Abstract: Ni base superalloys containing relatively large amounts of Al and/or Ti are known to be difficult to weld satisfactorily. As the Al and Ti content of the superalloy is increased to improve the strength, the weldability of the component is drastically reduced. It is concluded herein that reducing the ?? phase improves weldability. A stepwise, controlled heating and cooling process is described to be used in cooperation with a welding process to reduce the ?? present and thereby improve weldability.

Abstract: A spindle bolt structure is provided in a gas turbine engine, and includes a pilot region located within a bolt hole extending through a seal disk. The pilot region includes a circumferential pilot ridge located adjacent to a downstream axial face of the seal disk and a circumferential trough portion located between a bolt shoulder and the pilot ridge. The trough portion defines a trough diameter that is less than a diameter of the bolt shoulder and less than a diameter of the pilot ridge. The bolt shoulder and pilot ridge are formed with an applied compressive residual stress and are positioned for engagement with the seal disk.

Abstract: A superalloy component such as a gas turbine engine blade (40) having a ceramic thermal barrier coating (41) is repaired using a textured repair foil (30). A degraded region of the thermal barrier coating is removed (14) and the underlying superalloy material surface is prepared (16) for re-coating. The repair foil is includes a layer of boron-free braze material (34) and a layer of superalloy material (32) having a textured surface (36). The foil is brazed (18) to the prepared surface during a solution heat treatment effective to homogenize the braze (20). A new area of thermal barrier coating (46) is applied over the foil with a bond that is enhanced by the texturing of the foil surface.

Abstract: A method of making pre-sintered preforms using a mixture of base superalloy particles and titanium-containing boron and silicon free braze alloy particles, such as for the repair of superalloy gas turbine engine components. Alloy particles as large as 2 mm provide reduced shrinkage when compared to prior art preforms. Braze material compositions disclosed herein are boron and silicon free and may have melting temperature ranges as low as 10° C., and they include no element not already present in the composition of the superalloy component.

Abstract: A superalloy component such as a gas turbine engine blade (40) having a ceramic thermal barrier coating (41) is repaired using a textured repair foil (30). A degraded region of the thermal barrier coating is removed (14) and the underlying superalloy material surface is prepared (16) for re-coating. The repair foil is includes a layer of boron-free braze material (34) and a layer of superalloy material (32) having a textured surface (36). The foil is brazed (18) to the prepared surface during a solution heat treatment effective to homogenize the braze (20). A new area of thermal barrier coating (46) is applied over the foil with a bond that is enhanced by the texturing of the foil surface.

Abstract: Boron and silicon free braze alloys useful for structural repairs of superalloy gas turbine engine components. The braze alloy composition may contain only elements that are contained in the superalloy material to be repaired, and may have melting temperature ranges as low as 10° C. to facilitate producing a high strength homogenized braze joint during a solution heat treatment of the superalloy substrate material.

Abstract: Structural repair of cracks and other defects in superalloy components, such as steam or gas turbine blades in stationary or aero gas turbines, are performed by heating the blade substrate to an isothermal hold temperature below the substrate's incipient melting point and filling the crack with molten superalloy filler material. The molten filler solidifies into a casting and bonds with the component substrate at the isothermal hold temperature. Heat treatment processes are completed, so that the former crack is filled with cast superalloy material having identical or similar structural properties as the adjoining substrate superalloy material. The casting repair method may be utilized universally for all types of superalloy component defects, including those previously repaired by cosmetic, lower strength welding or brazing methods.

Abstract: Boron and silicon free braze alloys are useful for structural repair of superalloy gas turbine engine components. The braze alloy compositions include nickel, chromium, titanium, and at least one of zirconium and hafnium. All of the above elements are metallic and form ductile bonds within and across the braze interface when compared to non-metallic bonds of boron and silicon.

Abstract: A multi component braze filler alloy is described having a melting temperature less than about 1235 deg. C. and greater than about 1150 deg. C. This alloy can be processed by hot isostatic pressing (HIP) at a temperature above about 1065 deg. C. and is particularly suited for the repair of gas turbine blades and vanes, especially those made from alloy 247. The relatively low Ti content in the present braze alloy tends to form less MC carbides at the joint interface, particularly in comparison with other braze alloys high in Zr and/or Hf.

Abstract: A fatigue resistant turbine through bolt formed from a base material covered by a first surface modification and a second surface modification is disclosed. The first surface modification may be in contact with the base material and, in at least one embodiment, may be a low plasticity burnished layer that increases the residual compressive stresses on an outer surface of the turbine through bolt. The second surface modification may cover the first surface modification and, in at least one embodiment, may be a spinel oxide layer on the low plasticity burnished layer. The second surface modification may be positioned on the first surface modification or on the bare turbine through bolt contact surface without low plastiocity burnishing on the shaft of the turbine through bolt. The first and second surface modifications reduce the likelihood of fretting fatigue failures.

Abstract: Ni base superalloys containing relatively large amounts of Al and/or Ti are known to be difficult to weld satisfactorily. As the Al and Ti content of the superalloy is increased to improve the strength, the weldability of the component is drastically reduced. It is concluded herein that reducing the ?? phase improves weldability. A stepwise, controlled heating and cooling process is described to be used in cooperation with a welding process to reduce the ?? present and thereby improve weldability.

