• Journal of the Korean Society of Propulsion Engineers
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• v.13 no.6
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• pp.8-16
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• 2009

The development of high-pressure sub-scale combustion chambers is described. A total of four high-pressure sub-scale combustion chambers having either a detachable structure of the mixing head and the chamber or a single welded regenerative cooling structure have been developed. The sub-scale combustion chambers have a chamber pressure of 70 bar and propellant mass flow rate of 5.1~9.1 kg/s. The propellant mass flow rate and the recess number of the injector were changed for the improvement of combustion performance and they were validated through hot firing tests. The design and manufacturing techniques of regenerative cooling channel and film cooling to be applied to the full-scale combustion chamber were adopted through the present development and verified.

• Transactions of the Korean Society of Mechanical Engineers A
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• v.31 no.10
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• pp.1046-1052
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• 2007

Elastic-plastic structural analysis for regenerative cooling chamber of gas generator was performed. Uniaxial tension test was conducted for STS316L at room and high temperature conditions to get the material data necessary for the structural analysis of the chamber which was operated under thermal load and high internal pressure. Physical properties including thermal conductivity, specific heat and thermal expansion were also measured. The structural analysis for four different types of regenerative cooling chamber of gas generator revealed that increased cooling performance decreased the thermal load and strain of the cooling channel structure. The results propose that in order for the regenerative cooling gas generator chamber to have high structural stability with endurance to high mechanical and thermal loads, it is important for the chamber to be designed to have high cooling performance.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.32 no.4
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• pp.67-72
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• 2004

A regenerative cooling system has been designed through empirical 1-D analysis for a liquid rocket engine of 30-ton-level thrust. The hot-gas-side wall temperature from 1-D analysis shows 100K difference compared to 3D CFD analysis. Two variations of design with same cooling performance are suggested for different maximum channel widths i.e., 4mm and 2mm. The coolant pressure drop of the latter design is higher by 20%. The maximum liner temperature is about 700K when TBC and the thermal resistance of carbon deposit are considered. So film cooling is recommended to increase the cooling capacity as the present cooling capacity is insufficient

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.37 no.12
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• pp.1201-1208
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• 2009

Dome stretching tests and tension tests were performed to evaluate the formability of a copper alloy used for manufacturing the regenerative cooling chamber. The test specimens were prepared to investigate the effect of heat treatment and direction of specimens on the formability. The test results show that forming limit values are increased by the heat treatment of the material but the variation of the forming limit values by manufacturing direction is negligible compared to the heat treatment effect, and forming limit values are also different according to the test methods. These results indicate that the high temperature heat treatment of the material before bulging is a very important process to deform the inner cylindrical structure of the regenerative cooling chamber into a nozzle shape by the bulging process without necking or fracture and the test methods also have a great effect on a evaluation of the formability. The forming limit diagram obtained in this study would be utilized to the design of regenerative cooling chamber nozzles.

• Journal of the Korean Society of Propulsion Engineers
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• v.10 no.2
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• pp.53-61
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• 2006

Axisymmetric numerical thermal analysis for a 3-dimensional regenerative cooling system in a rocket engine is carried out. To predict the accurate heat transfer with the stiff temperature distribution, several tests have been conducted for the grid size, the properties variation of the coolant and the combustion gas depending on temperature. The axisymmetric heat flux model is defined using fin efficiencies and is designed to be equivalent to the heat flux of the 3-dimensional coolant channel. For comparison purpose, the 1-dimensional analysis using Bartz equation is also conducted. The performance of the present model in predicting the cooling characteristics of a 3-dimensional regenerative cooling system is compared with the 3-dimensional results of RTE(Rocket Thermal Evaluation). It is found that the present method predicts much closer results to those of RTE code than 1-dimensional analysis.

• Journal of the Korean Society of Propulsion Engineers
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• v.16 no.6
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• pp.16-22
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• 2012

Performance of a liquid rocket thrust chamber with regenerative cooling scheme has been numerically analyzed using in-house CFD code which can predict combustion/cooling performance and provide nozzle design parameters. This paper investigates trade-offs between combustion and cooling performance with varying amount of fuel directly injected into the chamber wall to form cooling films and mixture ratios for the peripheral injectors. Further efforts to verify/improve the simulation methodology including comparison with the firing test results are planned to make it a reliable tool to optimize the film cooling and other major design parameters.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2012.05a
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• pp.35-41
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• 2012

Performance of a liquid rocket thrust chamber with regenerative cooling scheme has been numerically analyzed using in-house CFD code which can predict combustion/cooling performance and provide nozzle design parameters. This paper investigates trade-offs between combustion and cooling performance with varying amount of fuel directly injected into the chamber wall to form cooling films. Also is analyzed the effect of varying mixture ratios for the peripheral injectors on combustion performance enhancement. Further efforts to verify/improve the simulation methodology including comparison with the firing test results are planned to make it a reliable tool to optimize the film cooling and other major design parameters.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.05a
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• pp.151-155
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• 2006

Elastic-viscoplastic structural analysis is performed for regenerative cooling chamber of liquid rocket thrust chamber using Bodner-Partom visco-plastic model. Strain rate test was also conducted for a copper alloy at various temperatures in order to get material constants of visco-plastic model used in the structural analysis. Material constants of visco-plastic model were obtained from strain rate test results and visco-plsstic model was incorporated into finite element program, Marc, by means of user subroutine. The structural analysis results indicate that the deformation of cooling channel is mostly caused by thermal loading rather than pressure loading and confirmed structural stability of the cooling channel under operating condition.

• Journal of the Korean Society of Propulsion Engineers
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• v.10 no.4
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• pp.69-76
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• 2006

Elastic-viscoplastic structural analysis has been performed for regenerative cooling chamber of liquid rocket thrust chamber using Bodner-Partom visco-plastic model. Strain rate test was conducted for a copper alloy at various temperatures in order to get material constants of visco-plastic model used in the structural analysis. Material constants of visco-plastic model were obtained from strain rate test results and visco-plastic model was incorporated into finite element program, Marc, by means of a user subroutine. The structural analysis results indicated that the deformation of cooling channel is mostly caused by thermal loading rather than pressure loading and confirmed structural stability of the cooling channel under the operating condition.

• Journal of the Korean Society of Propulsion Engineers
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• v.6 no.4
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• pp.25-35
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• 2002

A calculation procedure for designing LRE regenerative cooling system is introduced. In LRE thrust chamber, heat is transfered from the hot gas to the wall by convection and radiation, then conduction through the wall and finally convection to the liquid coolant. A cooling channel is designed on the basis of heat transfer rate calculated by using criterial method and integral method. The result is compared with existing Russian cooling channel design code. Also a design logic and quantitative effect of various parameters were introduced to help better understanding for those who is not familiar to LRE system.