Siemens Energy, Inc.

Siemens Aktiengesellschaft

Siemens Energy Global Gmbh & Co. Kg

Siemens Power Generation, Inc.

Abdullatif M Chehab

Kok-Mun Tham

David Allen Little

Ching-Pang Lee

Vincent Paul Laurello

Kevin L Hines

Jiping Zhang

Jose L Rodriguez

Kevin Michael Light

Gm Salam Azad

Ross H Peterson

Weidong Cai

Christopher K Rawlings

Kai Kadau

Yan Yin

Zhengxiang Pu

Mahmood K Silieti

Manjit Shivanand

Brian Harry Terpos

Barton M Pepperman

Shantanu P Mhetras

Amit K Paspulati

David A Kemp

John A Fussner

Bijay K Sultanian

Matthew D Bogdan

Scott T Waechter

A non-contact seal assembly for sealing a gap between relatively rotatable components in a gas turbine engine is presented. The non-contact seal assembly includes a primary seal having a radially movable seal shoe, a mid-plate, an aft secondary seal radially movable along with the seal shoe, a forward secondary seal and a U-shaped seal carrier for holding the components together using pins. The seal shoe includes a plurality of seal shoe segments. The aft secondary seal includes a plurality of aft secondary seal segments. Each aft secondary seal segment is attached to each seal shoe segment. Each aft secondary seal segment includes at least a notch at outer radial side to receive the pin for accommodating radial movement of the seal shoe segment.

Non-contact seal assembly in gas turbine engine

A gas turbine engine having a rotor centering cooling system for cooling struts within an exhaust diffuser and turbine case to reduce tip rub during hot restarts is disclosed. In particular, the rotor centering cooling system may be positioned within struts in the exhaust diffuser downstream from a turbine assembly for limiting thermal gradients between top and bottom struts to prevent the exhaust bearing body from becoming off-center during steady state operation as a result of the top struts becoming hotter than the bottom struts. The rotor centering cooling system may reduce the temperature at the exhaust diffuser and turbine case, thereby reducing the thermal gradient between the top and bottom struts and top and bottom of the turbine case. As such, the exhaust bearing body remains centered, thereby preventing a tighter blade tip clearance at the top of the turbine assembly than at the bottom of the assembly.

Gas turbine engine with rotor centering cooling system in an exhaust diffuser

A turbine engine including an intermediate space defined between outer and inner portions of the turbine engine. A flow energizer is provided including a flow body located within the intermediate space and including an inlet port, an outlet port and a flow passage extending within the flow body between the inlet and outlet ports. The inlet port receives a flow of a first medium located within the intermediate space and the flow body injects an energizing flow of a second medium to a portion of the first medium within the flow body to create an energized flow of a mixed medium from the outlet portion, the energized flow of mixed medium creates a flow of the first medium adjacent to the flow body within the intermediate space.

Temperature control within a cavity of a turbine engine

Turbine blade tip clearance is measured in a fully assembled turbine casing by mounting a non-contact displacement probe or sensor on a turbine blade that generates data indicative of sensor distance from the turbine casing that circumferentially surrounds the blade. The sensor is mounted on the blade with a sensor fixture, which includes a clamping mechanism and a sensor retention mechanism that retains and calibrates the sensor by selective movement of the sensor relative to the retention mechanism. Variations in sensor distance data are recorded when the turbine is operated in turning gear mode. Blade rotational position data are collected by a rotational position sensor. A data processing system correlates the distance and rotational position data with localized blade tip gap at angular positions about the turbine casing circumference. This method and apparatus facilitate assessment of turbine casing deformation impact on blade tip clearance and rotor/casing alignment.