Abstract: A braze foil (10) formed of a plurality of layers (12, 14, 16) of differing compositions wherein a combined melt of the foil has a desired braze composition, and wherein each layer is sufficiently ductile to be rolled into foil form, even though the desired braze composition is too strong or brittle to be fabricated as a foil Each interface (18,20) between layers may establish a near eutectic composition for initiating melting at the eutectic temperature, with the layer thicknesses selected so that as melting progresses away from the interfaces, the near eutectic composition is maintained within the melt puddles For certain nickel-based superalloy brazing applications, a foil having a layer of pure titanium, hafnium or zirconium may be sandwiched between respective alloy layers of 5-22% chrome-balance nickel

Abstract: A superalloy component, such as gas turbine blade or vane, is structurally welded by placing the component in an isolation chamber. Inert gas is introduced into the chamber. The substrate is welded in the chamber, creating a weld zone. Pressure is applied directly on the weld zone that is greater than atmospheric pressure. Application of such pressure increases the weld zone ductility and reduces likelihood of solidification cracking and strain age cracking, compared to weld zones formed at atmospheric pressure. In some embodiments an isostatic pressure chamber is used to apply isostatic pressure on the weld zone. In other embodiments the welding is performed by laser welding or cladding, TIG welding electron beam welding or autogenous welding.

Abstract: A method of repairing a superalloy component (22) wherein a section (24) of the component containing a plurality of service-induced cracks (18, 20) is removed, then a replacement section (26) of superalloy material is installed with a structural braze joint (28) containing no boron or silicon. The replacement section may have a textured surface ((38) to enhance bonding with an overlying thermal barrier coating (42). The replacement section may be pre-formed to standardized dimensions in expectation of a typical service-induced crack pattern, and the removed section excavated accordingly. The interface between the replacement section and the component may be shaped to provide a mechanical interlock there between.

Abstract: A ternary near eutectic alloy of Ni, Ti, Cr is described having a relatively low melting temperature of approximately 1230 deg. C. or less, suitable for fusing cracks in turbine blades and vanes without substantial risk of cracking during the repair process. Such an alloy is suitable for low temperature joining or repair of turbine blades since it contains the same components as typical turbine blades and vanes without foreign elements to lower the melting point of the repaired material or adversely affect the mechanical properties of the repaired component. Exclusion of boron eliminates the formation of brittle boron compounds, detrimental to the properties of the repair or seam.

Abstract: Ni base superalloy components containing relatively large amounts of Al and Ti are known to be difficult to build up by a weld build up process without cracking. As the Al and Ti content of the superalloy is increased to improve the strength, the susceptibility to cracking is increased. It is shown herein that reducing the ?? phase in the additive built up material improves robustness against cracking. A stepwise, controlled heating and cooling process is described to be used in cooperation with an additive build up process to reduce the ?? present and thereby reduce cracking.

Abstract: A multi component braze filler alloy is described having a melting temperature less than about 1235 deg. C. and greater than about 1150 deg. C. This alloy can be processed by hot isostatic pressing (HIP) at a temperature above about 1065 deg. C. and is particularly suited for the repair of gas turbine blades and vanes, especially those made from alloy 247. The relatively low Ti content in the present braze alloy tends to form less MC carbides at the joint interface, particularly in comparison with other braze alloys high in Zr and/or Hf.

Abstract: Structural repair of cracks and other defects in superalloy components, such as steam or gas turbine blades in stationary or aero gas turbines, are performed by heating the blade substrate to an isothermal hold temperature below the substrate's incipient melting point and filling the crack with molten superalloy filler material. The molten filler solidifies into a casting and bonds with the component substrate at the isothermal hold temperature. Heat treatment processes are completed, so that the former crack is filled with cast superalloy material having identical or similar structural properties as the adjoining substrate superalloy material. The casting repair method may be utilized universally for all types of superalloy component defects, including those previously repaired by cosmetic, lower strength welding or brazing methods.

Abstract: A durable blade, a method of manufacturing it and a method of repairing blades are described and claimed. Durable blade includes an airfoil having a pressure side and a suction side, a midspan shroud located on the airfoil, at least one recess on the midspan shroud and a wear pad attached to the recess of the midspan shroud.

Abstract: Article (e.g., turbine engine fan or compressor blade) comprising a titanium alloy has a first portion with alpha+beta microstructure and a second portion with martensitic or a bimodal microstructure. The modified microstructure of the second portion is provided by selectively heating, and immediately quenching, the second portion without substantially heating the first portion. An exemplary method includes providing a near net-shaped article having a first portion (e.g., an airfoil region) and a second portion (e.g., an unfinished dovetail region). Initially, the article comprises an alpha+beta microstructure throughout. Thereafter, the second portion is selectively heated, followed by immediate quenching, without substantially heating the first portion, to modify the microstructure of the second portion to a martensitic or bimodal microstructure without substantially modifying the microstructure of the first portion. Thereafter, the second portion may be processed to a final body dimension.