Abstract: A moving blade for a turbine or compressor stage of a gas turbine is provided, including a vane (10) and a radially outer shroud (20), for at least one contour point (PK) of an outer contour of a profile cross section (P) of the vane at a radial outer gas channel boundary, the wall thickness of the shroud at at least one edge point (PR) of the shroud, whose connecting line (V), including the contour point with a normal (N), which is perpendicular to the outer contour of the profile cross section in the contour point, encloses an angle (?), which is no more than 5°, is less than a wall thickness of the shroud at the contour point, the wall thickness of the shroud decreasing strictly monotonously along the connecting line.

Abstract: A method for manufacturing a blade or vane assembly having at least one hollow airfoil (10) for a gas turbine, profile sections (A, B, C) of this airfoil being configured (S20) on the basis of a predetermined desired torsion of the airfoil, and the airfoil being manufactured (S30) on the basis of the configured profile sections using an additive manufacturing process (S30).

Abstract: A moving blade for a turbine or compressor stage of a gas turbine is provided, including a vane (10) and a radially outer shroud (20), for at least one contour point (PK) of an outer contour of a profile cross section (P) of the vane at a radial outer gas channel boundary, the wall thickness of the shroud at at least one edge point (PR) of the shroud, whose connecting line (V), including the contour point with a normal (N), which is perpendicular to the outer contour of the profile cross section in the contour point, encloses an angle (?), which is no more than 5°, is less than a wall thickness of the shroud at the contour point, the wall thickness of the shroud decreasing strictly monotonously along the connecting line.

Abstract: A method for manufacturing a blade or vane assembly having at least one hollow airfoil (10) for a gas turbine, profile sections (A, B, C) of this airfoil being configured (S20) on the basis of a predetermined desired torsion of the airfoil, and the airfoil being manufactured (S30) on the basis of the configured profile sections using an additive manufacturing process (S30).

Abstract: A protective device (10) for a turbomachine, in particular a gas turbine, having a containment ring (12) which is disposed on the radially inner side of a casing element (14) of the turbomachine and which at least substantially prevents the casing element (14) from being penetrated by a fragment of a component of the turbomachine in the event of a structural failure of this component is provided. The containment ring (12) is composed, at least in the circumferential direction, of a plurality of ring segments (16) which are form-fittingly interconnected. A second aspect of the present invention relates to an aircraft engine.

Abstract: A component is disclosed. The component includes a base body with a first end section and a second end section. The first end section has a first locking protrusion with a first guide surface and a first guide section with a first support surface. The second end section has a second guide section with a second guide surface and a second locking protrusion with a second support surface. The first guide surface is contactable with the second guide surface with respect to a guide plane and the first support surface is contactable with the second support surface with respect to the guide plane such that the first locking protrusion is engageable with the second locking protrusion in a locked position.