Siemens Aktiengesellschaft

Siemens Energy, Inc.

Other

Siemens Energy Global Gmbh & Co. Kg

Jan H Marsh

Paul A Sanders

John Joseph Marra

Stephen John Messmann

Allister William James

Matthew J Golsen

Evan C Landrum

Carmen Andrew Scribner

Brian J Wessell

Daniel M Joo

Wayne J McDonald

Paul J Gear

Christian X Campbell, Jr

A method for attaching a first component composed of a ceramic matrix composite (CMC) to a second component composed of a metallic substructure is provided. The first component comprises at least two slots formed within a body of the first component and a configured to accommodate a thickness of a continuous CMC strap. The ends of the CMC strap are inserted into respective slots in the first component and then inserted into corresponding slots in a second component. The ends are secured to the second component by a fastening means thus securing the first component to the second component.

Method of securing a ceramic matrix composite (CMC) component to a metallic substructure using CMC straps

An airfoil for a turbine engine includes an array of features positioned in an interior portion of the airfoil. Each feature extends from a pressure side to a suction side. The array includes multiple radial rows (A-N) of features with the features in each row (A-N) being interspaced radially to define coolant passages therebetween. The radial rows (A-N) are spaced along a forward-to-aft direction toward an airfoil trailing edge. The coolant passages of the array are fluidically interconnected to lead a pressurized coolant toward the trailing edge via a serial impingement on to the rows of features. The coolant passages are geometrically configured to bias a coolant flow therethrough toward a first side in relation to a second side of the outer wall to effect a greater cooling of the first side than the second side.

Turbine airfoil with biased trailing edge cooling arrangement

An airfoil () includes at least one internal cooling channel (A-F) extending in the radial direction and adjoined on opposite sides by an airfoil pressure sidewall () and an airfoil suction sidewall (). An internal surface () of the airfoil pressure sidewall () and an internal surface () of the airfoil suction sidewall () define heat transfer surfaces in relation to a coolant flowing through the internal cooling channel (A-F). A flow splitter feature () is located in a flow path of the coolant in the internal cooling channel (A-F) between the pressure and suction sidewalls (). The flow splitter feature () is effective to create a flow separation region downstream of the flow splitter feature (), whereby coolant flow velocity is locally increased along the internal surfaces () of the pressure and suction sidewalls ().

Turbine airfoil with internal cooling channels having flow splitter feature

A turbine airfoil () includes a flow blocking body () positioned an internal cavity (). A first near-wall cooling channel () is defined between the flow blocking body () and an airfoil pressure sidewall (). A second near-wall cooling channel () is defined between the flow blocking body () and an airfoil suction sidewall (). A connecting channel () is defined between the flow blocking body () an internal partition wall () that connects the airfoil pressure () and suction () sidewalls. The connecting channel () is connected to the first () and second () near-wall cooling channels along a radial extent. Turbulating features (-) are located in the connecting channel () and are formed on the flow blocking body () and/or on the partition wall (). The turbulating features (-) are effective to produce a higher coolant flow rate through the first () and second () near-wall cooling channels in comparison to the connecting channel ().

Turbine airfoil with turbulating feature on a cold wall

The present disclosure provides a core structure comprising a trailing edge section including a plurality of rib-forming apertures () defined by a plurality of radially-extending channel elements () and axially-extending passage elements () and a radially outer low flow framing channel element () located adjacent to a radially outer edge (). The core structure may be used for casting a gas turbine engine airfoil (). The radially outer framing channel element () comprises a plurality of notches () extending radially inwardly from the radially outer edge (). A distal portion () of the notches () overlaps in an axial direction with the rib-forming apertures () of a first axially-aligned outer row (). A radial height of at least one of a first and a second axially-extending passage element () is greater than a prevalent radial height of other axially-extending passage elements () in the core structure.

Turbine blade trailing edge with low flow framing channel

A turbine airfoil () includes an impingement structure (A,B) comprising a hollow elongated main body () positioned in an interior portion () of an airfoil body (). The main body () extends lengthwise along a radial direction and defines coolant cavity () therewithin that receives a cooling fluid (). The main body () is spaced from a pressure side wall () and a suction side wall () of the airfoil body () and may be spaced from an airfoil tip (), to define respective passages () therebetween. A plurality of impingement openings () are formed through the main body () that connect the coolant cavity () with one or more of the respective passages (). The impingement openings () direct the cooling fluid () flowing in the coolant cavity () to impinge on the pressure and/or suction side walls () and/or the airfoil tip ().

Turbine airfoil with internal impingement cooling feature

A turbine airfoil () includes a flow displacement element (A-B,A'-B&#x2032;) positioned in an interior portion () of an airfoil body () between a pair of adjacent partition walls () and comprising a radially extending elongated main body (). The main body () is spaced from the pressure and suction side walls () and further spaced from one or both of the adjacent partition walls (), whereby a first near wall passage () is defined between the main body () and the pressure side wall (), a second near wall passage () is defined between the main body () and the pressure side wall () and a central channel () is defined between the main body () and a respective one of the adjacent partition walls (). The central channel () is connected to the near wall passages () along a radial extent. One or more radial ribs () are positioned in the central channel () that extend partially across the central channel () between the main body () and the respective adjacent partition wall ().

