Abstract: A method of bonding an insulation layer (26) to a CMC layer (22) by applying a compliant joining layer (24) between them in a series of steps (52-70) effective to bond the layers (22, 24, 26) with reduced differential shrinkage stresses during firing and reduced residual stresses. The CMC layer (22) and the compliant layer (24) may each be fired to an intermediate stage prior to applying the next layer (24 and 26 respectively), such that the compliant layer (24) has a remaining amount of curing shrinkage between that of the CMC layer (22) and the insulating layer (26) during a final firing stage. The insulation layer (26) may be a friable graded insulation (26F) cast as a composite of hollow ceramic spheres (26S) in a ceramic matrix (26M). The compliant layer (24) may form a checkerboard of cracks (72) oriented generally perpendicular to the layer surface that function to relax stress.

Abstract: An insulating material 14 adapted for use in a high temperature environment for coating a turbine component is provided. The insulating material comprises a plurality of geometric shapes 18. The insulating material further comprises a binder for binding together the geometric shapes. A plurality of discontinuous fibers is added to the binder. The discontinuous fibers are adapted to controllably affect one or more properties of the insulating material. For example, non-fugitive chopped fibers 50 may be added to affect a tensile strength property of the insulating material, and fugitive chopped fibers 52 may be added to affect a density property of the insulating material.

Abstract: A gas turbine ring segment (10) for use in gas turbine engines made from a ceramic matrix composite (CMC) material is disclosed. The ring segment includes a stacked multiplicity of CMC thin-sheet lamellae (25a, 25b) each comprising a peripheral surface collectively defining a cross-section profile of the ring segment. The lamellae collectively define a channel (11) formed in the center thereof for receiving a bow-tie member (27). The bow-tie member is disposed in the channel for holding together the stacked lamellae in a through thickness direction, and the in-plane strength of the bow-tie member is perpendicular to the in-plane strength of the lamellae. A stem portion (33) of the assembly may be further secured with a wrap (38) of CMC ribbon.

Abstract: A method of bonding an insulation layer (26) to a CMC layer (22) by applying a compliant joining layer (24) between them in a series of steps (52-70) effective to bond the layers (22, 24, 26) with reduced differential shrinkage stresses during firing and reduced residual stresses. The CMC layer (22) and the compliant layer (24) may each be fired to an intermediate stage prior to applying the next layer (24 and 26 respectively), such that the compliant layer (24) has a remaining amount of curing shrinkage between that of the CMC layer (22) and the insulating layer (26) during a final firing stage. The insulation layer (26) may be a friable graded insulation (26F) cast as a composite of hollow ceramic spheres (26S) in a ceramic matrix (26M). The compliant layer (24) may form a checkerboard of cracks (72) oriented generally perpendicular to the layer surface that function to relax stress.

Abstract: A pattern of depressions (36) in a sealing surface (34) on a CMC wall (32) of gas turbine ring segment (30) allows minimum clearance against turbine blades tips, and thus maximizes working gas sealing. An array of depressions (36) on the surface (34) increases abradability of the surface (34) by blade tip contact during zero clearance conditions and reduces blade tip damage. The depressions (36) are unconnected, preventing bypass of the working gas around the blade tips. A desired abradable surface geometry may be formed in a stacked laminate wall construction (40-43, 52) by staggered laminate edge profiles (50, 52) or by machining of depressions (36, 54) after construction.

Abstract: A gas turbine component (20) with a layer of ceramic material (22) defining a wear surface (21), in which an array of cross-shaped depressions (26A, 26B) formed in the wear surface (21) define a continuous labyrinth of orthogonal walls (28, 30) of the ceramic material, and reduce the area of the wear surface (21) by about 50%. Within a representative area (36) of the wear surface (21), the depressions (26A, 26B) provide a ratio of a lineal sum of depression perimeters (27) divided by the representative area (36) of the wear surface of least 0.9 per unit of measurement for improved abradability characteristics of the wear surface (21).

