• 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 KSME Conference
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• 2007.05b
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• pp.3068-3073
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• 2007

Aerodynamic forces and moments have been used to control rocket propelled vehicles. If control is required at very low speed, Those systems only provide a limited capability because aerodynamic control force is proportional to the air density and low dynamic pressure. But thrust vector control(TVC) can overcome the disadvantages. TVC is the method which generates the side force and roll moment by controlling exhausted gas directly in a rocket nozzle. TVC is classified by mechanical and fluid dynamic methods. Mechanical methods can change the flow direction by several objects installed in a rocket nozzle exhaust such as tapered ramp tabs and jet vane. Fluid dynamic methods control the flight direction with the injection of secondary gaseous flows into the rocket nozzle. The tapered ramp tabs of mechanical methods are used in this paper. They installed at the rear in the rocket nozzle could be freely moved along axial and radial direction on the mounting ring to provide the mass flow rate which is injected from the rocket nozzle. In this paper, the conceptual design and the performance study on the tapered ramp tabs of the thurst vector control has been carried out using the supersonic cold flow system and shadow graph. Numerical simulation was also performed to study flow characteristics and interactions between ramp tabs. This paper provides to analyze the location of normal shock wave and distribution of surface pressure on the region enclosed by the tapered ramp tabs.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.11a
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• pp.78-81
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• 2008

A numerical study of the unsteady flow in an over-expanded thrust optimized contour and compressed truncated perfect rocket nozzle is carried out in present paper. These rocket nozzles are subject to flow separation in transient phase at engine start-up and/or engine shut-down. The separation flow structures at different pressure ratios are observed. The start-up process exhibits two different shock structures such as FSS (Free Shock Separation) and RSS (Restricted Shock Separation). For a range of pressure ratios, hysteresis phenomenon occurs between these two separation patterns. A three-dimension compressible Navier-Stokes solver is used for the present study. One equation Spalart-Allmaras turbulence model is selected. The computed nozzle wall pressures show a good agreement with the experimental measurements. Present results have shown that present code can be used for the analysis of the transient flows in nozzle.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2005.11a
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• pp.487-490
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• 2005

The thrust Control in Solid Propellant Rocket is incomparably limited than that in Liquid Propellant Rocket. Because it is fixed that section to relate a combustion, that is a natural result. The control of a thrust directions in a Solid Propellant Rocket is not efficient for the purpose of a Solid Propellant Rocket. But it is a problem to solve that a weight on board should increase through the maximization of the thrust in a Solid Propellant Rocket.

• 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 for Aeronautical & Space Sciences
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• v.39 no.8
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• pp.744-749
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• 2011

In this study, determination method of initial optimal nozzle expansion in pintle rocket was investigated. The initial optimal initial nozzle expansion was decided by maximizing the mass-averaged thrust coefficient that is calculated from thrust coefficient of minimum and maximum chamber pressure. The determination of initial optimal initial nozzle expansion was equivalent to that of the minimum propellant mass which was required for obtaining the desired mission performance. The highest pressure, thrust turndown ratio and total impulse ratio effected on the initial optimal nozzle expansion. Among them, total impulse ratio had great influence on the initial optimal nozzle expansion.

• 한국전산유체공학회:학술대회논문집
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• 2008.03b
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• pp.547-550
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• 2008

A solid rocket motor has quite complex physical condition such exothermal chemical reaction in subsonic area and supersonic ex pansion in a converging-diverging nozzle. To introduce a simulation tool for compressible flow in supersonic range as well as incompressible chemical reaction zone in a whole rocket nozzle is a essential demand. Since the flow vary subsonic to super sonic, the convergence in computation becomes very low and unstable in a whole domain of rocket motor. This paper reports the 2-D Axisymmetric and simple 3-D solid propellant combustion and flow of gases in rocket motor by using a precondi tioning, shear stress turbulence modeling, AUSM(p). To simulate the simplified combustion process, Double base solid propellant is used to calculate reaction of solid propellant.

• Journal of Aerospace System Engineering
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• v.2 no.2
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• pp.28-34
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• 2008

This paper describes the performance characteristics of flow and structure in the post-chamber and nozzle. Using the computational fluid dynamics (CFD) technique, the stress and compressible flow fields in the downstream of the post chamber and nozzle were numerically calculated. Besides, the stress characteristics at the wall of post-chamber and nozzle were investigated under different thickness (2mm, 5mm and 10mm) of the nozzle wall. The stress pattern demonstrates that the strength of nozzle wall having 10mm is safer than that of 5mm or 2mm according to von Mises stress irrespective to the pressure field

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.1072-1076
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• 2017

In this paper, a numerical study has been performed to analyze flow characteristics of rocket propulsion. Through the ground spin test, combustion chamber pressure was measured. Based on the experimental results, numerical analysis was conducted under various nozzle pressure ratio conditions such as standard, operating and base pressure conditions. And it was compared with quasi-1D solution and experimental result. In addition, the difference in thrust characteristics according to the spin/non-spin of the flow conditions was confirmed at the same nozzle pressure ratio.

• Journal of the Korean Society of Propulsion Engineers
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• v.18 no.2
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• pp.80-85
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• 2014

Side jet thruster using nozzle closure separation device provides a solid rocket with a trajectory shift function. Side jet thruster consists of low combustion temperature propellant, neutral type propellant grain and nozzle closure separation device. If a trajectory shift is required, side jet thrust is generated on the rocket by separating some nozzle closures located in the opposite direction to thrust. After completing trajectory shift, the other nozzle closures located in the thrust direction are separated to cease side jet thrust. The operation process is verified through ground static test. The result in this study can be applied to changing rocket trajectory by controlling side jet thrust through nozzle closure separation.