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: 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 turbine airfoil usable in a turbine engine with a cooling system and a compliant dual wall configuration configured to enable thermal expansion between inner and outer layers while eliminating stress formation in the outer layer is disclosed. The compliant dual wall configuration may be formed a dual wall formed from inner and outer layers separated by a support structure. The outer layer may be a compliant layer configured such that the outer layer may thermally expand and thereby reduce the stress within the outer layer. The outer layer may be formed from a nonplanar surface configured to thermally expand. In another embodiment, the outer layer may be planar and include a plurality of slots enabling unrestricted thermal expansion in a direction aligned with the outer layer.

Abstract: A turbine airfoil usable in a turbine engine with a cooling system and a compliant dual wall configuration configured to enable thermal expansion between inner and outer layers while eliminating stress formation is disclosed. The compliant dual wall configuration may be formed a dual wall formed from inner and outer layers separated by a compliant structure. The compliant structure may be configured such that the outer layer may thermally expand without limitation by the inner layer. The compliant structure may be formed from a plurality of pedestals positioned generally parallel with each other. The pedestals may include a first foot attached to a first end of the pedestal and extending in a first direction aligned with the outer layer, and may include a second foot attached to a second end of the pedestal and extending in a second direction aligned with the inner layer.

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 mounting arrangement (10) for a multivane segment (12) of ceramic matrix composite (CMC) composition positioned between outer and inner metallic rings (14, 16). Selected ones of the vanes (18a) of the multivane segment surround internal struts (24) joining the outer and inner rings. Spring members (26, 28) accommodate differential thermal growth between the multivane segment and the outer and inner rings, and a compliant material (30) seals against gas leakage around the segments.

Abstract: A bushing (30, 31) in a hole (26) through a ceramic matrix composite structure (20) with a flange (34, 38) on each end of the bushing (30, 31) extending beyond and around the hole and pressing against opposed surfaces (22,24) of the CMC structure (20) with a preload that resists buckling of the composite structure fibers and resists internal CMC fiber separation. A connecting element (40), such as a bolt or pin, passes through the bushing (30, 31) for engagement with a supporting element (50). The bushing (31) may be formed in place as a single piece of ceramic, and cured along with the CMC structure (20), or it may be formed as two ceramic or metal parts (32, 36) that are joined together and preloaded by threads (33). The connecting element (40) may be a pin, or it may be a bolt with a shaft threaded into one part (32) of the bushing and a head (42) that pushes the second flange (38) toward the first flange (34).

Abstract: A stacked laminate component for a turbine engine that may be used as a replacement for one or more metal components is provided. The stacked laminate component can have a body formed by a process of stacking and laminating layers to define a radially inner surface along the hot gas path. The layers can be substantially orthogonal to the radially inner surface. The layers can be at least a first layer of a first material and a second layer of a second material. At least the first material is a ceramic matrix composite. The second material can have a higher thermal conductivity or a higher creep strength than the first material.

Abstract: The interlocking of two or more sections via the insertion of one or more out-of-plane features in one section or combination of sections through a void in one or more a complementary sections can result in a reinforced ceramic matrix composite article upon securing of the sections. The sections can be secured by friction between two tightly matched sections or by the use of a pin, hook, or clamp. The sections can be constructed from one or more CMC laminate sheets. The out-of-plane feature can be a loop or a flange and the void can be an orifice or a matched loop in the complementary sheet. The securing of the sections can result in a CMC article where the delamination between sheets is inhibited.

Abstract: The interlocking of two or more sections via the insertion of one or more out-of-plane features in one section or combination of sections through a void in one or more a complementary sections can result in a reinforced ceramic matrix composite article upon securing of the sections. The sections can be secured by friction between two tightly matched sections or by the use of a pin, hook, or clamp. The sections can be constructed from one or more CMC laminate sheets. The out-of-plane feature can be a loop or a flange and the void can be an orifice or a matched loop in the complementary sheet. The securing of the sections can result in a CMC article where the delamination between sheets is inhibited.

Abstract: Aspects of the invention relate to a construction system and method for components in high temperature environments, such as the hot gas path components of a turbine engine. Such a component can include a skeleton and a coating. The skeleton can be formed by a plurality of interconnected frame members, which can give the component its general shape. The frame members can be made of ceramic matrix composite. A coating can be provided around at least a portion of the skeleton. Preferably, the coating is a refractory material, such as refractory ceramic. Examples of turbine engine components that can be constructed according to aspects of the invention are airfoils with or without platforms, blade rings, combustor tiles and heat shields. A component according to aspects of the invention can be made using low cost fabrication and construction methods.

