Abstract: A seal arrangement for reducing the seal gaps within a rotary flow machine, preferably an axial turbomachine, has rotor blades and guide vanes, which are respectively arranged in at least one rotor blade row and guide vane row and have respective blade/vane roots (2,3) which protrude into fastening contours within the rotor blade and guide vane rows. A sealing element (4) of plastically deformable material is provided between at least two adjacent blade/vane roots (2,3) along a rotor blade row or guide vane row or between a blade/vane root (2,3) of a rotor blade or guide vane and a rotary flow machine component directly adjoining the blade/vane root.

Abstract: A seal arrangement for reducing the seal gaps within a rotary flow machine, preferably an axial turbomachine, having rotor blades and guide vanes, which are respectively arranged in at least one rotor blade row and guide vane row and have respective blade/vane roots (2,3) which protrude into fastening contours within the rotor blade and guide vane rows, is described. The invention is characterized in that a sealing element (4) in plastically deformable material is provided between at least two adjacent blade/vane roots (2,3) along a rotor blade row or guide vane row or between a blade/vane root (2,3) of a rotor blade or guide vane and a rotary flow machine component directly adjoining the blade/vane root.

Abstract: The invention discloses a process for treating a ceramic protective layer (3) which is applied to the surface (2) of a gas turbine part (1). The roughness of the ceramic protective layer (3) is reduced at at least one first location (5), and the original roughness is retained at at least one second location (6). The roughness is advantageously retained at locations (6) on a turbine blade or vane (1) which are at risk of detachment, while the roughness of the remaining surface (2) is reduced in order to reduce the heat transfer to the surrounding hot-gas flow.

Abstract: A cooled gas turbine blade with a shroud in which a cooling channel system is provided, which is closed off radially to the gas turbine blade with a cover plate is characterized in that the cover plate has a circumferential edge, along the entire extension of which the cover plate enters into a continuous shape-mated connection or a plurality of locally limited shape-mated connections with the shroud.

Abstract: A method for adjusting the size of cooling holes of a gas turbine component is disclosed in which the component is shielded with an easily removable and non- conductive shielding material without shielding the cooling holes themselves. The cooling holes are then coated by a galvanic process and the size of the cooling holes is restored to within manufacturing tolerances. The shielding material is removed and, subsequently, a heat treatment for securing a bonding of the metallic coating is applied.

Abstract: Described is a cooling device for a shroud of a gas turbine blade, having a cooling channel system provided in the shroud, which is provided radially to the gas turbine blade with cooling channels that are open on one side, and which is closed off with a cover plate. The invention is characterized in that, between the cover plate and the cooling channels that are open on one side, an impact cooling plate is provided in such a way that the impact cooling plate rests on the cooling channels that are open on one side, is pressed against them by force, and encloses a space with the cover plate.

Abstract: The invention relates to a process of repairing a MCrAlY-coating of an article, which has being exposed to the hot gases of, for example, a gas turbine. The MCrAlY-coating is examined and repaired only locally where it is needed and then, subsequently, on top of the MCrAlY-coating the article is aluminized and/or chromized, avoiding the stripping of the whole coating and re-coating over the entire surface of the article. This is for replenishing the coating of Al and/or Cr that become depleted during engine operation, in an easy, cost and time saving manner.

Abstract: The invention discloses a process for treating a ceramic protective layer (3) which is applied to the surface (2) of a gas turbine part (1). The roughness of the ceramic protective layer (3) is reduced at at least one first location (5), and the original roughness is retained at at least one second location (6). The roughness is advantageously retained at locations (6) on a turbine blade or vane (1) which are at risk of detachment, while the roughness of the remaining surface (2) is reduced in order to reduce the heat transfer to the surrounding hot-gas flow.

Abstract: A component such as a turbine blade (1) of a gas turbine is provided with an intermetallic felt (2). By providing the tip (11) of the turbine blade (1) with the intermetallic felt (2) and optionally a coating of a ceramic material (3), improved protection against thermal and mechanical effects and improved oxidation resistance can be achieved. Also conceivable would be an arrangement of the intermetallic felt (2) at the rotor (4, 4a) or stator (4, 4b) opposite from the turbine blade (1) or on the platform (12) of the turbine blade (1).

Abstract: Described is a cooling device for a shroud (1) of a gas turbine blade, having a cooling channel system provided in the shroud (1), which is provided radially to the gas turbine blade with cooling channels (4) that are open on one side, and which is closed off with a cover plate (2),

Abstract: Described is a cooled gas turbine blade with a shroud (2) in which a cooling channel system (K) is provided, which is closed off radially to the gas turbine blade with a cover plate (6).

Abstract: The invention relates to a process of repairing a MCrAlY-coating of an article, which has being exposed to the hot gases of, for example, a gas turbine. The MCrAlY-coating is examined and repaired only locally where it is needed and then, subsequently, on top of the MCrAlY-coating the article is aluminized and/or chromized, avoiding the stripping of the whole coating and re-coating over the entire surface of the article. This is for replenishing the coating of Al and/or Cr that become depleted during engine operation, in an easy, cost and time saving manner.

Abstract: A method for adjusting the size of cooling holes (4) of a gas turbine component (1) is disclosed in which the component (1) is shielded with an easy removable and non conduction shielding material (7) without shielding the cooling holes (4) itself. The cooling holes (4) are then coated by means of a galvanic process and thereby a manufacturing tolerance of the size of the cooling holes (4) is reached. The shielding material (7) is removed and, subsequently, a heat treatment for securing a bonding of the metallic coating is applied.