• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2005.11a
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• pp.162-166
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• 2005

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.

• 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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• 2012.05a
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• pp.123-128
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• 2012

According to the sequence of pintle operation, the performance characteristics of pintle nozzles are analyzed. The pintle movement was simulated using unsteady numerical techniques, the response lag and sensitivity at the chamber and nozzle are estimated for movable pintle. Three operation sequences of the pintle are considered for evaluating whether the rate of the chamber pressure increase and the operation sequence will have any significant impact towards the rocket performance. Three operation sequencies are as following; the pintle moves toward the nozzle throat, it turns instantly (case 1), stops at the nozzle throat for some time(0.5sec) (case 2), and stops at the nozzle throat (case 3). As a result, the dynamic characteristics according to the operation sequence and pintle shape were analyzed to take account the rocket performance.

• Journal of the Korean Society of Propulsion Engineers
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• v.3 no.4
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• pp.83-92
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• 1999

A numerical analysis of two phase flow in the solid rocket nozzle was conducted. Stoke number was defined over the various aluminum oxide($AI_2$$O_3$) particle sizes and particle trajectories were treated by Lagrangian approach. Particle stability was considered by the definition of Weber number in a rocket nozzle. Large particles are divided after the nozzle throat as the flow accelerates rapidly. The division of particles changes the particle distribution at the nozzle exit. From the above results, it was found that the nozzle converge section surface might be affected by aluminum oxide particles. Also, Mechanical erosion rate of nozzle surface was predicted for different materials.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.03a
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• pp.565-572
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• 2004

Numerical stability analysis of one-dimensional axial flow in solid rocket motors is performed based on the Euler equation coupled with an unsteady combustion equation of solid propellant. In order to check the numerical scheme, behavior of a standing wave in a closed tube is examined. A standing wave in solid rocket motor decays or grows depending on the total effect of propellant combustion, nozzle flow, and so on. The stability boundary of the fundamental mode standing wave is determined by changing one of the combustion parameters. In addition growth rates of the wave are calculated numerically in relatively low Mach number flow region for the motors with different port and nozzle throat diameters. The results obtained here agree well with the approximate solution. The same scheme is applied to a motor with shorter length and L＊-instability is observed.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.11a
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• pp.89-93
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• 2008

The effect of pintle shapes and position to the thrust performance of Solid Rocket Motor was studied by experimental-aided Computational Fluid Dynamic(CFD). Among the turbulent models for RANS in Fluent, Spalart-Allmaras model was better agreement with the nozzle wall pressure distribution attained by cold-flow test than other models. When nozzle throat area was decreased, magnitude of thrust was increased. The bigger pintle size was, the more thrust of pintle tip pressure was obtained. Meanwhile the more thrust of nozzle and chamber pressure decreased. Hence, total thrust of big pintle was less than a small pintle under same throat area condition.

• 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.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.36 no.8
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• pp.790-796
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• 2008

Typical solid rocket motors have a fixed propellant grain shape and nozzle throat size resulting in a fixed motor thrust. Pintle nozzle has been suggested as a means of providing variable thrust while maintaining the inherent advantage of solid rocket motors. In this study, the pintle shape effect on nozzle performance is investigated using experimental-aided Computational Fluid Dynamics(CFD). The pintle shape is modified by a principle of monotony. CFD analysis is performed using Fluent by applying the turbulent model. This analysis indicates that nozzle thrust and pintle load are influenced by change of nozzle shock pattern and flow separation due to pintle shape and there exists a high-performing pintle shape.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• v.y2005m4
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• pp.119-122
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• 2005

Ablation phenomena is very complicated because it includes momentum, energy and mass transfer, chemical reactions as well as phase change. In this paper, ablation at the rocket nozzle throat is modeled as unsteady one dimensional axi-symmetric with proper boundary conditions and field equation is solved numerically. Analytical results are compared with measured ablation data from firing experimental liquid rocket engine. Test variables are combustion pressure and mixture ratio. for low combustion pressure and low mixture ratio, the discrepancy between analysis and experiments are large but for the normal rocket operation range, two results show a simliar trend with maximum discrepancy of $100\%$.

• Journal of the Korean Society of Propulsion Engineers
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• v.21 no.4
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• pp.12-20
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• 2017

The ablation characteristics of graphite nozzle throat insert was analyzed for the use in solid rocket propulsion system. The propulsion system was composed of three types of conventional nozzles, such as De-Laval type, blast tube type, and submerged type. Various kinds of propellants were used in ten kinds of propulsion system that had different shapes with each other. Total forty eight tests were performed. From the results of the analysis, it was found that the ablation rate was increased for the higher average chamber pressure and the higher oxidizer mole fraction. A useful correlation for nozzle throat ablation rate was developed in terms of the chamber pressure, oxidizer mole fraction, and throat size. The calculated ablation rates from the correlation showed agreement within ${\pm}0.10mm/s$ with the experimentally determined values.