Abstract: A component can be formed by a hot isostatic pressing (HIP) process but it is necessary to reinforce intricate internal structures against collapse and deformation by the hot isostatic pressing process. The present method utilises a low melting point salt or alloy reinforcement within the structure which can be released when molten through a drain from the internal structure. The reinforcement may be molten as a result of the hot isostatic process or through achieving a temperature with the component which causes the reinforcement to become molten but without damaging the component itself. The remaining parts of the reinforcement may be removed by use of a solvent or simple washing with a corrosive agent to remove any reinforcement debris.

Abstract: An assembly for manufacture of a component by hot isostatic pressing comprises a first workpiece and a second workpiece. The first workpiece and second workpiece are assembled to define a cavity. Removable tooling pieces are provided to the assembly which further define the cavity and which are shaped to apply a forging load to a powder within the cavity during hot isostatic pressing.

Abstract: In order to avoid excessive internal stress and micro structure problems inherent with previous fusion welding techniques utilised for component structure fabrication, the present method creates a relatively thin in situ brazing alloy layer upon first and second component edges which are brought together in order to create a component joint. This in situ brazing alloy layer 6, 23, 26 is created by deposition of brazing elements, such as copper or nickel, from an electrical discharge cutting process electrode depletion utilised in order to cut the component edges. A subsequent brazing technique then creates through interstitial migration between that brazing alloy layer and the underlying material substrate of the components a robust component joint. Furthermore, the in situ brazing alloy layer penetrates the respective component cut edge surface to only a limited depth such that the geometric effect is similarly limited, and the properties of the underlying component material structure are maintained.

Abstract: A method of manufacturing a fan blade (26) for a gas turbine engine by powder metallurgy comprises the steps of forming a container (52) and placing at least one metal insert (52) at a predetermined position within the container (52) and filling the container (52) with metal powder (60). The at least one metal insert (62) has a predetermined pattern of stop off material (68,70) on at least one surface of the metal insert (64,66). The container (52) is evacuated and then sealed. The container (52) is hot pressed to consolidate the metal powder (60) into a consolidated metal powder preform (72). The container (52) is removed from the consolidated metal powder preform (72). The consolidated metal powder preform (72) is heated and a fluid is supplied to the predetermined pattern of stop off material (68,70) to hot form at least a portion of the consolidated metal powder preform (72) to form the hollow metal fan blade (26).

Abstract: A method of manufacturing a component (10,20) by consolidating a metal powder comprises preparing a metal powder, the metal powder comprising metal particles (30). Depositing a coating (32) containing at least one element on the surfaces (34) of the metal particles (30) of the metal powder. Applying heat and pressure to consolidate the metal particles (30) such that the at least one element of the coating (32) on the surfaces (34) of the metal particles (30) partially diffuses into the metal particles (30) and the coated metal particles (30) are diffusion bonded together to form a cellular structure (36).

Abstract: A method of manufacturing an article by consolidating powder material comprises forming a non-deformable tool (12) and a deformable tool (14). The non-deformable tool (12) comprises a plurality of tool parts (16,18). The non-deformable tool (12) defines a first cavity (20) having an open end (22) and the first cavity (20) corresponds in shape to that of the article. The non-deformable tool (12) is encapsulated in the deformable tool (14) and the non-deformable tool (12) and the deformable tool (14) define a second cavity (24) interconnected with the first cavity (20). The first and second cavities are filled with powder material (26). The deformable tool (14) is evacuated and sealed. Heat and pressure are applied to the deformable tool (14) and non-deformable tool (12) to consolidate the powder material (26) to form an article (30) in the first cavity (20) of the non-deformable tool (12). The method is used to make compressor blades, compressor vanes or compressor blisks for gas turbine engines.

Abstract: Forming a hollow structure having an internal coating includes the steps of placing a core shaped to form the internal surface of the structure in a mould, filling the mould with a material powder, hot isostatically pressing the powder about the mould to consolidate the powder, and removing the core from the hollow structure formed, wherein a coating is applied to the core prior to placement in the mould, which coating bonds to the hollow structure formed, during the hot isostatic pressing, to form the internal coating.

Abstract: A new nickel base superalloy suitable for compressor or turbine discs of gas turbine engines with fatigue crack propagation resistance equal to Waspaloy, tensile strength higher than Waspaloy and higher operating temperature than Waspaloy or UDIMET 720 family of alloys. The nickel base superalloy has a preferred composition by weight % of 14.0-19.0% cobalt, 14.35-15.15 Chromium, 4.25-5.25 Molybdenum, 1.35-2.15 tantalum, 3.45-4.15 titanium, 2.85-3.15 aluminium, 0.01-0.025 boron, 0.012-0.033 carbon, 0.05-0.07 zirconium, 0.5-1.0 hafnium, up to 1.0 rhenium, up to 2.0 tungsten, less than 0.5 niobium, up to 0.1 yttrium, up to 0.1 vanadium, up to 1.0 iron, up to 0.2 silicon, up to 0.15 manganese and balance nickel plus incidental impurities.

Abstract: A new nickel base superalloy suitable for compressor or turbine discs of gas turbine engines with fatigue crack propagation resistance equal to Waspaloy, tensile strength higher than Waspaloy and higher operating temperature than Waspaloy or UDIMET 720 family of alloys. The nickel base superalloy has a preferred composition by weight % of 14.0-19.0% cobalt, 14.35-15.15 Chromium, 4.25-5.25 Molybdenum, 1.35-2.15 tantalum, 3.45-4.15 titanium, 2.85-3.15 aluminium, 0.01-0.025 boron, 0.012-0.033 carbon, 0.05-0.07 zirconium, 0.5-1.0 hafnium, up to 1.0 rhenium, up to 2.0 tungsten, less than 0.5 niobium, up to 0.1 yttrium, up to 0.1 vanadium, up to 1.0 iron, up to 0.2 silicon, up to 0.15 manganese and balance nickel plus incidental impurities.