Abstract: A turbine component includes an airfoil-shaped core having an outer peripheral surface. A plurality of channels is formed in the core, each opening to the outer peripheral surface. The channels extend substantially radially to be elongated in the radial direction. A first platform is attached to the core. The component also includes an airfoil-shaped non-permeable skin having a hollow interior, an inner peripheral surface and an outer peripheral surface. The skin is sized so that the core can be received in the hollow interior of the skin. A second platform is attached to the skin. The core is received in the hollow interior of the skin such that the outer peripheral surface of the core engages the inner peripheral surface of the skin such that a plurality of generally radial cooling channels are formed between the channels in the core and the inner peripheral surface of the skin.

Abstract: A turbine component includes an airfoil-shaped core having an outer peripheral surface. A plurality of channels is formed in the core, each opening to the outer peripheral surface. The channels extend substantially radially to be elongated in the radial direction. A first platform is attached to the core. The component also includes an airfoil-shaped non-permeable skin having a hollow interior, an inner peripheral surface and an outer peripheral surface. The skin is sized so that the core can be received in the hollow interior of the skin. A second platform is attached to the skin. The core is received in the hollow interior of the skin such that the outer peripheral surface of the core engages the inner peripheral surface of the skin such that a plurality of generally radial cooling channels are formed between the channels in the core and the inner peripheral surface of the skin.

Abstract: A turbine airfoil usable in a turbine engine and formed from at least an outboard section and an inboard section such that an inner end of the outboard section is attached to an outer end of the inboard section. The outboard section may be configured to provide a tip having adequate thickness and may extend radially inward from the tip with a generally constant cross-sectional area. The inboard section may be configured with a tapered cross-sectional area to support the outboard section.

Abstract: A turbine airfoil usable in a turbine engine and formed from at least an outboard section and an inboard section such that an inner end of the outboard section is attached to an outer end of the inboard section. The outboard section may be configured to provide a tip having adequate thickness and may extend radially inward from the tip with a generally constant cross-sectional area. The inboard section may be configured with a tapered cross-sectional area to support the outboard section.

Abstract: A turbine airfoil has an airfoil-shaped core and a skin extending about the outer peripheral surface of the core. The core and/or the skin are configured such that a plurality of cooling channels is formed between them. In one embodiment, the cooling channels are formed by a plurality of radially extending channels in the core. In another embodiment, the cooling channels are formed by providing a wave-shaped skin such that, when the skin is attached to the outer peripheral surface of the core, the cooling channels can be formed in the space between the outer peripheral surface of the core and each wave of the skin. Such airfoil constructions can allow the inclusion of different materials in the airfoil. Further, such constructions can allow greater flexibility in the formation of cooling channels. Moreover, such constructions can help to achieve thinner outer walls and near surface cooling of the airfoil.

Abstract: An airfoil (10) as may be used in a gas turbine engine includes a ceramic matrix composite (CMC) element (12) that extends to define a leading edge portion (14) and chord portion (16) of the airfoil, and a separately formed but conjoined trailing edge element (18) that defines a desirably thin trailing edge of the airfoil without the need for using an excessively small bend radius for reinforcing fibers in the CMC element. The trailing edge element may include a plurality of interlock elements (26) that extend through the trailing edge attachment wall (22) of the CMC element and provide mechanical attachment there between. Alternatively, the trailing edge element may be adhesively bonded or sinter bonded to the CMC element. A cooling air insert (60) may be disposed within a cooling air cavity (64) of the CMC element and may include cooling tubes (66) that extend into the trailing edge element to deliver cooling air there through.

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: Aspects of the invention are directed to a multi-layer ring seal segment that can incorporate a plurality of material systems. The ring seal segment can include an inner layer, a central layer and an outer layer. The inner layer can be attached to one side of the central layer, and the outer layer can be attached to an opposite side of the central layer. The inner and outer layers can be made of a ceramic matrix composite, such as a hybrid oxide ceramic matrix composite or an oxide-oxide ceramic matrix composite. The central layer can be made of a material that has high shear strength relative to the inner and outer layers. The ring seal segment according to aspects of the invention can take advantage of the benefits of the different materials so as to better withstand the operational loads of the turbine.

Abstract: An airfoil (10) as may be used in a gas turbine engine includes a ceramic matrix composite (CMC) element (12) that extends to define a leading edge portion (14) and chord portion (16) of the airfoil, and a separately formed but conjoined trailing edge element (18) that defines a desirably thin trailing edge of the airfoil without the need for using an excessively small bend radius for reinforcing fibers in the CMC element. The trailing edge element may include a plurality of interlock elements (26) that extend through the trailing edge attachment wall (22) of the CMC element and provide mechanical attachment there between. Alternatively, the trailing edge element may be adhesively bonded or sinter bonded to the CMC element. A cooling air insert (60) may be disposed within a cooling air cavity (64) of the CMC element and may include cooling tubes (66) that extend into the trailing edge element to deliver cooling air there through.

Abstract: Aspects of the invention are directed to a multi-layer ring seal segment that can incorporate a plurality of material systems. The ring seal segment can include an inner layer, a central layer and an outer layer. The inner layer can be attached to one side of the central layer, and the outer layer can be attached to an opposite side of the central layer. The inner and outer layers can be made of a ceramic matrix composite, such as a hybrid oxide ceramic matrix composite or an oxide-oxide ceramic matrix composite. The central layer can be made of a material that has high shear strength relative to the inner and outer layers. The ring seal segment according to aspects of the invention can take advantage of the benefits of the different materials so as to better withstand the operational loads of the turbine.

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 crack-resistant vane segment assembly member for an industrial turbine engine is disclosed. The vane segment includes at least one internally-cooled body portion and inner and outer shroud portions. The body portions include end regions characterized by blending and transition zones that ensure heat is transferred at a non-crack-inducing rate from the shroud portions to the body portions. The vane segment assembly also includes a cooling arrangement that reduces the impact of thermal gradient-induced stresses acting at the interface between the body portions and shroud portions. The cooling arrangement cooperates with the transition and blending zones to produce synergistically-enhanced crack resistance properties.

Abstract: A crack-resistant vane segment assembly member for an industrial turbine engine is disclosed. The vane segment includes at least one internally-cooled body portion and inner and outer shroud portions. The body portions include end regions characterized by blending and transition zones that ensure heat is transferred at a non-crack-inducing rate from the shroud portions to the body portions. The vane segment assembly also includes a cooling arrangement that reduces the impact of thermal gradient-induced stresses acting at the interface between the body portions and shroud portions. The cooling arrangement cooperates with the transition and blending zones to produce synergistically-enhanced crack resistance properties.

Abstract: A turbo-machine having a means for attaching the blade ring to the casing at a plurality of alternative radial positions to provide for a plurality of alternative clocking angles for a stage of stationary airfoils. The blade ring is provided with a plurality of notches that may selectively be aligned with a pin inserted through the casing. By rotating the blade ring so that a selected notch aligns with the pin, a selected clocking position may be achieved. Assembly with a different clocking angle is achieved by aligning the pin with a different notch. The notches may be provided at any location around the blade ring so long as they are separated by a radial distance which is a non-integer multiple of the segment angle.