A vane assembly for use in a gas turbine engine includes an airfoil, a strut, and a support member. The airfoil is configured to interact with gases flowing through a primary gas path of the engine. The strut is located in an interior region of the airfoil and configured to carry loads that act on the airfoil by the gases. The support member is configured to compensate for relative thermal expansion between the components of the vane assembly caused by heat generated during use of the engine to locate the airfoil radially in the primary gas path.

The disclosure relates to a component for a turbine engine with a heated airflow and a cooling airflow. The component includes a wall separating the heated airflow from the cooling airflow. The wall can have a heated surface confronting the heated airflow and a cooled surface confronting the cooling airflow. The component can also include a baffle with a set of cooling holes.

A gas turbine engine comprising: a shaft about an axis; a first turbine assembly mounted to the shaft, a first flow path and a second flow path extending through first turbine assembly along the axis. The second flow path is located radially inward of the first flow path relative to the axis. A second turbine assembly is about the axis downstream of the first turbine assembly, with a gap defined between the first turbine assembly and the second turbine assembly, the gap in fluid communication with the first flow path and the second flow path. A washer is downstream of the first turbine assembly. The washer has an annular body including a deflector between the first turbine assembly and the second turbine assembly, the deflector obstructing the first flow path and extending toward the second flow path.

A pneumatically variable nozzle vane is disclosed that is capable of performing the same or similar function as a mechanically variable nozzle vane. Within its core, each pneumatically variable nozzle vane may comprise one or more cavities in fluid communication with one or more outlets to eject a gas from the nozzle vane into a flow path of working fluid through the nozzle. Each cavity may be shaped to match an internal pressure gradient to the external pressure gradient of the nozzle vane. The gas may be ejected as a curtain, substantially perpendicular to the flow path through the nozzle, to thereby manipulate the flow of a working fluid through the nozzle in a similar manner as a mechanically variable nozzle vane. In an embodiment, each nozzle vane may have two cavities supplying outlets on both the pressure-side and suction-side of the nozzle vane.

There is provided apparatuses and methods for a gas turbine engine. The embodiments include a core section with a flow path. The flow path includes a stator vane array having outlet vanes. Each outlet vane has a trailing edge. A strut has a leading edge that is upstream of the trailing edges. Alternatively, the flow path includes a stator vane array having inlet vanes. Each inlet vane has a leading edge. The strut has a trailing edge that is downstream of the leading edges. There also is increased spacing between adjacent vanes and rotor blades.

An insert system for an airfoil plenum includes a first insert and a second insert that include a plurality of impingement openings defined therein. The first insert includes a forward-facing inlet opening. The second insert includes a neck portion having a radial-facing inlet opening, an aft opening, and a cavity in flow communication between the radial-facing inlet opening and the aft opening. The second insert is sized for insertion into the plenum radially through a plenum inlet such that the neck portion is positioned in the plenum inlet. The first insert is sized for insertion into the second insert radially through the radial-facing inlet opening. When the neck portion is positioned in the plenum inlet, the first insert is moveable aftward through the aft opening into an installed position such that the forward-facing inlet opening opens into the cavity.

There is disclosed a fan assembly including a fan rotor including a hub and fan blades. The fan blades have a leading edge and a trailing edge. A fan stator downstream of the fan rotor relative to a direction of an airflow through the fan assembly. The fan stator includes vanes extending between radially inner ends and radially outer ends. A flow recirculation circuit has an inlet downstream of radially inner ends of the vanes of the fan stator and an outlet upstream of radially inner ends of the vanes. A recirculation stator has a plurality of stationary guide vanes circumferentially distributed around the axis and located in the flow recirculation circuit between the inlet and the outlet A method of operating the fan assembly is also disclosed.

There is disclosed a vane segment for a gas turbine engine. The vane segment includes vanes extending radially from inner ends to outer ends. The vanes have leading edges, trailing edges, and chords extending from the leading edges to the trailing edges. A platform circumferentially extends around the central axis. The platform has an inner face facing the central axis and an opposed outer face. The inner face of the platform is connected to the outer ends of the vanes. Ribs protrude away from the outer face and have lengths extending along the outer face.

Component for gas turbine engines are described. The components include an airfoil body having leading and trailing edges and pressure and suction sides. The airfoil has a leading edge cavity located proximate the leading edge defined between the leading edge and a separator rib and between the pressure side and the suction side. An insert member is installed within the leading edge cavity. The insert member has one or more metering flow apertures at an aft end and one or more impingement apertures at a forward end and at least one axially extending rib along an exterior surface thereof. At least one axial extending flow channel passage is defined along the axial extending rib between an exterior of the insert member and an interior of the airfoil body. Air flow through the metering flow apertures flows into the axial extending flow channel passage and flows forward toward the leading edge.

A turbine blade includes a root and an airfoil extending from a base, through which it is connected to the root, to a tip. The airfoil includes an intrados wall and an extrados wall, connected by a leading edge and by a trailing edge with a cooling circuit. The cooling circuit includes a conduit with a duct and a manifold prolonging this conduit. The conduit collects air from the blade root to supply the duct and the manifold that is located downstream from the duct and that supplies the slits in the trailing edge with air. The duct supplies air to one end of the manifold close to the tip. The manifold is separated from the duct by a partition including a portion close to the tip that is curved to be concave when seen from the trailing edge.