• Journal of the Korean Society of Propulsion Engineers
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• v.24 no.4
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• pp.55-65
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• 2020

In this paper, development and acceptance test results of 75-tonf class liquid rocket engine gas generators are described. Up to now, more than 330 times and cumulative time of 7,000 seconds gas generator autonomous tests have been carried out with 44 gas generator models. Through the tests it was verified that 75 tonf gas generator shows very reliable and reproducible characteristics in terms of chamber pressure, combustion efficiency, pressure loss, combustion stability, burnt gas temperature, and etc. 5 gas generators which are the last series of 75 tonf gas generator for the Korea Space Launch Vehicle II, will be manufactured until end of 2019 and their acceptance tests will be executed at the first half of 2020.

• Journal of the Korean Society of Propulsion Engineers
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• v.25 no.6
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• pp.45-52
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• 2021

A hot-firing test was conducted using gaseous-methane and gaseous-oxygen under highly fuel-rich condition as a prior study for the development of a liquid propellant small rocket engine. To compare combustion characteristics for various equivalence ratios, the oxygen flow rate was set to 12 g/s and the methane flow rate was changed according to the equivalence ratio. As a result, it was observed that the steady-state characteristic velocity obtained during the hot-firing test steeply rose in the latter part of each test: the difference between the former and the latter steady value was enhanced overall in proportion to the equivalence ratio. Based on this, the equivalence ratio range depending on the variational characteristics of the characteristic velocity could be divided into three combustion regimes.

• 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.

• The Korean Journal of Air & Space Law and Policy
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• v.19 no.1
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• pp.9-35
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• 2004

A launching State shall be absolutely liable to pay compensation for damage caused by its space object on the surface of the earth or to aircraft in flight. The compensation which the launching State shall be liable to pay for damage under "the Convention on International Liability for Damage caused by Space Objects" shall be determined in accordance with international law and the principles of justice and equity, in order to provide such reparation in respect of the damage as will restore the person, natural or juridical, State or international organisation on whose behalf the claim is presented to the condition which would have existed if the damage had not occurred. In the event of damage being caused elsewhere than on the surface of the earth to a space object of one launching State or to persons or property on board such a space object by a space object of another launching State, and of damage thereby being caused to a third State or to its natural or juridical persons, the first two States shall be jointly and severally liable to the third State, to the extent indicated by the following: If the damage has been caused to the third State on the surface of the earth or to aircraft in flight, their liability to the third State shall be absolute; If the damage has been caused to a space object of the third State or to persons or property on board that space object elsewhere than on the surface of the earth, their liability to the third State shall be based on the fault of either of the first two States or on the fault of persons for whom either is responsible. The Insurance requirements are satisfied for a launch or return authorised by a launch permit if the holder of the permit or authorisation is insured against any liability that the holder might incur to pay compensation for any damage to third parties that the launch or return causes; and the Commonwealth is insured against any liability that Commonwealth might incur, under the Liability Convention or otherwise under international law, to pay compensation for such damage. The liability for Damage caused by Space Objects should be regulated in detail in Korea.

• Applied Chemistry for Engineering
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• v.18 no.3
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• pp.227-233
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• 2007

Carbon/carbon (C/C) composites as unique materials possess exceptional thermal resistance with light weight, high stiffness, and strength even at high temperature. However, one serious obstacle for application of the C/C composites is their poor oxidation resistance in high temperature oxidizing environments. SiC coating has been employed to protect the composites from oxidation. This study explored combustion characteristics of 4-directional (4D) carbon/carbon composites using liquid fuel rocket engine to investigate ablative motion of the materials. C/C composites were made of coal tar pitch as a matrix precursor, and heat-treated at $2300^{\circ}C$. Throughout repeated densification process, the density of the material reached $1.903g/cm^3$. After machining 4D C/C composites, the nozzle surface was coated by a SiC layer by pack-cementation method to improve oxidation resistance. Erosion characteristics of SiC-coated C/C composites were measured as function of the ratio of oxygen to fuel. The morphological change of the composites after combustion test was investigated using SEM and erosion mechanism also was discussed.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.37 no.8
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• pp.759-765
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• 2013

