• Journal of Advanced Marine Engineering and Technology
• /
• v.11 no.2
• /
• pp.51-60
• /
• 1987

Recently, the problem of the axial vibration for the marine diesel engine shafting has become important due to the increased exciting forces resulting from high supercharging and large output, and the reduced natural frequencies resulting from long stroke and show speed. The effects of the axial vibration on the propulsion shafting induce cracks of the connecting point of crankpin and crankarm, the severe wear of thrust bearing, the fatigue failure of each fixing bolt and jointed parts, the hull and local hull vibrations, and also the wear and the noise due to intense hammering phenomena of thrust collar. Therefore, each classification society requires the calculation of natural frequencies and their amplitudes and also measurements of the forced damped axial vibration. At present, the technical and theoretical level is at the stage of estimating the resonant points and their maximum displacements, but the estimated displacements of the resonant points are not so reliable as the torsional one. In this study, induced stresses and amplitudes of the forced damped axial vibration are calculated. For this purpose, the equation of forced axial vibration with damping for the propulsion shafting is derived and its steady-state response is calculated by the mechanical impedance method. A computer program for above calculations is developed. The measured values are analyzed and the calculated results are compared with the measured ones. They show fairly good agreements and the reliability of developed program is confirmed.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.31 no.10
• /
• pp.66-72
• /
• 2003

The cooling mechanism for a liquid rocket engine of 10tf-thrust using kerosene as a fuel was studied from the viewpoint of both the regenerative and curtain cooling. Based on the concept of a highly-stratified gas flow in the combustion chamber, the cross section of the combustion chamber was spilt into 2 independent parts, core and exterior part. Additional fuel is injected into the exterior section and gas temperature can be reduced in the exterior section. Consequently, the heat flux into the coolant and wall temperature are reduced and the thermal stability of a liquid rocket en g i.ne could be improved.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.18 no.6
• /
• pp.88-95
• /
• 2014

Spinning process to inner wall has been applied for reducing the weight of regenerative cooling chamber of liquid propellent rocket engine. The fractures of the blanks of cylinder part and nozzle throat part have been observed during spinning processes. In order to overcome the problem, the mandrel and the blank shape have been modified, and the inner wall was successfully manufactured through the modifications. The manufactured spinning prototype of nozzle throat part was successfully bulged without cracking and necking, and it was confirmed to secure sufficient formability necessary for fabricating thrust chamber.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.20 no.4
• /
• pp.87-93
• /
• 2016

In a Direct Connect(DC) mode altitude engine test, a labyrinth seal is set up between an air intake duct and an engine. The labyrinth seal plays a key role in mechanically isolating them, which contributes to the accurate measurement of thrust and the other component forces. However, when high vertical temperature gradient is generated in the supplied air in the duct, the isolation breaks down. In this paper, a labyrinth seal control device is designed and installed in an effort to eliminate the issue. Test result shows the device successfully gets rid of the contact problem even when high vertical temperature gradient is produced.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.25 no.3
• /
• pp.1-13
• /
• 2021

Effects of the nozzle wall angles on the supersonic flow field in a thrust optimized nozzle were numerically investigated. The combustor and operating condition of 30-tonf rocket engine was selected to study the optimum nozzle shape. The nozzle flow of combustion products was realized by the shifting equilibrium calculation for the propellant of kerosene-LOx. The change of nozzle wall angles induced different developing patterns of the internal and secondary shock wave. The optimum nozzle was obtained when the internal shock was in a specific position at the nozzle outlet. The nozzle wall angles of the optimum nozzle were very similar to those of the optimum nozzle which does not consider the shock wave.

• Aerospace Engineering and Technology
• /
• v.13 no.1
• /
• pp.70-75
• /
• 2014

Programs of energy balance, mode analysis and transient analysis for a liquid rocket engine have been introduced. The analysis methods have been verified through comparison between the present results, and the results of the other program and experimental data. An energy balance analysis is used for engine system design at the early development phase. A mode analysis is used for decision of engine operation conditions and test conditions, and studying deviation of an engine performance. A transient analysis can predict a propellant flow rate, thrust, impulse at transient phase. It is essential to establish a startup/shut down sequence. The analysis programs will be used to develop the engines of KSLV-II.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2010.11a
• /
• pp.548-551
• /
• 2010

A 150 lbf-thrust class micro turbojet engine has been developed. The engine could be applied to power plant for small aviation vehicle such as UAV, decoy and anti-radar missile and was designed with concepts that has small size, low-cost and high performance. A prototype was manufactured and performed the ground static test and high altitude test. This paper outlines the features and layout of 150 lbf turbojet engine and also describes the design characteristics and test results of the engine and components.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2010.11a
• /
• pp.495-498
• /
• 2010

This test cell is developed to the institutes or laboratories research and study gas turbine engine for academic purpose with this test data to provide the fundamentals of operational mechanism and structural configuration, and further to verify thermodynamic calculation The test cell is installed to monitor and compare real-time data with reference engine model performance simulation data. using by NI DAQ(Data acquisition)device and LabVIEW program based on 80 lbf-micro turbojet engine.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.47 no.7
• /
• pp.517-524
• /
• 2019

A system design has been performed for a vacuum thrust 88 ton staged combustion cycle rocket engine. Previous research has been used to estimate the performance of the engine components. And the algorithm has been proposed to evaluate the converged engine system performance. The present methodolgy has been verified by comparing the published data for RD-180. The present work adopts the most of the previous KSLV-II engine heritage for both performance improvement and cost competitiveness. The combustion pressure has been decided as 12MPa considering manufacturing difficulty, cost and performance improvement, and as a result the vacuum specific impulse has increased by 23.4s.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.18 no.1
• /
• pp.50-57
• /
• 2014

A ground hot-firing test (HFT) was conducted to take out the optimal design configurations for the thrust chamber of 70 N-class liquid rocket engine under development. Monopropellant grade (purity: ${\geq}98.5%$) hydrazine was adopted as a propellant for the HFT, and three kinds of thrust chambers having characteristic lengths ($L^*$) of 2.79, 2.95, and 3.13 m were selected for their performance evaluation. It is revealed through the test and evaluation that the increase of the $L^*$ leads to a performance degradation in the test condition specified, and pulse response performance of the development model shows superior characteristics to commercialized hydrazine thrusters.