and injector were designed for 50 kgf thrust and burning time 2.5 sec. This rocket was loaded the data acquisition device for obtaining data of pressure and velocity during its flying and equiped an automatic ejector system using spring/motor and timer to collect the rocket more safely after launching. It was launched successfully, but found some problem that the rocket"s weight was heavier than expected and the thrust was not enough to reach the designed altitude and analyzed its flight path way.

과산화수소 폴리에틸렌 하이브리드 로켓 설계 및 평가

강신재(Shinjae Kang),이다혜(Dahae Lee),이은광(Eunkwang Lee),권세진(Sejin Kwon) 한국추진공학회 2015 한국추진공학회 학술대회논문집 Vol.2015 No.11

과산화수소는 단일추진제나 이원추진제 액체 로켓 엔진에 주로 사용되어왔다. 액체, 고체 로켓에 비해 높은 안정성을 갖기 때문에 상대적으로 적은 개발비용이 요구되는 하이브리드 로켓의 특성이 알려지면서 여러 연구가 진행되고 있다. 과산화수소는 높은 밀도와 낮은 독성을 가지고 상온에서 액체이며, 단일추진제 특성을 가지고 있어 하이브리드 로켓에 적용하기에 유리하다. 본 연구에서는 250 N급 과산화수소 폴리에틸렌 하이브리드 로켓을 설계하였다. 차례의 연소 시험 결과를 바탕으로 하이브리드 로켓의 설계를 평가하였다. 시험 결과, 폴리에틸렌은 축방향으로 일정하게 연소되었고, 특정 길이의 후연소실을 장착하였을 때 특성 속도 효율이 90%와 근접하거나 상회하였다. 또한 90 wt.% 과산화수소를 사용하였을 때 폴리에틸렌 고체 연료의 후퇴율 실험식을 제시하였다. Hydrogen peroxide was utilized as a monopropellant and oxidizer of bipropellant liquid rocket engine. Hybrid rocket has been interested of many research groups due to its low cost and enhanced safety characteristic comparing with its counterparts. Hydrogen peroxide has favorable characteristics for hybrid rocket such as high density, storable, low toxicity, and monopropellant characteristics. In this study, we designed 250 N hydrogen peroxide polyethylene hybrid rocket. With several hot fire tests, we evaluated the designed hybrid rocket. It showed uniform regression rate in axial position, characteristic velocity about 90% with certain length of post combustion chamber. We also proposed empirical equation of polyethylene solid fuel regression rate using 90 wt.% hydrogen peroxide.

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.

연소실 압력에 따른 하이브리드 로켓 성능 향상

이은광(Eunkwang Lee),윤용태(Yongtae Yun),강신재(Shinjae Kang),권세진(Sejin Kwon) 한국추진공학회 2016 한국추진공학회 학술대회논문집 Vol.2016 No.12

높은 안전성과 간단한 시스템을 갖는 하이브리드 로켓을 활용할 경우 가격 경쟁력을 갖는 발사체를 제작할 수 있다. 하이브리드 로켓의 성능 파라미터 중 하나인 고체 연료 후퇴율은 보통 질량 유속에 의해 결정된다고 설명되어 왔으며, 압력에 의한 영향은 고려되지 않았다. 이에 따라 본 연구에서는 250 N 급 과산화수소 폴리에틸렌 하이브리드 로켓을 활용하여 연소실 압력 40 bar 조건에서 연소시험을 실시하였다. 연소실 압력 20 bar 조건에서 얻어진 시험 결과와 비교를 통해 연소실 압력이 연료 후퇴율과 비추력 등 로켓 성능에 미치는 영향을 고찰하였다. 하이브리드 로켓 연소실 압력을 20 bar에서 40 bar로 상승시켰을 때 고체 연료 후퇴율이 약 7~10% 증가하였으며, 비추력은 평균 14 s 증가하였으나 비추력 효율은 감소하였다. Hybrid rocket-based launch vehicles may have the cost competitiveness due to their system simplicity and inherent safety. In general, fuel regression rate, one of the hybrid rocket performance parameter, has been investigated in terms of the mass flux, and the influence of the chamber pressure has been neglected. Thus, in this research, ground hot-fire tests were performed at the 40 bar chamber pressure using 250 N scale hydrogen peroxide polyethylene hybrid rocket. The effect of chamber pressure increase on the rocket performance was scrutinized by comparing the data with the result from hot-fire test at the 20 bar condition. As the chamber pressure doubled, the regression rate increased about 7~10%, and about 14 s higher average specific impulses were observed. However, the lower specific impulse efficiency was observed at the higher chamber pressure case.

추력 650 kgf 급 하이브리드 로켓 모터의 연소시험

이정표(Jungpyo Lee),김수종(Soojong Kim),김기훈(Gihun Kim),조정태(Jungtae Cho),김학철(Hakchul Kim),우경진(Kyongjin Woo),도규성(Gyu-Sung Do),소정수(Jungsoo So),오정수(Jung-soo OH),조민경(Mingyung Cho),문희장(Heejang Moon),성홍계(Honggye S 한국추진공학회 2009 한국추진공학회 학술대회논문집 Vol.2009 No.11

