A turbo fan engine includes a housing that supports a compressor section. At least one compressor hub is within the compressor section. A compressor shaft is arranged in the compressor section. An epicyclic gear train is coupled to the compressor shaft. At least one compressor hub and the compressor shaft are coupled at a common attachment point. The epicyclic gear train includes a carrier. A sun gear and intermediate gears are arranged about and intermeshing with the sun gear. The intermediate gears are supported by the carrier. A baffle is supported relative to the carrier and includes a lubrication passage near at least one of the sun gear and intermediate gears for directing a lubricant on at least one of the sun gear and the intermediate gears.

The present disclosure provides a blade positioning and support system for a gas turbine engine including a blade, the blade having a root and a tip, with the root having a surface oriented away from the tip, the surface having a forward end and an aft end, the forward end projecting farther away from the tip than the aft end. Further, the present disclosure provides a blade receiver, the blade receiver having a face and a facet, with the face being oriented away from the facet, the face having a forward end and an aft end, the aft end projecting farther away from the facet than the forward end.

A thrust reverser cascade for a gas turbine engine is disclosed. The thrust reverser cascade may comprise a plurality of turning vanes. One or more of the turning vanes may comprise a core formed from a polymer and a metallic coating applied to at least a portion of an outer surface of the core. The metallic coating may comprise nickel or a nickel alloy.

A gas turbine rotor and blade include a root portion, a platform and airfoil portion arranged along a span direction of the rotor blade, the platform located between the root and airfoil portion. The platform has an upstream and downstream side, side faces which extend from upstream to downstream side, an axial groove in each side face extends perpendicular to the span direction with a minor component of extension in span direction. A radial groove in each side face extends towards the axial groove with a component of extension in span direction and a component of extension perpendicular to the span direction. The radial groove has a first end that shows away from the axial groove and a second end that shows towards the axial groove. The second end is located a distance from the axial groove forming a groove free section between the second end and axial groove.

A tie rod includes a gusset which extends between a rod and a base.

A turbine exhaust case (28) comprises a fairing (120) defining an airflow path through the turbine exhaust case, and a multi-piece frame (100) disposed through and around the fairing to support a bearing load. The multi-piece frame comprises an inner ring (104), an outer ring (102), and a plurality of strut bosses (106). The outer ring is disposed concentrically outward of the inner ring, and has open bosses (126) at strut locations. The plurality of radial struts pass through the vane fairing, are secured to the inner ring via radial fasteners (108), and are secured via non-radial fasteners (114) to the open boss.

One aspect of the present disclosure includes a turbine vane assembly comprising a vane made from ceramic matrix composite material having an outer wall extending between a leading edge and a trailing edge and between a first end and an opposing second end; an endwall made at least partially from a ceramic matrix composite material configured to engage the first end of the vane; and a retaining region including corresponding bi-cast grooves formed adjacent the first end of the vane and a receiving aperture formed in the endwall; wherein a bond is formed in the retaining region to join the vane and endwall together.

A geared architecture includes a carrier that has spaced apart walls with circumferentially spaced mounts that interconnect the walls. The mounts provide circumferentially spaced apart apertures between the mounts at an outer circumference of the carrier. A sun gear and intermediate gears are arranged about and intermesh with the sun gear. The intermediate gears are supported by the carrier. The intermediate gears extend through the apertures to intermesh with a ring gear. A baffle is supported relative to the carrier and includes a lubrication passage near at least one of the sun gear and intermediate gears for directing a lubricant on the at least one of the sun gear and intermediate gears. The lubrication passage includes a primary passage that extends laterally between the walls and first and second passages and is in communication with the primary passage and respectively terminates near the sun gear and intermediate gears.

A fan exit guide vane assembly for a gas turbine engine includes a plurality of guide vanes having a pressure side and a suction side extending between a leading edge and a trailing edge. The vanes further include a span extending between a root and tip with a stagger angle defined as an angle between a longitudinal axis parallel to an engine axis of rotation and a line connecting the leading edge and the trailing edge that is less than about 15.

An arrangement for a turbomachine is provided. The arrangement includes a vane for directing a hot gas during the operation of the turbomachine, a stator ring for securing the vane, a heat shield for protecting the stator ring from the hot gas flow wherein the heat shield is arranged in downstream direction of the hot gas flow in front of the stator ring characterized in that the heat shield comprises a plurality of channels formed therein for directing a cooling air.