• Korean Journal of Air-Conditioning and Refrigeration Engineering
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• v.28 no.7
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• pp.275-282
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• 2016

MILD (Moderate and Intense Low-oxygen Dilution) combustion using oxygen as an oxidizer is considered as one of the most promising combustion technologies for high energy efficiency and for reducing nitrogen oxide and carbon dioxide emissions. In order to investigate the effects of nozzle angle and oxygen velocity conditions on the formation of oxygen-MILD combustion, numerical and experimental approaches were performed in this study. The numerical results showed that the recirculation ratio ($K_V$), which is an important parameter for performing MILD combustion, was increased in the main reaction zone when the nozzle angle was changed from 0 degrees to 15 degrees. Also, it was observed that a low and uniform temperature distribution was achieved at an oxygen velocity of 400 m/s. The perfectly invisible oxy-MILD flame was observed experimentally under the condition of a nozzle angle of $10^{\circ}$ and an oxygen velocity of 400 m/s. Moreover, the NOx emission limit was satisfied with NOx regulation of less than 80 ppm.

• Journal of the Korean Society of Propulsion Engineers
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• v.22 no.5
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• pp.125-131
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• 2018

Some propellants in a liquid rocket engine are burned in the pre-burner of a staged combustion cycle engine, resulting hot gas drives the turbine. The burned gas passing through the turbine is supplied to the combustor at high temperature and pressure. The form of the gas can be fuel rich or oxidizer rich dependent upon the mixture ratio or the engine scheme. When the cycle works at oxidizer-rich condition, the metal pipes composing the engine can be ignited or even exploded by an impact of very a small particle. In this study, we developed the powder combination and processes for an anti-oxidation coating through the analysis of various coating materials.

• Journal of the Korean Society of Propulsion Engineers
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• v.26 no.6
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• pp.34-42
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• 2022

A correlation between propellant supply condition and chamber pressure in GCH₄-LOx small rocket engine was explored and hot-firing tests were conducted to analyze the engine performance characteristics according to the equivalence ratio variation at a constant chamber pressure. Correlation studies have shown that chamber pressure is linearly proportional to oxidizer supply pressure. As a result of the test, the thrust, specific impulse and characteristic velocity that are the main performance parameters of a rocket engine, were found to be enhanced as the equivalence ratio starting from a fuel-lean condition approached the stoichiometric ratio, but the efficiencies of characteristic velocity and specific impulse were on the contrary, in their dependency on the equivalence ratio.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.35 no.12
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• pp.1108-1114
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• 2007

The effect of oxidizer injection angle on the combustion characteristics of end-burning hybrid combustor is numerically investigated. Besides the previously studied parameter(injector arrangement, port diameter and O/F ratio), three different injection angle are considered: parallel angle to fuel surface(Case 1), +30 degree inclined angle toward the fuel(Case 2) and 30 degree inclined angle toward the nozzle(Case 3). It is found that Case 2 has the best mixing pattern in the upstream area but has the worst combustion efficiency since non negligible amount of unburned fuel is expelled from the nozzle. In contrast, though Case 1 and Case 3 showed relatively low mixing effect than the Case 2, they had high combustion efficiency. The comparison of numerical results between Case 1 and Case 3 demonstrate that no major difference is encountered, however, Case 1 is expected to have the best combustion efficiency due to the low residence time of the Case 3 injector which heads toward the nozzle.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.489-494
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• 2017

This paper describes the combustion characteristics of aluminized paraffin fuel on the contents of micron-sized aluminum particles with nominal diameters of $8{\mu}m$. Aluminized paraffin fuels with mixture ratio of aluminum 0 wt%, 5 wt% and 10 wt% as fuel and GOx(Gaseous Oxygen) as oxidizer were used to perform the experiments. The experimental investigations were performed on the regression rate, the chamber pressure and the combustion efficiency. Increasing a content of micron-sized aluminum particles, the results of regression rate, chamber pressure and combustion efficiency show minor increase compared to those without particles.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.04a
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• pp.1-2
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• 2011

