Methods for material build-up on a tip of a blade of a gas turbine engine are provided. The method can include inserting a material supply and an inflatable bladder between the tip and a shroud such that the material supply is exposed to the tip and the inflatable bladder is positioned between the material supply and a shroud, inflating the inflatable bladder to force contact between the material supply and the tip, and causing relative movement between the material supply and the tip. The relative movement, in combination with the radial biased contact between the material supply and the tip, creates heat through friction. As such, the relative movement can frictionally weld new material from the material supply onto the tip of the blade. For example, the heat created can be sufficient to melt the surface of the material supply to transfer material from the material supply to the tip.

A method of additively manufacturing is provided. The method may include successively depositing and fusing together layers of a superalloy powder mixture comprised of a base material powder and a eutectic powder, to build up an additive portion, which eutectic powder has a solidus temperature lower than the solidus temperature of the base material powder. The method may also include heat treating the additive portion at a temperature greater than 1200° C. to heal cracks and/or fill pores and to homogenize the alloy of which the additive portion is comprised. The additive portion alloy has a chemistry defined by the superalloy powder mixture. The base material powder may be formed of a nickel-base superalloy with an aluminum content by weight of at least 1.5%. The eutectic powder may be a nickel-base alloy including by weight about 6% to about 11% chromium, about 5% to about 9% titanium, and about 9% to about 13% zirconium, with balance nickel as its primary components.

A method of repairing an airfoil is provided. The method includes providing an airfoil with a damaged section and removing the damaged section by machining or cutting an upper section of the airfoil. A replacement section is configured to mate with an upper surface of the airfoil. A presintered preform is provided to join the airfoil and the replacement sections through a resistance brazing process. The presintered preform is configured to mate with the upper surface of the airfoil and a lower surface of the replacement section and inserted between this upper surface and lower surface, creating a stacked airfoil comprising three mated sections in abutting contact. The stacked airfoil is resistance brazed such that only the braze material of the presintered preform melts and the upper surface of the airfoil and the lower surface of the replacement section remain below the grain boundary temperature of the material of the airfoil.

A nozzle or blade for a turbomachine includes an airfoil body including at least one first coolant passage, and an edge opening in a leading edge or a trailing edge of the airfoil body. The edge opening has an edge coupon retention member seat in or on an inner surface of the airfoil body. An edge coupon has a shape at least partially configured for coupling to the edge opening in the airfoil body. The edge coupon includes an edge coupon body, at least one second coolant passage in the edge coupon body configured for fluid communication with the at least one first coolant passage in the airfoil body, and a retention member extending from the edge coupon body for coupling to the edge coupon retention member seat in the airfoil body.

A method of forming a component includes mixing a powdered base material and a binder to define a mixture, forming the mixture into a desired shape without melting the base material, removing the binder from the desired shape to define a skeleton, the volume of the skeleton being between 80 percent and 95 percent base material, and infiltrating the skeleton with a melting point depressant material to define a finished component, the finished component having less than one percent porosity by volume.

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.

A dual walled component includes a spar comprising a plurality of pedestals; a coversheet attached to a first set of pedestals from the plurality of pedestals; and a repaired coversheet portion attached to a second set of pedestals from the plurality of pedestals and to the coversheet, where the repaired coversheet portion includes a braze material.

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 method for repairing a tip portion of a turbine component having a structural defect is provided. The method includes removing a damaged section in turbine component with a structural defect in a tip portion of the turbine component. A pre-sintered preform is provided including a first portion having a first composition and a second portion having a second composition. The pre-sintered preform is configured to mate with an upper surface of a remaining portion of the turbine component. The method also includes applying the pre-sintered preform to the upper surface, wherein the PSP comprises a superalloy material and a braze material. The PSP and the remaining portion of the turbine component are subjected to a brazing process to melt the braze material and fill in the structural defect.

A bonding probe is used for bonding a cover sheet to a core to form or repair a dual wall structure. The bonding probe includes a body, an attachment end at a distal end of the body, a head at an opposite end of the body, a tip extending from the head, and a cooling passageway. The attachment end couples the body to a resistance welder. The tip extends to form a proximate end of the body. The tip includes a contacting area having a predetermined three dimensional contoured surface. The contacting area aligns with a predetermined area of a three dimensional outer surface of a cover sheet of a dual walled structure. The cooling passageway provides a passageway for a flow of cooling fluid. The cooling passageway extends through the head, and into the tip.