• Proceedings of the Korean Society of Dyers and Finishers Conference
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• 2012.03a
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• pp.75-75
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• 2012

탄소섬유는 전구체의 종류에 따라 PAN계, 피치계 그리고 레이온계로 나뉘며 최종 탄소섬유의 특성에도 차이가 있는 것으로 알려져 있다. 최근에는 PAN계 탄소섬유가 세계 시장의 대부분을 차지하고 있으며, PAN계 탄소섬유의 초경량, 고강도, 고탄성, 내약품성 그리고 열안정성 등의 우수한 특성으로 최첨단 고기능성 제품의 복합재로 많이 이용되고 있다. 그러나 탄소섬유가 가지고 있는 높은 열전도성은 적용에 따라 단점으로 작용될 수도 있다. 예를 들면, 로켓 엔진의 노즐이나 원자로의 구조물 그리고 극한조건용 구조재료 등, 고강도 단열특성을 요하는 최첨단 복합재로 응용 범위를 넓히는데 한계로 작용한다. 레이온은 최초의 탄소섬유 전구체였으나 공정상 경제성이 떨어지는 이유로, 지금은 고탄성을 요구하는 특수 목적으로만 소량 생산되고 있다. 레이온의 주원료는 셀룰로오스이며 셀룰로오스는 지구상에서 가장 흔한 재료이므로 오늘날 셀룰로오스를 보강재로 이용하려는 연구가 활발히 진행되고 있다. 본 연구에서는 탄소섬유의 열전도도를 낮추기 위한 방법으로 안정화셀룰로오스를 첨가한 PAN용액을 출발물질로 탄소섬유를 제조하고 특성 연구를 진행하였다. PAN용액에 셀룰로오스의 분산성을 향상시키기 위해 셀룰로오스를 열처리하였다. 이 과정에서 얻어진 안정화 셀룰로오스의 수율을 높이기 위해 셀룰로오스를 난연 처리하였으며, 그 결과 안정화셀룰로오스의 수율을 향상시킬 수 있었다. 안정화셀룰로오스를 첨가시킨 PAN계 탄소섬유의 물리적, 기계적 그리고 열적 특성을 SEM, XRD, 만능 인장시험기, TGA 그리고 Laser Flash Method 등을 통해 주요 특성 및 변화를 관찰한 결과, 순수한 PAN계 탄소섬유의 특성과 유사한 결과를 얻었다. 향후 몇 가지 공정상의 문제점을 개선한다면 흥미로운 결과를 기대할 수 있을 것으로 본다.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.30 no.8
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• pp.65-70
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• 2002

The brazing of duplex stainless steels which is an essential process for rocket engine manufacturing has been investigated on bonding phenomena and shear strength. The UNS32550 was used for base metal, and the MBF-50 was used for insert metal. Brazing was carried out under the various conditions (brazing temperature : 1473K, 1498K, holding time : 0, 0.3, 0.9, 1.8 ks). There were various phases in the joint because of reaction between liquid insert metal and base metal, In the early stage, BN is formed in the bonded interlayer and base metal near the bonded layer. Cr nitride is formed in the bonded interlayer. The amount of BN and Cr nitride decrease with the increase of holding time. Superior shear strength of 550MPa is obtained by restraining the formation of nitride.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.33 no.6
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• pp.64-70
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• 2005

The overall results of hot firing tests of fuel-rich gas generator with impinging injector at design and off-design points are described. The gas generator consists of an injector head with impinging injector, a water cooled combustor wall, a turbulence ring to enhance mixing, an instrument ring measuring temperature and pressure and a nozzle. The combustion tests were successfully performed without damage of gas generator. Test results show that the outlet temperature is not dependent on residence time of hot gas within 4～6msec but dependent on chamber pressure. The relation between outlet temperature and combustion efficiency resulting from measured pressure, mass flow rate and area of nozzle throat is shown. The overall O/F ratio is the critical parameter to determine the outlet temperature and the linear correlation between two parameters is established.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.31 no.9
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• pp.103-109
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• 2003

