• Journal of Aerospace System Engineering
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• v.11 no.4
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• pp.15-21
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• 2017

In this study, the surface heat transfer coefficient of the 3D model of a thruster regulator in the high temperature and high pressure environment was estimated using the commercial CFD code. The thermal barrier coating (TBC) on the surface of the thruster regulator was modeled and the effect of the thickness of the TBC on the temperature of the thruster regulator was investigated. The thickness of the TBC was varied from $100{\mu}m$ to $500{\mu}m$. Results showed that the temperature of the surface and the inside the thruster regulator was lower for the thicker TBC case.

• Aerospace Engineering and Technology
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• v.3 no.2
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• pp.183-190
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• 2004

The performance analysis of a fuel pump for a liquid rocket engine has been performed numerically on its design condition. A commercial three-dimensional Navier-Stokes flow solver has been used for the computation. All of the fuel pump components - inducer, impeller, volute and secondary flow passages - are included in computation for the accurate estimation of the leakage flow rate which affects the performance and axial thrust. A pitchwise-averaged mixing plane method was used on the boundaries among the fuel pump components to save computational time. The predicted overall performance satisfied the design requirement. However, the axial thrust exceeded a permissible limit. In order to reduce the axial thrust, the secondary flow passage design has been changed. With this change, the axial thrust level has been reduced to 30% as compared with the original value.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.34 no.3
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• pp.237-243
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• 2010

In this study, the effect of the shape of a pintle(obstacle) on thrust-modulation performance and drag in a pintle rocket was investigated by a cold flow test and by computational fluid dynamics. Pintle movement caused a monotonic increase in the chamber pressure. Thrust generated by the pressure distribution on the pintle body was linearly changed to the chamber pressure, and this thrust was greater than that generated by the nozzle-wall pressure distribution. Because the shock pattern in the nozzle changes with the shape of the pintle body and pressure ratio, the thrust generated by the nozzle-wall pressure is not directly affected by chamber pressure. The drag due to the pintle(obstacle) can be minimized for a fully linear pintle shape, regardless of chamber pressure.

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

• Aerospace Engineering and Technology
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• v.7 no.2
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• pp.88-94
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• 2008

In this study, triaxial velocity is analyzed for COMS(Communication, Ocean and Meteorological Satellite) configuration, which is generated when thrusters are used to dump wheel momentum. Since COMS is designed to periodically change the thruster set in order to uniformly decrease the performance of thrusters, triaxial velocity would be different during the change of thruster set. So, the triaxial velocity generated due to the change of thruster set is optimized.

• Journal of Astronomy and Space Sciences
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• v.18 no.2
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• pp.137-144
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• 2001

This paper presents the REA thrusters performance estimation results for the KOREASAT F1&F2 launched in 1995 The satellite tracking data obtained from the ground system from end of 1999 to beginning of the 2000 are used to estimate the thruster performance. The estimation algorithm is derived from the least square estimation theory and designed to estimate the velocity change induced by the on-boarded thruster firing as well as the orbit parameter. The estimation results show that the Koreasat F1 thruster are in bad thruster condition of 64% performance for REA when it fires in on-pulse mode. Here, the performance is defined by the ratio of the resulted velocity change to that of planned. But, in the case o( the Koreasat F2, it is found that the performance approximately reaches up to 100%, even after the 5 years of the mission.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.11a
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• pp.215-218
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• 2007

Internal Ballistic modeling and performance prediction for solid propellant micro thruster was performed with heat loss to the chamber wall as an important factor of miniaturization. Simple l-D end-burner type thruster and general HTPB-AP type composite propellant were selected for computation model. The results showed that the performance loss with the heat loss to the surroundings becomes larger as the surface-to-volume ratio is increased. In this case, the total impulse was reduced about 3% of the case in adiabatic condition.

• Journal of Aerospace System Engineering
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• v.17 no.6
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• pp.63-71
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• 2023

In the present study, the effect of pitch amplitude on the unsteady aerodynamics of a NACA 0012 airfoil is numerically investigated. When the frequency ratio is equal to 1.0, airfoil pitching with 20 and 30 degrees of pitch amplitude shows almost small lift generation, but the lift is significantly increased in case of 10-degree pitch amplitude. When the frequency is 0.5, the lift coefficients have large values, and the lift increases with a decrease in pitch amplitude. When the frequency ratio is 1.0, the airfoil generates large thrust. The thrust decreases as the pitch amplitude decreases. When the frequency ratio is 0.5, drag is generated for the 30-degree pitch amplitude, but the thrust is generated for 10-degree pitch amplitude. In future, the effect of heave amplitude on the unsteady aerodynamics of the airfoil will be studied.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.44 no.7
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• pp.576-584
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• 2016

In this study, AP added propellant has been proposed as a method of enhancing the low specific impulse performance found for staged hybrid rocket engine. Experimental tests were carried out to analyze and evaluate the effect of AP added propellant on specific impulse performance as well as fuel-rich combustion characteristics in a staged hybrid rocket engine. Upper limit of AP content in propellant was set to be 15 wt% to maintain the hybrid rocket engine advantages. As a result, 15 wt% AP added propellant showed 3% higher specific impulse performance compared to 0 wt% AP added propellant. Moreover, AP addition proved to offer less injected oxidizer mass flow, less O/F variation, and less combustion pressure while producing fuel-rich gas of the same combustion temperature. Future studies will carry out more combustion tests with metal additives to further enhance specific impulse.

• Journal of the Korean Society of Propulsion Engineers
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• v.18 no.6
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• pp.19-26
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• 2014

A ground hot-firing test (HFT) was carried out to make a close examination into the pressure instability for the 70 N-class hydrazine thruster under development. Monopropellant grade hydrazine was adopted as a propellant for the HFT, and catalyst-bed was filled with $Ir/Al_2O_3$ catalyst. In order to investigate the effects of thrust-chamber diameter on combustion stability, evaluation tests for the development models were performed on three kinds of lower thrust chambers having the length-to-diameter ratio (L/D) of 1.03, 1.13, and 1.26. As results, it was found that low frequency instability (~ 50 Hz) was inherent in the models, and in addition, increase of the L/D and decrease of the operating pressure led to an amplification of pressure oscillation in the test condition specified.