Abstract: A stitching geometry and method for selective interlaminar reinforcement of a CMC wall (20A). The CMC wall is formed of ceramic fiber layers (22) individually infused with a ceramic matrix, stacked, and at least partially cured. A row of holes is formed in the wall, and a ceramic fiber thread (25) is infused with a wet ceramic matrix and passed through the holes to form stitches (28, 30, 31). The stitches are then cured, causing them to shrink more than any remaining wall shrinkage, thus tensioning the stitches and compressing the wall laminae together. The stitches may have through-wall portions (30, 31) that are angled differently in different wall areas as a function of interlaminar shear over interlaminar tension, optimizing wall reinforcement locally depending on magnitude and direction of shear. Alternate rows of stitches (54, 56) may have offsets in a stitch direction (34) and/or different through-wall angles (A1, A2).

Abstract: Methodology and tooling arrangements for increasing interlaminar shear strength in a ceramic matrix composite (CMC) structure are provided. The CMC structure may be formed by a plurality of layers of ceramic fibers disposed between a top surface and a bottom surface of the composite structure. A plurality of surface recesses are formed on the surfaces of the structure. For example, each of the surfaces of the composite structure may be urged against corresponding top and bottom surfaces of a tool having a plurality of asperities. The plurality of surface recesses causes an out-of-plane sub-surface fiber displacement along an entire thickness of the structure, and the sub-surface fiber displacement is arranged to increase an interlaminar shear strength of the structure.

Abstract: A method of forming an internal combustion engine component having a multi-panel outer wall. The multi-panel outer wall has an inner panel (16) with an inner surface (18) and an outer surface (37). The inner panel outer surface (37) has discrete pockets (23) formed by integral structural ribs (38). Each pocket (23) has a film cooling hole (31) between the pocket (23) and the plenum (20). The method includes: forming dimples (40) in the intermediate panel (22), at least one dimple (40) having a nozzle (29); securing the intermediate panel (22) to the inner panel outer surface (37), thereby enclosing at least one pocket (23); and ensuring a respective dimple (40) having a nozzle (29) protrudes into a respective enclosed pocket (24) and a respective nozzle (29) is configured to direct a respective jet (35) of cooling fluid onto the inner panel outer surface within the respective enclosed pocket (23).

Abstract: A method of forming and/or assembling a multi-panel outer wall (14) for a component (12) in a machine subjected to high thermal stresses comprising providing such a component (12) that includes an inner panel wall (16) having an outer surface, and an array of interconnecting ribs (38) on the outer surface of the component (12). An intermediate panel (22) is provided and preferably preformed to a general outer contour of the component (12), and is positioned over the inner panel (16). An external pressure force is applied across a surface area of the intermediate panel (22) against the outer surface of the component (12) to contour the intermediate panel (22) according to a geometric configuration formed by the ribs (38) thereby forming cooling chambers (24) between the outer surface and ribs (38) of the component (12) and the intermediate panel (22).

Abstract: A ceramic matrix composite (CMC) structure 12 includes a plurality of layers (e.g., 16, 18, 20) of ceramic fibers. The CMC structure 12 further includes a plurality of spaced apart objects 22 on at least some of the plurality of layers along a thickness of the composite structure. The inclusion of the objects introduces an out-of-plane fiber displacement arranged to increase an interlaminar shear strength of the structure.

Abstract: An arrangement (10) for conveying combustion gas from a plurality of can annular combustors to a turbine first stage blade section of a gas turbine engine, the arrangement (10) including a plurality of interconnected integrated exit piece (IEP) sections (16) defining an annular chamber (18) oriented concentric to a gas turbine engine longitudinal axis (20) upstream of the turbine first stage blade section. Each respective IEP (16) includes a first flow path section (40) receiving and fully bounding a first flow from a respective can annular combustor along a respective common axis (22) there between, and delivering a partially bounded first flow to a downstream adjacent IEP section (42). Each respective IEP further includes a second flow path section (112) receiving a partially bounded second flow from an upstream adjacent IEP (66) and delivering at least part of the second flow to the turbine first stage blade section.

