Abstract: The invention relates to a guide vane assembly of a turbomachine based on a modular structure, wherein the guide vane elements include at least an airfoil, an inner platform, an outer platform, wherein the guide vane airfoil and/or platforms have at its one ending provisions for connection of the guide vane elements among each other. The connections of the guide vane elements among each other are configured as a detachable, permanent or semi-permanent fixation with respect to the radial or quasi-radial extension of the airfoil compared to the rotor axis of the turbomachine. The assembling of the airfoil with respect to at least one platform is based on a force-fit and/or a form-fit connection, or on the use of a metallic and/or ceramic fitting surface, or on force closure means with a detachable, permanent or semi-permanent fixation.

Abstract: A method for machining a component on a machine tool includes the steps of: a) inserting the component into the machine tool; b) fixing the component in said machine tool with a surface to be machined being reachable by a tool of the machine tool; c) mapping the material composition on the surface of the inserted and fixed component via a spectroscopy tool; d) determine areas for removing material from or adding material to the component having regard to the mapped material composition; e) determine a tool path for removing material from or adding material to the component; and f) removing material from or adding material to the component in the determined areas along the determined tool path via machine tool.

Abstract: A thermal machine, especially a gas turbine, includes an annular combustor which is outwardly delimited by an outer shell and an inner shell (33) and through which a hot gas axially flows. The outer shell and inner shell (33) are each provided with a concentric cooling shroud (31) which is attached at a distance on their outer side, forming a cooling passage (32) through which cooling passage (32) cooling air flows in a direction which is opposite to the hot gas flow. The cooling of the combustor is improved by at least one of the cooling shrouds (31), on the side on which the cooling air enters the cooling passage (32), having an outwardly curved, rounded inlet edge (37) for improving the inflow conditions.

Abstract: In order to provide an annular combustion chamber (3) which can be heated without thermal distortion, expands without distortion, and is reliably closed, and for repairs can be opened and permanently closed again and can be reconditioned, the inner and/or outer shell (8, 7) of an annular combustion chamber (3) is produced from a rotationally symmetrical sheet metal. For assembly, the shells are cut open along a split line (16), and the lower half of the inner and of the outer shell (8, 7) is introduced into the lower casing half of a gas turbine. After the insertion of the rotor, the two upper shell halves can likewise be inserted and be connected to the lower halves by welding. In order to improve the service life of the split line weld seam (21), film cooling of the weld seam is routed by an indentation. This may be implemented, for example, by welded-in split line elements (15) which are also suitable for retrofitting.

Abstract: A thermal machine, especially a gas turbine, includes an annular combustor which is outwardly delimited by an outer shell and an inner shell (33) and through which a hot gas axially flows. The outer shell and inner shell (33) are each provided with a concentric cooling shroud (31) which is attached at a distance on their outer side, forming a cooling passage (32) through which cooling passage (32) cooling air flows in a direction which is opposite to the hot gas flow. The cooling of the combustor is improved by at least one of the cooling shrouds (31), on the side on which the cooling air enters the cooling passage (32), having an outwardly curved, rounded inlet edge (37) for improving the inflow conditions.

Abstract: The devices according to the invention serve to direct a cooling fluid 230 in a cooled blade 210 of a turbomachine, in particular a gas turbine. In this case, the blade 210 has at least one cooling channel 213 running in the blade. According to the invention, at least one drawer 220 is arranged in at least one slot 221 of the blade 210. In a preferred embodiment of the invention, the drawer 220 arranged in the slot 221 is directly adjacent to the cover wall 212 and/or at least one side wall 211 of the blade. Furthermore, arranged in the drawer is at least one flow channel 222, which is preferably formed from a groove arranged in the drawer 220 and from the directly adjacent cover wall 212 and/or side wall 211. The flow channel 222 is connected to the cooling channel 213 by means of at least one opening 223. In addition, the cooling fluid 230 fed to the flow channel 222 flows off through a passage opening 224 in the cover wall 212 and/or the side wall 211 into the main flow flowing around the blade.

Abstract: The arrangements according to the invention and the method according to the invention serve to cool platforms in turbomachines. To this end, in two platforms 210, 210′ arranged directly next to one another and having a separating gap 211 running between the platforms, at least one cooling passage 230 is arranged along the separating gap 211. In this case, the cooling passage 230 runs approximately parallel to the surface of the platforms 210, 210′. The cooling passage 230 is preferably designed as a slit-like recess in at least one side wall adjacent to the separating gap 211. If blades 220, 220′ are arranged on the platforms 210, 210′, the cooling passage 230 expediently runs approximately centrally between the blades. In addition, the cooling passage 230 preferably has a course similar to the course of the blade profile.