Abstract: The invention is directed to a method for cleaning a component from powder residues of an additive layering method using a cleaning device, wherein a machine plate and the component arranged thereon are excited during a cleaning process by a vibration actuator of the cleaning device with a set resonance frequency of the machine plate to carry out a mechanical vibration. It is provided that the machine plate is excited by predefined vibration movements of the at least one vibration actuator to the predefined mechanical vibration, wherein the predefined vibration movements of the at least one vibration actuator occur in parallel to a main plane of the machine plate. The invention also relates to a cleaning device for cleaning an at least partially additively manufactured component, in particular a component of a turbomachine.

Abstract: The invention is directed to a method for cleaning powder residues of an additive layer build-up method away from an at least partly additively manufactured component by a cleaning device, wherein a machine plate and the component arranged thereon are excited to mechanical oscillation during a cleaning process by a vibration actuator of the cleaning device with a set resonant frequency of the machine plate. According to the invention, before the cleaning process is carried out, a resonant frequency of the machine plate is set to the set resonant frequency by an arrangement of a mass element on a securing element of the machine plate.

Abstract: Described is a heat-protection element (50) for a gas turbine (10), in particular an aircraft gas turbine, the heat-protection element (50) being adapted to at least partially surround a bearing chamber (60) of the gas turbine (10) and having at least one connecting portion (52) which is disposed in an axially forward region (VB) and connectable or connected by a material-to-material bond to a protective element (54) of a seal carrier, in particular a seal carrier with a carbon seal, at least one supporting portion (58) which is disposed in an axially central region (MB) and adapted to support the heat-protection element (50) radially on the bearing chamber (60), an end portion (64) which is disposed in an axially rearward region (HB) and forms a free end (66) of the heat-protection element (50) and which is configured such that the end portion surrounds (64) the bearing chamber (60) in a contactless manner.

Abstract: A bearing chamber housing (20) for bearing a shaft (3) of a turbomachine (1), including a housing outer shell (21) that delimits an oil chamber (33) of the bearing chamber housing (20) radially outwardly in relation to a rotational axis (4) of the shaft (3), and a housing inner shell (22) for bearing the shaft (3). The housing inner shell (22) is radially connected to the housing outer shell (21) via support ribs (23) that in each case extend axially, at least in part, and the housing inner shell (22), the housing outer shell (21), and two support ribs (23) that are next-adjacent to one another jointly delimit a cavity (41) that is axially open at the rear, and thus lead into a rear opening (32). The rear opening (32), viewed in tangential sections, has a clearance (35) in each case that constitutes at least 50% of a circumferential distance (43) between the next-adjacent support ribs (23).

Abstract: The present invention relates to a rotating blade for a turbomachine, in particular a compressor stage or a turbine stage of a gas turbine, particularly of an aircraft engine, having a blade element for deflecting the flow, with a pressure side and a suction side, these sides being joined by a leading edge and a trailing edge, wherein a stacking axis of the blade element, in the radial direction over a radius r from a root of a blade element at r=0 to a tip of a blade element at r=H, has a course x(r) in a first downstream direction perpendicular to the radial direction and parallel to a principal axis of the turbomachine and has a course y(r) in a second direction perpendicular to the radial direction and to the first direction.

Abstract: The present invention relates to a rotating blade for a turbomachine, in particular a compressor stage or a turbine stage of a gas turbine, particularly of an aircraft engine, having a blade element (10) for deflecting the flow, with a pressure side and a suction side (11), these sides being joined by a leading edge and a trailing edge (12, 13), wherein a stacking axis (S) of the blade element (10), in the radial direction (R) over a radius r from a root of a blade element at r=0 to a tip of a blade element at r=H, has a course x(r) in a first downstream direction (X) perpendicular to the radial direction (R) and parallel to a principal axis of the turbomachine and has a course y(r) in a second direction (Y) perpendicular to the radial direction (R) and to the first direction (X).