Turbine blade-mounted sensor fixture for tip gap measurement

A gas turbine having rotor discs (), a disc cavity () and a stator stage () extending to the disc cavity (). Seal housing flanges () extend from a seal housing () of the stator stage (). Rotor flanges () extend from a rotor disk (-). An inner rotor flange () and first seal housing flange () are inward from a second seal housing flange (). One rotor flange () is outward from the second seal housing flange (). The inner rotor flange () and first seal housing flange () extend toward one another to limit movement of main gas flow (). An inlet () injects air () between the outward rotor flange () and second seal housing flange ().

Flow discourager integrated turbine inter-stage U-ring

The present invention comprises a gas turbine engine and a process for operating a gas turbine engine. A fluid structure receives compressed air from a compressor and extends toward a stationary blade ring in a turbine to discharge the compressed air directly against a surface of the blade ring such that the compressed air impinges on the blade ring surface. The compressed air then passes through at least one opening in the stationary blade ring and into cooling passages of a corresponding row of vanes.

Gas turbine engine with structure for directing compressed air on a blade ring

A cooling fluid air injection system for use in a gas turbine engine includes at an external cooling fluid source, at least one rotor cooling pipe, which is used to inject cooling fluid from the source into a rotor chamber, a piping system that provides fluid communication between the source and the rotor cooling pipe(s), a blower system for conveying the cooling fluid through the piping system and the rotor cooling pipe(s) into the rotor chamber, and a valve system. The valve system is closed during full load engine operation to prevent cooling fluid from the source from passing through the piping system, and open during less than full load engine operation to allow cooling fluid from the source to pass through the piping system.

External cooling fluid injection system in a gas turbine engine

A measuring assembly for use in a gas turbine engine includes an indicia portion that extends radially inwardly from an inner surface of a ring seal structure to a location radially inwardly from tips of blades mounted on a rotor. The indicia portion of the measuring assembly comprises a section that is abraded by the blades during rotational movement of the rotor to provide a visual indication of a distance between the tips of the blades and the inner surface.

Turbine tip clearance measurement

A system and method for actively managing blade tip clearances in a turbine engine, particularly under steady state operating conditions such as at base load, involves routing a portion of air from a rotor cooling air circuit to a vane carrier or other stationary support structure surrounding the turbine blades. Because the temperature of the air is less than the temperature of the stationary support structure, the stationary support structure will thermally contract when the air is passed in heat exchanging relation therewith. In one embodiment, the air can be passed through one or more passages extending through at least a portion of the stationary support structure. The contraction of the stationary support structure reduces the blade tip clearance because the blades do not contract. Thus, fluid leakage through the clearances is minimized, which in turn can increase engine performance.

Turbine blade tip clearance control

An inner mounting ring () for gas turbine flow path components such as shroud ring segments (). The inner ring () may be mounted to an outer ring () on radially slidable mounts () that maintain the two rings () in coaxial relationship, but allows them to thermally expand at different rates. This allows matching of the radial expansion rate of the inner ring () to that of the turbine blade tips (), thus providing reduced clearance () between the turbine blade tips () and the inner surface of the shroud ring segments () under all engine operating conditions. The inner ring () may be made of a material with a lower coefficient of thermal expansion than that of the outer ring ().

Inner ring with independent thermal expansion for mounting gas turbine flow path components

Aspects of the invention relate to a turbine engine system and method for actively managing blade tip clearances during part load operation of the engine. Aspects of the invention relate to extracting a portion of the combustion gases from the combustor section of the engine and routing these heated gases to the blade rings or other stationary structure surrounding the turbine blades. Upon exposure to the combustion gases, which can be mixed with compressor exit air, the stationary structure will thermally expand, causing blade tip clearances to increase. Thus, concerns of blade tip rubbing are minimized. Once the engines achieves steady state operation, the flow of the combustion gases to the blade rings can be substantially restricted so that only compressor exit air is supplied to the stationary structure. Consequently, the stationary structure will contract and the blade tip clearances will decrease, thereby increasing the efficiency of the turbine.

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Chuluota, FL, United States of America

Abdullatif M Chehab