• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.949-952
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• 2011

Study on differential pressure in the chamber of cannon by adjusting the mass flow of ignition-gas has been conducted using the 1-D interior ballistics numerical code called IBcode. In case of large-caliber cannon, high temperature ignition-gas is injected to the chamber through the side hole of the primer to ignite the propellant. Therefore, mass flow of injected ignition-gas affects the propellant combustion in the chamber. Mass flow of each side hole of the current primer was uniformly distributed. In this study, differences of propellant combustion with different mass flow of each side hole have been imposed. Results in case of the mass flow increase in the direction to the base show that the differential pressure decreases compared to the uniformed mass flow.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2012.05a
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• pp.106-109
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• 2012

It was designed and tested ignition that an oxidizer rich preburner for a staged combustion cycle liquid rocket engine propelled by kerosene and LOx. Operation conditions of the preburner are about 60 of OF ratio and 20 MPa of combustion pressure. Ignition characteristics were compared by propellants flowrate. As the results, the higher propellants flowrate, the shorter the ignition delay time and the higher ignition stiffness. The ignition delay time was affected by incoming the oxidizer flowrate through the refrigerative cooling channels. The oxidizer flowrate from the cooling channels decreased by inflow of combustion gas during initial ignition. The oxidizer flowrate of the cooling channels increases, it is rapid recovery by cooling effect, eventually the ignition delay time decreases.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.05a
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• pp.383-386
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• 2006

Several HTPB/AP and HTPB/AP/HMX propellants were investigated experimentally for ignition characteristics in subatmospheric pressure. The threshold ignition pressure was 4psia for HTPB/AP composite propellant. The partial replacement of AP in HTPB/AP composite propellant by $5\sim15%$ of HMX, HNIW showed improvements in the threshold pressure was below 0.4psia. This appears to be due to the exothermic dissociation characteristics of HMX and HNIW at lower temperature $(\sim220^{\circ}C)$ than that of AP. The ignition substance B/KNO3 was coated thinly on the propellant surface for better ignition effect. As a result, ignition delay time of 15% was improved. NC is applied to $B/KNO_3$ ignition substance as a secondary binder and $NC-B/KNO-3$ suspension solution is coated to the propellant surface.

• Journal of the Korean Society of Propulsion Engineers
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• v.24 no.4
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• pp.79-88
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• 2020

Hypergolic propellants have been widely used for space propulsion systems based their long-term storability and high ignition reliability. Since conventional hypergolic propellants are highly toxic and carcinogenic, handling and operating costs are significant. To overcome the drawbacks, numerous studies have been actively performed to develope new hypergolic propellants, ensuring that the combinations are high performance, low toxicity and low environmental impact. In the present study, a comprehensive survey was conducted to summarize the research and development of green hypergolic propellants involving hydrogen peroxide, nitric acid, and ionic liquids.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.04a
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• pp.225-228
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• 2007

The objective of this study is to characterize design parameters of rocket igniters for composite, double base and nitramine propellant. Arc image furnace and fiber optics surface reflectometer were used to measure ignition delay time and reflected optical energy of several compositions of composite, double base and nitramine base rocket propellant at different pressure levels each other. The order of ignitability was double base > composite > Nitramine propellants at initial pressure of over 75 psia. The highest ignition energy was needed to ignite nitramine propellant, however, as the pressure increased up to the range of $75{\sim}400$ psia as the ignition delay time decreased abruptly. The absorbtion of radiation energy could be increased by the addition of small amount of opacifiers as carbon black, ZrC, WC and burning catalyst.

• Journal of Aerospace System Engineering
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• v.14 no.3
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• pp.24-31
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• 2020

A study on the H2O2 based low toxic hypergolic propellant was conducted. The fuel candidates were chosen as a mixture of Amine solvent and reactive additive. The analytical performance was calculated via the NASA CEA code and 96% Isp of the NTO/UDMH was confirmed. The ignition delay measurement with drop test was performed and all candidates showed less than 10 ms in the best performance cases. Based on these results, the feasibility of high response H2O2 based low toxic hypergolic propellant was confirmed.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.05a
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• pp.211-214
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• 2009

Using the numerical code for the interior ballistics, the performance of the interior ballistics with the characteristic of the ignition injection has been investigated. When the maximum position of ignition injection is near the base, the pressure distribution at the chamber of the interior ballistics was uniform and the final projectile velocity is increased.

• Journal of the Korean Society of Propulsion Engineers
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• v.26 no.2
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• pp.20-27
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• 2022

Hypergolic propellant ignited spontaneously when fuel and oxidizer contact without ignition system. Due to this characteristic, the risk of accidents is high when new propellants are evaluated. Prevention of accidents is very important because the damage can be large when the accident occur. In this work, we proposed non-ignition evaluation method which can replace conventional ignition evaluation method by using microreactor. The reactor was fabricated by MEMS. The heat of reaction as according to fuel and NaBH₄ was estimated. At the condition of highest heat of reaction ignition was observed by drop test.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.05a
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• pp.138-141
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• 2010

A study to reduce the ignition peak pressure of gas generator for the missile launching system was accomplished. The igniter, as the energy release device for igniting the propellant, is aimed at simultaneous ignition of bundled 3-layered propellant grain without unstable burning. In case of our gas generator which must use the double-base propellant with low ignition property, the fast ignition of propellant and reduction of initial peak pressure should be required for the satisfaction of ejection velocity and acceleration condition. By applying MTV ignition charge for the igniter of gas generator, we accomplished all system performance requirements.

• Journal of Aerospace System Engineering
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• v.17 no.1
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• pp.51-58
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• 2023

Hypergolic propellants, which can ignite themselves without an ignition source, are difficult to handle due to their corrosiveness and toxicity. Therefore, it is necessary to develop green hypergolic propellants with little or no toxicity. In this study, basic research on green hypergolic ignition propellants was conducted. With 95% hydrogen peroxide as an oxidizer and CNU_HGFv1 as a fuel, ignition and combustion characteristics of propellants were evaluated through a drop test, an ignition test, and a combustion test. As a result of the drop test, the ignition delay time was 9.7 ms. It was 27 ms in the ignition test, which was fast enough to be used as a propellant. As a result of the combustion test, a combustion efficiency of 95.4~98.1% was achieved at about 11.7 bar. It was confirmed that fast and stable combustion was possible without hard start or combustion instability.