Abstract: A process for manufacturing a composite rotor blade includes manufacturing an oversized root region of a root region of a composite rotor blade; fixturing the CMC blade into a machining fixture at a primary Y and Z axis datum located at an attachment fillet radii of the root region; machining V-notches into the oversized root region to form a Y? and Z? axis datum of a sacrificial datum system in relation to the primary Y and Z axis datum; applying an oversized coating layer over the attachment fillet radii of the root region; fixturing the CMC blade into a machining fixture at the Y? and Z? axis datum of the sacrificial datum system; machining the oversized coating layer to a machined coating layer forming a Y? and Z? axis datum with respect to the Y? and Z? axis datum of the sacrificial datum system; fixturing the CMC blade into a machining fixture at the Y? and Z? axis datum; and machining off the sacrificial datum system removing the V-notches.

Abstract: A rotor blade for a gas turbine engine including an inner ply layer group includes a flared region and an airfoil region, a neck region between the flared region and the airfoil region; and a platform shell that comprises a platform shell root region, a platform shell platform region, and a platform shell neck region between the platform shell root region and the platform shell platform region, the platform shell root region sheathes the flared region, and the platform shell neck region flares outwardly away from the neck region to intersect the platform shell platform region at a platform shell neck region perimeter around an airfoil opening perimeter.

Abstract: A pre-diffuser for a gas turbine engine includes an exit guide vane ring having a multiple of exit guide vanes defined around an engine longitudinal axis; a hot fairing structure adjacent to the exit guide vane ring to define a multiple of diffusion passages around the engine longitudinal axis; an outer radial interface between a radial outer surface of the hot fairing structure and the exit guide vane ring, the outer radial interface being a full hoop structure; and an anti-rotation feature between the hot fairing structure and the exit guide vane ring, the anti-rotation features inboard of the multiple of diffusion passages.

Abstract: A semi-autonomous rapid response active clearance control system for a gas turbine engine includes an outer case, a carrier for a blade outer airseal positioned radially inward of the outer case, a blade outer airseal positioned radially inward of and mounted to the carrier, at least one electromechanical actuator attached to the outer case and selectively operable to move the carrier, and a sensing system. The sensing system includes a maneuver sensor housed within the actuator and configured to output a first signal, a position sensor mounted proximate to the blade outer airseal and configured to output a second signal, and a control means in communication with the maneuver sensor and the position sensor.

Abstract: An airfoil component includes an insert that has angled faces joined at an edge that provides an airfoil trailing edge. An outer CMC fiber layer overlaps the angled faces to provide a trailing edge portion of an airfoil.

Abstract: A vane includes a ceramic vane piece that defines first and second vane platforms and a hollow airfoil section that joins the first and second vane platforms. A structural platform adjacent the first platform defines a radial seal slot. A seal has a radial leg and an axial leg. The radial leg extends in the radial seal slot and the axial leg extends in an interface between the structural platform and the first vane platform.

Abstract: One exemplary embodiment of this disclosure relates to a system having a static segment with a circumferentially extending slot. The system further includes a retention clip partially received in the slot, and partially contacting a circumferential end of the static segment.

Abstract: A method of manufacturing a hydrostatic advanced low leakage seal configured to be disposed between relatively rotatable components is provided. The method includes manufacturing a plurality of separate, independent components of the seal. The method also includes assembling the plurality of separate, independent components to form the seal.

Abstract: Disclosed is a weaving method including placing a first section of a fill fiber between warp fibers, forming a pick, moving a base to reposition the warp fibers, and placing a second section of the fill fiber between the warp fibers to form a woven structure, wherein at least a portion of the warp fibers are introduced to the woven structure using a weave control grid. Also disclosed is a weaving assembly including a base, a base positional controller, a weave control grid, warp fiber arms, a warp fiber arm positional controller and a fill fiber wand.

Abstract: Disclosed is a ceramic matrix component having a fibrous core and a ceramic matrix composite shell surrounding at least a portion of the fibrous core. The ceramic matrix composite shell comprises a fibrous preform. The fibrous core has a greater porosity than the fibrous preform. A method of making the ceramic matrix component is also disclosed.

Abstract: Disclosed is a ceramic matrix component having a fibrous core and a ceramic matrix composite shell surrounding at least a portion of the fibrous core. The fibrous core has a three dimensional braided structure and cooling passages. A method of making the ceramic matrix component is also disclosed.

Abstract: A rotor blade for a gas turbine engine including an inner ply layer group that at least partially defines a base of a root region; a platform around the inner ply layer group that at least partially defines the base; and a platform over-wrap around the platform, the platform over-wrap at least partially defines the base.

Abstract: A gas turbine engine component includes a ceramic matrix composite (CMC) body that includes an interior surface that defines a blind cavity. A barrier coating is disposed on an exterior of the CMC body. A ceramic body in the blind cavity lines at least a portion of the interior surface of the CMC body. The ceramic body defines an open, interconnected network of pores.

Abstract: 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.

Abstract: A rotor disk assembly for a gas turbine engine includes a rotor disk that defines an axis; and a damper pivot support that extends from a rim of the rotor disk, the damper pivot support comprising an aperture that defines a pivot axis parallel to the axis. A method to dampen a rotor blade of a gas turbine engine, includes independently pivoting a first damper and a second damper of a damper assembly about the pivot axis to respectively contact a first blade and a second blade in response to centrifugal force below a platform of the respective first blade and second blade.

Abstract: A vane ring for a gas turbine engine includes a contoured stop that extends from the static flowpath wall, the contoured stop being of an airfoil shape such that a first side of the contoured stop matches a portion of a first side of the first variable vane along a chord length when the first variable vane is pivoted about the longitudinal axis to a first position.

Abstract: A flow path component assembly includes a carrier that has at least one radially extending tab. A flow path component has a base portion and a first wall that extends radially from the base portion. The first wall has an aperture. A pin extends through a hole in the radially extending tab and the aperture. The pin has an ellipsoid portion between a first end and a second end.

Abstract: An interface for a gas turbine engine. The interface includes a full-hoop vane ring around the engine axis, the full-hoop vane ring comprises a forward vane rail with a vane ring forward contact surface and a vane ring anti-rotation tab, the vane ring anti-rotation tab engaged with the anti-rotation case slot; and a multiple of BOAS segments around the engine axis, each of the multiple of BOAS segments comprise a BOAS aft engagement feature and a BOAS aft contact surface, the BOAS aft engagement feature engaged with the outer case and the anti-rotation case slot, the BOAS aft contact surface abuts the vane ring forward contact surface.

Abstract: A side plate seal assembly for a gas turbine engine includes a multiple of non-metallic side plate seals that are arranged about an axis of the gas turbine engine to form a full hoop seal, each of the multiple of side plate seals comprise a retention surface and a knife edge seal surface that extends at an angle therefrom.

Abstract: A side plate seal assembly for a gas turbine engine includes a multiple of non-metallic side plate seals that are arranged about an axis of the gas turbine engine to form a full hoop seal, each of the multiple of side plate seals comprise a retention surface and a knife edge seal surface that extends at an angle therefrom.