Abstract: A structural layer (30) may be bi-cast onto ligaments (62) extending from a porous cooling construction (20). The material of the structural layer may be optimized for high-temperature strength, while the material of the porous construction may be optimized for high thermal conductivity. A fugitive material (56) such as wax may be formed on the ligaments of the porous construction. A second fugitive material (58) such as ceramic may fill the remaining part of the porous construction. An investment casting shell (60) may be disposed around the porous construction and the fugitive materials. The first fugitive material may then be replaced with the material of the structural layer (30), and the second fugitive material may be removed to provide coolant paths (26). A second structural layer (52) may be bi-cast onto further ligaments (62) on a second side of the porous construction.

Abstract: A ceramic matrix composite wall structure (20A) constructed of interlocking layers (22A, 24A) of woven material with integral cooling channels (28A, 32A). The CMC layer closest to the hot gas path (41) contains internal cooling tubes (26A, 30A) protruding into a ceramic insulating layer (40A). This construction provides a cooled CMC lamellate wall structure with an interlocking truss core.

Abstract: A stem (34) extends from a second part (30) through a hole (28) in a first part (22). A groove (38) around the stem provides a non-threaded contact surface (42) for a ring element (44) around the stem. The ring element exerts an inward force against the non-threaded contact surface at an angle that creates axial tension (T) in the stem, pulling the second part against the first part. The ring element is formed of a material that shrinks relative to the stem by sintering. The ring element may include a split collet (44C) that fits partly into the groove, and a compression ring (44E) around the collet. The non-threaded contact surface and a mating distal surface (48) of the ring element may have conic geometries (64). After shrinkage, the ring element is locked onto the stem.

Abstract: A gas turbine CMC shroud plate (48A) with a vane-receiving opening (79) that matches a cross-section profile of a turbine vane airfoil (22). The shroud plate (48A) has first and second curved circumferential sides (73A, 74A) that generally follow the curves of respective first and second curved sides (81, 82) of the vane-receiving opening. Walls (75A, 76A, 77A, 78A, 80, 88) extend perpendicularly from the shroud plate forming a cross-bracing structure for the shroud plate. A vane (22) may be attached to the shroud plate by pins (83) or by hoop-tension rings (106) that clamp tabs (103) of the shroud plate against bosses (105) of the vane. A circular array (20) of shroud plates (48A) may be assembled to form a vane shroud ring in which adjacent shroud plates are separated by compressible ceramic seals (93).

Abstract: Methodology and tooling arrangements for increasing interlaminar shear strength in a ceramic matrix composite (CMC) structure are provided. The CMC structure may be formed by a plurality of layers of ceramic fibers disposed between a top surface and a bottom surface of the composite structure. A plurality of surface recesses are formed on the surfaces of the structure. For example, each of the surfaces of the composite structure may be urged against corresponding top and bottom surfaces of a tool having a plurality of asperities. The plurality of surface recesses causes an out-of-plane sub-surface fiber displacement along an entire thickness of the structure, and the sub-surface fiber displacement is arranged to increase an interlaminar shear strength of the structure.

Abstract: A cooling arrangement in a gas turbine system (120). The arrangement includes a plurality of flow network units (208) to transfer heat to cooling fluid, at least one unit including first (218), second (220), and third (222) flow sections between openings (64a) in a first wall (66) and an opening in a second wall (68) to pass cooling fluid through the walls. The first section includes first flow paths, between the openings in the first wall and the second section, extending to the second section. The third section includes third flow paths, between the second section and the opening in the second wall, to effect flow of cooling fluid. The second section includes one or more cooling fluid flow paths between the first section and the third section. The number of flow paths in the second section is fewer than the number of first flow paths and fewer than the number of third flow paths.

Abstract: A ceramic matrix composite (CMC) airfoil assembled from a pressure side wall (42) and a suction side wall (52) joined by interlocking joints (18, 19) at the leading and trailing edges (22, 24) of the airfoil to produce a tapered thin trailing edge. The trailing edge (24) is thinner than a combined thicknesses of the airfoil walls (42, 52). One or both of the interlocking joints (18, 19) may be formed to allow only a single direction of assembly, as exemplified by a dovetail joint. Each joint (18, 19) includes keys (44F, 54F, 56F, 46F) on one side and respective keyways (44K, 54K, 56K, 46K) on the other side. Each keyway may have a ramp (45) that eliminates indents in the airfoil outer surface that would otherwise result from the joint.

