A transition exit frame (10) for supporting a transition (12) extending downstream from a combustor (14) to a turbine assembly (16) in a turbine engine (18) and including one or more transition exit frame inserts (20) configured to reduce thermal distortion created during operation of the turbine engine (18) is disclosed. The transition exit frame (10) may be formed from one or more transition exit frame bodies (22). The transition exit frame body (22) may be formed from a first material (24) having a first coefficient of thermal expansion. The transition exit frame insert (20) may form at least a portion of the transition exit frame body (22). The transition exit frame insert (20) may be formed from a second material (26) having a second coefficient of thermal expansion that is different than the first coefficient of thermal expansion of the first material (24) to reduce distortion within the transition exit frame body (22) during operation of the turbine engine (18).

A ring seal apparatus for high temperature sealing includes a first ring including a pair of radial faces and a second ring including a second pair of radial faces, the second ring adapted to coact with the first ring. The first and second rings together define a pair of coacting mating faces. The mating faces are obliquely angled relative to the radial faces, such that each of the coacting mating faces is adapted to seal an interface of the two rings at an angle relative to their substantially parallel radial faces. The pair of coacting rings is adapted to seal a circumferential gap between a pair of components.

A heat shield assembly for an engine case of a gas turbine engine may include a heat shield having an annular shape. A first groove may be formed circumferentially along an inner surface of the heat shield. A support lock may have a second groove extending radially inward from a distal surface of the support lock. A retention ring may be configured to fit within the first groove of the heat shield and the second groove of the support lock.

A turbine vane assembly for use in a gas turbine engine includes an endwall, a flow path component, and a load-distribution system. The endwall is arranged around a central axis of the turbine vane assembly. The flow path component is configured to direct fluid flow through the turbine vane assembly. The load-distribution system is positioned between the endwall and the flow path component to distribute loads transmitted between the endwall and the flow path component.

A mounting arrangement for mounting a fluid cooler to an engine case of a gas turbine engine, the mounting arrangement comprises at least one fixture adapted to be mounted to the engine case and defining a circumferentially extending channel for receiving at least one circumferentially extending flange of a cooler body of the fluid cooler. The at least one flange being free to slide in the channel of the at least one fixture in a circumferential direction of the engine case. The fixture may further comprise a clamping arrangement to hold the fluid cooler while allowing the same to thermally grow.

Each fuel injector has a main nozzle and pilot nozzle. A main fuel manifold to supply fuel to the main nozzle of each fuel injector and pilot fuel manifold to supply fuel to the pilot nozzle of each fuel injector. The main fuel manifold includes plurality of flexible main fuel pipes. Also, plurality of cross-piece connectors including a stem and four arms. The stem of each connector mounted on one fuel injector. The first and third arms extend in opposite directions from the stem and second and fourth arms extend in opposite directions from the stem. Each fuel pipe interconnects the first arm of one connector with a third arm of an adjacent connector and each connector so the first and third arms are at an angle relative perpendicular to the axis of the annular combustion chamber casing so the main fuel manifold extends around the combustion chamber casing sinusoidally.

A turbine shroud assembly adapted for use in a gas turbine engine includes a blade track segment and a carrier. The blade track segment includes a runner that extends circumferentially partway around a central axis and an attachment post that extends radially from the runner. The carrier is coupled with the attachment post to support the blade track segment relative to the central axis.

A piston seal assembly for a gas turbine engine includes a seal composed of a nickel-based superalloy; a component in contact with the seal and defining a seal-counterface; and a coating on the seal at the seal-counterface, wherein the coating is a metal alloy binder phase and a hard particle phase distributed through the binder phase.

An assembly of at least two members. One of the members supports the other member, the assembly defining an axial direction, a radial direction, and a circumferential direction, an inner member of the at least two members being received radially inside an outer member of the at least two members, wherein the inner member and the outer member am attached to each other by a support arrangement, the support arrangement including at least one floating support assembly as a displaceable coupling between an inner member support point provided at the inner member and an outer member support point provided at the outer member. A displacement of support points in a radial direction results in an interrelated relative displacement of the support points in an axial direction end vice versa.

A heat shield assembly for an engine case of a gas turbine engine may include a heat shield and a support lock. The heat shield may have an annular shape. The heat shield may define an aperture extending through the heat shield. The support lock may have a tab extending radially outward from a distal surface of the support lock. The aperture in the heat shield may be configured to retain the tab of the support lock.