Abstract: The invention concerns an aircraft engine with an intake cone and a fan stage coupled thereto with a plurality of fan blades, wherein the fan blades are each connected to a drive shaft via a connecting means, wherein a sealing device for blocking air is arranged between a blocking air space and the interior of the intake cone, and during operation of the aircraft engine a leakage air stream flows out of the blocking air space via the sealing device, characterized by at least one outflow means for the leakage air stream in a component of the aircraft engine to allow a flow of the leakage air stream into the interior of the intake cone.

Abstract: A method for determining a shape of a mistuned blade of an impeller includes: providing a nominal blade with radially threaded-on profile elements; providing an initially mistuned blade with profile elements, the mistuned blade having a reduced or increased thickness at at least one of the profile elements compared with the nominal blade; determining the shape of the nominal blade in a hot state; determining the shape of the mistuned blade in the hot state; comparing the two shapes and determining a parameter of the mistuned blade which differs from the corresponding parameter for the nominal blade in the hot state; and changing the value of this parameter for the mistuned blade in the cold state such that the difference between the parameter for the mistuned blade and the nominal blade in the hot state no longer exists.

Abstract: A fan of a gas turbine engine, which has a fan disk with a multiplicity of fastening elements which are spaced apart in a circumferential direction and which project axially forwardly from the fan disk, and a nose cone which is arranged upstream of the fan disk and which is connected by means of the fastening elements to the fan disk. On an axially front side of the fan disk, there is formed a periphery which runs in encircling fashion in the circumferential direction and which runs radially at the inside in relation to the axially rear end region of the nose cone, wherein the periphery which runs in encircling fashion in the circumferential direction forms a concave indentation, in such a way that water which ingresses into a gap between the axially rear edge of the nose cone and the fan disk passes into the concave indentation.

Abstract: The invention concerns an aircraft engine with an intake cone and a fan stage coupled thereto with a plurality of fan blades, wherein the fan blades are each connected to a drive shaft via a connecting means, wherein a sealing device for blocking air is arranged between a blocking air space and the interior of the intake cone, and during operation of the aircraft engine a leakage air stream flows out of the blocking air space via the sealing device, characterized by at least one outflow means for the leakage air stream in a component of the aircraft engine to allow a flow of the leakage air stream into the interior of the intake cone.

Abstract: A gas turbine engine has an engine core, a fan arranged upstream of the engine core, and a hollow low-pressure shaft. The low-pressure shaft includes axially front and rear ends, wherein hot compressor air is applied to the axially rear end during operation. A valve is integrated into the interior of the low-pressure shaft, configured to open or close in accordance with the rotational speed of the low-pressure shaft, wherein the valve is closed from a predefined rotational speed and is open below this rotational speed, and wherein the valve, in the open state, allows hot compressor air to flow from the axially rear end of the low-pressure shaft to the axially front end of the low-pressure shaft and, in the closed state, prevents hot compressor air from flowing through the low-pressure shaft. A mechanism, when the valve is open, feeds hot compressor air outside of the fan disk.

Abstract: A gas turbine engine has an engine core, a fan arranged upstream of the engine core, and a hollow low-pressure shaft. The low-pressure shaft includes axially front and rear ends, wherein hot compressor air is applied to the axially rear end during operation. A valve is integrated into the interior of the low-pressure shaft, configured to open or close in accordance with the rotational speed of the low-pressure shaft, wherein the valve is closed from a predefined rotational speed and is open below this rotational speed, and wherein the valve, in the open state, allows hot compressor air to flow from the axially rear end of the low-pressure shaft to the axially front end of the low-pressure shaft and, in the closed state, prevents hot compressor air from flowing through the low-pressure shaft. A mechanism, when the valve is open, feeds hot compressor air outside of the fan disk.

Abstract: A fan of a gas turbine engine, which has a fan disk with a multiplicity of fastening elements which are spaced apart in a circumferential direction and which project axially forwardly from the fan disk, and a nose cone which is arranged upstream of the fan disk and which is connected by means of the fastening elements to the fan disk. On an axially front side of the fan disk, there is formed a periphery which runs in encircling fashion in the circumferential direction and which runs radially at the inside in relation to the axially rear end region of the nose cone, wherein the periphery which runs in encircling fashion in the circumferential direction forms a concave indentation, in such a way that water which ingresses into a gap between the axially rear edge of the nose cone and the fan disk passes into the concave indentation.