Abstract: A enhanced abradable friable graded insulator FGI results from the laser patterning of a coating where a series of top surfaces reside on a series of columns such that the walls of the columns are not significantly densified relative to the interior of the columns. Patterns can be generated where the columns are oriented independently normal to or at an acute angle to the top surfaces. The cross sections of the top surfaces are formed to conform to the average dimensions of the spheres of the FGI coating. The cross sections of the top surfaces can be more than 1.5 times the diameter of the spheres. Various patterns of top surfaces can be used including regular, random, quasiperiodic patterns. A gradient of abradability can be imposed on the coating.

Abstract: A ceramic matrix composite with an enhanced abradability has a patterned surface with an array of solid composite material and voids where the voids extend into but not through the composite. The flow of gas through the voids as the surface is traversed by an impinging component, such as a turbine blade tip, is inhibited by the shape and size of the voids which can be sealed by the passing blade tip. Separately or additionally the inhibition of gas flow can result from the filling of the voids with a ceramic material of higher abradability than the ceramic matrix composite.

Abstract: A ceramic matrix composite (CMC) is constructed from ceramic fabric sheets and a ceramic matrix to yield matrix rich zone of reduced dimensions between fabric sheets. This reduction in dimension decreases the occurrence of delamination between the fabric sheets. The fabric is optimized to reduce the size of matrix rich zones in the composite by increasing the number of fabric plies in a given thickness of composite.

Abstract: An insulated CMC structure (20A) formed of a CMC layer (22A), a thermal insulation layer (24A) applied to a front surface (30A) of the CMC layer (22A), and cooling channels (28A) formed along the interface (26A) between the CMC layer and the thermal insulation layer, thus directly cooling the thermally critical area of the interface. Embodiments include cooling channels in direct contact with both layers (FIG. 1); cooling channels in one layer and tangent to the other layer (FIGS. 4, 5 and 9); cooling channels in the CMC layer with an intervening wall (36D, 36E) that bulges into the thermal insulation layer for improved bonding thereof (FIGS. 6, 7); and cooling channels formed in ceramic tubes (38F of FIG. 8).

Abstract: A composite material (10) formed of a ceramic matrix composite (CMC) material (12) protected by a ceramic insulating material (14). The constituent parts of the insulating material are selected to avoid degradation of the CMC material when the two layers are co-processed. The CMC material is processed to a predetermined state of shrinkage before wet insulating material is applied against the CMC material. The two materials are then co-fired together, with the relative amount of shrinkage between the two materials during the firing step being affected by the amount of pre-shrinkage of the CMC material during the bisque firing step. The shrinkage of the two materials during the co-firing step may be matched to minimize shrinkage stresses, or a predetermined amount of prestress between the materials may be achieved. An aluminum hydroxyl chloride binder material (24) may be used in the insulating material in order to avoid degradation of the fabric (28) of the CMC material during the co-firing step.

Abstract: A thermal barrier layer (20) is formed by exposing an oxide ceramic material to a thermal regiment to create a surface heat affected zone effective to protect an underlying structural layer (18) of the material. The heat affected surface layer exhibits a lower strength and higher thermal conductivity than the underlying load-carrying material; however, it retains a sufficiently low thermal conductivity to function as an effective thermal barrier coating. Importantly, because the degraded material retains the same composition and thermal expansion characteristics as the underlying material, the thermal barrier layer remains integrally connected in graded fashion with the underlying material without an interface boundary there between.

Abstract: A ceramic article having improved interlaminar strength and a method of forming the article. The article may be a ceramic matrix composite article. The methods of forming the articles increase the interlaminar strength of the article by forming indentations in the article during processing. The indentations may be tabs that are formed such that they provide one or more beneficial features for ceramic articles, such as CMC articles and hybrid structures. The tabs may be any of a variety of shapes, orientations, spacings, and combinations. In an alternative embodiment, the indentations are formed by pulling one or more fibers from one side of the ceramic layer to the other side. The articles have increased surface area, which helps to increase the bonding strength between the ceramic layer and any thermal barrier coating layer and/or ceramic core in the ceramic article.

