• Journal of the Korean Society of Propulsion Engineers
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• v.5 no.2
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• pp.1-7
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• 2001

To protect combustion chamber from high temperature combustion gas, regenerative cooling is used for most liquid rocket engine. Although regenerative cooling is the most effective way to protect the chamber from high heat flux, realization of this system requires detail analysis, manufacturing technique and high cost. To demonstrate the possibility of applying regenerative cooling to a real rocket engine, the hot fire test has been carried out for the sub-scale liquid rocket with the water cooling system. The main purpose of the test is to identify the problem area of design, safety and cost effective manufacturing technique. The coolant passage was 3 mm in width and wall thickness was 1 mm with stainless steel. Maximum combustion time and pressure were 60 seconds and 400 psi, respectively. The flow rate of coolant was reduced gradually from 2 kg/s to 0.12 kg/s throughout firing test, combustion chamber was visually examined and no dwfect was observed.

• Transactions of the Korean Society of Mechanical Engineers A
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• v.41 no.1
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• pp.69-76
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• 2017

In this study, design of a small hybrid rocket is carried out using Failure Mode and Effects Analysis (FMEA) and Criticality Analysis(CA), which is a method for qualitative and semi-quantitative reliability analysis. In order to carry out FMEA, the structure of the hybrid rocket is divided into 31 parts and 72 potential failure modes. As a result of the FMEA, the relationship between potential failure modes, causes and effects, and their severity are evaluated qualitatively. Criticality analysis is followed for the failure modes, in which the criticality number is estimated using the failure rate information available from the handbook. Moreover, the failure modes with higher criticality and severity are chosen for improvement, and a series of design or material changes are made for the improvement of the hybrid rocket reliability.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.11a
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• pp.375-379
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• 2008

In this paper, a 10N class liquid propellant rocket utilizing a dissolving reaction of hydrogen peroxide is constructed and tested. Through a series of designs, seven orifices with a diameter of 200 ${\mu}m$ and a nozzle with a neck of 2.5mm in diameter and area ratio of 2.56 were made. The platinum coated on Isolite was used for catalyst. 90wt% peroxide pressed at 20 bar by nitrogen gas was used for performance evaluation. The length of the catalyst bed and the load of platinum was taken as the parameters for this experiment. For the catalyst support length of 4cm loaded on 5wt% platinum, satisfactory $c^*$ efficiency and stable thrust was observed. The light weight body of the rocket was composed of aluminum. Rocket rose about 10m with relatively constant velocity in launching test.

• Journal of the Korean Society of Propulsion Engineers
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• v.18 no.5
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• pp.54-61
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• 2014

In this study, a plume exhausted from rocket nozzle is investigated by using an unstructured 2-dimensional axisymmetirc DSMC code at various altitude. The small back-pressure to total-pressure ratio($P_b/P_o$) and large $P_b/P_o$ represent low and high altitude condition, respectively. At low altitude, the plume shows a typical complicated structure (e.g. Mach disk) of underexpanded jet while the high altitude plume experiences plain expansion. The various features of exhaust plume is discussed including density, translational/rotational temperature, Mach number and Knudsen number. The results shows that even at 20 km altitude where the freestream Knudsen number is small as $1.5{\times}10^{-5}$, the transitional and rarefied flow regimes can occur locally within the plume. It confirms the necessity of DSMC computation at low altitude.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2005.11a
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• pp.213-219
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• 2005

Small liquid rocket engine (SLRE) with calorimeter were developed and tested to evaluate cooling characteristics in the liquid rocket engine. Therefore, cooling performance analysis was performed to predict the heat transfer coefficient on gas side wall in 10 calorimeter channel. A heat transfer empirical formula was determined by results of firing test and computational simulation.

• Journal of the Korean Society of Propulsion Engineers
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• v.10 no.3
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• pp.32-40
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• 2006

Dynamic Radioscopy technique was developed to observe internal phenomena of motors during firings. For use of this technique, generator and image box consisting of conversion screen, mirror, and video camera were designed and fabricated at our laboratory, and static firing tests were conducted to small motors. Protection devices against vibration and noise induced from during combustion were also made and their performances were evaluated with reference to the vibration criteria. Test results have shown that the vibration levels at concerned points were measured less than vibration criteria, and the internal phenomena during firings were also observed clearly.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.04a
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• pp.370-373
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• 2007

In this study, a hybrid rocket motor with separable and detachable oxidizer tank from combustion chamber is developed. Initially, the measured thrust of the motor showed about 30% of the design thrust since the oxidizer supply was not enough. In order to solve this problem, application is made to expand the orifice diameter of oxidizer injector empirically, so that the mass flow rate of oxidizer was improved. The improved performance was about 60% of design thrust, 18kgf, and thrust-to-weight ratio was reasonable, compared with other sounding rockets.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.221-224
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• 2010

Recently, due to the enormous cost for sending a satellite into an orbit, small and more reliable satellites have been more demanded. An introduction of new binders(HTPB, GAP) and new oxidizers made great improvements of the large thrust modulation. In order to make cost reduction, one prefers to the low melting temperature thermoplastic propellant reforming the manufacturing process dramatically. Solid propellant rockets have been had a problem of the injection accuracy into orbit, but PBS(Post Boost Stage) using a liquid mono-propellant improves the injection accuracy. This paper also gives the direction of the advanced nozzle materials and the motor case.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.05a
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• pp.21-24
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• 2010

The cooling channel of the preburner for staged combustion engines was studied. The combustion pressure of the researched preburner is about 210 bar which is very high compared with the engines of the Korean Launch Vechicle and 30 ton class liquid rocket engines developed as a pre-research program. Also, the combustion is an oxygen rich process unlike the gas generators of open cycle kerosene engines. Thus the cooling process is very important to make the preburner stable. Many configurations for the preburner were developed and numerically analyzed. As a result, the pressure loss could be reached below the target.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.11a
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• pp.261-264
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• 2008

This study experimentally examines combustion characteristics of a hybrid rocket in which solid paraffin is used as a fuel, while oxidizer is pure oxygen. Especially, the experiment investigates the effects of chamber pressure and configuration of fuel grain. The pressure inside the combustion chamber is varied by changing a flow rate of oxidizer. The regression rate is observed to increase as the chamber pressure does. There also exists the effects of shape of fuel grain on thrust. Characteristic of paraffin hybrid rocket changes with shape of fuel grain. When there is a room near the injector, thrust increases. On the other hand, the room near the nozzle does not contribute to thrust increasement.