A hybrid vane ( 50 ) for a gas turbine engine having a ceramic matrix composite (CMC) airfoil member ( 52 ) bonded to a substantially solid core member ( 54 ). The airfoil member and core member are cooled by a cooling fluid ( 58 ) passing through cooling passages ( 56 ) formed in the core member. The airfoil member is cooled by conductive heat transfer through the bond (( 70 ) between the core member and the airfoil member and by convective heat transfer at the surface directly exposed to the cooling fluid. A layer of insulation ( 72 ) bonded to the external surface of the airfoil member provides both the desired outer aerodynamic contour and reduces the amount of cooling fluid required to maintain the structural integrity of the airfoil member. Each member of the hybrid vane is formulated to have a coefficient of thermal expansion and elastic modulus that will minimize thermal stress during fabrication and during turbine engine operation.

Ceramic matrix composite gas turbine vane

A turbine shroud apparatus is provided for a gas turbine engine having a central axis. The apparatus includes: (a) an annular support member (14); (b) a turbine shroud (12) disposed in the support member (14), the shroud (12) being a continuous ring comprising a low-ductility material and having opposed flowpath and back surfaces, and opposed forward and aft ends; and (c) a spring (32) mounted between the support member (14) and the shroud (12) and arranged to resiliently urge the shroud (12) to a concentric position within the support member (14).

Mounting apparatus for low-ductility turbine shroud

An abrasive article comprising a first group including a plurality of shaped abrasive particles overlying a backing, wherein the plurality of shaped abrasive particles of the first group define a first non-shadowing distribution relative to each other

Abrasive particles having particular shapes and methods of forming such particles

A component according to an exemplary aspect of the present disclosure includes, among other things a wall and a vascular engineered lattice structure formed inside of the wall. The vascular engineered lattice structure defines a hollow vascular structure configured to communicate a fluid through the vascular engineered lattice structure. The vascular engineered lattice structure has at least one inlet hole and at least one outlet hole that communicates the fluid into and out of the hollow vascular structure.

Gas turbine engine component having vascular engineered lattice structure

An engine has a blade stage (10) and a circumferential array of blade outer air seal segments (66). A support ring (70) carries the blade outer air seal segments (66). The support ring (70) has a low-CTE member (72) and a high-CTE member (90) intervening between the blade outer air seal segments (66) and the low-CTE member (72).

Blade clearance control using high-CTE and low-CTE ring members

An attachment method and flange for connecting a ceramic matrix composite (CMC) component, such as a gas turbine shroud ring ( 36, 68 ), to a metal support structure. A CMC flange ( 20 A) may be formed by attaching a wedge-shaped block ( 26 ) of a ceramic material to a CMC wall structure ( 22 ), and wrapping CMC layers ( 24 ) of the wall structure ( 22 ) at least partly around the block ( 26 ), forming the flange ( 20 A) with an inner oblique face ( 34 ) and an outer face ( 35 ) normal to the wall structure. An adjacent support structure, such as a metal support ring ( 40 A), may abut the outer face ( 35 ) of the CMC flange ( 20 A) and be clamped or bolted to the CMC flange ( 20 A).

Ceramic matrix composite attachment apparatus and method

A component ( 10 ) for a gas turbine engine formed of a stacked plurality of ceramic matrix composite (CMC) lamellae ( 12 ) supported by a metal support structure ( 20 ). Individual lamellae are supported directly by the support structure via cooperating interlock features ( 30, 32 ) formed on the lamella and on the support structure respectively. Mating load-transferring surfaces ( 34, 36 ) of the interlock features are disposed in a plane ( 44 ) oblique to local axes of thermal growth ( 38, 40 ) in order to accommodate differential thermal expansion there between with delta alpha zero expansion (DAZE). Reinforcing fibers ( 62 ) within the CMC material may be oriented in a direction optimized to resist forces being transferred through the interlock features. Individual lamellae may all have the same structure or different interlock feature shapes and/or locations may be used in different groups of the lamellae. Applications for this invention include an airfoil assembly ( 10 ) and a ring segment assembly ( 82 ).

Lamellate CMC structure with interlock to metallic support structure

Aspects of the invention are directed to a ring seal segment having a base portion and a frame portion. The base portion and the frame portion are a unitary structure. The frame portion is made of four unitary side walls: a forward wall, an aft wall and a pair of transverse side walls. The side walls can be arranged in a rectangular configuration. At least a part of the base portion can be coated with a thermal insulating material. A ring seal segment according to aspects of the invention is well suited to withstand the expected operational loads in the turbine section. The transverse side walls can provide bending strength to the base and can strengthen the forward and aft side walls. Transition regions between the side walls and the base portion are strengthened by three perpendicular planes that cooperate to provide structural strength to the ring seal segment.

Ceramic matrix composite turbine engine components with unitary stiffening frame

A CMC airfoil (20) formed with CMC stitches (37) interconnected between opposed walls (26, 28) of the airfoil to restrain outward flexing of the walls resulting from pressurized cooling air within the airfoil. The airfoil may be formed of a ceramic fabric infused with a ceramic matrix and dried, and may be partially to fully cured. Then holes (32, 34) are formed in the opposed walls of the airfoil, and a ceramic stitching element such as ceramic fibers (36) or a ceramic tube (44) is threaded through the holes. The stitching element is infused with a wet ceramic matrix before or after threading, and is flared (38) or otherwise anchored to the walls (26, 28) to form a stitch (37) there between. The airfoil and stitch are then cured. If the airfoil is cured before stitching, a pre-tension is formed in the stitch due to relative curing shrinkage.

Hollow CMC airfoil with internal stitch

A turbine airfoil (22A) is formed by a first process using a first material. A platform (30A) is formed by a second process using a second material that may be different from the first material. The platform (30A) is assembled around a shank (23A) of the airfoil. One or more pins (36A) extend from the platform into holes (28) in the shank (23A). The platform may be formed in two portions (32A, 34A) and placed around the shank, enclosing it. The two platform portions may be bonded to each other. Alternately, the platform (30B) may be cast around the shank (23B) using a metal alloy with better castability than that of the blade and shank, which may be specialized for thermal tolerance. The pins (36A-36D) or holes for them do not extend to an outer surface (31) of the platform, avoiding stress concentrations.

Jay A. Morrison

Papers

Ceramic matrix composite attachment apparatus and method

Lamellate CMC structure with interlock to metallic support structure

Ceramic matrix composite turbine engine components with unitary stiffening frame

Hollow CMC airfoil with internal stitch

Turbine Engine Airfoil and Platform Assembly