• Title/Summary/Keyword: 극저온 산화제 탱크

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위성 발사체 추진제 가압용 열교환기 기초 설계

  • 이희준;한상엽;정용갑;길경섭;하성업;김병훈
    • Bulletin of the Korean Space Science Society
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    • 2004.04a
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    • pp.74-74
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    • 2004
  • 액체추진제를 사용하는 위성 발사체의 경우 추진제탱크에 저장된 추진제를 추력을 발생하는 연소실에 공급하기 위하여 헬륨 등의 가압제를 사용한다. 본 연구에서는 액체추진제 로켓엔진의 산화제인 극저온의 액체산소를 저장하고 있는 탱크 내부에 설치된 별도의 탱크에 저장된 극저온/고압의 헬륨을 고온으로 열팽창 시켜 추진제 탱크로 재유입하여 추진제를 가압하는 시스템에 사용되는 가압제 열팽창용 열교환기의 개발을 위한 기초 설계를 수행하였다. (중략)

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Evaluation of Permeability Performance by Cryogenic Thermal Shock in Composite Propellant Tank for Space Launch Vehicles (우주 발사체용 복합재 산화제 탱크 구조물의 극저온 열충격에 따른 투과도 성능 평가)

  • Kim, Jung-Myung;Hong, Seung-Chul;Choi, Soo-Young;Jeong, Sang-Won;Ahn, Hyon-Su
    • Composites Research
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    • v.33 no.5
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    • pp.309-314
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    • 2020
  • Polymer composites were used to reduce the weight of the spacecraft's cryogenic propellant tank. Since these materials were directional, the permeability performance of the gas permeated or delivered in the stacking direction was an indicator directly related to performance such as tank stability and onboard fuel quantity estimation. In addition, the results of permeation measurements and optical analysis of the surface to verify the effect of the number of cycles exposed to the cryogenic-room temperature environment are included. As a result, the permeability was inversely proportional to the thickness and was proportional to the number of thermal shocks, and it was verified that the permeability performance was suitable for the cryogenic propellant tank material for the space launch vehicle.

Performance Test of an Oxidizer Tunnel-Type Pipe for Launch Vehicle (발사체 산화제 터널형 배관 성능시험)

  • Kil, Gyoung-Sub;Han, Sang-Yeop;Kho, Hyeon-Seok;Shin, Dong-Sun;Cho, In-Hyun
    • Proceedings of the Korean Society of Propulsion Engineers Conference
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    • 2009.11a
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    • pp.273-277
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    • 2009
  • An oxidizer tunnel-type pipe, which shall transport oxidizer from an oxidizer tank to a turbo-pump of an engine, studied is installed through a fuel tank located under an oxidizer tank. A tunnel-type pipe can save weight compared to a detour-type pipe, however may vary the temperature of fuel stored in a fuel tank because of a broad heat transfer area. Hence in this study the characteristics of main oxidizer pipe and thermal propagation from oxidizer to a fuel tank are monitored by a cryogenic performance test with a tunnel-type pipe. In addition, the possibility of adaptation of an oxidizer tunnel-type pipe to launcher system is also analyzed.

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Heating Apparatus Development for Cryogenic Gaseous Helium (극저온 헬륨가스 가열장치 개발)

  • Chung, Yong-Gahp;Kwon, Oh-Sung;Cho, Nam-Kyung;Cho, In-Hyun
    • Proceedings of the Korean Society of Propulsion Engineers Conference
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    • 2009.11a
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    • pp.363-367
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    • 2009
  • For the liquid rocket propulsion system using liquid oxygen as oxidizer, helium for pressurizing LOX is usually stored in the LOX tank with cryogenic temperature. For that kind of pressurizing system, cryogenic helium is discharged from the immerged pressurant cylinder and passes through the heat exchanger downstream of gas generator. During the process, helium pressurant is heated from cryogenic temperature to high one and supplied to the ullage of propellant tank. To develop the pressurizing system, a cryogenic heating apparatus is needed to simulate the heat exchanger. In this paper, the cryogenic heating apparatus for development of the pressurization system is presented along with its heating test results with cryogenic helium.

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Heating Apparatus Development and Tests for Cryogenic Gaseous Helium (극저온 헬륨가스 가열장치 개발 및 시험)

  • Chung, Yong-Gahp;Cho, Nam-Kyung
    • Journal of the Korean Society of Propulsion Engineers
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    • v.15 no.1
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    • pp.63-68
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    • 2011
  • For the liquid rocket propulsion system using liquid oxygen as oxidizer, helium for pressurizing LOX is usually stored in the LOX tank with cryogenic temperature. For that kind of pressurizing system, cryogenic helium is discharged from the immerged pressurant cylinder and passes through the heat exchanger downstream of gas generator. During the process, helium pressurant is heated from cryogenic temperature to high one and supplied to the ullage of propellant tank. To develop the pressurizing system, a cryogenic heating apparatus is needed to simulate the heat exchanger. In this paper, the cryogenic heating apparatus for development of the pressurization system is presented along with its heating test results with cryogenic helium.

