Abstract: A method for repairing or renewing cooling holes of a coated component of a gas turbine, includes the step of removing an old coating from an outer side of the component. The cooling hole, after removing the old coating, in a longitudinal section, has an old cross section which is larger than a nominal cross section which the cooling hole has in this longitudinal section in an original new state of the finished component. The method also includes applying a new coating to the component at least in the longitudinal section of the cooling hole so that the cooling hole, in the longitudinal section, has an interim cross section which is smaller than the nominal cross section. The method also includes partially removing the new coating inside the cooling hole so that the cooling hole, in the longitudinal section, has a new cross section which is about the same size as the nominal cross section.

Abstract: A component, especially a hot gas component of a turbomachine, has at least one passage (7, 7?), especially a cooling passage, which is embedded in an outer wall (5) of the component (1) of the turbomachine and basically extends parallel to the surface (6) of the component (1). The component (1) has a basic body (8) and at least one coating (9) which is applied to the basic body on the outside, and the passage (7, 7?) on one hand is formed by a cavity which is formed in the basic body (8), and on the other hand is closed off towards the surface (6) of the component (1) by the coating (9).

Abstract: A method for removing casting defects (5) from an article (1) with an oriented microstructure can include locating at least one casting defect (5) and melting the casting defect (5) locally by a heat source (7) to a depth at least as great as the casting defect (5) itself. The molten material can then be solidified epitaxially with respect to the surrounding oriented microstructure of the article (1) in a way that the resulting solidified area is substantially free of any defect.

Abstract: A gas turbine component consists of a superalloy base material with a single crystal structure and a protective MCrAlY-coating (6). The MCrAlY-coating (6) has a g/g? single crystal structure, which is epitaxial with the base material. It has be determined the critical factors for the successful epitaxial and crack-free growth of the MCrAlY-coating (6).

Abstract: In an emergency cooling system (17) for a component (1) which is subject to thermal load in operation, in, e.g., a turbine, the component (1) has a wall (3) which, in operation, is acted on by heat on a first wall side (14) and is acted on by a flow of cooling fluid (11) on a second wall side (15). The wall (3) has at least one emergency cooling opening (12) which is closed off by a plug (16) and through which cooling fluid flows from the second wall side (15) to the first wall side (14) when the plug (16) is absent. The plug (16) is designed so as to melt at a predetermined temperature. The plug (16) is a body which is produced separately from the component (1), the plug (16) being inserted into the emergency cooling opening (12), in which it is connected to the component (1).

Abstract: The present invention relates to an icosahedral, quasicrystalline compound or compound present in the form of an approximant having the nominal composition: TivCrwAlxSiyOz, in which v=60-65; w=25-30; x=0-6; Y=8-15; z=8-20; and in which the atom percent of oxygen is in the range of between 8 and 15%, and that of aluminum in the range of between 2 to 5%. Due to their layered structure and ceramic intermediate layers, compounds of this type exhibit excellent properties, in particular for use as coatings for gas turbine components, such as for example, rotor blades or guide vanes.

Abstract: A method of casting a directionally solidified (DS) or single crystal (SX) article with a casting furnace having a heating chamber (4), a cooling chamber (5), a separating baffle (3) between the both chambers includes a first step in which the shell mould (12) is filled with liquid metal (15), and the liquid metal (15) is directionally solidified by withdrawing the shell mould (12) from the heating to the cooling chamber (4, 5). An inert gas impinges from nozzles (8) arranged below the baffle (3) on the shell mould (12) and in steep transitions in outer surface area of the shell mould (12) the flow of the inert gas (9) is reduced or even stopped and when a protruding geometrical feature has passed the impingement area of the gas jets, the gas flow (9) is restored to a value adjusted to the geometry of the cast part presently passing the impingement area.

Abstract: A gas turbine component consists of a superalloy base material with a single crystal structure and a protective MCrAlY-coating. The MCrAlY-coating the MCrAlY has a ?/?-phase and single crystal structure, which is epitaxial with the base material.

Abstract: Nickel-based superalloys, for fabrication of monocrystalline turbine components to be used in industrial and aircraft turbine engines, having the following composition (in wt %): 5.6-8.1% Al, 4.1-14.1% Ru, 6.1-9.9% Ta, 3.6-7.5% Re, and the remaining balance Ni. The partitioning of alloying elements can be controlled to achieve a wide range of precipitate shapes and exceptional resistance to degradation under high temperature exposure conditions.

Abstract: A method is disclosed for recovering the properties of single crystal and directionally solidified Nickel based superalloys and articles made thereof, after use in a high temperature environment. The method comprises the steps of a stress relief treatment, a ?? rejuvenation heat treatment at a temperature below the ??-solvus temperature (Tsolvus, ??) of the superalloy, a precipitation heat treatment and an aging heat treatment.

