Siemens Energy, Inc.

Siemens Westinghouse Power Corporation

Siemens Aktiengesellschaft

David B Allen

Ramesh Subramanian

Brij B Seth

Gregg P Wagner

David W Hunt

Anand A Kulkarni

Allister William James

Andrew Jeremiah Burns

Wolfram Beele

Raymond Scott Nordlund

Daniel O Davies

Hubertus Edward Paprotna

Adam J Weaver

David G Sansom

Oran Bertsch

Anirudha Vaidya

Back scratcher

Voids, cracks or other similar defects in substrates of thermal barrier coated superalloy components, such as turbine blades or vanes, are filled with resin, without need to remove substrate material surrounding the void by grinding or other processes. The resin is cured at a temperature under 200&#xb0; C., eliminating the need for post void-filling heat treatment. The void-filled substrate and resin are then coated with a thermal barrier coating.

Turbine superalloy component defect repair with low-temperature curing resin

A system of layers in a protective coating () for a substrate (), including at least an outer thermal barrier layer () and a tungsten bronze structure ceramic underlayer () that reduces spalling of the outer layer. The range of materials for the underlayer includes ceramics of the form BaReTiO, where 0&#x3c;m&#x3c;1.5, and Re is any rare earth element or mixture thereof. These underlayer materials reduce spalling of the thermal barrier layer, and thus extend the life of the coating system. In some embodiments, materials for the outer thermal barrier layer may include Yttria-stabilized Zirconia (YSZ) or ceramics with lower thermal conductivity than YSZ. A segmented YSZ layer () is provided in some embodiments for additional thermal expansion compliance.

Thermal barrier coating system with porous tungsten bronze structured underlayer

A squealer tip preform having a coating including an abrasive is disclosed. The squealer tip may be separately formed from a generally elongated blade to which the squealer tip preform is configured to be attached. The squealer tip may be configured such that an outer surface of the squealer tip preform is generally aligned with an outer side surface of the turbine blade defining a cross-sectional profile of the turbine blade. Forming the squealer tip preform with an abrasive coating separate from the turbine blade greatly reduces costs and improves efficiency as compared with conventional systems.

Abrasive coated preform for a turbine blade tip

A method for welding to a superalloy material without causing cracking of the material. The method includes the steps of depositing a weld strip () of a weldable material onto a superalloy substrate material () using a spray deposition process () and then forming a weldment () to the weld strip. None or a controlled amount of the substrate material () is melted during the weld in order to maintain the concentration of strengthening elements in the local melt within a zone of weldability. The spray deposition process may be a thermal process such as HVOF or a cold spray process. A groove () may be formed in a surface () of the superalloy substrate material to receive the weld strip. A diffusion heat treatment step may be used to improve adhesion between the weld strip and the superalloy material.

Sprayed weld strip for improved weldability

Fabricating a turbine component () by casting a core structure (), forming an array of pits () in an outer surface () of the core structure, depositing a transient liquid phase (TLP) material () on the outer surface of the core structure, the TLP containing a melting-point depressant, depositing a skin () on the outer surface of the core structure over the TLP material, and heating the assembly, thus forming both a diffusion bond and a mechanical interlock between the skin and the core structure. The heating diffuses the melting-point depressant away from the interface. Subsurface cooling channels () may be formed by forming grooves () in the outer surface of the core structure, filling the grooves with a fugitive filler (), depositing and bonding the skin (), then removing the fugitive material.

Sprayed skin turbine component

A method of measuring a thickness of a coating on a substrate material. A first pulse of monochromatic light having a predetermined first wavelength is emitted toward the coating and substrate material. A second pulse of monochromatic light having a predetermined second wavelength is emitted toward the coating and substrate material, the second wavelength being different than the first wavelength. A first elapsed time is measured from emission of the first pulse of light to reception of a reflection of the first pulse of light from a surface of the substrate material at an interface with the coating. A second elapsed time is measured from emission of the second pulse of light to reception of a reflection of the second pulse of light from an outer surface of the coating. A thickness of the coating is determined as a function of a difference between the first and second elapsed times.

Method of measuring coating thickness using infrared light

A bimetallic bond layer () for a thermal barrier coating or TBC () on a superalloy substrate () for a high temperature environment. An interlayer () is applied on the substrate. A bond coat () comprising a CoNiCrAlY or NiCoCrAlY alloy is applied on the interlayer. A ceramic TBC () such as 8YSZ is applied on the bond coat. The interlayer () is an alloy that is compatible with the substrate and the bond coat, and that blocks or delays diffusion of aluminum from the bond coat into the substrate at high operating temperatures. This preserves aluminum in the bond coat that maintains a beneficial alumina scale () between the bond coat and the TBC. This delays spalling of the TBC, and lengthens the coating and component life.

