• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.44 no.8
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• pp.668-674
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• 2016

In this paper the development directions of the next generation launch vehicle program following KSLV-II has been discussed, which are to be executed after year 2020 according to the Medium and Long Term Plan for National Space Development. Also, several areas of technology advancement have been identified for the successful development of the LVs. The next generation LV must aim for not only the high performance but also for low cost as well as high reliability in order to compete against global commercial launch service providers. To this end, the next generation LVs program shall capitalize on many anticipated accomplishments of the KSLV-II program such as the 75 ton class LOX/kerosene rocket engine.

• Journal of the Korean Society of Propulsion Engineers
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• v.15 no.6
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• pp.26-37
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• 2011

A 75-tonf-class LRE(liquid rocket engine) thrust chamber with ablative material for technology demonstration was manufactured on the basis of development technologies of 30-tonf-class LRE. Hydraulic characteristics of the thrust chamber were examined through cold flow test and ignition test of low flow condition. Test result showed that hydraulic function was good. Side ignition method with igniter ring also showed a fine function of ignition in operating ways of static condition. But a close review is required to understand the phenomena of generation and extinction of specific frequencies showed in dynamic characteristics ways. To achieve these, a large combustion test facility which is capable of performing combustion test at design condition of the 75-tonf-class thrust chamber should be constructed as soon as possible.

• Journal of the Korean Society of Propulsion Engineers
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• v.21 no.4
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• pp.1-11
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• 2017

TPTF (Turbopump Real Propellant Test Facility) at Naro Space Center has used alcohol burner system to simulate the gas flow of gas generator of liquid rocket engine. During the test at TPTF, the temperature and pressure at turbine inlet were smoothly increased while those of the gas generator of engine were constant. Present research developed a simulation code for the burner and the piping system and applied to the system. The calculation results were in good agreement with the test, and confirmed quantitatively that the non-steadiness is due to the heat transfer of the pipe. While the insulation of the pipe is ineffective, the length has a large impact on the turbine inlet condition. The present research clarified the empirically estimation of test condition, and can be applied to determination of the following test conditions.

• 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.

• Journal of the Korean Society of Propulsion Engineers
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• v.24 no.1
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• pp.64-77
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• 2020

An electric-pump cycle, which is a propellant supply system for driving pumps of a liquid rocket engine using an electric motor, has the advantages of simple system configuration and easy control of supply flow rate and pressure. This paper investigates and analyzes the overseas research trends of the electric-pump cycle. In addition, the research and development country, performing organization, application, engine thrust, pump pressure increase, motor power, and rotation speed are summarized. Among them, the design variables of the overseas research that applied the upper-stage propulsion system with the thrust range of 0.445~2.2 kN could be used in the study of a similar electric-pump cycle in Korea.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.31 no.9
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• pp.103-109
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• 2003

The development of a pulse gun used for the combustion stability rating test(SRT) has been domestically tried out since the SRT should be accompanied in the process of the development of LREs. At the beginning of the development, a rupture problem of the pulse gun body had been solved by changing its material and increasing its thickness. Also, the sealing test was undertaken for a pulse gun equipped with a membrane before conducting a explosion test and by using a specially designed jig after a explosion test. Since it is appropriate to reuse the holder of a pulse gun cavity for an actual SRT, a series of experiments for several variants has been taken to make it reusable. After all, the solution for a reusable holder of a pulse gun has been found and applied for complementing its specification.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.44 no.8
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• pp.695-701
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• 2016

Spacecraft propulsion system is a kind of rocket engine that has been developed from the end of 1950s for attitude control and orbit maintenance of satellite. Since the spacecraft propulsion system has to be used for a relatively long time, therefore, stability of propellant and life of thruster could be very important factor for propulsion system design. Recently, green propellant propulsion and all electrical propulsion system have became very important issue, and we also need a development according to well organized plan. In this paper, we will introduce the development status, key technologies and development prospect of spacecraft propulsion system.

• Journal of the Korean Society of Propulsion Engineers
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• v.14 no.5
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• pp.8-14
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• 2010

For the development of a liquid rocket engine, hot-firing tests of a regeneratively cooled thrust chamber were performed at chamber pressures of approximately 30 and 60 bars. In the paper, pressure fluctuation data, which were obtained from the dynamic pressure transducers installed in propellant manifolds and combustion chamber, were analyzed. Compared to the data at chamber pressure of 60 bar, the results at chamber pressure of 30 bar showed low-frequency oscillations around 150 Hz in the combustion chamber. The low-frequency waves in the combustion chamber were coupled with those in the manifolds. However, the RMS values of the chamber pressure fluctuations at chamber pressure of 30 bar were only 0.8% of the chamber pressures. Thus, it can be inferred that the thrust chamber operates in the stability boundary even at low chamber pressure.

• 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.

• 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.