• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.163-170
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• 2011

Applicability of regenerative cooling in 2,500 N-class bipropellant thruster using hydrogen peroxide and kerosene was considered for improvement performance and application in various missions. Calculation was performed by one dimensional approach using hydrogen peroxide as a coolant. In designed regenerative cooling thruster, heat flux at nozzle throat was estimated at 18 ~ 20 $MW/m^2$. Designed cooling channel width and height were 2.5 mm and 0.5 mm, respectively. Based on designed cooling channel configuration, flat plate model was manufactured and tested for estimation of pressure drop in cooling channel, and CFD analysis was compared with the test result. The maximum error between CFD analysis and experimental result was approximately 13% and average error was approximately 5%.

• Journal of the Korean Society of Propulsion Engineers
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• v.17 no.6
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• pp.120-130
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• 2013

A study was conducted for the performance characteristics of methane taking recently the limelight in the world as a next-generation propellant, with the survey for state of the art in the development of methane/oxygen rocket engine being accompanied. Liquid methane as a rocket fuel has the favorable characteristics such as non-toxic, low cost, regenerative cooling capability, and potential for in-situ resource utilization (ISRU). The combination of liquid methane and liquid oxygen also provides the excellent performance including high specific impulse and low system mass. For these reasons, many researches have been actively carried out on the methane/oxygen engine, nevertheless, its technology readiness level is not that high enough just yet. Therefore, it is judged that it is the time to mitigate the technical gap with the space technology of advanced countries through a swift onset of the development of methane rocket engine.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.787-790
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• 2017

This study is a preliminary research on characterization of methane - oxygen combustion used in bipropellant thruster. The limit of combustion stability and flame shape of methane - oxygen non-premixed flame injected by shear coaxial injector in the model combustion chamber Experimental studies have been carried out. A direct image of the flame was photographed using a DSLR camera, and combustion characteristics and flame length were quantified through image post-processing. As a result, it was confirmed that the stabilized flame was generated at the stoichiometric ratio as the oxidizer Reynolds number ($Re_o$) was increased, and the length of the turbulent flame was increased under the same injector diameter condition.

• Journal of the Korean Society of Propulsion Engineers
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• v.20 no.4
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• pp.68-76
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• 2016

This research is carried out for the performance evaluation of the injector that is one of the critical components of bipropellant-rocket-engine. Spray characteristics are investigated in detail according to the recess length and injection pressure on the swirl-coaxial-injector using gaseous methane and liquid oxygen as propellants. A visualization is conducted by the Schlieren photography that is composed of a light source, concave mirrors, knife, and high-speed-camera. A hollow-cone-shape is identified in the liquid spray that is spread only by inner injector and the spray angle is decreased due to the diminution of swirl strength in accordance with the increase of the length of injector orifice. When the injector sprays the liquid through the inner injector with the aid of gas through the outer injector, the spray angle in external mixing region tends to increase with rise of the recess length, while in internal mixing region, it is decreased. It is also confirmed that the same tendency of the spray angle with recess length appears irrespective of the injection pressure of liquid spray.

• Journal of the Korean Society of Propulsion Engineers
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• v.21 no.1
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• pp.63-71
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• 2017

Liquid methane ($CH_4$)/oxygen ($O_2$) bipropellants have been recently considered as a next-generation propellant due to eco-friendly and non-toxic properties, low cost and high performance. In this study, the combustion characteristics of gaseous $CH_4/O_2$ nonpremixed coflow flames in a model combustor with variation of internal pressure are investigated through measuring the combustion stability limits and visualizing flames. Results show that the combustion stability limits are extended and the reaction region is widened with increasing internal pressure of the combustion chamber for fuel-rich condition.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2003.10a
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• pp.131-134
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• 2003

There has been a renewed interest in the use of hydrogen peroxide as an oxidizer in bipropellant liquid rocket engines as well as in hybrid rocket engines. This is because hydrogen peroxide is a propellant of low toxicity and enhanced versatility. The present paper details the features of the designed engine of l00N thrust and its facility. Also explained is the arrangement of the distillation unit to be used to prepare rocket-grade hydrogen-peroxide propellant. Results of the simulated "cold" tests are presented.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.36 no.5
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• pp.491-496
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• 2008

New concept ramjet propulsion system with liquid bipropellant rocket using "Green Propellant" hydrogen peroxide for launch stage is proposed. In this novel concept, hydrogen peroxide gas generator produces hot oxygen at launch stage and kerosene injects to this jet in combustor. For basic study of this new concept ramjet system, investigation of auto-ignition characteristics and combustion of decomposed hydrogen peroxide and kerosene was conducted. In various test cases, auto-ignition and stable combustion was verified. The combustion temperature of 400°C and Fuel/Oxidizer mixture ratio of 0.6 were the limit of auto ignition. Through the experiment results, the possibility of novel concept combined propulsion system using hydrogen peroxide gas generator is ascertained.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2005.11a
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• pp.191-195
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• 2005

Propellant remains as outage at engine shutdown contributes no useful impulse to the rocket and produces an unwanted increase in burnout weight. Minimization of outage, is therfore is a basic consideration in attaining the maximum performance capability of my bipropellant liquid rocket. This paper present the calculation procedures of outage and optimum loading propellant mass. And some control methods and measurement techniques for outage are presented.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.1061-1067
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• 2017

Since the 1960s hydrazine is used as a propellant to power rocket, satellites or deep space missions. Due to hydrazine's high toxicity and operating cost, the request for Green Propellant as energetic ionic liquids(HAN, ADN), nitrous oxide blends is growing. Nitrous Oxide Fuel Blend(NOFB) having advantage of a bipropellant performance as well as the advantage of a mono-propellant in respect to the simple propellant tank and feed system. It is worth replacing traditional hydrazine based propellant system if handled and designed properly.

• Journal of the Korean Society of Propulsion Engineers
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• v.21 no.2
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• pp.1-8
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• 2017

In order to analyze combustion characteristics of methane-oxygen diffusion flame in a model combustor, combustion experiments were carried out under various spray conditions of propellant scrutinizing combustion stability limit and flame shapes. As the propellant approached the theoretical equivalence ratio condition, a stable detached flame was observed even under high oxygen Reynolds number. And the length of the visible flame increased and the lift-off distance of the flame exhibited a tendency toward decrease. Due to the swirl effect of the propellant by the swirl-coaxial injector, a wide and short flame was produced. Thus, it may be appropriate to employ the swirl-coaxial injector in thrusters having a limited physical dimension.