• Aerospace Engineering and Technology
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• v.7 no.2
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• pp.103-109
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• 2008

A study has been conducted on the bulging process of regenerative cooling nozzle which is essential for the manufacturing of liquid rocket thrust chamber. Tension tests have been performed for the material to be used for the development of the bulging process and mechanical properties are obtained by the test. Two or three bulging tools were required to complete the bulging process. The necking of the material was a major failure encountered in the bulging process and a research has revealed that grain size of the material has considerable effect on its occurrence. The presently developed bulging process with a controlled grain size material has been successfully applied to the manufacturing of subscale and 30-tonf full scale regeneratively cooled nozzle while demonstrating the applicability and usefulness of the presently developed bulging process.

• Proceedings of the Korean Society For Composite Materials Conference
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• 2002.10a
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• pp.100-105
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• 2002

This paper predicts the material properties of spatially reinforced composites (SRC) and analyzes the thermo-elastic behavior of a kick motor nozzle manufactured from that material. To find the appropriate SRC structure for the nozzle throat that satisfies given design conditions, the equivalent material properties of the SRC are predicted using the superposition method for those of rod and matrix. Studied are the elastic behavior, temperature distribution, and thermo-elastic behavior of a kick motor nozzle composed of carbon/carbon SRC as a throat part. The elastic deformation of the nozzle composed of 3D carbon/carbon SRC shows asymmetry in a circumferential direction. However, 4D carbon/carbon SRC nozzle shows uniform deformation in the circumferential direction. Stress concentration in connecting parts of the kick motor nozzle is ultimately high due to the high temperature gradient in each connecting part. The thermo-elastic deformations of both the 3D and the 4D SRC nozzles are uniform in the circumferential direction due to the isotropy of CTE of each SRC. The deformation of the 3D SRC nozzle is a slightly smaller than that of the 4D SRC nozzle in the nozzle throat, which is favorably effective on rocket thrust. The circumferential stress is the most critical component of the kick motor nozzle. The 4D SRC nozzle having 1,1,1,1.7 diameters in each direction has the smallest circumferential stress among several SRC nozzles.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.521-528
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• 2017

In this paper, the numerical analysis for heat conduction of silica/phenolic composite material, which is used for solid rocket nozzle liner or insulator, was conducted. 1-D Finite Difference Method for the analysis of silica/phenolic during the firing of solid rocket motor was used to calculate the heat conduction considering the surface ablation and the thermal decomposition. The boundary condition at the nozzle wall took into account the convective heat transfer, which was obtained by integration equation. The numerical results of the surface ablation and char depth were compared with the results of test motor that is TPEM-10. It was found that the result of calculation is favorably agreed with the thermal response of test motor.

• Journal of the Korean Society of Propulsion Engineers
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• v.22 no.4
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• pp.76-84
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• 2018

In this paper, the numerical analysis for heat conduction of silica/phenolic composite material, used for solid rocket nozzle liners or insulators, is conducted. A 1-dimensional finite difference method for the analysis of silica/phenolic during the firing of a solid rocket motor is used to calculate heat conduction, considering surface ablation and thermal decomposition. The boundary condition at the nozzle wall, considering the convective heat transfer, is obtained via integration equations. The numerical results of the surface ablation and char depth are compared with the results of a TPEM-10 test motor, finding that the result of calculation agrees with the thermal response of the test motor.

• Journal of the Korean Society of Propulsion Engineers
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• v.9 no.2
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• pp.25-31
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• 2005

The theory developed in a preceding paper [1] for poroelastic composite material behavior under thermal and gas diffusion is applied to thermo-chemical decomposition of a carbon-phenolic composite rocket nozzle liner under typical operating conditions. Specifically, the structural component simulated is the cowl ring for which distributions of pressure in the material pores, temperature and across-ply stress are presented. The results for particular composite designs show that across-ply failure occurs due to tensile stress in the material which is indicative of plylift. This prediction corroborates observations of plylift in a nozzle cowl. Simulations suggest designs to avoid plylift in the cowl zone.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.05a
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• pp.113-122
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• 2010

A two-dimensional thermal response and ablation analysis code for predicting charring material ablation and shape change on solid rocket nozzle is presented. For closing the problem of thermal analysis, Arrhenius' equation and Zvyagin's ablation model are used. The moving boundary problem are solved by remeshing-rezoning method. For simulation of complicated thermal protection systems, this method is integrated with a three-dimensional finite-element thermal and structure analysis code through continuity of temperature and heat flux.

• Journal of Astronomy and Space Sciences
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• v.3 no.1
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• pp.29-40
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• 1986

It is the aims of this study to choose the materials and determine the material thickness of laminated Rocket Nozzle Wall operating at high pressure and high temperature. The heat conduction analysis of each layer was performed by Crank Nicolson method changing the thickness and the materials for the imput data of Tungsten, Graphite, Alumina, Aluminum, Molybdenum, Plastic laminate. The results of the study for pressure of 93.5kg/$cm^2$ and temperature of $3000^{circ}C$ in the nozzle dia of 40cm are as follows.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.05a
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• pp.311-314
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• 2009

A two-dimensional thermal response and ablation simulation code for predicting charring material ablation and shape change on solid rocket nozzle is presented. For closing the problem of thermal analysis, Arrhenius' equation and Zvyagin's ablation model are used. The moving boundary problem and endothermic reaction in thermal decomposition are solved by rezoning and effective specific heat method. For simulation of complicated thermal protection systems, this method is integrated with a three-dimensional finite-element thermal and structure analysis code through continuity of temperature and heat flux.

• Journal of the Korean Society of Propulsion Engineers
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• v.22 no.1
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• pp.36-44
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• 2018

The thermal response of carbon/phenolic used in a solid rocket nozzle liner was analyzed. In this paper, the numerical analysis of the thermal response of carbon/phenolic consists of (1) the integration equation of the boundary layer to obtain the convective heat transfer coefficient of the combustion gas on the rocket nozzle wall and (2) 1-D finite difference method for heat conduction of carbon/phenolic to calculate the ablation, char, and temperature. The calculated result was compared with the result of a blast-tube-type test motor. It is found that the calculated result shows good agreement with the thermal response of the test motor, except at the vicinity of the throat insert.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.11a
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• pp.114-118
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• 2007

Design and fabrication of a 30-tonf-class full-scale regenerative cooling combustion chamber of a liquid rocket engine for a ground hot firing test are described. It has chamber pressure of 60 bar and nozzle expansion ration of 12 and manufactured to have a single welded structure of· the mixing head and the chamber. The material of the mixing head is STS316L which has excellent mechanical property in cryogenic condition. The chamber comprise of the cylinder, nozzle throat, and 1st/2nd nozzle parts. The material of the inner jacket is copper alloy/STS329J1/STS316L and that of the outer jacket is STS329J1. The components of· the combustor were manufactured by mechanical processing including lathing, milling, MCT, rolling and pressing. The machined components were integrated to a single body by means of general welding, electron beam welding(EBW), and brazing.