• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2012.05a
• /
• pp.574-577
• /
• 2012

Thrust measuring is one of the crucial factor to decide the performance of a liquid rocket engine when the engine development test, especially for the combustion chamber, is implemented. Calculating the thrust from a combustion pressure is used when direct measuring the thrust is impossible, but direct measuring the thrust is necessary and various methods for doing it more precisely should be considered. This paper introduces the preliminary design concept about the new thrust measurement system for the vertical firing test stand, which is introduced domestically for the first time, of a liquid rocket engine combustion chamber.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.43 no.4
• /
• pp.304-310
• /
• 2015

Pintle thrusters use pintle stroke to change nozzle throat area, and this controls thrust. Using MATLAB, one-dimensional simulation has been investigated and the results are compared to those of cold flow tests and computational fluid dynamics for the pintle thruster of Chungnam National University. The prediction based on one-dimensional flow theory shows good agreement with measurements for chamber pressure, but deviates for thrust, partly because of nozzle wall separation. Computational results show that nozzle wall separation occurs at an early stage of nozzle expansion, near the design nozzle throat, for the course of pintle strokes. Empirical thrust prediction incorporates nozzle wall separation, and thus 1-D simulation using empirical thrust prediction showed good results for an early stage of pintle stroke.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.18 no.6
• /
• pp.68-74
• /
• 2014

During the life firing test of 1 N-class thruster development model, pulse mode performance and performance changes were examined. The deviation of pulse mode response time according to thruster feed pressure was relatively small and the resultant ignition delay, response time, tail-off time were 32-35 ms, 86-91 ms, 89-98 ms, respectively. For the stabilized pulse region the impulse bit revealed the outstanding reproducibility of 1.41, 1.32, 2.10% at $3{\sigma}$. During the life firing test, the impulse bit was decreased with limited amounts, therefore the pulse mode performance could be considered to be maintained. The thrust centroid was also maintained during the life firing test.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.16 no.2
• /
• pp.1-9
• /
• 2012

Applicability of regenerative cooling in 2,500 N-class bipropellant thruster using hydrogen peroxide and kerosene was considered for improvement of performance and application in various missions. Calculation was performed by one dimensional approach using hydrogen peroxide as a coolant. The heat flux of thruster 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
• /
• 2012.05a
• /
• pp.100-103
• /
• 2012

The blending method that is an addition of small quantity of fuel was used to increase the performance of green propellant thruster. 90 wt.% hydrogen peroxide as a green propellant was selected, and ethanol was used as a blended fuel. The o/f ratio was chosen as 50 which has higher theoretical performance than 98 wt.% hydrogen peroxide. The chamber temperature of blended hydrogen peroxide was higher than adiabatic chamber temperature of hydrogen peroxide. Therefore, performance can be improved by ethanol blending. Several catalyst and its support were compared to find appropriate catalyst for decomposition and combustion of ethanol blended hydrogen peroxide. As a experimental results, Pt was suitable, but $MnO_2$ had a chamber instability when it was reused. The ${\alpha}-Al_2O_3$ which is high heat-resistant support showed very unstable performance in both Pt and $MnO_2$ catalyst since it has low decomposition performance.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2007.04a
• /
• pp.135-140
• /
• 2007

An analysis has been conducted of the optimal condition to maximize the specific impulse for a liquid rocket engine with film cooling. The present engine performance has been compared with the published conceptual design to be verified satisfactorily accurate. The optimal combination of film coolant flow rate and the regenerative cooling capacity has been found for maximum specific impulse. The optimal fuel pump pressure increases and the optimal film coolant flow decreases for a larger thrust engine. Higher turbine inlet temperature increases both the fuel pump pressure and the film coolant flow rate as the optimal condition. The coking temperature has the same qualitative effect as the turbine inlet temperature.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.18 no.4
• /
• pp.1-8
• /
• 2014

Overexpanded flow of an axisymmetric thruster nozzle is numerically simulated to investigate effects of nozzle pressure ratio (NPR) on the shock structure and thrust performance. The Reynolds-averaged Navier-Stokes equations with k-${\omega}$ SST turbulence model are solved utilizing FLUENT solver. As the NPR is raised, thrust performance monotonically increases with the shock structure and flow-separation point being pushed toward the nozzle exit. It is also discussed that the flow structure at nozzle-exit plane which is immediately affected by a position of nozzle-interior shocks and expansion waves, has strong influence upon the thrust performance of thruster nozzle.

• Aerospace Engineering and Technology
• /
• v.4 no.2
• /
• pp.203-208
• /
• 2005

Electronic components for space launch vehicles are exposed to a severe vibrational environment at launch and flight. The structural reliability of each component can be verified using mathematical approaches. In order to verify the structural reliability, an important parameter is the natural frequency of PCB(Printed Circuit Board) assembly mounted electronic components on and housing mounted PCB assembly in. In this paper, in order to find natural frequencies of PCB assemblies and the housing of hydrazine TCU(Thruster Control Unit), FEM(Finite Element Method) is adapted. The analytical result of FEM is verified by experimental method.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.32 no.3
• /
• pp.88-94
• /
• 2004

The KSR-III rocket is a liquid propellant sounding rocket and thrust vector control actuators and cold gas thrusters are used to control pitch and yaw, roll attitude respectively during thrusting phase. In this paper, the structure of designed attitude controller and gain scheduling, results of stability analysis for KSR-III rocket are presented. The attitude controller is implemented with flight software in the domestically developed INS and successfully performed its function in the flight test. The flight data are coincident with simulation results.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.38 no.8
• /
• pp.834-841
• /
• 2010

The STSAT-3 (science and technology satellite) is the first satellite whose entire structure was made of composite materials in Korea and it will be launched later in 2011. As like other small satellites, it is also equipped with several advanced instruments whose major objectives focused on the scientific tests in space. The HPS (hall thruster propulsion system) using xenon gas as a propellant has been developed and its overall ground tests were conducted. This research emphasizes on the technologies and procedure applied to the development of the entire HPS and its function and environment tests.