Siemens Energy, Inc.

Siemens Westinghouse Power Corporation

Westinghouse Electric Corporation

Siemens Aktiengesellschaft

Siemens Power Generation, Inc.

Raymond Scott Nordlund

Adam J Weaver

Jody W Wilson

Richard Colin Charron

William E North

Kent Goran Hultgren

Zachary Sinnott

Jay Alan Morrison

Frank J Cunha

Clifford Hatcher

David Allen Little

Steffen Udluft

Matthew D Montgomery

Siegmund Düll

David B Allen

Marc Christian Weber

Kai Heesche

Daniel O Davies

Adam Matthew Foust

Gerard McQuiggan

James Michael Zborovsky

Daniel J Pierce

Robert Kenmer Scott

Steven Douglas Ward

Leslie R Southall

Steven Wayne Dowman

Barry Lee McKinley

Nitin Chhabra

Ernie B Campbell

Anthony L Schiavo, Jr

Brian Charles Owen

Thomas Malow

Ricky Lee Smith

A method of controlling a complex system and a gas turbine being controlled by the method are provided. The method comprises providing training data, which training data represents at least a portion of a state space of the system; setting a generic control objective for the system and a corresponding set point; and exploring the state space, using Reinforcement Learning, for a control policy for the system which maximizes an expected total reward. The expected total reward depends on a randomized deviation of the generic control objective from the corresponding set point.

Randomized reinforcement learning for control of complex systems

An optical based system and method for monitoring turbine engine blade deflection during engine operation. An optical camera is coupled the exterior of an engine inspection port, so that its field of view captures images of a rotating turbine blade, such as the blade tip. The camera's image capture or sampling rate matches blade rotation speed at the same rotational position, so that successive temporal images of one or more blades show relative movement of the blade tip within the image field of view. The captured successive images are directed to a blade deflection monitoring system (BDMS) controller. The controller correlates change in a blade's captured image position within the camera field of view between successive temporal images with blade deflection. The BDMS alarms or trips engine operation if the monitored blade deflection falls outside permissible operation parameters.

Optical based system and method for monitoring turbine engine blade deflection

A turbine engine assembly for a generator including a turbine engine having a compressor section, a combustor section and a turbine section. An air bleed line is in communication with the combustor section for receiving combustor shell air from the combustor section and conveying the combustor shell air as bleed air to a plurality of stages of the turbine section. Bleed air is controlled to flow through the air bleed line when an operating load of the turbine engine assembly is less than a base load of the engine to bypass air exiting the compressor section around a combustor in the combustor section and effect a flow of high pressure combustor shell air to the stages of the turbine section.

Cooling of turbine components using combustor shell air

A wall assembly () for separating a first fluid at a highest pressure and lowest temperature outside () the wall assembly from a second fluid at a lowest pressure and highest temperature inside () the wall assembly. The wall assembly () having: a structural cold wall () for exposure to the first fluid and partly defining a first cavity (), and a structural cold wall aperture () for creating a first pressure drop (); a structural middle wall () partially defining the first cavity () and partially defining a second cavity (), and a structural middle wall aperture () for creating a second pressure drop (); and a floating wall () for exposure to the second fluid and partially defining the second cavity (), and a floating wall aperture () for creating a third pressure drop ().

Thermally isolated wall assembly

An arrangement () for conveying combustion gas from a plurality of can annular combustors to a turbine first stage blade section of a gas turbine engine, the arrangement () including a plurality of interconnected integrated exit piece (IEP) sections () defining an annular chamber () oriented concentric to a gas turbine engine longitudinal axis () upstream of the turbine first stage blade section. Each respective IEP () includes a first flow path section () receiving and fully bounding a first flow from a respective can annular combustor along a respective common axis () there between, and delivering a partially bounded first flow to a downstream adjacent IEP section (). Each respective IEP further includes a second flow path section () receiving a partially bounded second flow from an upstream adjacent IEP () and delivering at least part of the second flow to the turbine first stage blade section.

Assembly for directing combustion gas

An arrangement for delivering gasses from can combustors of a can annular gas turbine combustion engine to a turbine first stage section including a first row of turbine blades, the arrangement including a flow-directing structure for each combustor, wherein each flow-directing structure includes a straight path and an annular chamber end, wherein the annular chamber ends together define an annular chamber for delivering the gas flow to the turbine first stage section, wherein gasses flow from respective combustors, through respective straight paths, and into the annular chamber as respective straight gas flows, and wherein the annular chamber is configured to unite the respective straight gas flows along respective shear planes to form a singular annular gas flow, and wherein the annular chamber is configured to impart circumferential motion to the singular annular gas flow before the singular annular gas flow exits the annular chamber to the first row of blades.

Modular transvane assembly

A transition duct () for a gas turbine engine () having improved cooling and reduced stress levels. The transition duct may be formed of two panels (() joined together with welds () disposed remote from the bent corner regions () of the panels. Cooling channels () extending longitudinally in the direction of flow of the hot combustion gas carried by the duct are formed within each panel, including the corner regions. Because the entire annular width (W) of the transition duct is cooled, the gap (G) separating adjacent ducts around the inlet to the turbine () may be reduced when compared to prior art designs. Two-panel construction with welds remote from the corner regions is facilitated by maintaining the minimum bend radius in the corners (R) and in the direction of flow (R) to be greater than in prior art designs.

