Chul-Sung Ryu

Structural Analysis of Gas Generator Regenerative Cooling Chamber

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.

Structure design of regenerative cooling chamber of liquid rocket thrust chamber

Elastic-plastic structural analysis for regenerative cooling chamber of liquid rocket thrust chamber is performed. Uniaxial tension test is also conducted for the copper alloy in order to get material data necessary for the structure analysis. The results of uniaxial tension test reveal that copper alloy become ductile after brazing process and flow stress becomes lower as temperature becomes higher. As a result of structural analysis using the material data, the deformation of cooling channel is more increased by thermal load than by internal pressure of cooling fluid. Therefore, the results of analysis show that structural stability and cooling performance of combustion thrust chamber which is designed to endure mechanical load and minimized a channel thickness are improved by decreased thermal load as possible.

Development of Spinning Process for Manufacturing Liquid Rocket Engine Thrust Chamber

ABSTRACT Spinning process to inner wall has been applied for reducing the weight of regenerative cooling chamber of liquid propellent rocket engine. The fractures of th e blanks of cylinder part and nozzle throat part have been observed during spinning processes. In or der to overcome the problem, the mandrel and the blank shape have been modified, and the inner w all was successfully manufactured through the modifications. The manufactured spinning prototype of nozzle throat part was successfully bulged without cracking and necking, and it was confirmed to se cure sufficient formability necessary for fabricating thrust chamber.초 록 액체로켓 엔진 재생냉각챔버의 무게 감량을 위하여 내피에 스피닝 공정을 적용하였다. 스피닝 도중 실린더부와 노즐목부의 블랭크에 파손이 발생하였다. 이 문제를 극복하기 위해 맨드럴 및 블랭크 형상을 수정하였으며 이러한 수정을 통해 스피닝을 사용하여 내피를 제작할 수 있었다. 제작된 노즐목부 스피닝 시제품은 균열이나 네킹없이 성공적으로 벌징이 되어 연소기 제작에 적용할 수 있는 충분한 성형성을 확보한 것을 확인하였다.Key Words: Chromium Copper Alloy(크롬 구리합금), Spinning(스피닝), Inner Wall(내피), Combustion Chamber(연소실), Regenerative Cooling Chamber(재생냉각 챔버)

Thermal Barrier Coating Durability Testing Trends for Thrust Chamber of Liquid-propellant Rocket Engine

․ Chul-Sung Ryu** ․ Byoung-Jik Lim* ․ Hwan-Seok Choi*ABSTRACT Durability testing method trends of the thermal barrier coati ng(TBC) for the combustion chamber of the liquid-propellant rocket engine have been investigated. Many types of the durability testing method such as the mechanical tests to measure surface cohesion force, the thermal fatigue tests with laser, furnace, burner or plasma, the small scale combustion te sts using injectors, and the thermo-mechanical fatigue tests were observed. The TBC with suf ficient durability can be selected for the use of combustion chamber through such specimen-level tests and the durability can be verified by the tests using the real scale combustion chambers.초 록 액체로켓 연소기에 사용되는 열차폐코팅(TBC)의 내구성 시험 기술동향을 조사하였다. 표면 접합력 측정을 위한 기계적 시험, 레이저나 가열로, 버너나 플라즈마 등을 이용한 열피로 시험, 분사기를 이용한 소형 연소시험, 열적 기계적 피로시험 등의 많은 내구성 시험들이 있었다. 연소기에 사용하기 위해 이러한 시편 수준의 시험을 통해 내구성이 확보된 TBC를 결정할 수 있으며, 실제 연소시험을 통해 내구성을 검증할 수 있다.Key Words: Combustion Chamber(연소기), Thermal Barrier Coating(열차폐코팅), Bending Test(굽힘시험), Durability(내구성), Thermal Fatigue(열피로), Combustion Test(연소시험)

A Technical Trend of Manufacturing and Materials of Nozzle Extension for Thrust Chamber of Liquid Rocket

The combustion chamber and nozzle of a liquid rocket engine should be protected from the high temperature combustion gas generated by the chamber. An upper-stage nozzle extension has a large expansion ratio, therefore, The light-weight refractory materials have been used since the weight impact on the launcher performance is crucial. Gas film cooling method was used before, but was not applicable nowadays. Ablative cooling method and radiative cooling method with niobium alloy, Ni-based superalloy and ceramic based composite have been used to this day.

Formability Evaluation of a Copper Alloy for Regenerative Cooling Thrust Chamber

The dome stretching tests and tension tests have been performed to obtain a forming limit curve(FLC) for the copper alloy which is used for manufacturing the regenerative cooling thrust chamber. For experimental investigation of the forming limit curve, we have used in-plane tension specimen to obtain tension-compression strain state as well as out-of-plane specimen to obtain tension-tension strain state through dome stretching test. All specimens were divided into longitudinal and transverse directions according to the orientation of test specimen. The test results showed that in the tension-tension region, copper alloy revealed a maximum major strain of 62.3% and a maximum minor strain of 58.6%. In the tension-compression region, the maximum major strain and the maximum minor strain were measured to be 60.5% and 25.8%, respectively.

