An excellent abrasive blade tip is provided by a two-layer coating system consisting of a brazing solder coating and an NiCoCrAlY coating containing cBN (cubic Boron Nitride).

A method for applying an abrasive comprises: applying, to a substrate, the integral combination of: a self-braze material; and an abrasive embedded in the self-braze material; and securing the combination to the substrate.

A gas turbine engine component may include an airfoil extending radially from a base to a blade tip, the airfoil including a pressure sidewall and a suction sidewall each extending between a leading edge and a trailing edge opposite the leading edge, an internal cooling circuit extending from the base to the blade tip; and a squealer tip. The squealer tip may include a first and a second squealer tip rail adjacent to the first squealer tip rail, and a squealer tip cap extending between the first and second squealer tip rails. The blade tip, the first and second squealer tip rails, and the squealer tip cap may define an internal squealer tip cooling channel. The blade tip may define a supply aperture that fluidly connect the internal cooling circuit and the internal squealer tip cooling channel.

A method for repairing a damaged leading or trailing edge region of a metallic turbine blade, which has a suction-side blade wall, a pressure-side blade wall and a leading edge and a trailing edge, includes: removing the damaged leading or trailing edge region, with a cutout being formed; providing a replacement part which can be inserted in form-fitting fashion into the cutout and which restores the intended geometry of the turbine blade; inserting the replacement part in form-fitting fashion into the cutout in a predetermined insertion direction, and cohesively connecting the replacement part to the turbine blade, wherein mechanical connection of the replacement part and turbine blade using separate fastening elements is performed in addition to the cohesive connection.

A method of applying a braze component to a honeycomb structure may comprise: applying at least a partial vacuum within a chamber, the chamber defined at least partially by a vacuum device and a cover, the honeycomb structure disposed within the chamber, the braze component disposed between the honeycomb structure and the cover; pulling the cover towards the braze component in response to applying the partial vacuum; and pulling the braze component into a plurality of hexagonal cells defined by the honeycomb structure in response to pulling the cover towards the braze component.

An assembly for a dual-walled component of a gas turbine engine and methods of forming and repairing a dual-walled component. The assembly includes a cold section part having an outer surface that defines a plurality of impingement apertures, a hot section part including a pre-sintered preform, the hot section part positioned over the outer surface of the cold section part, and a plurality of support structures including the pre-sintered preform, the plurality of support structures positioned between the hot section part and the cold section part, the plurality of support structures separating the hot section part from the cold section part to define at least one cooling channel therebetween.

A method of furnace-less brazing of a substrate is provided. The method includes providing a substrate having a braze region thereon; disposing braze precursor material containing a nickel powder, an aluminum powder, and a platinum group metal powder on the braze region; and initiating an exothermic reaction of the braze precursor material such that the exothermic reaction produces a braze material that reaches a braze temperature above the solidus temperature of the braze material. A braze precursor material is also provided.

A method may include removing a portion of a base component adjacent to a damaged portion of the base component to define a repair portion of the base component. The base component may include a cobalt- or nickel-based superalloy, and the repair portion of the base component may include a through-hole extending from a first surface of the base component to a second surface of the base component. The method also may include forming a braze sintered preform to substantially reproduce a shape of the through-hole. The braze sintered preform may include a Ni- or Co-based alloy. The method additionally may include placing the braze sintered preform in the through-hole and heating at least the braze sintered preform to cause the braze sintered preform to join to the repair portion of the base component and change a microstructure of the braze sintered preform to a brazed and diffused microstructure.

A blade for a gas turbine includes a removed portion space, and further includes an airfoil portion defining the removed portion space, the airfoil portion formed from a base material, and a replacement component formed to fill the removed portion space. The replacement component is formed from a material that includes 50%-80% base material, 0%-30% braze material, and 0%-8% aluminum. A braze joint is formed between the airfoil portion and the replacement component to attach the replacement component to the airfoil portion and fill the removed portion space.

A closure element for an internal passage in a component, and a related method and turbine blade or nozzle are disclosed. The closure element includes a spherical body made of a first superalloy, and a plurality of extensions extending from a surface of the spherical body. The plurality of extensions made of the same, similar or different material other than the first superalloy. Subjecting the component to at least one thermal cycle causes a braze material to form a metallurgical bond with the spherical body, the plurality of extensions and the passage wall to seal the internal passage.