The invention relates to a turbine engine separate flow mixer centered on a longitudinal axis, comprising an exhaust housing, a shroud directly connected to the exhaust housing and intended to mix the flows originating in the turbine engine, said shroud comprising a metal sheet formed by a succession of first and second longitudinal strips distributed circumferentially around the longitudinal axis of the mixer by circumferentially placing the second strips on either side of the first strips, the first and second strips being configured to form the shroud when at rest and grooves in the shroud when operating, the grooves being defined by an alternation of internal lobes and of external lobes.

A ring assembly for holding a double-skin combustor liner has an inner skin of ceramic matrix composite and an outer skin. The assembly comprises an annular end body, a first projecting portion projecting relative to the annular end body at least partially in an axial direction, the first projecting portion configured for supporting the inner skin, a second projecting portion projecting relative to the annular end body at least partially in the axial direction, the second projecting portion configured for supporting the outer skin. The first and second projecting portions are radially spaced apart to provide a radial spacing between the inner skin and the outer skin. An expansion joint is between the ring assembly and the inner skin configured to permit motion resulting from a variation in thermal expansion between the inner skin and the ring assembly.

A cooling structure includes: a housing including an inner cylindrical portion through which a shaft is inserted; a coolant flow path formed in the housing and opened on one side of the housing in a rotational axis direction, the coolant flow path being located radially outside the inner cylindrical portion; a lid member disposed in an opening of the coolant flow path, the lid member located radially outside the inner cylindrical portion and adjacent to the coolant flow path; a first end portion on the lid member, the first end portion contacting, in the radial direction, an inner circumferential surface of the coolant flow path on one side in the radial direction; and a second end portion on the lid member, the second end portion contacting, in the rotational axis direction, an abutment surface on the housing on the other side in the radial direction.

An attachment system for an exhaust component is disclosed. In various embodiments, the attachment system includes a radial attachment flange of the exhaust component; and a radial ring having at least one of a radially outer surface configured for engagement with a radially inner surface of the radial attachment flange or a radially inner surface configured for engagement with a radially outer surface of the radial attachment flange.

An airfoil includes a ceramic matrix composite airfoil core that defines an airfoil portion and a root portion. The ceramic matrix composite airfoil core is subject to core thermal growth. A platform includes a ceramic matrix composite that wraps around the root portion. The platform is subject to platform thermal growth. A buffer layer is located between the root portion and the platform. The buffer layer absorbs a mismatch between the core thermal growth and the platform thermal growth.

Combustion sections and sealing systems for fuel ignition assemblies of gas turbine engine combustion sections are provided. For example, a sealing system comprises a ferrule positioned on an outer surface of a ceramic matrix composite (CMC) combustor liner an aperture defined in the CMC liner; a sleeve positioned within an adapter of the fuel ignition assembly such that an inner end portion of the sleeve is in contact with the ferrule, the sleeve having an end wall that forms an inner boundary of a cavity defined by the sleeve; and a biasing member positioned within the cavity. The biasing member extends between a bushing and the end wall of the sleeve. The biasing member continuously urges the sleeve into contact with the ferrule to seal the aperture against fluid leakage therethrough. The exemplary sealing system may be part of a fuel ignition assembly of a gas turbine engine combustion section.

A turbine shroud assembly includes a carrier assembly, a blade track assembly, and a first biasing member. The carrier assembly includes a carrier segment having an outer wall, a first mount flange and a second mount flange having a chordal seal. The blade track assembly includes a blade track segment and a first mount pin. The first biasing member is arranged axially between and engaged with a retainer plug of the first mount pin and a pin segment of the first mount pin such that a portion of the pin segment engages an attachment feature of the blade track segment so as to bias the attachment feature into engagement with the chordal seal of the second mount flange.

An aircraft engine includes a core gas path through which a core gas flow flows, an exhaust duct receiving the core gas flow, an air-cooled heat exchanger disposed in a heat exchanger duct having an air inlet providing cooling air to the air-cooled heat exchanger and an air outlet in fluid communication with the exhaust duct, and a cover disposed in the heat exchanger duct downstream of the air-cooled heat exchanger. The cover is movable between an open position, in which the cover allows the cooling air to flow through the air outlet into the exhaust duct, and a closed position, in which the cover substantially blocks the air outlet. The cover is operable to move from the open position to the closed position at a predetermined temperature of the cover.

An assembly adapted for use with a gas turbine engine includes a static component and a metering band. The static component is fixed relative to an axis. The metering band is arranged to extend circumferentially at least partway about the axis and is coupled with the static component. The metering band defines at least a portion of a cooling passageway for air to flow through.

A turbine vane assembly adapted for use in a gas turbine engine includes a support strut and a vane. The support strut is made of metallic materials. The vane is made of composite materials and is arranged around the support strut to insulate the metallic materials of the support strut during use of the turbine vane assembly.