Cooled transition duct for a gas turbine engine

A transition-to-turbine seal () including an upstream portion () adapted to engage a flange () of a transition (). The upstream portion () may be U-shaped in cross-sectional profile and include a primary wall () that includes a proximal section () and a distal section () relative to a hot gas path (). The proximal section () includes a plurality of recesses () which are spaced apart and separated by intervening wall (). In each recess () is provided one or more outlets () of cooling fluid holes (). The outlets () communicate via the cooling fluid holes () with a supply of compressed cooling fluid, such as compressed air that is provided from the compressor. During operation the outlets () release this fluid into the respective recesses (). The flow of cooling fluid provides for a more uniform cooling effect that includes impingement and convective cooling.

Transition-to-turbine seal apparatus and transition-to-turbine seal junction of a gas turbine engine

One embodiment of a transition-to-turbine seal () comprises a first, flattened section () adapted to be received in a peripheral axial slot () of a transition (), and a second, generally C-shaped section (). The generally C-shaped section () comprises a flattened portion () near the first, flattened section (), and a curved portion () extending to a free edge (). A fiber metal strip component () may be attached to the flattened portion () to define a first engagement surface adapted to engage an upstream side () of an outer vane seal rail (), and a second engagement surface, adjacent the free edge (), provides an opposed wear surface adapted to engage a downstream side () of the outer vane seal rail (). System embodiments also are described, in which such transition-to-turbine seal () is isolated from a hot gas path () by provision of a plurality of cooling apertures () in the transition ().

Slidable spring-loaded transition-to-turbine seal apparatus and heat-shielding system, comprising the seal, at transition/turbine junction of a gas turbine engine

A replaceable section () of force-cooling tubing assembly () for attachment to a transition piece () of a gas turbine is comprised of two ends () fashioned for attachment by removable unions () to adjoining parts of a force-cooling tubing assembly. When assembled thereto to complete the assembly, the replaceable sectionprovides for fluid communication between a manifold () and the transition piece () for either the supply or return of cooling fluid. A transition piece () in combination with two such assemblies, one a supply assembly (), the other a return assembly (), comprises a field-installable transition piece assembly () that provides for rapid and easy installation. The features of the replaceable section () include a relatively inflexible bracing zone such as a bracing member () having a support structure such as a lateral plate () that extends to the transition piece (), a formed tubing bend () typically forming a U-bend, and optionally a second flexible component comprised of an inline flexible coupling ().

Multi-zone tubing assembly for a transition piece of a gas turbine

Aspects of the present invention relate to hollow elongated turbine engine components, such as transition ducts, having a cooling system. The component can have first and second ends, and can be defined by an outer peripheral wall. The cooling system extends longitudinally within the wall between the first and second ends. While configured to cool the entire component, the cooling system is particularly configured to reduce thermal gradients that can occur in regions proximate to the first and second ends of the component caused by higher loads in those regions. The cooling system can include at least one pair of cooling channels. Each channel pair includes a first channel and a second channel. The first and second channels can be arranged such that a coolant supplied to each channel can exchange heat with itself in at least the region substantially proximate to one of the ends.

Transition duct cooling system

A turbine vane assembly which includes an outer endcap having a plurality of generally straight passages and passage segments therethrough, an inner endcap having a plurality of passages and passage segments therethrough, and a vane assembly having an outer shroud, an airfoil body, and an inner shroud. The outer shroud, airfoil body and inner shroud each have a plurality of generally straight passages and passage segments therethrough as well. The outer endcap is coupled to the outer shroud so that outer endcap passages and said outer shroud passages form a fluid circuit. The inner endcap is coupled to the inner shroud so that the inner end cap passages and the inner shroud passages from a fluid circuit. Passages in the vane casting are in fluid communication with both the outer shroud passages and the inner shroud passages. Passages in the outer endcap may be coupled to a cooling system that supplies a coolant and takes away the heated exhaust.

Cooled turbine vane with endcaps

A closed-loop cooling scheme for cooling stationary combustion turbine components, such as vanes, ring segments and transitions, is provided. The cooling scheme comprises: (1) an annular coolant inlet chamber, situated between the cylinder and blade ring of a turbine, for housing coolant before being distributed to the turbine components; (2) an annular coolant exhaust chamber, situated between the cylinder and the blade ring and proximate the annular coolant inlet chamber, for collecting coolant exhaust from the turbine components; (3) a coolant inlet conduit for supplying the coolant to said coolant inlet chamber; (4) a coolant exhaust conduit for directing coolant from said coolant exhaust chamber; and (5) a piping arrangement for distributing the coolant to and directing coolant exhaust from the turbine components. In preferred embodiments of the invention, the cooling scheme further comprises static seals for sealing the blade ring to the cylinder and flexible joints for attaching the blade ring to the turbine components.

Cooling scheme for turbine hot parts

A gas turbine that has turbine vanes that can be cooled with either steam or air includes a compressor that compresses air, a combustor enclosed within a shell and in flow communication with the compressor, a turbine in flow communication with the combustor and receiving hot gas from the combustor and turbine vanes. The turbine vanes having an airfoil, a shroud and a cooling circuit that cools the airfoil and the shrouds with either steam or air.

Cooling system for gas turbine vane

An apparatus for cooling the trailing edge portion of a gas turbine vane. Two radially extending passages connected to the outer shroud direct cooling fluid to a plenum formed about mid-span adjacent the trailing edge. Two arrays of cooling fluid passages extend from the plenum. One array extends radially outward toward the outer shroud while the other array extends radially inward toward the inner shroud. The plenum distributes the cooling fluid to the two arrays of passages so that it flows radially inward and outward to manifolds formed in the inner and outer shrouds. The manifolds direct the spent cooling fluid to a discharge passage.

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Orlando, FL, United States of America

Raymond Scott Nordlund