• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.31 no.9
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• pp.32-37
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• 2003

The trend of miniaturization has been applied to the research on micro rockets resulting in prototype rockets fabricated by MEMS processes. In this paper, the development of three-dimensional micro rockets using micro milling as well as the results of combustion and flight tests are discussed. The body of rocket was made of 6061 aluminum cylinder. The three-dimensional micro nozzles were fabricated on brass by micro endmill with 127${\mu}m$ diameter. Two different micro nozzles were fabricated, one with 1.0mm of throat diameter and the other with 0.5mm. The total mass of rocket was 7.32g and that of propellant was 0.65g. The thrust-to-weight ratio was between 1.58 and 1.74, and the flight test with 45 degree launch angle form the ground resulted in 46m-53m of horizontal flight distance

• Journal of the Korean Society of Propulsion Engineers
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• v.3 no.4
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• pp.58-66
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• 1999

The flow characteristics of supersonic ejectors is often subject to compressibility, unsteadiness and shock wave systems. The numerical works carried out thus far have been of one-dimensional analyses or some Computational Fluid Dynamics(CFD) which has been applied to only a very simplified configuration. For the design of effective ejector-pump systems the effects of secondary mass flow on the supersonic ejector flow should be fully understood. In the present work the supersonic ejector-pump flows with a secondary mass flow were simulated using CFD. A fully implicit finite volume scheme was applied to axisymmetric compressible Navier-Stokes equations. The standard two-equation turbulence model was employed to predict turbulent stresses. The results obtained showed that the flow characteristics of constant area mixing tube types were nearly independent of the secondary flow rate, but the flow fields of ejector system with the second-throat were strongly dependent on the secondary flow rate due to the effect of the back pressure near the primary nozzle exit.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.629-634
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• 2011

A prestudy on expendable turbine engine for high-speed vehicle was conducted. The two possible mission profiles were established to decide the engine requirements and Design Point, and Design Point analysis was performed with the values of design parameter which were obtained from similar class engines and technical references. The results showed that Specific Net Thrust is 2599.4 ft/s and Specific Fuel Consumption is 1.483 lb/($lb^*h$) at the flight condition of Sea Level, Mach 1.2. It was also found through the performance analysis on the two possible mission profiles that major design parameters for determining Net Thrust were Turbine Inlet Temperature for low supersonic flight speed and Compressor Exit Temperature for high supersonic flight speed. In addition, simple turbojet engine with axial compressor, straight annular combustor, axial turbine and fixed throat area converge-diverge exhaust nozzle was proposed as the configuration of simple low cost light engine.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.163-170
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• 2011

Applicability of regenerative cooling in 2,500 N-class bipropellant thruster using hydrogen peroxide and kerosene was considered for improvement performance and application in various missions. Calculation was performed by one dimensional approach using hydrogen peroxide as a coolant. In designed regenerative cooling thruster, heat flux at nozzle throat was estimated at 18 ~ 20 $MW/m^2$. Designed cooling channel width and height were 2.5 mm and 0.5 mm, respectively. Based on designed cooling channel configuration, flat plate model was manufactured and tested for estimation of pressure drop in cooling channel, and CFD analysis was compared with the test result. The maximum error between CFD analysis and experimental result was approximately 13% and average error was approximately 5%.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.807-812
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• 2011

Fuel-side cold-flow tests were performed on the technology demonstration model of a 75 ton-class liquid rocket engine combustion chamber for the first stage of the Korea space launch vehicle II. Pressure drop in the cooling channels of the combustion chamber was measured by changing fuel mass flow rate through a pressure regulating system. Pressure drop in each segment of the chamber could be obtained and a lot of pressure drop was caused by high flow velocity in the nozzle throat segment. The accuracy of a hydraulic analysis method for calculating a pressure loss in cooling channels could be verified by applying it to the cold-flow test conditions.

• Journal of the Korean Society of Propulsion Engineers
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• v.19 no.1
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• pp.68-75
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• 2015

In this paper, unsteady characteristics of pressure in solid rocket motor were analyzed by using response of pintle actuation, pressure and thrust data from ground test. Pressure and thrust in solid rocket motor can be controlled in real time by varying nozzle throat area with pintle, installed in the valve. Unsteady characteristics of pressure can be observed in this system occurred by various reasons. Two critical reasons, error of pintle actuation and ablation of center tube, are found and effects of each reason can be analyzed individually by re-prediction of pressure with response of pintle actuation and analyzing thrust to pressure ratio.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.503-506
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• 2009

In this study, we presented the results of static firing tests on the PE/LN2O hybrid rocket motor, which has a thrust of 650 kgf level. Through the early tests, we found that the combustion chamber pressure and the thrust were lower than design values because an actual oxidizer flow rate was less than that expected. In order to complement this result, the methods of decrease of nozzle throat and the increase of oxidizer mass flow rate were conducted in the next experiment, and we studied the combustion phenomena with the experimental results. Also we compared and analyzed a difference of combustion characteristics on scale effect. It show that a sub-scale motor regression rate was a little less than that of a lab-scale motor with the same oxidizer mass flux. Results of this study might be used as a basic data for development of hybrid sounding rocket.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.573-577
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• 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
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• 2009.05a
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• pp.88-92
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• 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.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.37 no.2
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• pp.139-148
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• 2013

In this study, an ejector was designed to recirculate the anodic off-gas of SOFC, and a parametric study of the system performance was conducted at various ejector entrainment ratios. Aspen Plus, a chemical engineering program, was used to calculate the operational conditions of the ejector. To minimize the calculation load of the CFD and to ensure the global optimum, a genetic algorithm and Kriging model were used for the optimization. The optimization results showed that the dominant design variables of the sonic ejector are the throat diameter and the first flow nozzle position. The designed ejector has enough flexibility for different operating conditions of a 1-kW SOFC system. When the ejector was applied to the SOFC, it reduced 56% of the steam and 8.4% of the fuel compared to the reference case.