• Title/Summary/Keyword: 접촉점화 추진제

Search Result 13, Processing Time 0.032 seconds

A Review of the Technical Development on Green Hypergolic Propellant (친환경 접촉점화 추진제 연구 개발 동향)

  • Park, Seonghyeon;Kang, Hongjae;Park, Youngchul;Lee, Jongkwang
    • Journal of the Korean Society of Propulsion Engineers
    • /
    • v.24 no.4
    • /
    • pp.79-88
    • /
    • 2020
  • Hypergolic propellants have been widely used for space propulsion systems based their long-term storability and high ignition reliability. Since conventional hypergolic propellants are highly toxic and carcinogenic, handling and operating costs are significant. To overcome the drawbacks, numerous studies have been actively performed to develope new hypergolic propellants, ensuring that the combinations are high performance, low toxicity and low environmental impact. In the present study, a comprehensive survey was conducted to summarize the research and development of green hypergolic propellants involving hydrogen peroxide, nitric acid, and ionic liquids.

Non-ignition Evaluation Method for Hypergolic Propellant Using Microreactor (마이크로 반응기를 이용한 접촉점화 추진제의 비점화 평가 방법)

  • Lee, Kyounghwan;Park, Seonghyeon;Kang, Hongjae;Lee, Jongkwang
    • Journal of the Korean Society of Propulsion Engineers
    • /
    • v.26 no.2
    • /
    • pp.20-27
    • /
    • 2022
  • Hypergolic propellant ignited spontaneously when fuel and oxidizer contact without ignition system. Due to this characteristic, the risk of accidents is high when new propellants are evaluated. Prevention of accidents is very important because the damage can be large when the accident occur. In this work, we proposed non-ignition evaluation method which can replace conventional ignition evaluation method by using microreactor. The reactor was fabricated by MEMS. The heat of reaction as according to fuel and NaBH4 was estimated. At the condition of highest heat of reaction ignition was observed by drop test.

Evaluation of Ignition Performance of Green Hypergolic Propellant (친환경 접촉점화 추진제 점화 성능 평가)

  • Sunjin Kim;Minkyu Shin;Jeongyeol Cha;youngsung Ko
    • Journal of Aerospace System Engineering
    • /
    • v.17 no.1
    • /
    • pp.51-58
    • /
    • 2023
  • Hypergolic propellants, which can ignite themselves without an ignition source, are difficult to handle due to their corrosiveness and toxicity. Therefore, it is necessary to develop green hypergolic propellants with little or no toxicity. In this study, basic research on green hypergolic ignition propellants was conducted. With 95% hydrogen peroxide as an oxidizer and CNU_HGFv1 as a fuel, ignition and combustion characteristics of propellants were evaluated through a drop test, an ignition test, and a combustion test. As a result of the drop test, the ignition delay time was 9.7 ms. It was 27 ms in the ignition test, which was fast enough to be used as a propellant. As a result of the combustion test, a combustion efficiency of 95.4~98.1% was achieved at about 11.7 bar. It was confirmed that fast and stable combustion was possible without hard start or combustion instability.

Experimental Ignition Delay Assessment of H2O2 Based Low Toxic Hypergolic Propellants with Variation of Reactive Additive Concentration (반응성 첨가제 농도에 따른 과산화수소 기반 저독성 접촉점화성 추진제의 점화지연 시험평가)

  • Rang, Seongmin;Kim, Kyu-Seop;Kwon, Sejin
    • Journal of Aerospace System Engineering
    • /
    • v.14 no.3
    • /
    • pp.24-31
    • /
    • 2020
  • A study on the H2O2 based low toxic hypergolic propellant was conducted. The fuel candidates were chosen as a mixture of Amine solvent and reactive additive. The analytical performance was calculated via the NASA CEA code and 96% Isp of the NTO/UDMH was confirmed. The ignition delay measurement with drop test was performed and all candidates showed less than 10 ms in the best performance cases. Based on these results, the feasibility of high response H2O2 based low toxic hypergolic propellant was confirmed.

Research Studies of Impingement Characteristics for Hypergolic Propellant (접촉 점화성 추진제의 충돌형 혼합 특성 연구 사례)

  • Kim, Kyu-Seop;Kim, Yehyun;Jung, Sangwoo;Jeong, Junyeong;Kwon, Sejin
    • Journal of the Korean Society of Propulsion Engineers
    • /
    • v.23 no.5
    • /
    • pp.90-100
    • /
    • 2019
  • Hypergolic thrusters have been extensively researched and applied to spacecraft propulsion based on their simplicity and high reliability of ignition. Research on the impingement characteristics of $N_2O_4$/amine has been profoundly carried out since the 1960s in advanced countries, especially the United States. Recently, enhancements to advanced hypergolic thrusters using MON/MMH have been planned by NASA to improve compactness and high performance. In this work, technical studies were investigated on the mixing of hypergolic propellant and its combustion instabilities such as reactive separation flow and popping.