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.

Abstract: A method of manufacturing a hybrid structure (100) having a layer of CMC material (28) defining an interior passageway (24) and a layer of ceramic insulating material (18) lining the passageway. The method includes the step of casting the insulating material to a first thickness required for effective casting but in excess of a desired second thickness for use of the hybrid structure. An inner mold (14) defining a net shape desired for the passageway remains in place after the casting step to mechanically support the insulating material during a machining process used to reduce the thickness of the insulating material from the as-cast first thickness to the desired second thickness. The inner mold also provides support as the CMC material is deposited onto the insulating material. The inner mold may include a fugitive material portion (20) to facilitate its removal after the CMC material is formed.

Abstract: Aspects of the invention are directed to a ceramic matrix composite ring seal segment. The ring seal segment according to aspects of the invention includes a relatively simple body that is circumferentially curved. At least a portion of the hot gas path surface of the ring seal segment can be coated with a thermal insulating. material. In one embodiment, each ring seal segment can be operatively connected to a stationary support structure, such as by way of isolation rings. The ring seal segments and/or the isolation rings can be configured so as to restrain the ring seal segments in the axial, radial and/or circumferential directions. The ring seal segments can be attached to the isolation rings so that the support points act opposite the operating pressure loads. Thus, the ring seal segments carry these loads in compression, a strong direction of the CMC fibers.

Abstract: A hybrid structure (50) and method of manufacturing the same including a structural ceramic matrix composite (CMC) material (42) coated with a layer of ceramic insulating tiles (24). Individual ceramic tiles are attached to a surface (22) of a mold (20). The exterior surface (32) of the tiles may be subjected to a mechanical process such as machining with the mold in place to provide mechanical support for the tiles. A layer of CMC material is then applied to bond the tiles and the CMC material together into a hybrid structure. The mold may include a fugitive material portion (26) to facilitate removal of the mold when the hybrid structure has a complex shape. Tiles located in different regions of the structure may have different compositions and/or dimensions. The gaps between adjacent tiles may be filled from the outside before the CMC material is applied or from the inside after the mold is removed.

Abstract: An article comprising a ceramic material having a ceramic matrix composite backing adapted for use in a gas turbine engine is provided. The article comprises a structural ceramic material having a hot side facing toward a high temperature environment and a cold side facing away from the high temperature environment; and a ceramic matrix composite composition having a strength greater than the strength of the ceramic material attached to the back of the cold side of the ceramic material, whereby crack initiation and propagation are inhibited by the ceramic matrix composition to a greater degree than by the ceramic material.

Abstract: An airfoil (44) formed of a plurality of pre-fired structural CMC panels (46, 48, 50, 52). Each panel is formed to have an open shape having opposed ends (54) that are free to move during the drying, curing and/or firing of the CMC material in order to minimize interlaminar stresses caused by anisotropic sintering shrinkage. The panels are at least partially pre-shrunk prior to being joined together to form the desired structure, such as an airfoil (42) for a gas turbine engine. The panels may be joined together using a backing member (30), using flanged ends (54) and a clamp (56), and/or with a bond material (36), for example.

Abstract: A stacked ceramic matrix composite lamellate assembly (10) including shear force bearing structures (48) for resisting relative sliding movement between adjacent lamellae. The shear force bearing structures may take the form of a cross-lamellar stitch (50), a shear pin (62), a warp (90) in the lamellae, a tongue (104) and groove (98) structure, or an inter-lamellar sealing member (112), in various embodiments. Each shear force bearing structure secures a subset of the lamellae, with at least one lamella being common between adjacent subsets in order to secure the entire assembly.