Development of 2-ton thrust-level sub-scale calorimeter

A calorimeter of 2-ton thrust level rocket engine chamber has been developed to measure the wall heat flux. The liner of the chamber is made of copper-chromium alloy to maximize the heat transfer performance and structural strength. 1-D design code based on empirical correlations has been used for the prediction of the global thermal characteristics while 3-D CFD has been applied for the verification of local cooling performance. The predicted average wall heat flux at the throat is 43 for the combustion chamber pressure of 53 bar. The chamber structure is confirmed to be safe at the pressure of 150 bar through 2-D stress analysis and measurement of the strain of the test species. Finally, the test of pressurizing the calorimeter chamber has been performed with water at the pressure of 150 bar in room temperature environment. No thermal damage has been detected after the hot-fire test in the test nozzle of same cooling performance with the developed calorimeter though the measured throat heat flux is higher than the design value by 10%.

Developing Trends of Spinning Process for Manufacturing Thrust Chamber of Launch Vehicle

ABSTRACT Spinning process is generally used for manufacturing axisymme trical, thin-walled thickness and hollow circular cross-section parts. Traditional spinning techn ology is classified to conventional spinning and power spinning(shear spinning and flow forming). L iterature surveys of spinning application for regenerative cooling chamber and divergent nozz le of liquid propellent rocket thrust chamber have been conducted. Most spinning technology has been used mandel for manufacturing chamber and nozzle. Recently, hot spinning has been used much c ompared to traditional cold spinning.초 록 스피닝 공정은 축대칭의 얇은 두께를 가지고 있는 속이 빈 실린더형 단면을 가지고 있는 부품에 일반적으로 사용된다. 전통적인 스피닝 제작 기술은 컨벤셔널 스피닝과 파워 스피닝(전단 스피닝과 유동성형)으로 구분된다. 액체추진로켓의 연소기의 재생냉각 챔버와 확대노즐부에서 적용된 스피닝에 대한 문헌조사를 수행하였다. 연소실과 노즐의 제작에 사용되는 스피닝은 대부분 맨드럴을 사용하였다. 최근에는 전통적인 냉간 스피닝에 비해 열간 스피닝도 많이 사용되고 있었다.Key Words:Spinning(스피닝), Combustion Chamber(연소실), Divergent Nozzle(확대노즐), Liquid- propellant Rocket(액체추진로켓)Received 2 June 2015 / Revised 22 September 2015 / Accepted 29 September 2015

Experimental Study on the Physical and Mechanical Properties of a Copper Alloy for Liquid Rocket Combustion Chamber Application

Mechanical and physical properties of a copper alloy for a liquid rocket engine(LRE) combustion chamber liner application were tested at various temperatures. All test specimens were heat treated with the condition they might experience during actual fabrication process of the LRE combustion chamber. Physical properties measured include thermal conductivity, specific heat and thermal expansion data. Uniaxial tension tests were preformed to get mechanical properties at several temperatures ranging from room temperature to 600. The result demonstrated that yield stress and ultimate tensile stress of the copper alloy decreases considerably and strain hardening increases as the result of the heat treatment. Since the LRE combustion chamber operates at higher temperature over 400, the copper alloy can exhibit time-dependent behavior. Strain rate, creep and stress relaxation tests were performed to check the time-dependent behavior of the copper alloy. Strain rate tests revealed that strain rate effect is negligible up to 400 while stress-strain curve is changed at 500 as the strain rate is changed. Creep tests were conducted at 250 and 500 and the secondary creep rate was found to be very small at both temperatures implying that creep effect is negligible for the combustion chamber liner because its operating time is quite short.

Optimum Performance Analysis of KSR-III LRE

To understand the each performance parameter correlation of flight type liquid-propellant rocket engine for KSR-III(Korea Sounding Rocket-III), the analysis of engine stand-alone combustion test results was carried out. Considering the variation of ablative material combustion chamber caused by erosion, linear regression analysis that ignores oxidizer/fuel ratio effect and two-variable 2nd-order polynomial regression analysis that considers oxidizer/fuel ratio change were performed. It can be described that linear regression analysis is simple and very practical method, and can predict the performance within 1% error inside analyzed region. And two-variable 2nd-order polynomial regression analysis can predict with very high accuracy inside region and shows that KSR-III engine`s optimum oxidizer/fuel ratio for thrust(or specific impulse) is 2.22 and that for combustion chamber pressure(or characteristic velocity) is 2.17.