Ignition Test of an Oxidizer Rich Preburner (산화제과잉 예연소기 점화시험)

  • Moon, Il-Yoon;Moon, In-Sang;Yoo, Jae-Han;Jeon, Jae-Hyoung;Lee, Seon-Mi;Hong, Moon-Geun;Ha, Seong-Up;Kang, Sang-Hun;Lee, Soo-Young
    • Proceedings of the Korean Society of Propulsion Engineers Conference
    • /
    • 2011.11a
    • /
    • pp.869-872
    • /
    • 2011
  • Ignition tests of an oxidizer rich preburner for a staged combustion cycle liquid rocket engine were performed to evaluate combustion performance. Design operation conditions of the tested oxidizer rich preburner are about 60 of OF ratio and 20 MPa of combustion pressure. The entire kerosene and some LOx injected into the mixing head is burned in combustion chamber and the remaining LOx injected through center holes of combustion chamber is vaporized. Full flow ignition method with hypergolic fuel was used. Each propellant was supplied in two stages for soft ignition. Test results, low frequency oscillation was occurred in low flow rate conditions under 45% of design flow rate. Stable ignition in the course of design combustion pressure was able to induce by minimization of low flow rate ignition region to escape low frequency oscillation.

  • PDF

Helium Quantity Estimation for LOx Tank Pressurization of a Restartable Pressure-fed Propulsion System (재 점화가 있는 가압식 추진기관의 액체산소 탱크 가압 헬륨량 산정)

  • Cho, Gyu-Sik;Jung, Young-Suk;Oh, Seung-Hyub
    • Journal of the Korean Society of Propulsion Engineers
    • /
    • v.16 no.3
    • /
    • pp.77-81
    • /
    • 2012
  • In a cryogenic propellant tank the pressurant is contracted due to heat loss and the propellant itself evaporates. On a restartable propulsion system such phenomena are more intensive because the propellant contacts with the pressurant on the larger surface during the coast flight. Such heat and mass transfer phenomena should be considered for estimating the amount of pressurant. On the hypothesis that the heat and mass transfer quasi-equilibrium is achieved during the coast flight, the calculation process of the equilibrium pressure is presented. On the process the amount of loaded helium on the Falcon-1 second stage is calculated.

Helium Quantity Estimation for LOx Tank Pressurization of a Restartable Pressure-fed Propulsion System (재 점화가 있는 가압식 추진기관의 액체산소 탱크 가압 헬륨량 산정)

  • Cho, Gyu-Sik;Jung, Young-Suk;Oh, Seung-Hyub
    • Proceedings of the Korean Society of Propulsion Engineers Conference
    • /
    • 2011.11a
    • /
    • pp.201-205
    • /
    • 2011
  • In a cryogenic propellant tank the pressurant is contracted due to heat loss and the propellant itself evaporates. On a restartable propulsion system such phenomena are more intensive because the propellant contacts with the pressurant on the larger surface during the coast flight. Such heat and mass transfer phenomena should be considered for estimating the amount of pressurant. On the hypothesis that the heat and mass transfer quasi-equilibrium is achieved during the coast flight, the calculation process of the equilibrium pressure is presented. On the process the amount of loaded helium on the Falcon-1 second stage is calculated.

  • PDF

Development of 500 Kgf Thrust Liquid Propellant Rocket Engine (추력 500 Kgf 액체추진제 로켓엔진 개발)

  • 정동호;조용재;정규상
    • Proceedings of the Korean Society of Propulsion Engineers Conference
    • /
    • 1997.04a
    • /
    • pp.3-10
    • /
    • 1997
  • 본 연구에서는 추력 500 Kgf의 액체 추진기관을 설계, 제작 및 연소시험을 수행하여 연소 특성을 살펴보았다. 추진제로는 우주발사체 Booster용으로 폭넓게 사용되는 탄화수소계 연료인 kerosene과 산화제로 취급이 용이하고 저장 특성을 지닌 98 % White Fuming Nitric Acid(WFNA)를 사용하였고, 엔진 점화를 위해 WFNA와 접촉 발화성 (Hypergolic)을 갖는 Furfuryl Alcohol/Aniline 혼합액을 사용하였다. 로켓엔진은 20 Kgf/$cm^2$의 연소실 압력으로 500 Kgf의 평균 추력을 내도록 설계되었고, 연소실벽을 고온 연소가스로 부터 보호하기 위해 Film Cooling 방식을 적용하였다.

  • PDF

Study on the Ignition and Burning Characteristic of Single Aluminum Particle with Thermal Radiation (열복사에 의한 단일 알루미늄 입자 점화-연소특성 측정)

  • Lim, Ji-Hwan;Yoon, Woong-Sup;Lee, Do-Hyung
    • Proceedings of the Korean Society of Propulsion Engineers Conference
    • /
    • 2010.05a
    • /
    • pp.450-459
    • /
    • 2010
  • 고체추진제의 첨가제 또는 연료로써 주로 사용되는 알루미늄 단일 입자 연소시험 장비를 제작하고 연소 실험을 수행하였다. 산화 알루미늄으로 피복된 금속입자는 약 30~100 ${\mu}m$의 크기를 사용하였다. 단일 입자는 Electrodynamic Balance (EDB) 방법에 의해 공중 부양된 상태로, 중력에 의한 영향이 배제되어 금속입자 고정용 또는 측정용 장치들의 접촉에 의한 열손실을 제거시켜 실험 정확도를 높였다. Standard Hyperbolic Electrodynamic Levitator (SHEL) 내에서 부양된 입자에 $CO_2$ 레이저를 사용하여 점화시킨 후, 입자로부터 방사되는 열복사를 이용한 two wavelength pyrometry를 적용하여 알루미늄 입자 크기에 따른 연소시간, 평균 화염온도, 점화온도, 점화시간을 획득하였으며, 단일 알루미늄 입자의 점화-연소특성을 평가하였다.

  • PDF