Siemens Power Generation, Inc.

Siemens Westinghouse Power Corporation

Other

United Technologies Corporation

General Railway Signal Corporation

Siemens Energy, Inc.

Harry A Albrecht

Yevgeniy P Shteyman

Jay Alan Morrison

Gary Brian Merrill

Steven C Butner

Christian X Campbell

Jay Edgar Lane

Steven James Vance

Daniel George Thompson

Scott M Widrig

Thomas Barrett Jackson

Daniel W Van Alstine

Chris Campbell

Norman F Fowle

Yevgeuiy Shteyman

Edwin E Knight

Michael Diver

Yevgenly Shteyman

A means () for structurally stiffening or reinforcing a ceramic matrix composite (CMC) gas turbine component, such as an airfoil-shaped component, is provided. This structural stiffening or reinforcing of the airfoil allows for reducing bending stress that may be produced from internal or external pressurization of the airfoil without incurring any substantial thermal stress. The stiffener is disposed on a CMC wall and generally extends along a chord length of the airfoil.

Ceramic matrix composite vane with chordwise stiffener

A turbine airfoil support system for coupling together a turbine airfoil formed from two or more components, wherein the support system is particularly suited for use with a composite airfoil. In at least one embodiment, the turbine airfoil support system may be configured to attach shrouds to both ends of an airfoil and to maintain a compressive load on those shrouds while the airfoil is positioned in a turbine engine. Application of the compressive load to the airfoil increases the airfoil's ability to withstand tensile forces encountered during turbine engine operation.

Support system for a composite airfoil in a turbine engine

An airfoil () formed of a plurality of pre-fired structural CMC panels (). Each panel is formed to have an open shape having opposed ends () 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 () for a gas turbine engine. The panels may be joined together using a backing member (), using flanged ends () and a clamp (), and/or with a bond material (), for example.

CMC component and method of fabrication

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.

Cooling systems for stacked laminate CMC vane

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 platform cooling system usable in a turbine engine together with an airfoil for preheating cooling fluids before the cooling fluids enter a cooling system in the airfoil in a turbine engine. The platform cooling system includes cooling channels in either the ID or OD platforms, or both, of the airfoil. The channels transfer heat to the cooling fluids flowing through the platform cooling system and thereby heat the cooling fluids. The preheated cooling fluids are particularly useful with cooling composite ceramic airfoils, which are susceptible to damage from large temperature gradients developed between combustion gases outside the airfoil and cooling fluids inside the airfoil. The platform cooling system may be combined with an airfoil cooling system to create a serial cooling system in which cooling fluids may enter the platform and flow through the platform and airfoil without being supplemented with additional cooling fluids along the flow path.

Cooling fluid preheating system for an airfoil in a turbine engine

A method of manufacturing a composite structure uses a layer of an insulating material () as a mold for forming a substrate of a ceramic matrix composite (CMC) material (). The insulating material may be formed in the shape of a cylinder () with the CMC material wound on an outer surface () of the cylinder to form a gas turbine combustor liner (). Alternatively, the insulating material may be formed in the shape of an airfoil section () with the CMC material formed on an inside surface () of the insulating material. The airfoil section may be formed of a plurality of halves () to facilitate the lay-up of the CMC material onto an easily accessible surface, with the halves then joined together to form the complete composite airfoil. In another embodiment, a box structure () defining a hot gas flow passage () is manufactured by forming insulating material in the shape of opposed airfoil halves () joined at respective opposed ends by platform members (). A layer of CMC material () is then formed on an outside surface of the insulating material. A number of such composite material box structures are then joined together to form a vane ring () for a gas turbine engine.

Composite structure formed by CMC-on-insulation process

An airfoil () having a continuous layer of ceramic matrix composite (CMC) material () extending from a suction side () to a pressure side () around a trailing edge portion (). The CMC material includes an inner wrap () extending around an inner trailing edge portion () and an outer wrap () extending around an outer trailing edge portion (). A filler material () 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 () or a protrusion (). The filler material may include an upper layer () and a lower layer () separated by an intermediate layer () that extends to between the inner wrap and the outer wrap along the suction and/or pressure sides.

Ceramic matrix composite airfoil trailing edge arrangement

A multi-layer ceramic matrix composite structure ( ) having a plurality of fiber-reinforced cooling passages ( ) formed therein. The cooling passages are formed by the removal of a fugitive material ( ). The fugitive material is part of a wrapped fugitive material structure ( ) containing a layer of reinforcing ceramic fibers ( ) that is used to lay-up the multi-layer structure. An intermediate layer of ceramic fabric may be placed alternately over and under the wrapped fugitive material structure to separate the cooling passages into alternating upper ( ) and lower ( ) cooling passages. The transversely oriented fibers surrounding the cooling passages serve to increase the interlaminar strength of the structure when compared to prior art designs. An airfoil member ( ) incorporating such reinforced integral cooling passages ( ) is provided.

Ceramic matrix composite structure having integral cooling passages and method of manufacture

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 ceramic matrix composite material (CMC) vane for a gas turbine engine wherein the airfoil member ( ) and the platform member ( ) are formed separately and are then bonded together to form an integral vane component ( ). Airfoil member and the platform member may be bonded together by an adhesive ( ) after being fully cured. Alternatively, respective joint surfaces ( ) of the green body state airfoil member and platform member may be co-fired together to form a sinter bond ( ). A mechanical fastener ( ) and/or a CMC doubler ( ) may be utilized to reinforce the bonded joint ( ). A matrix infiltration process ( ) may be used to create or to further strengthen the bond.

Ceramic matrix composite turbine vane

Apparatus for operating railroad switches, including operation in the trailing mode, and a switch machine having a throw bar and throw rod for enabling the throwing of switch points into selected positions so as to control the switching of rolling stock. A trailing device, including a trailing rod portion of the throw rod, is coupled to the throw bar, said device including an arrangement responsive to a wrong direction load impressed on said switch points so as to move said throw rod, for enabling non-damaging movement of said throw rod.

Trailing device for a railroad switch layout

The engine exhaust flow diverter is formed as a duct, extending from the exhaust section of a gas turbine powerplant, having a closed end which is contoured externally as a nozzle plug. First exit means for the flow from the powerplant extends around the duct just forwardly of the contoured plug and second exit means for the flow extends around the duct midway between the first exit means and the end of the exhaust section of the powerplant. An annular door is axially movable around the duct between an open and closed position to control the flow through the second exit opening and an annular shroud is axially movable around the duct between an open and closed position to control the flow through the second exit means. The annular shroud cooperates with the contoured nozzle plug to form a conventional nozzle. An actuating system is connected to both the annular door and annular shroud to actuate them so that when one is open the other is closed. Means insure that each door or shroud is sealed when it is in its closed position. A bellcrank lever device is connected to the annular shroud so that its movement can be varied with respect to the movement of the annular door, this is done to control the respective areas of the openings as they move between their open and closed positions to maintain a desired effective area for optimum engine operation.

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Hobe Sound, FL, United States of America

Harry A Albrecht