PSI - Issue 60

Sarath Chandran Nair S. et al. / Procedia Structural Integrity 60 (2024) 564–574 Sarath Chandran Nair S./ Structural Integrity Procedia 00 (2023) 000 – 000

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1. Introduction In a semi-cryogenic rocket engine, combustion of fuel and oxidiser mixture takes place in the combustion chamber or thrust chamber. The combustion temperature in a semi-cryogenic engine is approximately 3700K at which most of the materials melt. Combustion chamber has to be cooled properly to take care of this problem. Most effective cooling method is regenerative cooling and it is suited for aerospace applications. This will ensure the inner wall temperature of the engine never cross the temperature limit of the material of engine construction especially the material used at the region where combustion takes place. This demands a double walled construction combustion chamber design with milled channel at the inner shell of a high thermal conductive material - high strength Copper alloy and brazed outer shell of high strength stainless steel to form the coolant channel passages. The ribs provided on the inner shell act as the load transfer member to the outer wall and as conductive fins to augment the heat transfer. A special protective coating called Thermal Barrier Coating (TBC) is provided on the inner wall of combustion chamber of semi-cryogenic engine to enhance its reusability. This coating is applied on the inner wall of the engine at convergent and a certain portion of divergent. A detailed thermal analysis is carried out to map the temperature distribution in the engine. This will be an input for the coating design. The designed coating undergoes various tests for evaluating its integrity for its reusability. Thermal barrier coating is designed for semi cryogenic engine for one of the developmental launch vehicles of Indian Space Research Organisation. The coating forms an integral part of the inner wall and subject to a complex pattern of loads involving both thermal and pressure loads. A detailed literature survey is carried out on thermal barrier coating design and its integrity evaluation. Details of different types of TBC implemented in Gas turbine and Aero engines are given in technical/journal papers by Nitin P. Padture et al. (2002) and Brian Gleeson (2006). Design of TBC is given by Kunal Mondal et al. and Mohit Gupta (2014). Details of durability assessment and life prediction of TBC are given by Robert Eriksson (2013) and Vishnu Sankar (2014). A failure analysis of Plasma sprayed TBC is carried out by NASA (1984). A life prediction model for thermal barrier coating was developed by J. DeMasi et al. (1989). From the literature survey, could not find any technical paper related to design and qualification of thermal barrier coating in rocket engine combustion/thrust chamber. Thermal barrier coating discussed in the above said papers are not similar to semi-cryogenic rocket engine which is discussed here. This paper gives the details of finite element analysis carried out to determine the stresses and strains on the coating under thermal and pressure loads and tests carried out to determine the coating integrity through specimen level testing. Test specimens are prepared with thermal barrier coating as in the actual thrust chamber of semi-cryogenic engine. These specimens are tested under tensile and compressive loading that will induce the expected strains obtained from finite element analysis. The integrity of the coating is assessed by means of Non-Destructive Test after each cycle of loading and based on the observation, modifications were made for better performance. The modified coating was also tested for the same load conditions by which coating integrity is evaluated. Tests conducted in TBC coated specimen give a lot of information during thermal and mechanical loads. Thermal cycling is done to simulate the thermal load whereas mechanical load is applied to simulate pressure load condition. Specimen design, test method, qualification and its results are discussed in detail in the following sections. 2. Necessity of Thermal Barrier Coating in combustion chamber of a semi-cryogenic engine In the semi-cryogenic engine, which is being developed by Indian Space Research Organisation, Isrosene – a kerosene derivative is used as fuel and liquid Oxygen (LOX) is used as oxidiser. Combustion occurs as fuel and oxidiser mix at the combustion chamber. The combustion temperature is approximately 3700 K at which most of the materials will melt. Combustion chamber has to withstand this high temperature during the entire engine operation. In addition to this, combustion chamber has also to withstand pressure and thrust loads generated inside the chamber during engine operation. To address these critical points, combustion chamber has to be cooled efficiently and effectively. Most effective cooling method is a regenerative cooling combined with film cooling. This method is adopted to manage the chamber inner wall temperature within the bearable limit of the material used for realizing the combustion chamber. To enable regenerative cooling, combustion chamber is designed as double walled construction as shown in fig.1.