• Journal of the Korean Society of Propulsion Engineers
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• v.22 no.5
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• pp.66-72
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• 2018

A numerical study was conducted to investigate the performance characteristics of a STED with various pressure ratios (PRs) and cone shapes. Due to momentum loss, the pressure in vacuum chamber increased with cone angle for a PR of 75. Also, the STED is started between PRs of 36 and 37 in the case of a cone angle of $15^{\circ}$ and a blockage ratio (BR) of 15%. The results for various PRs and cone shapes are presented, and the optimal cone shape is found to have a cone angle of between $5{\sim}20^{\circ}$ and a BR of between 15~40%.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.34-37
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• 2009

The present study presents a numerical methodology for early conceptual trade-off study between propulsive performance, cooling efficiency, weight and size, in which combustion and cooling precesses in regeneratively cooled rocket thrust chamber are interactively simulated. To address the capabilities and reliability of the design tool, some application results are given involving contour design, performance analysis, and wall cooling prediction as well as a systematic design evaluation.

• Transactions of the Korean Society of Mechanical Engineers
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• v.12 no.2
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• pp.328-337
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• 1988

Experimental study was carried out to investigate the combustion characteristics of the hero-valved pulsating combustor with maximum operating capacity of 56kW. The pressure, the ion current, and the temperature fluctuations were simultaneously measured and statistically analyzed to identify the combustion process, the reignition and the mixing process of the reactants. It was found that the pulse combustion process was intermittent and the reignition of the reactants was caused by a direct contact and rapid mixing with the previous hot residuals. The analysis of the measured data indicated that the combustion process consisted of there stages in the combustion chamber; the preheating of the reactants in the vicinity of the air inlet pipe, the explosive combustion in the central region and the afterburning in the vicinity of the tailpipe. Wile the inflow of the fresh air occurred during the negative period of the pressure in the mechanical valved system, it occurred during the rising period of the pressure in the aero-valved system.

• 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.

• Journal of the Korean Society of Propulsion Engineers
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• v.24 no.1
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• pp.9-16
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• 2020

A performance evaluation of the 200 N-class GCH₄/LOx small rocket engine was performed through ground hot-firing test. As a result, the combustion pressure and thrust raised with the increase of the oxidizer supply pressure, and thus the specific impulse, characteristic velocity, and their efficiency increased. The characteristic velocity was measured at about 90% performance efficiency. The change of chamber aspect ratio did not affect the performance of the rocket engine in the test condition specified. In addition, uncertainty evaluation was conducted to ensure the reliability of the test results.

• Journal of the Korean Society of Propulsion Engineers
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• v.24 no.1
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• pp.1-8
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• 2020

In this study, a configuration design of a CH₄/LOx small rocket engine was made and test system was established for the performance evaluation. A coaxial swirl injector was chosen because of its remarkable atomization performance and low combustion instability. Three aspect ratios for the combustion chamber configuration, i.e., 1.5, 1.8, and 2.1 were also set for the comparison of the combustion efficiency. The reliability of the thrust measurement rig was enhanced by pre-and post-calibration process. From the preliminary ground hot-firing test, the measured thrust and specific impulse values were 89.2 N and 181.8 s, respectively, which were 21.6% lower than the ideal values. In addition, the efficiency of characteristic velocity was measured as 84.2%.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.35 no.10
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• pp.1083-1088
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• 2011

The effects of an impeller and floating ring seals on the performance of centrifugal pumps are investigated on the basis of their test results using water. The pumps are single-staged centrifugal pumps developed for 30-ton- and 75-ton-class liquid rocket engines, and are components of a turbopump that supplies propellants (liquid oxidizer and kerosene) to the combustion chamber. The exit width of the impellers and the numbers and exit angles of the impeller blades are found to have influences on the pump heads. In addition, the pumps have different efficiencies according to the gaps between the floating ring seals and the impellers, whereas the pump size seems to have less effect on the efficiency.

• Journal of the Korean Society of Propulsion Engineers
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• v.18 no.2
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• pp.73-79
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• 2014

The sandwich shell with Egg-box core has been used for the combustion chamber case of air breathing propulsion system. The alteration on pitch length and thickness of Egg-box core was required to be lighter and save manufacturing time and cost of combustion chamber case. In this paper, the finite element analysis method which simulated bending test was used to predict the equivalent shear modulus which affect structural stability of sandwich shell in short time. The result of FE calculation on sandwich panel with homogeneous material, H130-foam core, showed a good agreement with the values available in the reference. The equivalent shear modulus of Egg-box core according to the variation of pitch length and thickness can be obtained.

• Journal of the Korean Society of Propulsion Engineers
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• v.18 no.1
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• pp.8-15
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• 2014

The ablation characteristics of graphite nozzle throat insert is analyzed for the use in solid rocket propulsion system. The propulsion system is composed of three types of conventional nozzles, such as De-Laval type, blast tube type, and submerged type. Various kinds of propellants are used in the thirteen kinds of propulsion system that has different shapes of each other. Total thirty seven tests are performed. From the results of the analysis, it is found that the ablation rate is higher for the higher average chamber pressure and the higher concentration of oxidizing species in combustion gas.

• Transactions of the Korean Society of Mechanical Engineers
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• v.11 no.5
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• pp.846-855
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• 1987

The experimental study was carried out to investigate the performance characteristics of the aero-valved pulsating combustor designed to increase the practical applications of the system. The geometric effect on the stable condition and the dynamic behavior of the system is identified. The equivalence ratio, the inflammability limit, the operating frequency, and thrust were also measured when the system oscillated stably. It is found that while the operating condition is sensitive to the diameter of the inlet pipe and the length of the tailpipe, the maximum value of the turn down ratio was obtained up to 3.2. The measured air flow rate shows that the equivalence ratio increases monotonously with the increasing fuel flow rate and decreasing air inlet diameter and tailpipe length. The measured operating frequency can be approximated by the simple linear equation and the discrepancy is within five percent. The system produced the maximum total thrust of 14N and the minimum specific fuel consumption of 0.155 Nm$^{3}$/h.N when the total thrust was 13N.