• Journal of the Korean Society of Propulsion Engineers
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• v.4 no.2
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• pp.6-11
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• 2000

In the pressurized propellant feed system of liquid rocket, feed pressure is decided chamber pressure of normal combustion state. However, during ignition period the initial chamber pressure is atmosphere. So, it may have overflow, hard-start and even critical damage of engine. This paper proposes an improved propellant feed system for the stable combustion of liquid rocket. Hot test were already performed to verify the presented propellent feed system. The proposed propellant feed system uses two steps - pre and main combustion - to prevent large pressure increase and uses cavitating venturis for stable flow rate in whole combustion. This system feeds the flow rate lesser than the designed flow rate, so combustion pressure reached pre-combustion pressure. Cavitating venturis offer unique flow control capabilities at normal and abnormal combustion state, because flow rate is solely dependent on upstream absolute pressure and fluid properties, but independent on downstream condition.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.34 no.4
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• pp.443-448
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• 2010

In this paper, the changes in performance parameters, e.g., the combustor pressure, turbine power, engine mixture ratio, temperature of gas generator, and product gas, of a liquid rocket engine employing gas generator cycle with the variations in propellant-supply pressure have been described. Engine performance is numerically calculated using the 13 major system-level variables of the rocket engine. The combustor pressure and turbine power increase with an increase in the oxidizer-supply pressure and decrease with an increase in fuel-supply pressure. The lower mixture ratio of gas generator for increased fuel mass flow rate decreases the gas generator gas temperature and deteriorates the gas material properties as the turbine working fluid. The turbine power decreases with an increase in fuel-supply pressure; this results in a decrease in the main-combustor pressure, which is directly proportional to engine thrust.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 1999.10a
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• pp.4-4
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• 1999

압축가스를 이용하여 추진제를 액체 로켓 엔진에 공급하는 경우, 공급압력은 정상 연소상태의 연소압을 기준으로 하여 설계한다. 그러나 연소초기의 연소실 압력은 대기압 상태이므로 과도한 유량이 공급되어 이로 인해 hard-start 가 발생하며, 최악의 경우 엔진의 파손을 가져온다.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2005.11a
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• pp.247-250
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• 2005

In the case of appling a unielement injector developed for a full scale liquid rocket combustor, a operating condition or configuration of the injector is changed by combustion pressure, arrangement and injector quantity of a full scale liquid rocket combustor. In order to verify application, swirl coaxial injectors propelled by jet-A1 and liquid oxygen are tested at different conditions of a combustion pressure, a flowrate and an injector length. As a test result, the application of the present swirl coaxial injectors is excellent because an efficiency of a characteristic velocity is increased at the each test condition beyond that variation of dynamic pressure intensity is small.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.94-98
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• 2017

Research for developing 9 tonf-class staged combustion cycle engine which is high performance upper stage engine has been conducted. Technical demonstration model of the staged combustion cycle engine was installed in combustion test facility of Naro space center. Combustion tests of Power-pack which consists of pre-burner and turbopump without main combustion chamber and combustion tests of engine have been conducted. The vacuuming process of hydraulic lines is introduced and the characteristics of turbopump and hydraulic lines of engine combustion test are described in this paper.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• v.y2005m4
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• pp.265-268
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• 2005

A series of ignition tests had been conducted for a thrust chamber propelled by Jet A-1 and liquid oxygen with a chamber pressure of 52.5 bara and a thrust of 30 tonf. The chamber ignited by a hypergolic fluid, TEAL, keeps its first constant pressure low at $63\%$ of the design value by $61\%$ of a liquid oxygen mass flow rate and $67\%$ of fuel for 0.25 sec. The operating O/F ratio of the chamber was kept at lower values than that of the design operating condition throughout the whole ignition procedure. Surge of the chamber pressure is below $6\%$ of the design value.

• Journal of the Korean Society of Propulsion Engineers
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• v.10 no.3
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• pp.60-66
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• 2006

Notable are the burning rate characteristics of solid propellant burning at extremely high pressure(10000-20000 psia). The burning rate test using closed bomb shows the discontinuous increment around 4000 psia so that the exponent of burning rate(n) is almost double, from 0.4 to 0.8. The pressure-increasing rate of the test motor is about 300 times as high as that of the motor operating at the conventional pressure, less than 2000 psia, is, therefor the burning rate is augmented about 5-50 times. The performance prediction reflecting the pressure-change-rate effect are fairly comparable with the test data at various test conditions.

• Aerospace Engineering and Technology
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• v.12 no.1
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• pp.200-206
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• 2013

A performance sensitivity of liquid rocket engine to propellant density or supply pressure change was studied. The analysis program was verified to have 1% error comparing with the measured data of a turbopump-gas generator system. The engine combustion pressure decreases as fuel supply pressure increases due to decreased mixture ratio which reduces the turbine power. The engine combustion pressure increases as fuel density increases because the total propellant flow rate is increased substantially even though mixture ratio is slightly decreased. The engine combustion pressure increases when the oxidizer density or supply pressure increases.

• Journal of the Korean Society of Propulsion Engineers
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• v.17 no.1
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• pp.89-96
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• 2013

Combustion pressures were measured to study combustion stability for an oxygen rich preburner by both of static and dynamic pressure sensors. The resolutions of each static and dynamic pressure sensor are the 1,000 Hz and 25,600 Hz, respectively. The nominal combustion pressure of the preburner was 200 bar but 80 bar was used at the several initial tests for the safety reason. Two stage ignition was applied to reduce the ignition impact for every tests including the tests with 200 bar combustion pressure. The tests lasted for 10 sec. max. and a little fluctuations of pressure was observed during the main mode. The measured pressures were studied by FFT analysis and no noticeable frequency coupling was found. Thus the preburner can be regarded as stable and it can be utilized for further study on staged combustion cycle liquid rocket engine.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.919-921
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• 2011

Variable nozzles are used to enhance the effectiveness of aircraft engines at various altitudes. Unsynchronized movements of variable nozzle flaps affect the direction of thrust in case the variable nozzle consists of many flaps. A synchronization test system was developed to verify the synchronization characteristics of variable nozzle mechanism including flaps. The test system has a capability to simulate combustion pressure in variable nozzle space. The test system was used to qualify the synchronization characteristics of a variable nozzle flaps affected by magnitude and uniformity of simulated combustion pressure, and time delay of each nozzle actuators.