Abstract: A gas turbine engine composite vane assembly and method for making same are disclosed. The method includes providing at least two gas turbine engine airfoil composite preform components. The airfoil composite preform components are interlocked with a first locking component so that mating faces of the airfoil composite preform components face each other. A filler material is inserted between the mating surfaces of the airfoil composite preform components.

Abstract: A blade and method for producing the blade for a gas turbine engine are described herein. The blade may include a composite airfoil. The airfoil may comprise a ceramic material, and a distal end. A tip may extend from the distal end of the airfoil. The tip of the airfoil may comprise a substantially porous structure and may comprise infiltrated material extending from an airfoil preform to a tip preform to join the airfoil preform and the tip preform.

Abstract: A blade and method for producing the blade for a gas turbine engine are described herein. The blade may include a composite airfoil. The airfoil may comprise a ceramic material, and a distal end. A tip may extend from the distal end of the airfoil. The tip of the airfoil may comprise a substantially porous structure and may comprise infiltrated material extending from an airfoil preform to a tip preform to join the airfoil preform and the tip preform.

Abstract: A gas turbine engine composite vane assembly and method for making same are disclosed. The method includes providing at least two gas turbine engine airfoil composite preform components. The airfoil composite preform components are interlocked with a first locking component so that mating faces of the airfoil composite preform components face each other. A filler material is inserted between the mating surfaces of the airfoil composite preform components.

Abstract: A blade for a gas turbine engine comprises a blade portion having a first end and a second end and an engagement portion including a first surface coupled to the second end of the blade portion and a second surface coupled to the second end of the blade portion, the first and second surfaces arranged to extend divergently away from one another. The engagement portion is adapted for coupling to a wheel included in a gas turbine engine wheel.

Abstract: A gas turbine engine is disclosed with a turbine section having at least one turbine rotor with a plurality of turbine blades, a plurality of blade tracks positioned circumferentially around the turbine blades, at least one dovetail shaped connecting member extending radially outward from each blade track, and a hanger connected to a structural member of the gas turbine engine and configured to releasably couple with the at least one dovetail shaped connecting member of a corresponding blade track.

Abstract: A turbine shroud or blade track assembly adapted to extend around a turbine wheel assembly is disclosed. The turbine shroud includes a carrier and a blade track coupled to the carrier. The blade track is movable between a radially-inward position having a first inner diameter and a radially-outward position having a second inner diameter larger than the first inner diameter.

Abstract: A turbine shroud for a gas turbine engine includes a carrier and a blade track segment. The carrier is formed to include an inwardly opening blade track channel and the blade track segment is positioned in the blade track channel to couple the blade track segment with the carrier.

Abstract: A vane assembly for a gas turbine engine is disclosed in this paper. The vane assembly includes an inner platform, an outer platform, and a ceramic-containing airfoil. The ceramic-containing airfoil extends from the inner platform to the outer platform. A clamp mechanism couples the inner platform and the outer platform to the ceramic-containing airfoil.

Abstract: A blade and method for producing the blade for a gas turbine engine are described herein. The blade may include a composite airfoil. The airfoil may comprise a ceramic material, and a distal end. A tip may extend from the distal end of the airfoil. The tip of the airfoil may comprise a substantially porous structure and may comprise infiltrated material extending from an airfoil preform to a tip preform to join the airfoil preform and the tip preform.

Abstract: A gas turbine engine composite vane assembly and method for making same are disclosed. The method includes providing at least two gas turbine engine airfoil composite preform components. The airfoil composite preform components are interlocked with a first locking component so that mating faces of the airfoil composite preform components face each other. A filler material is inserted between the mating surfaces of the airfoil composite preform components.

Abstract: An apparatus and method for controlling turbine blade tip clearance is disclosed herein. A blade track assembly can include a blade track carrier having a plurality of slots and rails defined by paths that vary in radial position as a function of circumferential location. An expansion ring can be operably coupled with the slots of the blade track carrier. A plurality of blade track segments can be operably coupled with the expansion ring and engageable with the rails of the blade track carrier such that expansion and contraction of the expansion ring causes radially outward movement and radially inward movement, respectively, of the blade track segments.

Abstract: A turbine shroud or blade track assembly adapted to extend around a turbine wheel assembly is disclosed. The turbine shroud includes a carrier and a blade track coupled to the carrier. The blade track is movable between a radially-inward position having a first inner diameter and a radially-outward position having a second inner diameter larger than the first inner diameter.

Abstract: A technique is directed to launching an unmanned aerial vehicle (UAV). The technique involves positioning a UAV launcher over a ground location. The technique further involves installing, after the UAV launcher is positioned over the ground location and prior to launching the UAV, an anchor into the ground location to anchor the UAV launcher to the ground location. The technique further involves operating, after the UAV launcher is anchored to the ground location, the UAV launcher to impart launching force onto the UAV to launch the UAV, the anchor holding the UAV launcher substantially in place at the ground location to minimize energy loss as the UAV launcher imparts launching force onto the UAV. In some arrangements, the UAV launcher is capable of pivoting while remaining anchored between launches to accommodate changes in wind direction while maintaining substantial connection to the ground location for enhanced consistency and performance.

Abstract: A gas turbine engine airflow member including a blade core portion, a shroud tip portion extending from the blade core portion, and an airfoil portion formed exteriorly to the blade core portion, where the blade core portion and the shroud tip portion are constructed as a first unitary structure and the airfoil portion is constructed as a second structure. A method of forming a gas turbine engine component is also disclosed.

Abstract: A gas turbine engine is disclosed with a turbine section having at least one turbine rotor with a plurality of turbine blades, a plurality of blade tracks positioned circumferentially around the turbine blades, at least one dovetail shaped connecting member extending radially outward from each blade track, and a hanger connected to a structural member of the gas turbine engine and configured to releasably couple with the at least one dovetail shaped connecting member of a corresponding blade track.

Abstract: A gas turbine engine is disclosed with a turbine section having at least one turbine rotor with a plurality of turbine blades, a plurality of blade tracks positioned circumferentially around the turbine blades, at least one dovetail shaped connecting member extending radially outward from each blade track, and a hanger connected to a structural member of the gas turbine engine and configured to releasably couple with the at least one dovetail shaped connecting member of a corresponding blade track.

Abstract: In one embodiment, a gas turbine engine component includes a foam based core and a composite skin member. Both the foam based core and the composite skin member can be used to structurally support the gas turbine engine component. The composite skin member can be a CMC material and is used to partially encapsulate the foam core. The gas turbine engine component can take the form of an airfoil member such as a blade or a vane, a combustor liner, etc. A first portion of the composite skin member includes a first surface extending past an edge of the component creating a step approximate an edge section. In another embodiment, composite skin members can be used to form a continuous shape for the edge section such that the foam core forms part of a gas path surface.

Abstract: The present disclosure provides for a turbine wheel having a blade platform disposed to compliantly secure ceramic turbine blades to a rotatable disk. The platform includes opposing ends to engage a portion of an airfoil of each turbine blade and radial extensions to engage a portion of a root of each turbine blade.

Abstract: One aspect of the present disclosure includes a ceramic matrix composite airfoil having a shaped ceramic matrix composite core, a cooling fluid flow network having ceramic foam components, and a ceramic lamina skin disposed about the shaped core and the cooling fluid flow network.