Turbine airfoil having flow displacement feature with partially sealed radial passages

A turbine airfoil () includes a flow displacement element () occupying a space between a pair of adjacent partition walls () in an interior portion () of a generally hollow airfoil body (). The flow displacement element () includes an elongated main body () extending lengthwise in a radial direction and a pair of connector ribs () respectively connecting the main body () to pressure and suction sides () of the airfoil (). A pair of adjacent radial flow passes (--) of symmetrically opposed flow cross-sections are defined on chordally opposite sides of the flow displacement element (). The radial flow passes (--) conduct cooling fluid in opposite radial directions and are connected in series via a chordal passage () defined in the interior portion () between the flow displacement element () and a radial end face () of the airfoil body (), to form a serpentine cooling path ().

Internally cooled turbine airfoil with flow displacement feature

A turbine airfoil including a central cavity defined by an outer wall including pressure and suction sides extending between and joined at leading and trailing edges, and a chordal axis extends generally centrally between the pressure and suction sides. Rib structures located in the central cavity define radial central channels extending across the chordal axis. Radial near wall passages are defined between the rib structures and each of the pressure and suction sides of the outer wall. The radial near wall passages are each open to an adjacent central channel along a radial extent of both the near wall passages and the adjacent central channel to define a radial flow pass associated with each central channel. The flow passes are connected in series to form a serpentine cooling path extending in the direction of the chordal axis.

Airfoil cooling with internal cavity displacement features

A cooling system for a turbine engine for directing cooling fluids from a compressor to a turbine blade cooling fluid supply and from an ambient air source to the turbine blade cooling fluid supply to supply cooling fluids to one or more airfoils of a rotor assembly is disclosed. The cooling system may include a compressor bleed conduit extending from a compressor to the turbine blade cooling fluid supply that provides cooling fluid to at least one turbine blade. The compressor bleed conduit may include an upstream section and a downstream section whereby the upstream section exhausts compressed bleed air through an outlet into the downstream section through which ambient air passes. The outlet of the upstream section may be generally aligned with a flow of ambient air flowing in the downstream section. As such, the compressed air increases the flow of ambient air to the turbine blade cooling fluid supply.

Cooling system with compressor bleed and ambient air for gas turbine engine

A turbine airfoil cooling system including a low pressure cooling system and a high pressure cooling system for a turbine airfoil of a gas turbine engine is disclosed. In at least one embodiment, the low pressure cooling system may be an ambient air cooling system, and the high pressure cooling system may be a compressor bleed air cooling system. In at least one embodiment, the compressor bleed air cooling system in communication with a high pressure subsystem that may be a snubber cooling system positioned within a snubber. A delivery system including a movable air supply tube may be used to separate the low and high pressure cooling subsystems. The delivery system may enable high pressure cooling air to be passed to the snubber cooling system separate from low pressure cooling fluid supplied by the low pressure cooling system to other portions of the turbine airfoil cooling system.

Turbine airfoil cooling system with cooling systems using high and low pressure cooling fluids

A vane assembly () having: an airfoil () and a shroud () held together without metallurgical bonding there between; a channel () disposed circumferentially about the airfoil (), between the airfoil () and the shroud (); and a seal () disposed in the channel (), wherein during operation of a turbine engine having the vane assembly () the seal () has a sufficient ductility such that a force generated on the seal () resulting from relative movement of the airfoil () and the shroud () is sufficient to plastically deform the seal ().

Ductile alloys for sealing modular component interfaces

A turbine airfoil usable in a turbine engine and having at least one snubber with a snubber cooling system positioned therein and in communication with an airfoil cooling system is disclosed. The snubber may extend from the outer housing of the airfoil toward an adjacent turbine airfoil positioned within a row of airfoils. The snubber cooling system may include an inner cooling channel separated from an outer cooling channel by an inner wall. The inner wall may include a plurality of impingement cooling orifices that direct impingement fluid against an outer wall defining the outer cooling channel. In one embodiment, the cooling fluids may be exhausted from the snubber, and in another embodiment, the cooling fluids may be returned to the airfoil cooling system. Flow guides may be positioned in the outer cooling channel, which may reduce cross-flow by the impingement orifices, thereby increasing effectiveness.

Turbine airfoil with laterally extending snubber having internal cooling system

A turbine airfoil usable in a turbine engine and having at least one ambient air cooling system is disclosed. At least a portion of the cooling system may include one or more cooling channels configured to receive ambient air at about atmospheric pressure. The ambient air cooling system may have a tip static pressure to ambient pressure ratio of at least 0.5, and in at least one embodiment, may include a tip static pressure to ambient pressure ratio of between about 0.5 and about 3.0. The cooling system may also be configured such that an under root slot chamber in the root is large to minimize supply air velocity. One or more cooling channels of the ambient air cooling system may terminate at an outlet at the tip such that the outlet is aligned with inner surfaces forming the at least one cooling channel in the airfoil to facilitate high mass flow.

Growing community of inventors

Orlando, FL, United States of America

Jan H Marsh