• Journal of the Korean Society of Propulsion Engineers
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• v.19 no.6
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• pp.91-97
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• 2015

The limit values of mechanical properties(MPs) of HTPB/AP/Al Solid Propellant was reviewed according to the rocket motor development procedures and the in-process values of MPs were analyzed by the tool of Process Capability Index. Based on finding the dependency among MPs, the optimization is proposed for reducing the properties defects and for improving the rocket grain safeties.

• Journal of the Korean Society of Propulsion Engineers
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• v.21 no.4
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• pp.21-27
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• 2017

In this paper, in order to control the burning rate and pressure exponent of HTPB/AP/Al based propellant for the improvement of performance, the effect of the size ratio of AP particles and various contents of Butacene as burning catalyst on combustion properties was investigated. In the propellant formulation with both $28{\mu}m$ Al of 23% and Butacene of 3%, the burning rate and pressure exponent were increased with increasing the contents of $9{\mu}m$ AP particles. And the burning rate was increased with increasing the contents of Butacene with showing the relatively low pressure exponent in the propellant containing Butacene. However, the significant variations of pressure exponent by contents of Butacene were not observed.

• Journal of the Korean Society of Propulsion Engineers
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• v.13 no.5
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• pp.1-6
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• 2009

A development of a composite solid propellant is carried out for an application to gas generators as an energy source of rocket system. With HTPB as a propellant binder which has 80% of particle loading ratio, a favorable rheology, and moderate curing properties at the range of $-50^{\circ}{\sim}70^{\circ}C$, AN is selected as the first kind of oxidizer having the characteristics of a low flame temperature, minimal particle residual as well as nontoxic products. AP is the second oxidant for ballistic property control. A series of experiments for the improvement of physical properties were conducted and resulted in the propellant formulation having 30% of strain rate at 8 bar of max. stress.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.11a
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• pp.283-287
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• 2008

The heat released during the external frictional motion is a factor responsible for initiating energetic materials under all types of mechanical stimuli including impact, drop, or penetration. We model the friction-induced ignition of HMX, RDX and AP/HTPB propellant using the BAM friction apparatus and one-dimensional time-to-explosion apparatus whose results are used to validate the friction ignition mechanism and the deflagration kinetics of energetic materials, respectively. The ignition times for each energetic sample due to friction are presented.

• Journal of the Korean Society of Propulsion Engineers
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• v.4 no.4
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• pp.50-58
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• 2000

Burning rate of the matrix propellant($r_{sb}$) and burning rate along the metal wire($r_w$) were measured and analyzed for the HTPB/AP/Al propellant embedded with Ag wire($\phi$0.15mm) according to weight % of RDX(0~20%). Variation of burning rate increment ratio($r_w$/$r_{sb}$) and pressure exponent(n) was studied for the nitramine propellant having 10% RDX embedded with three kinds of metal(Ag, Cu, and Ni-Cr) of which diameter range is between 0.1~0.6mm. Maximum burning rate increment ratio of the nitramine propellant embedded with Ag wire($\phi$0.1mm) was 5.94 at $20^{\cire}C$, 1000 psia, 16.4% faster than that of HTPB/AP propellant, it is because that autoignition temperature of the nitramine propellant was higher than that of HTPB/AP propellant. Standard deviation of absolute ($r_{wc}$/$r_{we}$)/$r_{we}$ calculated by using new empiracal equation composed of four dimensionless groups, is 6.11% less than that calculated by using original empirical equation composed of three dimensionless group. The new empiracal equation is derived from Buckingham pi theorem using the parameters such as thermal diffusivity, melting temperature. wire diameter, propellant sample diameter, frame temperature, autoignition temperature and matrix burning rate which influence on $r_w$.

• 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.

• Journal of the Korean Society of Propulsion Engineers
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• v.9 no.2
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• pp.46-53
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• 2005

We have analyzed the temperature and humidity aging test results of a composite solid propellant. The temperature aging test was performed to evaluate the storage life of a propellant, while the humidity aging test could provide the hygroscopicity of Ammonium Perchlorate(AP) exposed to .elative humidity (RH) 10, 30, $50\%$ environment. A specimen was used in the temperature test, and a block of propellant from the actual motor was used in the humidity test. We report that the 4-month storing at 60 degree is equivalent to the 10-year 60 degree condition. The composite solid propellant with HTPB binder showed signs of hardening with time lapse but the effect of humidity up to RH $50\%$ was not noticeable.

• Journal of the Korean Society of Propulsion Engineers
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• v.10 no.4
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• pp.77-84
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• 2006

With the development of micro/nano spacecraft, concepts of micro propulsion are introduced for orbit transfer and drag compensation as well as attitude control. Micro solid propellant thruster has been attention as one of possible solution for micro thruster. In this paper, micro solid propellant thruster is introduced and research on basic components of a micro solid propellant thruster is reported. Micro Pt igniter was fabricated through negative patterning and quantitative effect of geometry was estimated. The characteristic of HTPB/AP solid propellant was investigated to measure the homing velocity. A combustion chamber was fabricated by means of anisotropic etching of photosensitive glass. Finally, micro solid propellant thrusters having various geometries were fabricated and tested.

• Journal of the Korean Society of Propulsion Engineers
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• v.12 no.4
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• pp.1-6
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• 2008

The burning characteristics of HTPB/AP solid propellants were measured by closed bomb of internal volume of 200 cc and 700 cc up to 30,000 psia. The burning rates of closed bomb method showed good agreement with those of strand burner method between 1,000 psia and 5,000 psia, and the sharp increment of pressure exponent(n) around 6,000 psia as a result of testing in accordance with loading densities. The burning rate measured in 200 cc and 700 cc of internal volume of closed bomb agreed well without the relation of internal volume size.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 1998.10a
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• pp.25-25
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• 1998

최근의 선진국의 전략미사일에 사용되는 추진제 개발 방향은 고에너지, 저민감, 무연 및 무공해 추진제를 목표로 하고 있다. 현재까지 가장 널리 사용되어온 AP는 성능이나 기계적 특성에서 우수한 산화제이나 로켓 모타 배출기체 중에 HC1을 생성함으로써 2차 연기를 생성하며, HC1 자체가 인체에 유해한 물질이라는 단점들을 가지고 있다. 또한 RDX나 HMX는 에너지가 높은 반면 쇽(shock)에 대해 매우 민감한 단점을 가지고 있다. AN은 이러한 HC1을 생성하는 AP의 단점과 쇽에 민감한 RDX와 HMX의 단점을 해결할 수 있는 대체 산화제로 사용할 수 있으나, 성능이 대외적으로 떨어지며 상변이와 흡습성이 높은 단점이 있다. ADN은 -100~+10$0^{\circ}C$범위에서 상 변이가 없고 밀도는 1.801g/㎤, Hf는 -290cal/g이다. ADN을 산화제로 사용하면 HTPB/AP 추진제와 동등하거나 그 이상의 성능을 갖는 추진제를 만들 수가 있으며, 유해 기체인 HC1을 배출하지 않음으로 인하여 2차 연기를 없애는 장점이 있다.