Siemens Energy, Inc.

Siemens Aktiengesellschaft

United Technologies Corporation

Siemens Power Generation, Inc.

Siemens Energy Global Gmbh & Co. Kg

Robert W Sunshine

Kok-Mun Tham

Todd A Ebert

Kevin Michael Light

Hans-Thomas Bolms

Daniel E Kane

Ralph J Thompson

Yan Yin

Gennadiy Afanasiev

Christian Kowalski

John N Strain

Burkhard Voss

Uwe Lohse

Robert D Barker

Ernie B Campbell

Fan Zhang

Dieter Brillert

Wayne Giddens

Harald Hoell

Juergen Hermeler

Sam Baghdadi

A cooling system configured to cool aspects of the turbine engine between a compressor and a turbine assembly is disclosed. In at least one embodiment, the cooling system may include one or more mid-frame cooling channels extending from an inlet through one or more mid-frame torque discs positioned downstream of the compressor and upstream of the turbine assembly. The inlet may be positioned to receive compressor bleed air. The mid-frame cooling channel may be positioned in a radially outer portion of the mid-frame torque disc to provide cooling to outer aspects of the mid-frame torque disc such that conventional, low cost materials may be used to form the mid-frame torque disc rather than high cost materials with capacity to withstand higher temperatures. The cooling fluid routed through the mid-frame cooling channel in the mid-frame torque disc may be exhausted into a cooling system () for the downstream turbine assembly.

Compressor bleed cooling system for mid-frame torque discs downstream from a compressor assembly in a gas turbine engine

A seal assembly () for a gas turbine engine for controlling air flow between a diffuser () and rotor disks comprising first and second annular flange ends () and an annular seal mid-section () between and operatively connected to the flange ends (). The first and second annular flange ends () abut respective outer frame members () of the diffuser, whereby a fluid flow path is formed between the seal assembly () and the rotor disks (). The first and second end flanges () are composed of a material having a coefficient of thermal expansion that is substantially the same as a coefficient of thermal expansion of the material of the outer frame members (). In addition, the material of the seal mid-section () has a coefficient of thermal expansion that is different than that of the materials of the annular flange ends () and outer frame members ().

Seal assembly for controlling fluid flow

A turbine section having a plurality flow path components forming a plurality of guide vane rings and ring segments arranged in axial succession to define a boundary of a hot gas duct. A vane carrier located around the gas duct, and sealing elements extend radially between circumferentially extending grooves in the vane carrier and respective grooves in the flow path components. The sealing elements include radially inner and outer edges, and at least one axially facing side defining a chamfered portion extending to one of the edges to accommodate axial movement of the sealing element about the one edge within a respective groove.

Gas turbine engine sealing structure

Embodiments of the invention relate to a system and method for controlling the size of gaps in a turbine engine. In many instances, it is desirable to minimize the size of the gaps between neighboring rotating and stationary components in a turbine engine, such as between a disc cover plate and a proximate pre-swirler. According to embodiments of the invention, each component can be provided with a sealing surface. The sealing surfaces can be angled relative to the axis of rotation. The sealing surfaces are spaced from each other so as to form a gap therebetween. The sealing surfaces may or may not be substantially parallel. As a result of such configuration, the size of the gap can be controlled by axial and radial movement of the components. For example, the gap between the cover plate and the pre-swirler can be adjusted by axially movement of the rotor.

Gap control system for turbine engines

A turbine shroud segment for use in a gas turbine engine includes a serpentine channel along at least one axial edge of the segment. Various construction details are developed that disclose a channel for efficiently flowing cooling fluid through an axial edge of a shroud segment. In a particular embodiment, a turbine shroud segment includes a leading edge serpentine channel having a bend passage which includes a purge hole to avoid separating flow in the bend passage.

Turbine shroud segment with serpentine cooling channels having a bend

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Hobe Sound, FL, United States of America

Robert W Sunshine