Siemens Power Generation, Inc.

Westinghouse Electric Corporation

Siemens Westinghouse Power Corporation

Siemens Energy, Inc.

Daniel George Thompson

Jay Alan Morrison

Gary Brian Merrill

Jay Edgar Lane

Steven James Vance

Yevgeniy P Shteyman

Michael Anthony Burke

Harry A Albrecht

Christian X Campbell

Douglas A Keller

Marc Tertilt

Daniel Hofmann

Chris Campbell

Eric V Carelli

Joseph H Hoffman

Stefan Hoffmann

Christine Taut

Vijay Parthasarathy

Siu-Kee Chan

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.

Stacked laminate gas turbine component

A thermal barrier layer () 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 () 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. This invention is particularly advantageous when embodied in an apparatus formed of an oxide-oxide ceramic matrix composite (CMC) material wherein reinforcing fibers () are anchored in the underlying load-carrying portion and extend into the non-structural thermal barrier portion to provide support and to function as surface crack arrestors. In one embodiment an airfoil () is formed of a stacked plurality of CMC plates having such a heat-affected thermal barrier layer formed thereon.

In-situ formed thermal barrier coating for a ceramic component

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.

Ceramic material having ceramic matrix composite backing and method of manufacturing

A component () for a gas turbine engine formed of a stacked plurality of ceramic matrix composite (CMC) lamellae () supported by a metal support structure (). Individual lamellae are supported directly by the support structure via cooperating interlock features () formed on the lamella and on the support structure respectively. Mating load-transferring surfaces () of the interlock features are disposed in a plane () oblique to local axes of thermal growth () in order to accommodate differential thermal expansion there between with delta alpha zero expansion (DAZE). Reinforcing fibers () 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 () and a ring segment assembly ().

Lamellate CMC structure with interlock to metallic support structure

A stacked ceramic matrix composite lamellate assembly () including shear force bearing structures () for resisting relative sliding movement between adjacent lamellae. The shear force bearing structures may take the form of a cross-lamellar stitch (), a shear pin (), a warp () in the lamellae, a tongue () and groove () structure, or an inter-lamellar sealing member (), 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.

Stacked lamellate assembly

Embodiments of the invention relate to a turbine vane having a fail safe cooling system. According to embodiments of the invention, the vane can have multiple concentric layers of radial cooling holes extending about the vane; each layer being fluidly connected to the adjacent layer or layers. Such fluid communication can occur through one or more plenums in the vane or in the shrouds bounding the radial ends of the vane. Coolant can initially be supplied to the innermost layer of cooling holes. From there, the coolant can sequentially progress through successive outer layers. Between two adjacent layers, the coolant can flow in opposite directions. Not only does such a system provide needed cooling to the vane, but the multilayer redundant cooling system can avoid or delay catastrophic failures that can occur if the vane surface is damaged, such as by impact.

Fail safe cooling system for turbine vanes

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. The CMC material can permit the inclusion of additional features such as cooling passages, ribs, spars, and thermal coatings, without compromising the strength characteristics of the material.

Stacked laminate CMC turbine vane

A hybrid vane ( ) for a gas turbine engine having a ceramic matrix composite (CMC) airfoil member ( ) bonded to a substantially solid core member ( ). The airfoil member and core member are cooled by a cooling fluid ( ) passing through cooling passages ( ) formed in the core member. The airfoil member is cooled by conductive heat transfer through the bond (( ) between the core member and the airfoil member and by convective heat transfer at the surface directly exposed to the cooling fluid. A layer of insulation ( ) bonded to the external surface of the airfoil member provides both the desired outer aerodynamic contour and reduces the amount of cooling fluid required to maintain the structural integrity of the airfoil member. Each member of the hybrid vane is formulated to have a coefficient of thermal expansion and elastic modulus that will minimize thermal stress during fabrication and during turbine engine operation.

Ceramic matrix composite gas turbine vane

A vane assembly for a turbine assembly includes an inner endcap, an outer endcap, and a body. The body includes a metallic core assembly, a ceramic shell assembly and a support assembly. The metallic core assembly is coupled to the inner and outer endcaps and bears most of the mechanical loads, including aerodynamic loads. The ceramic shell bears substantially all of the thermal stress placed on the vane assembly. The support assembly is disposed between the metallic core assembly and said ceramic shell assembly and is coupled to the metallic core assembly.

Ceramic composite vane with metallic substructure

Shrink fit sleeves 30 interposed between the hub 21 and the shaft 10 of a large rotating apparatus adapted for operation in an environment subject to severe cyclical loading as occasioned by large thermal transients, are provided with the property of having greater axial shear flexibility than radial flexibility and/or greater axial flexibility than that of the shrunk on component 21.

Shrink fit sleeve for rotating machinery

A hub assembly for a centrifugal fan subjected to large transient thermal stresses is provided in which a pair of opposed hubs sandwich the inner peripheral margin of a disc-shaped fan plate, with the inner peripheral margin of the plate including a thickened ring portion to provide oppositely projecting annular shoulders at the outer periphery of the thickened ring portion, with the opposing hubs having annular recesses to receive the thickened ring portion and with the outer peripheries of the recesses in closely abutting relation with the shoulders to provide a shoulder-recess bearing area and with a plurality of fastener means extending through the hubs and through the thickened ring portion so that the shoulder-recess bearing area, which is located radially outwardly from the bearing stress areas at the bolt holes, largely receives the transient thermal stresses and therefore reduces the stress in the bolts and plate.

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Pittsburgh, PA, United States of America

Daniel George Thompson