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: A ceramic ring segment for a turbine engine that may be used as a replacement for one or more metal components. The ceramic ring segment may be formed from a plurality of ceramic plates, such as ceramic matrix composite plates, that are joined together using a strengthening mechanism to reinforce the ceramic plates while permitting the resulting ceramic article to be used as a replacement for components for turbine systems that are typically metal, thereby taking advantage of the properties provided by ceramic materials. The strengthening mechanism may include a bolt or a plurality of bolts designed to prevent delamination of the ceramic plates when in use by keeping the ceramic plates in compression.

Abstract: A ceramic matrix composite (CMC) material (10) with increased interlaminar strength is obtained without a corresponding debit in other mechanical properties. This is achieved by infusing a diffusion barrier layer (20) into an existing porous matrix CMC to coat the exposed first matrix phase (19) and fibers (12), and then densifying the matrix with repeated infiltration cycles of a second matrix phase (22). The diffusion barrier prevents undesirable sintering between the matrix phases and between the second matrix phase and the fibers during subsequent final firing and use of the resulting component (30) in a high temperature environment.

Abstract: A ceramic ring segment for a turbine engine that may be used as a replacement for one or more metal components. The ceramic ring segment may be formed from a plurality of ceramic plates, such as ceramic matrix composite plates, that are joined together using a strengthening mechanism to reinforce the ceramic plates while permitting the resulting ceramic article to be used as a replacement for components for turbine systems that are typically metal, thereby taking advantage of the properties provided by ceramic materials. The strengthening mechanism may include a bolt or a plurality of bolts designed to prevent delamination of the ceramic plates when in use by keeping the ceramic plates in compression.

Abstract: A method to form an improved hybrid ceramic matrix composite structure comprises providing a ceramic matrix composite (CMC) with open features at a surface and applying a thermal barrier coating to the surface. The thermal barrier coating is preferably a friable graded insulation coating (FGI) having hollow ceramic spheres. For acceptance of a FGI coating the open features have a center-to-center separation distance in one or more directions that is between about 100 and about 1,000 percent of the diameter of the hollow ceramic spheres in the FGI and a depth that is between about 20 percent and about 200 percent of the diameter of the hollow ceramic spheres in the FGI. The open feature can result from the winding of tows about a mandrel; the lay-up of fabric including a surface fabric having an open weave; or the use of a 3-D weave, or a 3-D braid.

Abstract: A ceramic matrix composite (CMC) material (10) with increased interlaminar strength is obtained without a corresponding debit in other mechanical properties. This is achieved by infusing a diffusion barrier layer (20) into an existing porous matrix CMC to coat the exposed first matrix phase (19) and fibers (12), and then densifying the matrix with repeated infiltration cycles of a second matrix phase (22). The diffusion barrier prevents undesirable sintering between the matrix phases and between the second matrix phase and the fibers during subsequent final firing and use of the resulting component (30) in a high temperature environment.

Abstract: A hybrid ceramic matrix composite (CMC) structure 10 and method for fabricating such an structure are provided. A CMC substrate 12 includes layers 16, 18, 20 of ceramic fibers. Fugitive objects 22 are disposed on at least one of the plurality of layers prior to laying a subsequent layer of ceramic fibers. An outer surface of the subsequent layer influences a shape of the outer surface of the substrate by defining protuberances 24 on the outer surface of the substrate where respective cavities 26 are formed beneath respective protuberances upon dissipation of the fugitives. A liquefied ceramic coating 34 is deposited on the outer surface of the ceramic substrate to fill the cavities. When the ceramic coating is cured to a solidified state, the cavities containing the solidified coating constitute an anchoring arrangement between the ceramic substrate and the ceramic coating.

Abstract: A hybrid ceramic matrix composite structure and method for fabricating such an structure are provided. A ceramic matrix composite substrate 12 includes a plurality of layers of ceramic fibers. A plurality of spaced apart imprints 22 is disposed in at least one of the plurality of layers. An outer surface of a subsequent layer disposed over the layer with the imprints to influence a texture of the outer surface of the substrate by defining a plurality of indent regions 32 on the outer surface of the substrate. A ceramic coating 14 is deposited on the surface of the substrate. The plurality of indent regions 32 constitutes a bond-enhancing arrangement between the surface of the substrate and a corresponding boundary of the coating.

Abstract: A ceramic ring segment for a turbine engine that may be used as a replacement for one or more metal components. The ceramic ring segment may be formed from a plurality of ceramic plates, such as ceramic matrix composite plates, that are joined together using a strengthening mechanism to reinforce the ceramic plates while permitting the resulting ceramic article to be used as a replacement for components for turbine systems that are typically metal, thereby taking advantage of the properties provided by ceramic materials. The strengthening mechanism may include a ceramic matrix composite overwrap or plurality of overwraps designed to help prevent delamination of the ceramic plates when the ceramic article is in use by placing the plates in compression.