Abstract: A ceramic matrix composite (CMC) structure (50) with first (26) and second (28) CMC walls joined at an intersection (34) containing continuous fibers (53). A gusset (52) is formed in the intersection by an inward bending of some or all ceramic fibers (53) of the intersection, resulting in a diagonal brace between the first and second CMC walls. This creates a depression (54) or void (59) in the intersection. One or more ceramic reinforcement devices fill or span the depression to prevent distortion of the gusset. The reinforcement devices may include a ceramic filler (60) or core (61), a CMC rod or cord (56), and a CMC tape (62). The ceramic filler (60) may be continuous with a ceramic insulation layer (36) on an outer surface of the first CMC wall.

Abstract: 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.

Abstract: A turbine vane assembly includes an airfoil extending between an inner shroud and an outer shroud. The airfoil can include a substructure having an outer peripheral surface. At least a portion of the outer peripheral surface is covered by an external skin. The external skin can be made of a high temperature capable material, such as oxide dispersion strengthened alloys, intermetallic alloys, ceramic matrix composites or refractory alloys. The external skin can be formed, and the airfoil can be subsequently bi-cast around or onto the skin. The skin and the substructure can be attached by a plurality of attachment members extending between the skin and the substructure. The skin can be spaced from the outer peripheral surface of the substructure such that a cavity is formed therebetween. Coolant can be supplied to the cavity. Skins can also be applied to the gas path faces of the inner and outer shrouds.

Abstract: A ceramic matrix composite (CMC) anchor (20, 100) joining a metal substrate (40) and a ceramic thermal barrier (38). The CMC anchor extends into and interlocks with the ceramic barrier, and extends into and interlocks with the metal substrate. The CMC anchor may be a honeycomb (20) or other extending-into-and-interlocking geometry. A CMC honeycomb may be formed with first (22) and second (24) arrays of cells (26) with open distal ends (28) on respective opposite sides of a sheet (30). The cells may have walls (32) with transverse passages (36). A metal (40) may be deposited into the cells and passages on one side of the sheet, forming a metal substrate locked into the honeycomb. A ceramic insulation material (38) may be deposited into the cells and passages on the opposite side of the sheet, forming a layer of ceramic insulation locked into the honeycomb.

Abstract: A structure for use in high temperature applications is provided. The structure may include an inner ceramic matrix composite (CMC) material (12). At least a portion of this CMC material includes waves that define a first wavy surface (140 and an opposed second wavy surface (16). A ceramic insulation material (18) may be bonded with the first wavy surface and includes a distal surface (20) for exposure to a high temperature environment. A core material (22) is bonded with at least a portion of the second wavy surface. One or more cooling channels (24) are disposed in the core material. An outer CMC material (26) may be joined to a portion of the inner CMC material. The core material is a material different than a matrix material of the inner CMC material.

Abstract: A ceramic matrix composite (CMC) anchor (20, 100) joining a metal substrate (40) and a ceramic thermal barrier (38). The CMC anchor extends into and interlocks with the ceramic barrier, and extends into and interlocks with the metal substrate. The CMC anchor may be a honeycomb (20) or other extending-into-and-interlocking geometry. A CMC honeycomb may be formed with first (22) and second (24) arrays of cells (26) with open distal ends (28) on respective opposite sides of a sheet (30). The cells may have walls (32) with transverse passages (36). A metal (40) may be deposited into the cells and passages on one side of the sheet, forming a metal substrate locked into the honeycomb. A ceramic insulation material (38) may be deposited into the cells and passages on the opposite side of the sheet, forming a layer of ceramic insulation locked into the honeycomb.

Abstract: A ceramic matrix composite wall structure (20A) constructed of interlocking layers (22A, 24A) of woven material with integral cooling channels (28A, 32A). The CMC layer closest to the hot gas path (41) contains internal cooling tubes (26A, 30A) protruding into a ceramic insulating layer (40A). This construction provides a cooled CMC lamellate wall structure with an interlocking truss core.