A gas turbine engine generally includes a fan and a core arranged in flow communication with one another. Additionally, the core of the gas turbine engine general includes, in serial flow order, a compressor section, a combustion section, a turbine section, and an exhaust section. In operation, air is provided from the fan to an inlet of the compressor section where one or more axial compressors progressively compress the air until it reaches the combustion section. Fuel is mixed with the compressed air and burned within the combustion section to provide combustion gases. The combustion gases are routed from the combustion section to the turbine section. The flow of combustion gasses through the turbine section drives the turbine section and is then routed through the exhaust section, e.g., to atmosphere.
More commonly, non-traditional high temperature materials, such as ceramic matrix composite (CMC) materials, are being used as structural components within gas turbine engines. For example, given an ability for CMC materials to withstand relatively extreme temperatures, there is particular interest in replacing components within the combustion section of the gas turbine engine with CMC materials. For example, typical combustion sections include an inner liner, an outer liner, and a dome together defining a combustion chamber. More commonly, at least the inner and outer liners are being formed of CMC materials.
The inventors of the present disclosure have found that it may be also beneficial to form other components of the combustor assembly of CMC materials. However, presently problems exist with joining multiple components defining the combustion chamber of CMC materials. Accordingly, a combustor assembly capable of utilizing multiple components formed of CMC materials would be useful. Specifically, a combustor assembly capable of effectively joining multiple components formed of CMC materials would be particularly beneficial.