A double-skin liner for a combustion chamber of a gas turbine. The double-skin liner includes an inner layer, an outer layer with an expansion gap provided on the outer layer, a gap covering element, a sealing element, a plurality of internal cooling holes provided on the inner layer, and a plurality of external cooling holes provided on the outer layer. When the inner layer expands due to temperature increase, a proximal section of the outer layer and a distal section of the outer layer distance from each other through lengthening the expansion gap.

A gas turbine engine and combustor assembly are provided, the combustor assembly including a first liner and a second liner together defining at least in part a combustion chamber, wherein the first liner and the second liner are separated by a gap along the longitudinal direction, and wherein the first liner is forward of the second liner relative to a flow of fluid through the combustion chamber along the longitudinal direction, and wherein the gap is extended along the circumferential direction.

A seal assembly to seal a gas turbine hot gas path flow at an interface of a combustor liner and a downstream component, such as a stage one turbine nozzle, in a gas turbine. The seal assembly including a piston ring seal housing, defining a cavity, and a piston ring disposed within the cavity. The piston ring disposed circumferentially about the combustor liner. The piston ring is responsive to a regulated pressure to secure sealing engagement of the piston ring and outer surface of the combustor liner. The seal assembly includes at least one of one or more sectional through-slots, bumps or channel features to provide for a flow therethrough of a high-pressure (P.sub.high) bypass airflow exiting a compressor to the cavity. The high-pressure (P.sub.high) bypass airflow exerting a radial force on the piston ring.

A leakage management assembly for an orifice configured to limit a flow of at least one of a leakage fluid and a cooling fluid past an instrument positioned within the orifice are provided. The leakage management assembly includes a labyrinth ferrule including an annular ferrule body having a central bore extending therethrough. The labyrinth ferrule further includes an annular assembly cone formed at a first end of the labyrinth ferrule. The annular assembly cone includes an annular convergent lip and is configured to facilitate directing the instrument into the central bore. The labyrinth ferrule also includes a plurality of labyrinthine annular restrictions extending from a radially inner surface of the annular body toward the central bore. A combustor and a gas turbine engine including such a seal assembly are also provided.

A combustion liner assembly includes metal liner having a hot side and a cold side, a ceramic matrix composite (CMC) liner tile configured to provide a heat shield for the metal liner, the CMC liner tile having a different thermal conductivity than the metal liner, and a connection device configured to attach the CMC liner tile to the metal liner. The connection device accommodates the different thermal conductivity of the CMC liner tile and the metal liner. The connection device is free from a radial fastener exposed to hot gases on the hot side of the metal liner, and the connection device allows radial movement and axial movement between the metal liner and the CMC liner tile.

There are provided transition piece cooling holes which make NOx reduction and combustion performance improvement possible while effectively cooling the transition piece end frame and the first-stage stator vane end wall. The transition piece cooling holes include a transition piece which guides combustion gas from a combustor to a turbine, a transition piece end frame which is installed on a turbine-side outlet of the transition piece and is disposed so as to face a first-stage stator vane end wall of the turbine with a predetermined gap being interposed, and a seal member which is fitted on the transition piece end frame and is fitted into the first-stage stator vane end wall so as to seal cooling air which is supplied into the gap. The cooling holes are made in the transition piece end frame so as to directly supply the cooling air to the first-stage stator vane end wall.

An integrated combustor nozzle includes an inner liner segment; an outer liner segment; and a panel extending radially between the inner and outer liner segments. The panel includes a forward end, an aft end, and a side walls extending axially from the forward end to the aft end. The aft end defines a turbine nozzle having a trailing edge circumferentially offset from the forward end. The inner liner segment has a pair of sealing surfaces, each of which defines a first continuous curve in the circumferential direction. The outer liner segment has a pair of sealing surfaces, each of which defines a second continuous curve in the circumferential direction. In some instances, the curves are monotonic in the circumferential direction. A segmented annular combustor including an array of such integrated combustor nozzles is also provided.

A method of manufacturing a nested combustor liner includes manufacturing a nested combustor liner into a green state including a plurality of annular interior walls radially adjacent to one another and circumferentially surrounding an exhaust duct aperture and a plurality of annular exterior walls radially adjacent to one another and radially spaced apart from and circumferentially surrounding the plurality of annular interior walls and an ignitor wall attached to a first annular interior wall at a first interior end, extending radially toward and attached to a first annular exterior wall at a first exterior end. The method includes assembling the plurality of annular interior walls and the plurality of annular exterior walls, forming an assembled combustor liner. The method includes densifying the assembled combustor liner.

An assembly is provided that includes a slider seal having a washer and a tubular puck. The washer includes an inner washer surface and an outer washer surface. The inner washer surface extends around an axis of the slider seal and has a non-circular cross-sectional geometry. The inner washer surface forms a washer bore axially through the washer. The outer washer surface extends around the axis and has a cross-sectional geometry with a shape that is different than a shape of the non-circular cross-sectional geometry. The tubular puck projects axially through the washer bore and includes an outer puck surface. The outer puck surface extends around the axis. The outer puck surface is sealingly engaged with and configured to slide axially along the inner washer surface.

A turbomachine includes a combustion chamber with a radially extending downstream flange at its downstream end. A distributor is disposed downstream of the combustion chamber and includes a platform from which at least one vane extends radially. The platform has an upstream edge extending radially and delimiting, with the downstream flange disposed opposite, a gap opening into the combustion chamber at its radially inner end and closed at its radially outer end by sealing means fixed to the distributor. The downstream flange of the combustion chamber has at least one rectilinear cooling orifice passing through the flange and opening into the gap opposite the platform of the distributor.