Hybrid rockets have lately attracted attention as a strong candidate of small, low cost, safe and reliable launch vehicles. A significant topic is that the first commercially sponsored space ship, SpaceShipOne vehicle chose a hybrid rocket. The main factors for the choice were safety of operation, system cost, quick turnaround, and thrust termination. In Japan, five universities including Hokkaido University and three private companies organized "Hybrid Rocket Research Group" from 1998 to 2002. Their main purpose was to downsize the cost and scale of rocket experiments. In 2002, UNISEC (University Space Engineering Consortium) and HASTIC (Hokkaido Aerospace Science and Technology Incubation Center) took over the educational and R&D rocket activities respectively and the research group dissolved. In 2008, JAXA/ISAS and eleven universities formed "Hybrid Rocket Research Working Group" as a subcommittee of the Steering Committee for Space Engineering in ISAS. Their goal is to demonstrate technical feasibility of lowcost and high frequency launches of nano/micro satellites into sun-synchronous orbits. Hybrid rockets use a combination of solid and liquid propellants. Usually the fuel is in a solid phase. A serious problem of hybrid rockets is the low regression rate of the solid fuel. In single port hybrids the low regression rate below 1 mm/s causes large L/D exceeding a hundred and small fuel loading ratio falling below 0.3. Multi-port hybrids are a typical solution to solve this problem. However, this solution is not the mainstream in Japan. Another approach is to use high regression rate fuels. For example, a fuel regression rate of 4 mm/s decreases L/D to around 10 and increases the loading ratio to around 0.75. Liquefying fuels such as paraffins are strong candidates for high regression fuels and subject of active research in Japan too. Nakagawa et al. in Tokai University employed EVA (Ethylene Vinyl Acetate) to modify viscosity of paraffin based fuels and investigated the effect of viscosity on regression rates. Wada et al. in Akita University employed LTP (Low melting ThermoPlastic) as another candidate of liquefying fuels and demonstrated high regression rates comparable to paraffin fuels. Hori et al. in JAXA/ISAS employed glycidylazide-poly(ethylene glycol) (GAP-PEG) copolymers as high regression rate fuels and modified the combustion characteristics by changing the PEG mixing ratio. Regression rate improvement by changing internal ballistics is another stream of research. The author proposed a new fuel configuration named "CAMUI" in 1998. CAMUI comes from an abbreviation of "cascaded multistage impinging-jet" meaning the distinctive flow field. A CAMUI type fuel grain consists of several cylindrical fuel blocks with two ports in axial direction. The port alignment shifts 90 degrees with each other to make jets out of ports impinge on the upstream end face of the downstream fuel block, resulting in intense heat transfer to the fuel. Yuasa et al. in Tokyo Metropolitan University employed swirling injection method and improved regression rates more than three times higher. However, regression rate distribution along the axis is not uniform due to the decay of the swirl strength. Aso et al. in Kyushu University employed multi-swirl injection to solve this problem. Combinations of swirling injection and paraffin based fuel have been tried and some results show very high regression rates exceeding ten times of conventional one. High fuel regression rates by new fuel, new internal ballistics, or combination of them require faster fuel-oxidizer mixing to maintain combustion efficiency. Nakagawa et al. succeeded to improve combustion efficiency of a paraffin-based fuel from 77% to 96% by a baffle plate. Another effective approach some researchers are trying is to use an aft-chamber to increase residence time. Better understanding of the new flow fields is necessary to reveal basic mechanisms of regression enhancement. Yuasa et al. visualized the combustion field in a swirling injection type motor. Nakagawa et al. observed boundary layer combustion of wax-based fuels. To understand detailed flow structures in swirling flow type hybrids, Sawada et al. (Tohoku Univ.), Teramoto et al. (Univ. of Tokyo), Shimada et al. (ISAS), and Tsuboi et al. (Kyushu Inst. Tech.) are trying to simulate the flow field numerically. Main challenges are turbulent reaction, stiffness due to low Mach number flow, fuel regression model, and other non-steady phenomena. Oshima et al. in Hokkaido University simulated CAMUI type flow fields and discussed correspondence relation between regression distribution of a burning surface and the vortex structure over the surface.

• 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.

• Aerospace Engineering and Technology
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• v.2 no.1
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• pp.164-170
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• 2003

KSR-III engine with film-cooled baffle was tested. The purpose of this test is to verify the effect of ablative baffle on avoiding combustion instability which occurred in the acoustic cavity case. The engine had expansion ratio of 5.04 and the test condition was design condition(oxidizer mass flow rate 42.04, and fuel 17.95 kg/s). In the test, combustion instability did not occur. So, the effect of film-cooled baffle on avoiding combustion instability was verified.

• Journal of Mechanical Science and Technology
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• v.20 no.6
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• pp.874-881
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• 2006

Combustion stability characteristics in actual full-scale combustion chamber of a rocket engine are investigated by experimental tests with the model (sub-scale) chamber. The present hot-fire tests adopt the combustion chamber with three configurations of triplet impinging-jet injectors such as F-O-O-F, F-O-F, and O-F-O configurations. Combustion stability bound-aries are obtained and presented by the parameters of combustion-chamber pressure and mixture (oxidizer/fuel) ratio. From the experimental tests, two instability regions are observed and the pressure oscillations have the similar patterns irrespective of injector configuration. But, the O-F-O injector configuration shows broader upper-instability region than the other configurations. To verify the instability mechanism for the lower and upper instability regions, air-purge acoustic test is conducted and the photograph or the flames is taken. As a result, it is found that the pressure oscillations in the two regions can be characterized by the first impinging point of hydraulic jets and pre-blowout combustion, respectively.

• Advances in aircraft and spacecraft science
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• v.1 no.3
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• pp.273-289
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• 2014

Accuracy of a reconstruction technique assuming a constant characteristic exhaust velocity ($c^*$) efficiency for reducing hybrid rocket firing test data was examined experimentally. To avoid the difficulty arising from a number of complex chemical equilibrium calculations, a simple approximate expression of theoretical $c^*$ as a function of the oxidizer to fuel ratio (${\xi}$) and the chamber pressure was developed. A series of static firing tests with the same test conditions except burning duration revealed that the error in the calculated fuel consumption decreases with increasing firing duration, showing that the error mainly comes from the ignition and shutdown transients. The present reconstruction technique obtains ${\xi}$ by solving an equation between theoretical and experimental $c^*$ values. A difficulty arises when multiple solutions of ${\xi}$ exists. In the PMMA-LOX combination, a ${\xi}$ range of 0.6 to 1.0 corresponds to this case. The definition of $c^*$ efficiency necessary to be used in this reconstruction technique is different from a $c^*$ efficiency obtained by a general method. Because the $c^*$ efficiency obtained by average chamber pressure and ${\xi}$ includes the $c^*$ loss due to the ${\xi}$ shift, it can be below unity even when the combustion gas keeps complete mixing and chemical equilibrium during the entire period of a firing. Therefore, the $c^*$ efficiency obtained in the present reconstruction technique is superior to the $c^*$ efficiency obtained by the general method to evaluate the degree of completion of the mixing and chemical reaction in the combustion chamber.