Abstract: A method of casting a directionally solidified (DS) or single crystal (SX) article with a casting furnace having a heating chamber (4), a cooling chamber (5), a separating baffle (3) between the both chambers includes a first step in which the shell mould (12) is filled with liquid metal (15), and the liquid metal (15) is directionally solidified by withdrawing the shell mould (12) from the heating to the cooling chamber (4, 5). An inert gas impinges from nozzles (8) arranged below the baffle (3) on the shell mould (12) and in steep transitions in outer surface area of the shell mould (12) the flow of the inert gas (9) is reduced or even stopped and when a protruding geometrical feature has passed the impingement area of the gas jets, the gas flow (9) is restored to a value adjusted to the geometry of the cast part presently passing the impingement area.

Abstract: It is a method of repairing a ceramic coating of an article after use of this article in a high temperature environment disclosed. The ceramic coating is removed locally at spalled areas, at the areas where the ceramic coating is removed locally a mixture including a powder of zirconia stabilized with one or a combination of yttria, calcia, scandia, magnesia, ceria and oxides of the rare earth group, and hydrated metallic halides as a binder is applied and the applied ceramic powder is dried.

Abstract: A method for removing casting defects (5) from an article (1) with an oriented microstructure can include locating at least one casting defect (5) and melting the casting defect (5) locally by a heat source (7) to a depth at least as great as the casting defect (5) itself. The molten material can then be solidified epitaxially with respect to the surrounding oriented microstructure of the article (1) in a way that the resulting solidified area is substantially free of any defect.

Abstract: A gas turbine component consists of a superalloy base material with a single crystal structure and a protective MCrAlY-coating. The MCrAlY-coating the MCrAlY has a &ggr;/&bgr;-phase and single crystal structure, which is epitaxial with the base material.

Abstract: A gas turbine component consists of a superalloy base material with a single crystal structure and a protective MCrAlY-coating (6). The MCrAlY-coating (6) has a g/g′ single crystal structure, which is epitaxial with the base material. It has be determined the critical factors for the successful epitaxial and crack-free growth of the MCrAlY-coating (6).

Abstract: The present invention relates to an emergency cooling system (17) for a component (1) which is subject to thermal load in operation, in particular a component belonging to a turbine. The component (1) has a wall (3) which, in operation, is acted on by heat on a first wall side (14) and is acted on by a flow of cooling fluid (11) on a second wall side (15). The wall (3) has at least one emergency cooling opening (12) which is closed off by a plug (16) and through which cooling fluid flows from the second wall side (15) to the first wall side (14) when the plug (16) is absent. The plug (16) is designed so as to melt at a predetermined temperature.

Abstract: A tip material for a turbine blade or for repairing a damaged tip of a turbine blade having a metallic coating, the material used for the tip is equivalent in composition to the metallic coating material used for the turbine blade. Also disclosed is a method of manufacturing or repairing a tip of a turbine blade having a metallic coating by reforming the blade tip with a material equivalent in composition to the metallic coating material used for the turbine blade by bonding the material to the damaged tip.

Abstract: The present invention relates to an icosahedral, quasicrystalline compound or compound present in the form of an approximant having the nominal composition: TivCrwAlxSiyOz, in which v=60-65; w=25-30; x=0-6; Y=8-15; z=8-20; and in which the atom percent of oxygen is in the range of between 8 and 15%, and that of aluminum in the range of between 2 to 5%. Due to their layered structure and ceramic intermediate layers, compounds of this type exhibit excellent properties, in particular for use as coatings for gas turbine components, such as for example, rotor blades or guide vanes.

Abstract: A tip material for a turbine blade or for repairing a damaged tip of a turbine blade having a metallic coating, the material used for the tip is equivalent in composition to the metallic coating material used for the turbine blade. Also disclosed is a method of manufacturing or repairing a tip of a turbine blade having a metallic coating by reforming the blade tip with a material equivalent in composition to the metallic coating material used for the turbine blade by bonding the material to the damaged tip.

Abstract: A process of brazing cracks and gaps in a single crystal article which takes places isothermally under the following conditions: the temperature of the isothermal solidification is between TLiqidus, Braze+5*(wt-%BBraze) and (Tsolidus, base material−70*(wt-%BBraze)), while (wt-%B*wt-%Cr) is between 15 and 40 and (Tsolv.&ggr;′, base material−TLiqidus, Braze) is above 140° C. This results in an homogeneous &ggr;/&ggr;′-microstructure of the isothermal solidified, brazed joint with mechanical properties similar to those of the base material.