Siemens Aktiengesellschaft

Siemens Energy, Inc.

Other

Nicholas F Martin, Jr

Ching-Pang Lee

Gm Salam Azad

Kok-Mun Tham

Vincent Paul Laurello

David James Wiebe

Gary Brian Merrill

Zhihong Gao

David G Sansom

Neil Hitchman

Manjit Shivanand

Ramesh Subramanian

Eric Joseph Schroeder

William W Pankey

Walter Marussich

Kenneth K Landis

Erik Johnson

Marco Claudio Pio Brunelli

Beatriz Suarez-Crosse

Sameer A Khan

A component is provided and includes a core including a ceramic matrix composite material, one or more cooling channels formed about the core, an outer metal shell disposed about the core and the one or more cooling channels and a protective material between the core and the outer metal shell. The one or more cooling channels are formed about the core as an array of cooling channels in the protective material.

Hybrid component with cooling channels and corresponding process

A flexible damper () for turbine blades () includes a plurality of segments () positioned together in a substantially linear pattern, each segment () having a first side (), a second side (), a top side (), a bottom side (), a length (), a width (), and a thickness ().

Flexible damper for turbine blades

A process for forming a component is provided. The process includes providing a cooling channel flow definition at least partially about a core including a ceramic matrix composite material. A metal material is cast about the core and the cooling channel flow definition to form an outer metal shell. In addition, a cooling channel is formed from the cooling channel flow definition in the component.

A removably attachable snubber assembly for turbine blades includes a turbine blade airfoil including a trailing edge and a leading edge joined by a pressure side and a suction side to provide an outer surface extending in a radial direction to a tip. At least one snubber attachment platform is integrally formed onto the outer surface of the turbine blade airfoil. The at least one snubber attachment platform includes an interlocking mechanism. A snubber is removably attachable to the at least one snubber attachment platform, the snubber including a first end, a second end, a trailing edge, a leading edge, a snubber length, and a snubber width. The snubber also includes a removable attachment mechanism on at least one of the first end and the second end that connects with the interlocking mechanism on the at least one snubber attachment platform.

Removably attachable snubber assembly

Turbine and compressor casing/housing abradable component embodiments for turbine engines, have abradable surfaces with asymmetric forward and aft ridge surface area density. The forward ridges have greater surface area density than the aft ridges to compensate for greater ridge erosion in the forward zone during engine operation and reduce blade tip wear in the aft zone. Some abradable component embodiments increase forward zone ridge surface area density by incorporating wider ridges than those in the aft zone.

Turbine abradable layer with compound angle, asymmetric surface area ridge and groove pattern

A turbine airfoil is provided with at least one insert positioned in a cavity in an airfoil interior. The insert extends along a span-wise extent of the turbine airfoil and includes first and second opposite faces. A first near-wall cooling channel is defined between the first face and a pressure sidewall of an airfoil outer wall. A second near-wall cooling channel is defined between the second face and a suction sidewall of the airfoil outer wall. The insert is configured to occupy an inactive volume in the airfoil interior so as to displace a coolant flow in the cavity toward the first and second near-wall cooling channels. A locating feature engages the insert with the outer wall for supporting the insert in position. The locating feature is configured to control flow of the coolant through the first or second near-wall cooling channel.

Turbine airfoil having near-wall cooling insert

A seal assembly between a hot gas path and a disc cavity in a turbine engine includes a non-rotatable vane assembly including a row of vanes and an inner shroud, a rotatable blade assembly axially adjacent to the vane assembly and including a row of blades and a turbine disc that forms a part of a turbine rotor, and an annular wing member located radially between the hot gas path and the disc cavity. The wing member extends generally axially from the blade assembly toward the vane assembly and includes a plurality of circumferentially spaced apart flow passages extending therethrough from a radially inner surface thereof to a radially outer surface thereof. The flow passages each include a portion that is curved as the passage extends radially outwardly to effect a scooping of cooling fluid from the disc cavity into the flow passages and toward the hot gas path.

Outer rim seal assembly in a turbine engine

Turbine and compressor casing abradable component embodiments for turbine engines, with composite grooves and vertically projecting asymmetric non-parallel walls or trapezoidal cross section ridges that reduce, redirect and/or block blade tip airflow leakage downstream into the grooves rather than from turbine blade airfoil high to low pressure sides. In some embodiments at least one angularly oriented first groove formed in the ridge plateau is adapted for angular orientation upstream a turbine blade rotation direction to resist blade tip airflow leakage and the ridges are separated by second grooves that are skewed relative to the respective ridge plateaus and the substrate that are also adapted for orientation upstream the turbine blade rotation direction to resist blade tip airflow leakage.

Turbine abradable layer with asymmetric ridges or grooves

Turbine and compressor casing abradable component embodiments for turbine engines, with composite grooves and vertically projecting alternating rows of first and second height ridges in planform patterns, to reduce, redirect and/or block blade tip airflow leakage downstream into the grooves rather than from turbine blade airfoil high to low pressure sides. The first ridges have a first ridge height greater than that of the second ridges. These ridge or rib embodiments have first lower and second upper wear zones. The lower zone, at and below the second ridge height, optimizes engine airflow characteristics, while the upper zone, between tips of the second and first ridges, is optimized to minimize blade tip gap and wear by being more easily abradable than the lower zone.

Turbine abradable layer with progressive wear zone multi level ridge arrays

A ring segment for a gas turbine engine includes a panel and a cooling system. The cooling system is provided within the panel and includes a cooling fluid supply trench having an open top portion and extending radially inwardly from a central recessed portion of the panel. The cooling system further includes a plurality of cooling fluid passages extending from the cooling fluid supply trench to a leading edge and/or a trailing edge of the panel. The cooling fluid passages receive cooling fluid from the cooling fluid supply trench, wherein the cooling fluid provides convective cooling to the panel as it passes through the cooling fluid passages.

Ring segment with cooling fluid supply trench

A seal assembly between a hot gas path and a disc cavity in a turbine engine includes a non-rotatable vane assembly including a row of vanes and an inner shroud, a rotatable blade assembly adjacent to the vane assembly and including a row of blades and a turbine disc that forms a part of a turbine rotor, and an annular wing member located radially between the hot gas path and the disc cavity. The wing member extends generally axially from the blade assembly toward the vane assembly and includes a plurality of circumferentially spaced apart flow passages extending therethrough from a radially inner surface thereof to a radially outer surface thereof. The flow passages effect a pumping of cooling fluid from the disc cavity toward the hot gas path during operation of the engine.

Turbine and compressor casing abradable component embodiments for turbine engines, with nested loop pattern abradable surface ridges and grooves. The nested loops comprise nested fully closed loops or a spiraling maze pattern loop. Some embodiments include distinct forward upstream and aft downstream composite multi orientation groove and vertically projecting ridges planform patterns, to reduce, redirect and/or block blade tip airflow leakage downstream into the grooves rather than from turbine blade airfoil high to low pressure sides. Ridge or rib embodiments have first lower and second upper wear zones. The lower zone optimizes engine airflow characteristics while the upper zone is optimized to minimize blade tip gap and wear by being more easily abradable than the lower zone.

Growing community of inventors

York, SC, United States of America

Nicholas F Martin, Jr