• Aerospace Engineering and Technology
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• v.3 no.2
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• pp.50-55
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• 2004

Performance predictions of the propulsion system were conducted for a 900㎏ class Canard Rotor/Wing vehicle. The main components of the propulsion system are turbojet engine, exhaust ducts and nozzles. The internal flow of the duct was considered as one-dimensional, compressible and viscous flow. Adequate governing equations including centrifugal force effect were applied to the analysis of the duct flows. Results such as available power, available thrust, engine throttle, mass flow rates, rotor RPM and cruise nozzle area were presented for rotary-wing mode and transition mode.

• Journal of the Korean Society of Propulsion Engineers
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• v.10 no.1
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• pp.30-37
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• 2006

A thermal resistance estimation of carbon/carbon composites used as the nozzle throat insert of solid rocket motor was performed using TPEM motor. Three types of TPEM motor and two types of propellant were employed. The ablation rate is higher for the higher chamber pressure and also higher for the higher concentration of oxidizing species in combustion gas, but it is lower for the higher material density.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 1995.11a
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• pp.207-213
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• 1995

로켓트 노즐 재료로 사용되는 그라파이트의 내삭마 특성을 향상시키기 의해 다층코팅에 관한 연구를 하였으며 $AI_2$$O_3$/PT/pack-SiC/Graphite와 CVD-$Si_3$$N_4$/CVD-SiC/pack-SiC/Graphite의 두 다층코팅 구조를 제작하여 내삭마 효과를 알아 보았다. $AI_2$$O_3$와 Pt충은 스퍼터링 증착방식을, $Si_3$$N_4$와 SiC층의 CVD는 저압화학증착방식을 사용하였다. 코팅층의 내삭마 특성 평가는 직접 노즐에 코팅층을 입혀 소형고체 추진기관 연소모타에 장착, 지상연소시험을 통해 수행하였으며 코팅층이 없는 표준모타에 비해 각각 60%, 80% 이상의 내삭마 특성의 향상을 보였다.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.37 no.4
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• pp.392-398
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• 2009

Theoretical thrust equations for the diverse nozzle expansion condition were derived. By using the obtained thrust equations, parametric studies were carried out to estimate the effect of pressure exponent, minimum operation pressure, ambient pressure and extinguishment pressure on thrust modulation performance in pintle-nozzle solid rocket motors. Analysis results showed that thrust turndown ratio can be easily attained by small nozzle-throat area variation at high pressure exponent, low minimum operation pressure, high ambient pressure and high extinguishment pressure condition. At those conditions, the highest chamber pressure to obtain the intended thrust turndown ratio can be minimized.

• Journal of the Korean Society of Propulsion Engineers
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• v.16 no.2
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• pp.42-50
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• 2012

The goal of heat transfer studies is the accurate prediction of temperature and heat flux distribution on material boundaries. To this purpose, general-purpose computational fluid dynamics(CFD) code is used : FLUENT. Mass fluxes and pressure ratio are calculated for two types of nozzle. The comparative studies reveal that the computational results are in agreement with the experimental data. Also, heat transfer coefficients from FLUENT for one type of nozzle are very similar and agree well with the experimental data in the diverging part of the nozzle, but the calculated results are large in the converging part. The heat transfer coefficients from Bartz equation are over-predicted. We can consider various reasons for these differences, i.e., laminarization by the highly accelerated flow in the nozzle, turbulent flow model and grid generation.

• Journal of the Korean Society of Propulsion Engineers
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• v.8 no.3
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• pp.53-60
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• 2004

In order to investigate the nozzle - rotor interactions and the effect of partial admission, the flows in supersonic turbine rotor cascades with a nozzle have been computed. Extensive computations of partial admission supersonic turbines provide the shock structures and flow patterns in the nozzle and rotor. The governing equations were discretized with Euler implicit method in time and 2nd-order upwind scheme of FVM in space. The $\kappa$-$\varepsilon$ turbulence model was utilized to describe the turbulent flow field. It is clearly shown that the nozzle flow is highly affected by the shocks or expansion waves propagated from the rotor leading edge. And the rotor flow is also affected by the shocks or wakes originated from the nozzle.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.33 no.6
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• pp.71-75
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• 2005

An experimental study was carried out to investigate the effect of film cooling in the lab-scale dump-cooled liquid rocket engine using LOX and kerosene as propellants. The nozzle of the rocket engine was film cooled with water as coolant. A special film cooling adapter was fabricated to introduce the film-coolant into the thrust chamber. The flow rates of film coolant was approximately 15~19 percent of the total propellant. The nozzle heat flux was determined from the measured temperature rise and flow rate of the coolant(water). Large reductions in the nozzle heat flux was resulted when film cooling adapter located directly upstream of the nozzle.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.206-212
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• 2011

Hysteresis phenomena in fluid flow systems are frequently encountered in many industrial and engineering applications and mainly appear during the transient processes of change of the pressure ratio. Shock-containing flow field in supersonic nozzles is typically subject to such hysteresis phenomena, but associated flow physics is not yet understood well. In the present study, experimental work has been carried out to investigate supersonic nozzle flows during the transient processes of change in the nozzle pressure ratio. Time-dependent surface wall pressures were measured by a multiple of pressure transducers and the flow field was visualized using a nano-spark Schlieren optical method. The results obtained show that the hysteresis phenomenon is strongly dependent on the nozzle geometry as well as the time scale of the change of pressure ratio.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.575-578
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• 2011

The thrust vectoring concept has been used for use in new advanced supersonic aircraft. This study presents the performance characteristics of the thrust vectoring nozzle by visualizing the shock behaviors with Schlieren method. We performed experimental tests to see the geometrical effects of the thrust vector nozzle by changing pitch angle and length of pitch flaps. From this study we could understand the supersonic flow characteristics of the thrust vector nozzle. The total thrust of thrust vector nozzle is diminished by increasing the flap angle. But there is an optimum flap length ratio for attaining the highest thrust level and proper pitch effect.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.116-121
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• 2011

Performance of Secondary Injection Thrust Vector Control system is investigated under various secondary injection operating conditions. 3-dimensional converging-diverging nozzle having 8 secondary injection nozzles is used in this numerical study. Total pressure of flow inside the nozzle is about 70bars, and total temperature set to 300K for cold flow simulation. Effect of secondary injection flow rate and injection nozzle configuration is considered in this research. Simulation is conducted with commercial CFD code Ansys Fluent v13. Spalart-Allmaras(1-equation)model is used for turbulence modeling with AUSM+ scheme. Various performance factors as Axial thrust, side force, system specific impulse ratio are considered and explained for system performance evaluation.