• Journal of the Korean Society of Propulsion Engineers
• /
• v.14 no.4
• /
• pp.46-58
• /
• 2010

Hydrogen peroxide(HP) and liquid methane have deserved renewed considerations as green propellants in recent years, because main design concerns in the development of the new generation propulsion system for spacecrafts are concentrated on low operation cost and environmental cleanness. Although HP has a long history of application to aerospace propulsion systems due to high density, mono-propellant characteristics and low toxicity, it had been replaced by hydrazine and liquid oxygen due to extreme performance requirement during the cold war. But HP has received a renewed interest due to its increased stability and many researches have been conducted to develop high performance LREs(Liquid Rocket Engines) using HP. Liquid methane has also received a new interest in rocket propulsion system for the future space exploration according to its possibility of ISRU(In-Situ Resource Utilization).

• Journal of the Korean Society of Propulsion Engineers
• /
• v.17 no.2
• /
• pp.122-127
• /
• 2013

The paper includes the analytic results of pressure fluctuation data from the combustion of an oxidizer-rich preburner applicable to high-performance, closed-cycle liquid rocket engine systems. The combustion experiments went through two different steps of a chamber pressure during single run. Self-excited pressure fluctuations with a frequency of 78 Hz were observed only at the relatively low chamber pressure condition. These pressure fluctuations are regarded as a bulk mode. The intensity of pressure fluctuations by a root-mean-square value is 13.3% normalized by the chamber static pressure and no pressure excitation was observed at the design pressure condition. The bulk mode has an identical phase across the inside of the chamber and reveals the similar characteristics to the Helmholtz resonator.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.17 no.4
• /
• pp.25-31
• /
• 2013

A blisk with rotor shroud is usually adopted in LRE turbine to maximize its performance. However it experiences the severe thermal load and resulting damage during engine stating and stop. Shroud splitting is devised to relieve the thermal stress on the turbine rotor. Structural analysis confirmed the reduction of plastic strain at the blade hub and tip. However, split gap at the rotor shroud entails additional tip leakage and results performance degradation. In order to assess the effect of shroud split on the turbine performance, tests have been performed for various settings of shroud split. For the maximum number of shroud splitting, measured efficiency reduction ratio was 2.65% to the value of original shape rotor.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.24 no.6
• /
• pp.78-84
• /
• 2020

The performance of pintle thruster is analyzed by using the pintle thruster performance analysis model which integrating the element models introduced in Part I. To verify the performance analysis, the results of the developed program are compared with the experimental data of kerosene/hydrogen peroxide liquid pintle thrusters. Based on the results, the characteristics of the pintle thruster are analyzed. The sensitivity analysis is performed to investigate the effect of thruster shape and operation parameters on performance characteristics using both OAT and scatter plot methods. The four performance parameters such as droplet diameter, film flow rate, O/F ratio, and nozzle throat diameter are evaluated to investigate their effects on characteristic speed, combustor pressure, and specific thrust.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.24 no.4
• /
• pp.55-65
• /
• 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
• /
• v.24 no.6
• /
• pp.69-77
• /
• 2020

A quasi one-dimensional multi-phase reaction flow analysis code is developed for the performance analysis of liquid pintle thrusters. Unsteady flow field, droplet evaporation, finite reaction and film cooling models are composed as the major models of the performance analysis. The droplet vaporization takes account of Abramzon's vaporization model, and the combustion employs a flamelet model based on detail chemical reactions. Shine's model is applied for the film cooling calculation. To verify each model, the Sod shock tube, single droplet vaporization, kerosene droplets combustion, and film length are evaluated.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.25 no.6
• /
• pp.45-52
• /
• 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
• /
• v.17 no.1
• /
• pp.51-58
• /
• 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.

• Applied Chemistry for Engineering
• /
• v.18 no.3
• /
• pp.227-233
• /
• 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
• /
• v.37 no.8
• /
• pp.759-765
• /
• 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.