This study proposes the characteristics of the pressure drop in a compressible fluid through porous media for application to a porous injector in a liquid rocket engine in order to improve the uniformity of the drop size distribution and the mixing performance of shear coaxial injectors. The fluid through the porous media is a Non-Darcy flow that shows a Nonlinear relation between the pressure drop and the velocity at high speed and high mass flow rate. The pressure drop of the Non-Darcy flow can be derived using the Forchheimer equation that includes the losses of viscous and inertia resistance. The permeability and Ergun coefficient represented as a function of the pressure drop and pore size can be applied to the porous injector, where the fluid through the porous media is compressible. A generalized correlation between the pressure drop in relation to the pore size was derived.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.30 no.8
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• pp.120-125
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• 2002

In the preceding tests using the KSR-III Sub.(I) engines, it was observed that the heat resistant capability of the engines was not enough for the mission. So Sub.(I) Mod. engines were designed and tested. The Sub.(I) Mod. engines have three major design parameters - the arrangement of main injectors, the impinging angle of main injectors and thermal barrier coating. More than twenty experiments were carried on to evaluate engine performance and heat resistance capability with respect to design parameters. In this study, the test results are introduced. Analysing the result of Sub.(I) engine tests, it is found that decreasing the impinging angle, adopting the H-type arrangement(rather than radial type arrangement) and adopting the thermal barrier coating can increase heat resistance capacity substantially. Also, engine performance evaluation is conducted using specific impulse and characteristic velocity parameter. The results show that the performance variation is small(about 5%) and the performance is better in the case of radial arrangement. It is suspected that these phenomena are caused by the change of flame structure atomization mixing characteristic of sprays and the distortion of recirculation zone. Also from the low frequency instability point of view, it is observed that reducing the impinging angle and adopting the H type arrangement can increase the instability characteristics.

• Journal of the Korea Academia-Industrial cooperation Society
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• v.20 no.10
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• pp.100-106
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• 2019

Ground and high altitude simulated combustion experiments were conducted using a liquid rocket engine with hydrogen peroxide and kerosene as the propellant. A ground and high altitude simulated combustion test facility was constructed by installing a high altitude model diffuser and TMS (Thrust Measuring System) on a vertical combustion test bench. The thrust characteristics according to altitude were investigated using the combustion test equipment. The diffuser was designed on a 1:4.8 scale to verify the characteristics of the high diffusing diffuser and starting pressure. The cold flow tests were conducted using nitrogen gas, and the performance characteristics and starting characteristics of the scale down diffuser were verified. A diffuser and TMS were installed on the vertical combustion test bench, and the thrust correction equations for the system resistance were derived. The thrust correction equations were derived from the step test and vacuum step test before the actual hot firing test. Nozzles with an operating altitude of 10km were designed. Hot firing tests were conducted to analyze the thrust characteristics according to the operating altitude changes. The actual thrust was calculated using each correction equation with the thrust value measured by the TMS.

• Journal of Aerospace System Engineering
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• v.16 no.5
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• pp.17-25
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• 2022

The pintle injector has many advantages in the key characteristics of a liquid rocket engine, such as combustion stability, combustion efficiency, and wide range of comprehensive thrust control, design and manufacture, and test fired under supercritical conditions. The pintle injector is manufactured with a rectangular, single-row orifice for thrust control and production considerations. In order to verify the combustion performance of the pintle injector and its potential as a commercial injector, the combustion characteristics were analyzed by varying the TMR (Total Momentum Ratio) and BF (Blockage Factor). The result of the hot firing test showed that the heat flux increased as TMR increased, and it confirmed that the characteristic velocity efficiency was more affected by BF than TMR. Suppose a single-row pintle injector with efficiency characteristics insensitive to changes in TMR can achieve high efficiency at low fuel differential pressure conditions. In that case, the variable pintle injector's design flexibility can be increase.