Bimetallic bond layer for thermal barrier coating on superalloy

A quantity of particles () may be cold-sprayed toward the distal surface () of a turbine blade () at a temperature below a melting point of the particles at a velocity sufficiently high to cause at least a portion of the particles to adhere to the surface to form squealer ridge () with a desired cutting profile. The cutting profile may be defined by a face () and a land () disposed at a relief angle () relative to a direction of motion () of a cutting edge (). The ridge may be built by multiple passes of a cold spray gun depositing sequential layers () of varying widths. An in-service blade having worn distal surface may be repaired using this technique.

Cold spray method for producing gas turbine blade tip

A method of forming an interstage seal including removing a diaphragm seal box () from a gas turbine compressor assembly () and removing a labyrinth sealing member () from the diaphragm seal box (). An abradable material layer () may be deposited on the diaphragm seal box (). A spray gun may be mounted in relation to an engine disk () of the gas turbine compressor assembly () for cold-spraying a quantity of particles toward the engine disk (). The particles may be sprayed at a velocity sufficiently high to cause at least a portion of the quantity of particles to adhere to the engine disk (). The spray gun may be controlled to deposit a quantity of particles on the compressor disk () to form a geometry () that will abrade the abradable material layer () during operation of the gas turbine compressor assembly (). The geometry () abrading the abradable material layer () forms an interstage seal. Rotation of the engine disk () within the gas turbine compressor assembly () may be controlled during the step of controlling the spray gun. The diaphragm seal box () may be reinstalled within the gas turbine compressor assembly () so that the geometry () aligns with the abradable material layer ().

Method for forming turbine seal by cold spray process

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

This invention relates to an abradable coating system for use in axial turbine engines. When coated onto a turbine ring seal segment the coating system may allow formation of an individualized seal between turbine blade disks and the surrounding ring seal without causing excessive wear to the blade tips. The abradable coating system includes columns of an abradable material. Thus, interference between the blades and the abradable coating system causes the individual columns to break off at the base. This abrasion mechanism may reduce blade wear and spalling of the coating system when compared to conventional coatings.

Abradable coating system

Aspects of the invention are directed to a system for improving engine performance by reducing vane tip clearances in the compressor section of a turbine engine. According to aspects of the invention, an abrasive material can be attached to a rotor disc. Thus, if a vane tip contacts the rotor disc during engine operation, the vane tip is worn away by the abrasive material. The system can reduce the risk of substantial component damage resulting from a vane tip rubbing event. The abrasive material can be provided in the form of an insert that can be removably attached to the rotor disc. Thus, the insert can be readily removed and replaced when necessary.

Rotor disc assembly with abrasive insert

A conformal seal () for sealing air flow between a cooling airflow path and a hot gas flow path within a combustion turbine engine. The conformal seal () may be fitted within cooperating side slots of adjacent vane segments () within the combustion turbine engine. The conformal seal () may include an elongated metallic substrate () forming an upper surface and a lower surface. A conformal coating () may be deposited over one or both surfaces of the substrate (). The conformal coating () may be deposited to a depth so that a point contact between the conformal coating () and respective interior walls of the side slots wears the conformal coating () to establish surface area contact there between. The surface area contact improves a sealing function between the conformal coating () and the respective interior walls during operation of the combustion turbine engine.

Thermally sprayed conformal seal

An abradable thermal barrier coating material () formed of a highly defective fluorite ceramic matrix () having a desired degree of porosity () created in part by the addition of a fugitive material (). The ceramic material has a concentration of a stabilizer sufficiently high that the oxygen vacancies created by the stabilizer interact within the matrix to form multi-vacancies, thereby improving the sintering resistance of the material. Such a concentration of stabilizer results in a material that is softer than prior art materials having lower concentrations of stabilizer, and that will be more resistive to sintering than prior art materials. Embodiments include a fluorite matrix of zirconia stabilized by at least 30 wt. % yttria, or stabilized by at least 30 wt. % ytterbia, and with porosity of 10-40%. In one embodiment, a metallic gas turbine seal ring segment is coated with a bond coat layer, then with a layer of porous 8 wt. % YSZ material, and finally with a layer of 33 mole % YbSZ (61.3 wt. % YbSZ) material having porosity of 10-40%.

Growing community of inventors

Oviedo, FL, United States of America

David B Allen