The development of a pulse gun used for the combustion stability rating test(SRT) has been domestically tried out since the SRT should be accompanied in the process of the development of LREs. At the beginning of the development, a rupture problem of the pulse gun body had been solved by changing its material and increasing its thickness. Also, the sealing test was undertaken for a pulse gun equipped with a membrane before conducting a explosion test and by using a specially designed jig after a explosion test. Since it is appropriate to reuse the holder of a pulse gun cavity for an actual SRT, a series of experiments for several variants has been taken to make it reusable. After all, the solution for a reusable holder of a pulse gun has been found and applied for complementing its specification.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.44 no.8
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• pp.695-701
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• 2016

Spacecraft propulsion system is a kind of rocket engine that has been developed from the end of 1950s for attitude control and orbit maintenance of satellite. Since the spacecraft propulsion system has to be used for a relatively long time, therefore, stability of propellant and life of thruster could be very important factor for propulsion system design. Recently, green propellant propulsion and all electrical propulsion system have became very important issue, and we also need a development according to well organized plan. In this paper, we will introduce the development status, key technologies and development prospect of spacecraft propulsion system.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.44 no.8
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• pp.668-674
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• 2016

In this paper the development directions of the next generation launch vehicle program following KSLV-II has been discussed, which are to be executed after year 2020 according to the Medium and Long Term Plan for National Space Development. Also, several areas of technology advancement have been identified for the successful development of the LVs. The next generation LV must aim for not only the high performance but also for low cost as well as high reliability in order to compete against global commercial launch service providers. To this end, the next generation LVs program shall capitalize on many anticipated accomplishments of the KSLV-II program such as the 75 ton class LOX/kerosene rocket engine.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.05a
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• pp.135-140
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• 2006

In the experimental study of $N_2$ purge cold flow test of impinging(FOOF) injector for determining of instability region, the whistling sound which has a specific frequency is generated. The frequency of whistling is proportional to the gas flow velocity in part of the oxidizer orifice and due to the coupling of the vibrating gas column and the natural frequency of pipe-orifice shape, the discontinuous jumping phenomena arises. The whistling phenomena have no effect on the combustion instability. Compared the damping factor of 1T1L mode with the hot fire test, the instability region of $N_2$ purge cold flow test is very much like that. It means that flow instability by impinging or mixing of jet is the main reason of combustion instability of impinging injector(FOOF) in the hot firing test.

• Journal of the Korean Society of Propulsion Engineers
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• v.19 no.4
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• pp.85-93
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• 2015

Liquid rocket engines of KSLV-II employ a turbopump feed system for propellants. A turbopump real-propellant test facility based on liquid oxygen and kerosene has been constructed for the experimental verification of the turbopump performance using the real media of propellants(i.e., LOX/Kerosene). The verification tests of sub-systems were performed such as LOX/kerosene feed system and alcohol burner system. Finally, the performance of the whole system was executed and verified through a sets of validation tests with the development model of the KSLV-II turbopumps. It has been confirmed that the test facility satisfies the operating conditions and time of the turbopump at the design and off-design performance test using real-propellant.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.515-519
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• 2011

Structural tests for the mixing head of a 75tonf class thrust chamber were performed to verify structural stability. The mixing head of a thrust chamber is loaded by high pressure with regeneratively cooled fuel and cryogenic liquid oxygen(LOx) as well as it transfers thrust load generated by liquid rocket engine. Therefore structural stability of mixing head is a very important factor to work without any plastic deformation or structural failure. In this study, two mixing heads were manufactured using different welding methods, Tungsten Inert Gas(TIG) welding and Electron Beam Welding(EBW) and evaluated a structural stability. The results of structural tests showed that the mixing head assembled by EBW can withstand the applied design load without any structural failures and be structurally more stable than that of TIG welding.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.869-872
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• 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.