• Journal of the Korean Society of Propulsion Engineers
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• v.6 no.1
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• pp.30-38
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• 2002

Bullet impact tests for two solid rocket motors were performed and its results were compared and analyzed. One was loaded with the existing propellant with decreased weight content of burning rate catalyst and added high density additives to improve mechanical properties and the other was loaded with the existing propellant with decreased weight content of burning rate catalyst to improve its insensitivity as well as to maintain the ballistic performance. The composite cases were used for both motors.

• Composites Research
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• v.19 no.4
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• pp.31-38
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• 2006

In this study, effects of curing temperature and autofrettage pressure on a Type 3 cryogenic propellant tank, which is composed of composite hoop/helical layers and a metal liner, were investigated by thermo elastic analysis and composite/aluminum ring specimen tests. Temperature field of a Type 3 tank was obtained from solving the heat transfer problem and, in turn, was used as nodal temperature boundary conditions during the elastic analyses for curing temperature and autofrettage pressure effects. As a result, it was shown that the higher curing temperature was, the more residual compressive stress and tensile stress were induced in composites and metal liner, respectively. On the contrary, autofrettage pressure brought the reduction of these residual thermal stresses caused by cryogenic environments to the tank structure. This tradeoff for curing temperature and autofrettage pressure must be considered in the design and manufacturing stages for a Type 3 cryogenic tank.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.199-204
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• 2010

The objective of this paper is to present new binder systems that can be applied in composite rocket propellants, to improve properties of these propellants not only for better performance, but also to reduce waste and pollution. These novel systems are based on the thermoplastic elastomer (TPE) binders, which consists of copolymers with the addition of a plasticizer, and additives. The effect of the novel TPE binder systems on the burning rate and mechanical properties of AP based propellants was studied. The results show that propellants based on the novel TPE binders have a better energy performance than today's workhorse hydroxyl terminated polybutadine/ammonium perchlorate propellant, exhibit a similar range of burning rate, possess appropriate mechanical properties, and exhibit good processing and aging characteristics at low cost.

• Journal of the Korean Society of Propulsion Engineers
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• v.5 no.4
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• pp.40-49
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• 2001

The basic research on AKM(Apogee Kick Motor) for space launch vehicle was carried out. AKM which will be used as 3rd stage solid rocket motor in 3-stage Korean Sounding Rocket(III) has been developing. KM is a solid rocket motor using composite propellant based on HTPB and is composed of composite motor case and submerged nozzle. To develop KM rocket motor satisfing a given set of requirement, firstly the full-scale KM with diameter 520mm was designed, then sub-scale motors reduced about 60% were manufactured and tested. Full-scale ground firing test is accomplished two times.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 1997.11a
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• pp.15-15
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• 1997

본 연구는 전방점화 방식(the head-end ignition)을 채택하고 있는 composite 고체 추진기관(구룡 1모타)을 이용하여 후방점화 방식(the after-end ignition)에 의한 점화 가능성을 검토하였으며, 점화 방식 차이에 따른 추진기관의 초기 연소거동의 차이점을 고찰하고자 한 실험 연구로서, 후방 점화장치를 설계·제작하여 지상연소시험을 수행하였다. 점화장치는 착화장치(initiation system)와 에너지 방출장치(energy release system), 구조물(Hardware)로 구성되는데, 착화장치는 기존의 K2 squib를 사용하였고, 에너지 방출장치는 FRP튜브에 MTV pellet 점화제를 사용하였으며, 점화기를 후방에 부착시키는 방법으로는 flexible finger 형태의 locking sleeve를 설계하여 노즐목에 고정하였다.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.35 no.7
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• pp.592-600
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• 2007

Nowadays, researches for replacing material systems for cryotanks by composites have been being performed for the purpose of lightweight launch vehicle. In this paper, a type 3 propellant tank, which is composed of the composite developed for cryogenic use and an aluminum liner, was fabricated and tested considering actual operating environment, that is, cryogenic temperature and pressure. For this aim, liquid nitrogen (LN2) was injected into the fabricated tank and in turn, gaseous nitrogen (GN2) was used for pressurization. During this test procedure, strains and temperatures on the tank surface were measured. The delamination between hoop layer and helical one, was detected during the experiment. Several attempts were followed to investigate the cause analytically and experimentally. Thermo-elastic analysis in consideration of the progressive failure was done to evaluate the failure index. Experimental approach through a LN2 immersion test of composite/aluminum ring specimens suitable for simulating the Type 3 tank structure.

• Journal of the Korean Society of Propulsion Engineers
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• v.7 no.1
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• pp.18-27
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• 2003

To perform combined tests with propellant feeding system and engine, which were developed for KSR-III launcher, vertical test stand was organized and a series of hot-fire combustion tests were carried out with engines of several injector faceplate types. In hot-fire tests in vertical installation, combustion instabilities occurred right after ignition with an engine without baffle, and such combustion instabilities did not occur at ignition add during mainstage operation for an engine with STS or composite baffle. 1.regular and temporary pressure pulsations(popping) were detected during steady operation with a baffle engine, however a development to combustion instabilities with resonant mode was highly suppressed by baffle. With a series of tests, it was confirmed that the last developed engine, which has composite baffle, was operated successfully in KSR-III flight propulsion system.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.957-967
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• 2017

Firing test of the fuel grain for solid fuel ramjet with additives and ignition support material was conducted. Fuel grain consist of HTPB mixed with AP particle 15 wt.%, Boron particle 5 wt.%. To cause the short ignition delay, ignition support consist of $NC/BKNO_3$ and composite propellant was coated to the fuel grain. An oxidant gas having a controlled temperature, pressure and oxygen composition close to the air condition in the ramjet combustor was supplied using the Ethanol blended $H_2O_2$ gas generator. Gas was set to flow at a mass flow rate of 150 g/s and mass flux of $200kg/m^2s$ in the grain port. Through the test, ignition support operated well and ignition delay of 0.5. During the test, stable chamber pressure with 8 bar and high combustion efficiency of 0.86 was confirmed.

• International Journal of Aeronautical and Space Sciences
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• v.18 no.3
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• pp.512-521
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• 2017

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.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.05a
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• pp.196-199
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• 2009

The most important thing is to analyze the Fast Cook Off problem of the solid motor case exposed to direct flame is a heat transfer analysis. Heat causes degradation and ignition of the propellant. To archive an acceptable reaction level in Fast Cook Off, the rocket motor case generally must fail structurally prior to propellant ignition. We investigate the responses of the solid motor case exposed to Fast Cook Off by using finite element method for the thermal analysis.