Abstract: A multiple laser sight system for an archery bow or the like configured so that the multiple laser systems can be calibrated together and having features such that the user can use one laser system during the day and one laser system during the light. The laser sight is further configured to not interfere with the optional use of conventional sighting pins and the use of evening infrared systems, like the prior art use of night vision goggles.

Abstract: A multiple laser sight system for an archery bow or the like configured so that the multiple laser systems can be calibrated together and having features such that the user can use one laser system during the day and one laser system during the light. The laser sight is further configured to not interfere with the optional use of conventional sighting pins and the use of evening infrared systems, like the prior art use of night vision goggles.

Abstract: A multiple laser sight system for an archery bow or the like configured so that the multiple laser systems can be calibrated together and having features such that the user can use one laser system during the day and one laser system during the light. The laser sight is further configured to not interfere with the optional use of conventional sighting pins and the use of evening infrared systems, like the prior art use of night vision goggles.

Abstract: A multiple laser sight system for an archery bow or the like configured so that the multiple laser systems can be calibrated together and having features such that the user can use one laser system during the day and one laser system during the light. The laser sight is further configured to not interfere with the optional use of conventional sighting pins and the use of evening infrared systems, like the prior art use of night vision goggles.

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 method for forming interphase layers in ceramic matrix composites. The method forms interphase layers in ceramic matrix composites thereby enabling higher matrix densities to be achieved without sacrificing crack deflection and/or toughness. The methods of the present invention involve the use fugitive material-coated fibers. These fibers are then infiltrated with a ceramic matrix slurry. Then, the fugitive material is removed and the resulting material is reinfiltrated with an interphase layer material. The ceramic matrix composite is then fired. Additional steps may be included to densify the ceramic matrix or to increase the strength of the interphase layer. The method is useful for the formation of three dimensional fiber-reinforced ceramic matrix composites envisioned for use in gas turbine components.

Abstract: A method for forming interphase layers in ceramic matrix composites. The method forms interphase layers in ceramic matrix composites thereby enabling higher matrix densities to be achieved without sacrificing crack deflection and/or toughness. The methods of the present invention involve the use fugitive material-coated fibers. These fibers are then infiltrated with a ceramic matrix slurry. Then, the fugitive material is removed and the resulting material is reinfiltrated with an interphase layer material. The ceramic matrix composite is then fired. Additional steps may be included to densify the ceramic matrix or to increase the strength of the interphase layer. The method is useful for the formation of three dimensional fiber-reinforced ceramic matrix composites envisioned for use in gas turbine components.

Abstract: Embodiments of the invention relate to various cooling systems for a turbine vane made of stacked ceramic matrix composite (CMC) laminates. Each airfoil-shaped laminate has an in-plane direction and a through thickness direction substantially normal to the in-plane direction. The laminates have anisotropic strength characteristics in which the in-plane tensile strength is substantially greater than the through thickness tensile strength. Such a vane construction lends itself to the inclusion of various cooling features in individual laminates using conventional manufacturing and forming techniques. When assembled in a radial stack, the cooling features in the individual laminates can cooperate to form intricate three dimensional cooling systems in the vane.

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 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: 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 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).

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: Embodiments of the invention relate to a robust turbine vane made of stacked airfoil-shaped CMC laminates. Each laminate has an in-plane direction and a through thickness direction substantially normal to the in-plane direction. The laminates have anisotropic strength characteristics in which the in-plane tensile strength is substantially greater than the through thickness tensile strength. Thus, the laminates can provide strength in the direction of high thermal gradients and, thus, withstand the associated high thermal stresses. The laminates are relatively weak in through thickness (interlaminar) tension, but, in operation, relatively low through thickness tensile stresses can be expected. The laminates can be strong in through thickness compression; accordingly, the laminate stack can be held in through thickness compression by one or more fasteners.

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 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.

Abstract: A joining method for assembling components with complex shapes from CMC elements of simpler shapes. A first CMC element (30) is fabricated and fired to a selected first cured state. A second CMC element (36) is fabricated and left in a green state, or is fired to a second partially cured state that is less complete than that of the first cured state. The two CMC elements (30, 36) are joined in a mating interface that captures an inner joining portion (38) of the second element (36) within a surrounding outer joining portion (32) of the first element (30). The assembled elements (30, 36) are then fired together, resulting in differential shrinkage that compresses the outer joining portion (32) onto the inner joining portion (38), providing a tightly pre-stressed joint. Optionally, a refractory adhesive (42) may be used in the joint. Shrinkage of the outer joining portion (32) avoids shrinkage cracks in the adhesive (42).

Abstract: An airfoil (30) having a continuous layer of ceramic matrix composite (CMC) material (34) extending from a suction side (33) to a pressure side (35) around a trailing edge portion (31). The CMC material includes an inner wrap (36) extending around an inner trailing edge portion (38) and an outer wrap (40) extending around an outer trailing edge portion (42). A filler material (44) is disposed between the inner and outer wraps to substantially eliminate voids in the trailing edge portion. The filler material may be pre-processed to an intermediate stage and used as a mandrel for forming the outer trailing edge portion, and then co-processed with the inner and outer wraps to a final form. The filler material may be pre-processed to include a desired mechanical feature such as a cooling passage (22) or a protrusion (48).

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: A composite structure (62) having a bond enhancement member (76) extending across a bond joint (86) between a ceramic matrix composite (CMC) material (80) and a ceramic insulation material (82), and a method of fabricating such a structure. The bond enhancement member may extend completely through the CMC material to be partially embedded in a core material (84) bonded to the CMC material on an opposed side from the insulation material. A mold (26) formed of a fugitive material having particles (18) of a bond enhancement material may be used to form the CMC material. A two-piece mold (38, 46) may be used to drive a bond enhancement member partially into the CMC material. A compressible material (56) containing the bond enhancement member may be compressed between a hard tool (60) and the CMC material to drive a bond enhancement member partially into the CMC material.