본 연구에서는 추력 650 kgf 급의 PE/N2O 하이브리드 로켓 모터의 지상 연소시험을 수행하였다. 초기 실험에서 산화제 유량이 작게 유입됨으로 인해 연소실 압력 및 추력이 설계치 만큼 확보되지 못함을 확인 하였다. 이를 보완하기 위해 노즐목 감소 및 산화제 유량을 증대하여 실험을 수행하였고, 실험에서 발생하는 연소현상을 분석하였다. 또한 sub-scale과 lab-scale의 실험결과를 통해 scale에 따른 연소특성 변화를 비교? 분석 하였고, 동일 산화제 유속에서 sub-scale의 후퇴율이 lab-scale의 후퇴율보다 차이는 작지만 낮게 나타남을 확인했다. 본 연구의 결과를 통해 실제 하이브리드 사운딩 로켓 개발을 위해 고려되어야 할 사항을 파악할 수 있었다. In this study, we presented the results of static firing tests on the PE/LN2O hybrid rocket motor, which has a thrust of 650 kgf level. Through the early tests, we found that the combustion chamber pressure and the thrust were lower than design values because an actual oxidizer flow rate was less than that expected. In order to complement this result, the methods of decrease of nozzle throat and the increase of oxidizer mass flow rate were conducted in the next experiment, and we studied the combustion phenomena with the experimental results. Also we compared and analyzed a difference of combustion characteristics on scale effect. It show that a sub-scale motor regression rate was a little less than that of a lab-scale motor with the same oxidizer mass flux. Results of this study might be used as a basic data for development of hybrid sounding rocket.

LES를 이용한 하이브리드 로켓 주 연소실의 스월유동 분석

김정은(Jungeun Kim),이창진(Changjin Lee) 한국추진공학회 2016 한국추진공학회 학술대회논문집 Vol.2016 No.5

하이브리드 로켓의 단점으로는 낮은 밀도 비추력과 연소율 감소이다. 연료의 연소 특성을 향상시키기 위하여 스월 인젝터를 사용함으로써 내부유동의 특성을 변화시켰다. 본 논문에서는 LES를 이용하여 하이브리드 로켓 주 연소실에서의 스월 유동을 분석하였다. 그 결과 스월 유동에 의해 열전달과 원심력에 의한 연소율이 향상되었다. 또한 스월 유동을 통하여 유동이 안정성을 분석하였으며 이를 통하여 연소의 안정성을 예측하였다. Disadvantages of Hybrid Rocket Engine are low Density Specific Impulse and low Regression Rate. Therefore, Swirl injector has been applied in order to enhance the regression rate of the hybrid rocket. In this study, Swirl flow was analyzed by LES study in the Hybrid Rocket Engine main combustion. As a result, Swirl flow improved the regression rate by the heat transfer and centrifugal force. Also, the stability of the flow was analyzed by the swirl flow and predicted the stability of the combustion.

하이브리드 로켓 추진 시스템의 혼합 연소실 적용에 따른 Trade-off 평가

김학철,문근환,문희장,김진곤,Kim, Hakchul,Moon, Keunhwan,Moon, Heejang,Kim, Jinkon 항공우주시스템공학회 2016 항공우주시스템공학회지 Vol.10 No.3

The intermediate mixing chamber is one of various methods for improving the regression rate and combustion efficiency of the hybrid rocket. The mixing chamber with its non-combustible material makes the propulsion performance increase, but it leads to a low fuel-loading density in the combustion chamber; therefore, this performance-related trade-off between the mixing chamber and the low fuel-loading density was studied. In this study, the trade-off was conducted by comparing the intermediate-mixing-chamber case with a w/o-mixing-chamber case. The small hybrid-sounding rocket is designed with internal ballistics for comparing the rocket length to the weight. In addition, an external ballistic analysis was conducted for comparing the performances of the w/- and w/o-mixing-chamber cases. As a result, the intermediate-mixing-chamber case shows that the length and the weight were decreased to 12 % and 8 %, respectively; furthermore, when compared with the normal cases, the estimated altitude result of the w/-mixing-chamber case was increased to approximately 75 m.

하이브리드 로켓에서 고체연료 후퇴율 측정기법에 대한 연구

박종원(Jong Won Park),구건우(Kun Woo Ku),홍정구(Jung Goo Hong),이충원(Choong Won Lee) 대한기계학회 2010 대한기계학회 춘추학술대회 Vol.2010 No.11

Hybrid rocket had many advantage with compared to solid and liquid rockets. However, the engines have not yet been used in practical rocket systems, due mainly to the disadvantage of hybrid combustion, such as low fuel regression rate. In this study, lab-scale hybrid motor was designed and manufactured. And the methods of regression rate improvement were considered. The regression rate of solid fuel was observed with ultra sonic sensor and digital video camera and compared each other. It is found that the results using ultra sonic sensor was more accurate. The regression rates showed a strong dependency on gaseous oxygen mass flux, grain port diameter, and grain length. The frequency analysis technique of the bulk-mode oscillation of motor was applied to a hybrid rocket motor and was based on the principle that this frequency was inversely proportional to the square root of the chamber volume.

RATO(Rocket-Assisted Take Off) 시스템 적용을 위한 하이브리드 로켓 비단공형 연료 그레인 기초 연소특성 연구

김수진,고수한,김설희,김경모,이성근,한예찬,문희장 한국항공운항학회 2022 한국항공운항학회지 Vol.30 No.4

In an attempt to apply hybrid rocket to the RATO (Rocket-Assisted Take Off) system, combustion characteristics of the non-circular grain were figured out in this study. Having larger combustion area, it was reconfirmed that the non-circular grain has advantages over regression rate, characteristic velocity and chamber pressure in which all gave higher values. Experiments were performed to understand the effect of the non-circular grain geometry over time where local regression rates depending on grain location were analyzed. It was found that the regression rate of five distinct locations were different. Partial conclusion driven was that these differences are due to the heat transfer caused by dissimilar distances from the flame layer. Besides, as combustion duration increased, the fuel port became circular, and the regression rate converged to a single value over the whole grain.