Performance Test of PSD Oxidizer Drain Valve for KSLV-II (한국형발사체 PSD 산화제 배출밸브 성능시험)

  • Chung, Yonggahp;Han, Sangyeop;Kim, Suengik
    • Proceedings of the Korean Society of Propulsion Engineers Conference
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    • 2017.05a
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    • pp.1171-1175
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    • 2017
  • Cryogenic helium gas is used as the pressurant for the oxidizer pressurization of DR(Damper Receiver) sphere in the PSD(Pogo Suppression Device) system and liquid oxygen is used as the oxidizer for the propellant in Korea Space Launch Vehicle-II. The helium gas is stored in pressurant cylinders inside the cryogenic liquid oxygen tank and liquid oxygen is stored in the oxidizer tank. In this study, the performance test of PSD liquid oxygen drain valve for KSLV-II was considered.

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Thermal Analysis of Prelaunch Transients in Cryogenic Oxidizer Tank of Liquid Propulsion Rocket (발사대기 중인 액체추진 로켓의 극저온 산화제 탱크 내 비정상 열해석)

  • Kim, Kyoung-Hoon;Ko, Hyung-Jong;Kim, Kyoung-Jin;Cho, Kie-Joo;Oh, Seung-Hyub
    • Journal of the Korean Society of Propulsion Engineers
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    • v.12 no.4
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    • pp.33-41
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    • 2008
  • The prelaunch thermal transients in the cryogenic oxidizer tank of liquid propulsion rocket subjected to uniform heat flux from outside are numerically analyzed through thermodynamic equations and heat and mass transfer relations. The prelaunch stage is assumed to be composed of five idealized sub-stages including pressurization process by helium gas injection. The Peng-Robinson equation of state is utilized in the lumped analysis of ullage gas. The liquid region is divided into a number of horizontal layers of uniform properties to account for the thermal stratification. The computational result for the typical case shows that the temperature rise of liquid oxidizer is less than 1K and the adsorbed helium into the liquid is approximately 10g.

Numerical Flow Analysis for Anti-Vortex Device (AVD) in Oxidizer Tank (산화제 탱크의 와류방지장치 유동해석)

  • Jang, Je-Sun;Han, Sang-Yeop;Kil, Gyoung-Sub;Cho, In-Hyun
    • Aerospace Engineering and Technology
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    • v.9 no.2
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    • pp.168-175
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    • 2010
  • During feeding oxidizer to the engine, vortices are occurred at lower dome of oxidizer tank inside by various working environments and external forces for liquid propellant feeding system of space launch vehicle. To eliminate the vortices or swirls Anti-Vortex Devices(AVD) shall be installed at inside lower oxidizer tank. Using the numerical analysis, we have confirmed the performance of AVD and analyzed the mass flow rate by feeding time and magnitudes of swirls on the free surface of oxidizer or exit surface according to the AVD number and length. Then we could derive the optimal design of the AVD number and length.

터보펌프식 액체 로켓의 추진제 공급시스템 설계

  • 조기주;이한주;정영석;임석희;김지훈;오승협
    • Bulletin of the Korean Space Science Society
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    • 2003.10a
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    • pp.89-89
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    • 2003
  • 로켓엔진의 연소에 필요한 추진제를 안정적으로 공급하기 위한 추진제 공급시스템의 주요 구성과 설계 주요 인자를 정리하였다 공급시스템은 추진제 주입/배출 장치, 추진제탱크 가압 및 배기 장치, 추진제 공급 주/분기 배관, 극저온 산화제 온도 유지 장치 등으로 구성되어 있다. 주요 설계 제한 조건으로는 터보 펌프 입구에서의 추진제 압력 및 온도, 필요 추진제 공급 유량 및 온도 그리고 추진제 충진 및 비상 배출 허용 시간 등이며 이는 각 로켓의 해당 임무에 따라 적절히 결정된다. 발사체로부터 할당된 중량값 이내에서 고신뢰도의 작동성, 안정성이 보장되는 시스템을 설계하여야 하며 초기 설계 단계에서 개발 및 수급 가능성을 동시에 고려하여야 할 것이다. 또한 고추력 생성을 위해 엔진 클러스터링이 수행되어야 할 경우 각 엔진으로의 균등한 추진제 배분 공급이 설계의 중요한 요구 조건이 된다. 이러한 공급시스템의 개념은 액체산소와 케로신 조합의 액체 로켓인 100kg급 소형 위성 발사체(KSLV-Ⅰ)에 적용될 예정이다.

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Investigation of the Cryogenic Oxidizer Tank Inner Phenomena of Pump-fed Liquid Rocket Engine Propulsion System (터보펌프식 액체추진기관에서의 극저온 산화제 탱크 내부 현상 고찰)

  • 조남경;권오성;정용갑;조인현;김영목;조기주;정영석
    • Proceedings of the Korean Society of Propulsion Engineers Conference
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    • 2003.10a
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    • pp.238-241
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    • 2003
  • In case of liquid rocket using turbopump, the inner pressure of liquid oxygen tank is maintained low, so vaporization of LOX is generally occurred. This vaporization tendency increases as the inlet helium gas temperature is higher. For estimating the amount of helium in the rocket system, the LOX vaporization phenomena should be carefully considered. In this paper, Inner process of LOX tank is analyzed by two phase flow modeling. the vaporization rate and required Helium mass is investigated with varying inlet helium temperature and heat transfer coefficient.

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