• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.11a
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• pp.255-263
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• 2006

In order to measure the acoustic amplification factor of an Al/HTPB propellant, T-burner tests using pulsed DB/AB method were conducted. In the experiment, powdered aluminum content was varied to a certain extent. Simultaneous ignition on the internal surface of a propellant was achieved by the use of a fast ignition disk. From the experimental data, the damping factor for a non-zero aluminum content could not be calculated due to the fast attenuation of perturbed pressure. Therefore, the addition of aluminum particle was more than sufficient to stabilize pressure-coupled instability.

• Journal of the Korean Society for Aviation and Aeronautics
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• v.14 no.4
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• pp.17-23
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• 2006

This paper describes the preliminary design procedure for the hybrid propulsion system. For a given mission defined by velocity increment, the design of a polyethylene/LOX hybrid rocket was implemented. In addition, Seven-cluster multi-port fuel-grain was considered. After determining the system size including the combustion chamber, the performance parameters such as specific impulse, thrust, characteristic velocity, and thrust coefficient can be predicted by using empirical regression rate correlation, though most of preliminary design code assume constant regression rate. The results of the performance prediction indicated that besides the widely used HTPB/LOX, polyethylene/LOX hybrid motor can be a viable alternative to the more widely used SRMs.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 1996.05a
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• pp.131-137
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• 1996

HTPB/AP/Al이 주성분인 2종의 혼합형 추진제를 대상으로 4종의 금속선(Ag, Cu, Al, Ni-Cr)윰 각각 삽입하여 금속선의 직경(0.1~0.8mm)과 연소 압력에 따른 금속선에 인접한 고체 추진제의 연소 속도 증가비($r_w$/$r_sb$)와 압력 지수(n)의 변화를 고찰하였다. 금속선을 추진제에 삽입함으로써 추진제의 $r_w$/$r_sb$ 는 크게 중가하였고, 1000 psia의 압력에서의 최대 6.59배 증가하였다. 그러나 녹는 점이 비교적 낮은 Al 선이나 Ag선은 금속선 직경이 작아짐에 따라 한계 직경에서 $r_w$가 급격히 감소하는 최대값이 존재하였다. 자연 발화 온도와 금속선으로 전달되는 열원인 연소 기체 불꽃 온도로 구성된 무차원군을 Buckingham pi 정리에 의해 구해진 $r_w$/$r_sb$ 에 대한 기존의 무차원 실험식에 추가로 적용하여 계산해 본 결과, 자연 발화 온도는 고려하지 않고 금속선의 녹는 점과 연소 기체의 불꽃 온도를 무시한 Hsing 의 실험식에 의해 계산된 결과보다 표준 편차가 45%이상 줄어듬을 알 수 있었다.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2003.05a
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• pp.147-150
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• 2003

The purpose of this research was to develope components of micro solid propellant thruster. Micro solid propellant thruster had four basic components: combustion chamber, nozzle, solid propellant and micro heater for ignition. A performance of micro heater and characteristic of solid propellant was investigated. Micro heater was fabricated by conventional MEMS process and Platinum layer was used for heating element. Effect of geometry parameters on micro heater was tested. The temperature responses of heater with respect to each parameters was compared for a given electrical power. The characteristic of solid propellant(HTPB/AP) was investigated to obtain burning velocity in small chamber. Additionally, a capacity of filling propellant with high viscosity in small chamber was checked to guarantee for the micro fabrication.

• Journal of the Korean Society of Propulsion Engineers
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• v.6 no.2
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• pp.19-27
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• 2002

The smokeless liner, was required to propulsion system of gas generator, was developed by appling oxamide as a filler and HTPB system as a binder. Relative quantities of smoke produced during burning of liner by propellant combustion was measured by using smoke chamber. The smoke of the smokeless liner containing 50% of oxamide as filler was reduced to 1/10 as compared with conventional liner containing silica as filler. This result was showed by ground test of gas generator.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.03a
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• pp.886-898
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• 2008

A survey is offered of the present status of microaluminized propellants industrially used worldwide in most space applications, but new directions are also pointed out making profitable use of the nanoaluminized propellants currently tested in many laboratories. Different industrial- and research-type of solid rocket propellants, mainly but not only, of the well-known family oxidizer/Al/HTPB(oxidizer being AP, AN or a mixture of the two) were experimentally analyzed at the Space Propulsion Laboratory of Politecnico di Milano. In general, they feature the same nominal composition but implement different grain size distributions of the oxidizer or metal fuel. The basic properties of all formulations were compared to that of a standard propellant already certified for flight.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.03a
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• pp.431-436
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• 2008

T-burner tests of an Al/HTPB propellant in conjunction with a Pulsed DB/AB Method were conducted to find an acoustic amplification factor. Aluminum-free and aluminum-heavy propellants were examined. Instant surface ignition was successfully made by the use of a supplementary propellant of fractionally higher reaction rate. With the presence of higher aluminum concentration in the propellants, the pressure perturbations were promptly damped down and the pressure fluctuations were no longer dispersive. Addition of aluminum particles into the propellant was advantageous for stabilizing pressure-coupled unstable waves.

• Journal of the Korean Chemical Society
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• v.51 no.3
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• pp.240-243
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• 2007

The eight coordinated yttrium(III) complex Y(tpb)3(H2O)2 (Htpb=4,4,4-Trifluoro-1-phenyl-1,3-butanedione) has been synthesized and structurally characterized by X-ray diffraction method. The coordination polyhedron of Y(tpb)3(H2O)2 has a dodecahedron. The angle between two trapezia, Y-O2-O1-O5-O6 and Y-O4-O3-O8-O7, is 89.59°. The O1-O5 and O3-O8 distances are 2.965 and 2.995 A whereas the O2-O6 and O4-O7 distances are 4.256 and 4.403 A.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2005.11a
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• pp.307-310
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• 2005

This Study was performed for the Insensitive Munition Technology Program contract between Roxel and ADD. Two Slow Coo-off(SCO) tests and One fast Cook-off(FCO) test have been made based on MIL-STD-2105C. SCO and FCO tests were made in order to evaluate the behaviour of the hybrid rocker motor with insensitive igniter and two types of propellants of which burning rates were 9.8 mm/s and 21.2 mm/s @ 7 MPa each other. The Reaction level of the two rocker motors to SCO test was classified as type IV and that of FCO test was classified as type V.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.352-358
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• 2011

This paper describes on the study of mitigation technique in which a pyrosensor is automatically sensing the rate of risk of fire or explosion of solid rocket motor exposed to an unexpected fire and makes the rocket motor burn itself safely. SCO test was carried out with a rocket motor loaded with HTPB propellant, in which a thermal pyrosensor igniter was installed. The rocket motor in SCO test was located in an oven at $50^{\circ}C$ for 7 hours. The temperature was regulated to be elevated at the rate of $3.3^{\circ}C$ per hour. Results showed Type V(Burning) reaction in this SCO test.