• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2011.11a
• /
• pp.540-543
• /
• 2011

This work is devoted to an experimental investigation on conceptual design of dual consecutive stage micro plasma thruster (${\mu}PT$). Optimization study on the thruster configuration has been performed for various electrode gap distances from 1 mm to 2 mm and the hole diameter from 0.3 mm to 2 mm depending on desired operating conditions and corresponding nozzle design requirement. The operation of ${\mu}PT$ at low pressure from $10^{-1}$ Torr to $10^{-4}$ Torr and at various argon flow rates ranging from 5 sccm to 300 sccm has been studied to understand the physic of plasma and the gas dynamics in details. The specific impulse can reach up to 3000-4000 seconds at low power consumptions from 1 to 5 W. Image of exhaust plume from ${\mu}PT$ will be provided and electrical characteristics is also mentioned in this paper.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.3 no.1
• /
• pp.9-14
• /
• 1999

The effect of operating condition was studied experimently on the characteristics of twin sprays ejected from two airblast atomizers, within the range of the mass air-fuel ratio 1.36∼3.54. Water and nitrogen gas were used as test fluids for the experiments. Spray characteristics of liquid spray were measured with measurement of mass distribution and instantaneous image of the spray cone. Experimental results show that the maximum specify of the distribution were lowered but distributed over the larger area when the ROA ratio increased, Center of mass position did not change with increasing water mass flow, Increase of the nozzle distance has an small effect on mass distribution of interaction area but distributed over the larger area. It was also conformed that the effect of interaction near central point of collision decreased with the increase of the ROA ratio on interaction area from comparison using superposition method

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2006.11a
• /
• pp.399-402
• /
• 2006

The purpose of this study is to propose a method for protecting injector face for prolonged combustion time and heat flux measurement technique at the injector face. To obtain basic design data and verify the performance of the proposed method, a regenerative cooling injector face was designed and manufactured for the hot firing test. Due to the safety reason, hot fire test were performed 3, 10, 30, 60 and 120 seconds time step. The discrepancy between analytical results adapting to combustion and nozzle and experimental results is believed due to the over estimation of the convection heat transfer calculation. for the injector face, flow velocity is almost negligible, therefore radiation is more important than convection. Consecutive hot firing test during 10, 30, 60 and 120 seconds combustion time shows good repeatability.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2008.05a
• /
• pp.133-137
• /
• 2008

Results of combustion tests performed for a regenerative cooling combustor of a 30 tonf-class liquid rocket engine were described. The combustion chamber has chamber pressure of 60 bar, propellant mass flow rate of 89 kg/s, and nozzle expansion of 12. The combustion chamber is composed of mixing head, baffle injector, and regenerative cooling chamber. The hot firing tests were performed at design and off-design points. The test results show that the combustion characteristic velocity is in the range of 1738${\sim}$1751 m/sec and the specific impulse of the combustion chamber is in the range of 253${\sim}$270 sec. The peak of combustion characteristic velocity and specific impulse for this combustor is shown at mixture ratio of 2.35 and 2.5, respectively.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2009.05a
• /
• pp.263-266
• /
• 2009

The mainly objectives of this study was a combustion dynamics and instability characteristics in a model gas turbine dump combustor which is the scale down of GE 7FA+e DLN 2.6 gas turbine combustor. Model gas turbine injector has 2-stage swirl vane and it’s reduced 1/3 size of the original one. The shape of plenum and combustor were designed for similar acoustic characteristics. Inlet air was preheated to $200{\sim}400^{\circ}C$. The flow velocity at mixing nozzle was 30 to 75 m/s and equivalent ratio was 0.4 to 1.2. The combustor length was varied for different acoustic characteristics to $375{\sim}700\;mm$. As the result, this research have been show the combustion instability was observed at lower equivalence ratios ($\Phi$ < $0.5{\sim}0.6$) and higher equivalent ratios ($\Phi$ > $1.1{\sim}1.2$).

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2009.11a
• /
• pp.573-577
• /
• 2009

Dynamic behavior simulation of supersonic engine was performed and PI control algorithm was studied for the buzz control in the inlet and the thrust control. Firstly, required thrust was tracked according to the fuel flow control and then inlet pressure was regulated through the nozzle throat area control so that the buzz margin has the positive all the time. The control was performed according to the change of flight Mach number, altitude and angle of attack. The proportional gain and the integral gain for regulating the buzz margin was induced and simulated. In the results, it was confirmed and satisfied that control target in the operating area was changed the angle of attack from $0^{\circ}$ to $10^{\circ}$ at the flight Mach number of 2.1~3.0.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2009.05a
• /
• pp.88-92
• /
• 2009

Total pressure recovery ratio in intake is crucial factor to the operational characteristics of supersonic propulsion system because it does not compress inlet air mechanically by compressor, but does compress inlet air by ram compression. As the result of that the dynamic characteristic analysis of engine was performed before the controller was designed, it could be ascertained when the AoA of flight vehicle increases, the buzz margin decreases so that the shock wave produced outside intake in the specified area according to flight operation's characteristics. Therefore the PID control algorithm was designed to be controlled buzz margin that the characteristic of shock wave could meet the requirement of performance in intake. The PID controller was designed that the buzz margin value is being positive number using the control variables; fuel flow and nozzle throat area.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2009.05a
• /
• pp.71-75
• /
• 2009

Test results of combustion chamber to verify the operation and the combustion performance at low pressure, design and off-design conditions for 30ton-class liquid rocket engine were described. The combustion chamber has nominal chamber pressure of 60 bar, propellant mass flow rate of 89 kg/s, and nozzle expansion of 12. Effects of chamber pressure on combustion characteristic velocity are largely affected by mixture ratio. The specific impulse of combustion chamber is proportional to the chamber pressure regardless of the mixture ratios. The present results can be used as the base to predict the combustion performance of large sized chamber at high pressure while demonstrating the possibility of low pressure firing test of large sized chamber.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2009.11a
• /
• pp.69-73
• /
• 2009

Combustion test results of 1/2.5-scale thrust chamber for 75tonf-class liquid rocket engine were described. The thrust chamber has chamber pressure of 60 bar, propellant mass flow rate of 89 kg/s, and nozzle expansion ratio of 12. The combustion tests were conducted to verify the combustion performance, the regenerative cooling performance and the durability of thrust chamber at design point condition, and then were performed to confirm the operation and the combustion performance at low combustion pressure condition. All the tests had been successfully executed without the damage of the hardware. These test results present a possibility of hot firing test at low combustion pressure condition, and can be used as fundamental data to predict the combustion performance at design point condition for 75 tonf thrust chamber.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.21 no.5
• /
• pp.88-96
• /
• 2017

As a part of the development procedures of scramjet engine with a regenerative cooling system, this experiment was performed using air-assist type injectors for scramjet engine. Two types of injectors were used in this experiment with the 90 and 60 degrees of the injection angle to the main flow. Mie-scattering was used for spray visualization and PDPA was used for the measurement of the atomization characteristics. It was found that increasing the pressure of supplied gas and the distance from nozzle tip led to the enhancement atomization characteristics and the injector with 60 degrees injection angle has better atomization characteristics than 90 degrees injector.