• Journal of the Korean Society of Propulsion Engineers
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• v.19 no.2
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• pp.1-8
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• 2015

Dual pulse rocket motor is a variant of solid rocket motor with two propellant grain separated by a pulse separation device. The major performance of such a rocket motor is influenced by the change in the hole area of pulse separation device to nozzle throat area ratio. In this study, we performed flow analysis to investigate the internal flow characteristics according to the pulse separation device hole area to nozzle throat area ratio change. Gases used flow analysis were used combustion gas of HTPB/AP composite propellant and nitrogen gas. Flow analysis results of the dual pulse rocket motor were validated by comparison with experimental results of pneumatics. Commercial CFD code ANSYS FLUENT 14.5 is used in this study to simulate flow analysis.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.31-37
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• 2011

Detail flow structure in a large solid rocket motor with two inhibitors has been investigated using 3D Large Eddy Simulation and Proper Orthogonal Decomposition(POD) analysis. Vortex shedding frequencies periodically occurred by inhibitors are coupled with flow acoustics induced by the impinging of vorticity on nozzle head. As a result of 3D analysis, it was observed that the nozzle exit flow causes roll-torques from the vortex being decomposed in unbalanced shape for the impinging of vorticity on the nozzle head.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2003.05a
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• pp.127-130
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• 2003

This paper presents the numerical thermal analysis for regeneratively cooled rocket thrust chambers. An integrated numerical model incorporates computational fluid dynamics for the hot-gas thermal environment, and thermal analysis for the liner and coolant channels. The flow and temperature fields in rocket thrust chambers is assumed to be axisymmetric steady state which is presumed to the combustion liner. The heat flux computed from nozzle flow is used to predict the temperature distribution of the combustion liner. As a result, we present the wall temperature of combustion liner and the temperature change of coolant.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.05a
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• pp.215-218
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• 2009

A design of a subscale solid propellant rocket motor was conducted to do the similitude experiments for the large scale rocket motor. One of the main factor to subscale was the mach number of the solid propellant flume through a nozzle exit The analysis of the flume flow was done to obtain the mach number for the large and subscale rocket motor. The flume shapes on the non dimensional axises by the nozzle exit diameter was matched each other. The propellant grain of a subscale solid rocket motor was designed by the profile of pressure vs time obtained by the mach number of the flume shape. Some analyses of the theoretical solution were compared with the results of the ground static test.

• Journal of the Korean Society of Propulsion Engineers
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• v.15 no.3
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• pp.31-39
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• 2011

The flowfield in a solid rocket motor was simulated at the wall-injection test model, which has a fin-slot grain and submerged nozzle, and visualized by a smoke-wire. The high speed CCD camera captured the visualized images around the nozzle inlet through the grain center port. The vortical tube structure and circumferential flow patterns at the nozzle throat were visualized. The radial momentum transfer caused by the shear-interactions of slot-outlet flow, fin-base flow and grain port flow from upstream worked as the source of these phenomena.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.32 no.9
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• pp.35-40
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• 2004

The high-speed wind tunnel test of rocket model was performed to investigate the effect of skirt configuration on aerodynamic load characteristics of nozzle. Test parameters were the length and diffusing angle of skirt. Test results showed that the gimbals actuator power could be reduced to 1/10 of that without skirt. The normalized test result was proposed to be used as database for skirt design.

• 한국전산유체공학회:학술대회논문집
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• 2001.10a
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• pp.99-105
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• 2001

Flow computations have been conducted to study the impingement flowfield over the KSR-III flume deflector To validate Euler solver for the jet impingement flowfieid, the jet flow over a double wedge deflector have been calculated and showed reasonable agreement with experimental data. The transient flow behavior of flume over deflector have been investigated and the flume from the rocket nozzle proved to be getting out of the deflector safely and the thermal effect on the base region of rocket was not considerable.

• Journal of the Korean Society of Visualization
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• v.9 no.2
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• pp.61-65
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• 2011

Several previous works on rocket nozzle flows have revealed the existence of the transition from FSS to RSS and the occurrence of asymmetric flow associated with the boundary layer separation, which can cause excessive side-loads of the propulsion system. Thus, it is of practical importance to investigate the asymmetric flow behaviors of the propulsion nozzle and to develop its control method. In the present study, the asymmetric flow control method using a cavity system was applied to supersonic nozzle flow. Time-dependent asymmetric flow was experimentally investigated with the rate of change of the nozzle pressure ratio. The results obtained showed that the cavity system installed on nozzle wall would be helpful in fixing the unsteady motions of the boundary layer separation, consequently reducing the possibility of the occurrence of the asymmetric flow.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.03a
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• pp.165-172
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• 2004

LE-7A is the main engine of the H-IIA launch vehicle. Under its development, the nozzle suffered from two troubles during startup and shutdown transients of the engine. One is a large side load, which damages the actuator of the nozzle, and the other is damage on regenerative cooling tubes due to high heat load. It has been considered that these problems are caused by a peculiar separation pattern called Restricted Shock Separation (RSS). RSS is observed in several rocket nozzles, for example, LE-7A nozzle, Vulcain nozzle and so on. Their contours are not conventional truncated perfect (TP) nozzle - LE-7A nozzle is a compressed truncated perfect (CTP) nozzle and Vulcain nozzle is a thrust optimized (TO) nozzle. Although it is believed that the occurrence of RSS is affected by the nozzle contour, the mechanisms are not clarified sufficiently yet. In the present paper, a parametric numerical study is carried out to investigate the mechanisms of the occurrence of RSS in CTP nozzles during startup transient. The results show that RSS is caused by the adverse pressure gradient downstream of the Mach disk. The adverse pressure gradient is caused by the interaction of the pressure wave and Mach disk. The method to avoid the occurrence of RSS is also examined. A small step inside the nozzle affects the position of the separation point and prevents RSS. The result shows that the possibility that RSS can be suppressed by controlling the position of the separation point.

• Journal of the Korean Society of Propulsion Engineers
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• v.19 no.1
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• pp.42-49
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• 2015

The pressure oscillation induced by inhibitor in a solid rocket motor has been investigated by 3D large eddy simulation(LES) and proper orthogonal decomposition(POD). The vortex generation and breakdown at inhibitor are periodically observed between the inhibitor and the nozzle by flow-acoustic coupling mechanism. The excitation of pressure oscillation occurs as the flow impinges on the submerged nozzle head which recirculate in the cavity in rear dome of the motor chamber. The vortex generation frequency is closely related with the shedding frequencies of the detached vorticities at the inhibiter, which fairly compared with the experimental data.