Abstract: It is described a containment casing, in particular a containment casing of a gas turbine engine of an aircraft engine, which at least regionally has a structure selected from at least one of the structure types beam structure, cylinder structure, strips cage structure, foam structure, honeycomb structure, corrugated structure or net structure, each of which is formed from SMA. Furthermore, a gas turbine engine including an impactor containment casing is proposed.

Abstract: A fan housing of a turbofan engine that forms an internal space surface at the inner side, delimiting a flow path through the fan of the turbofan engine radially outside, wherein the fan housing has a beginning of the housing that is arranged upstream. It is provided that a divergent cross-sectional surface extension of the flow path is realized by the internal space surface of the fan housing where it directly adjoins the beginning of the housing, and that the internal space surface of the fan housing is suited for continuously extending an inlet diffuser of an engine inlet, which is arranged upstream of the fan housing, into the area of the fan housing. The invention further relates to an engine assembly with a fan housing and an engine inlet.

Abstract: A flow guiding device in a turbomachine with at least one rotor module that is arranged on a shaft is provided. The rotor module has at least two rotor discs between which respectively one rotor cavity is arranged. The rotor cavities are connected to each other in an air-guiding manner for an airflow by means of openings inside the rotor discs and at least one pressure sink device for creating a localized negative pressure is arranged inside the last rotor cavity as viewed in the main flow direction of the turbomachine.

Abstract: A gas turbine having a rotor which includes a turbine rotor, a shaft and a compressor rotor, the turbine rotor having at least one rotor disk and a rotor cone leading from the or a rotor disk to the shaft, the downstream end of the shaft being rotatably supported in a bearing having a bearing chamber, the interior space of the shaft being designed as a flow channel for bearing-chamber sealing air, and the space surrounding the rotor cone upstream of the same being designed as a flow space for cooling air is disclosed. In the region of the rotor connection, the shaft exhibits an expanded portion, at whose upstream end, openings are provided to allow cooling air to enter, and, at whose downstream end, openings are provided to allow cooling air to exit into the space between the bearing chamber and the rotor cone, a wall separating the streams of the cooling air and of the sealing air in the shaft interior from one another.

Abstract: A damping arrangement for guide vanes, in particular for guide vanes of a gas turbine or an aircraft engine, is disclosed. Referring to this damping arrangement, the radially external ends of the guide vanes of a guide vane grid or a guide vane ring are mounted to a housing. The radially internal ends of the guide vanes form an inner shroud. At least one seal bearing is mounted to the inner shroud of the guide vanes. At least one spring element is installed between the inner shroud of the guide vanes and the, or each, seal bearing. The, or each, spring element is configured as a leaf spring.

Abstract: A gas turbine having a rotor which includes a turbine rotor, a shaft and a compressor rotor, the turbine rotor having at least one rotor disk and a rotor cone leading from the or a rotor disk to the shaft, the downstream end of the shaft being rotatably supported in a bearing having a bearing chamber, the interior space of the shaft being designed as a flow channel for bearing-chamber sealing air, and the space surrounding the rotor cone upstream of the same being designed as a flow space for cooling air is disclosed. In the region of the rotor connection, the shaft exhibits an expanded portion, at whose upstream end, openings are provided to allow cooling air to enter, and, at whose downstream end, openings are provided to allow cooling air to exit into the space between the bearing chamber and the rotor cone, a wall separating the streams of the cooling air and of the sealing air in the shaft interior from one another.

Abstract: A method provides fire protection for titanium components of an aircraft gas turbine. To enable the use of titanium components on aircraft gas turbines without the risk of titanium fire, a surface-near, nonflammable, intermetallic phase of titanium and at least one further metal, selected from the group of aluminum, nickel, vanadium and chromium, including combinations thereof, is produced in the surface of the component by a diffusion process.

Abstract: A rotor of a turbine machine having several rotor blade rings positioned in succession in the axial direction, is disclosed. Each rotor blade ring has a rotor base body with several rotor blades anchored by blade footings in axial grooves of the rotor base body that extend essentially in the axial direction and, in their anchored position in the rotor base body, are secured against displacement in the axial direction. The rotor blades of at least one rotor blade ring have hook-like securing elements, which, in the anchored position on a low-pressure side of the rotor blade ring, abut the rotor base body. The hook-like securing elements, between the rotor base body and a sealing element, which extends between the rotor base body and a rotor base body of a rotor blade ring that connects on the low-pressure side, are fixed in the axial direction on both sides.

Abstract: A damping arrangement for guide vanes, in particular for guide vanes of a gas turbine or an aircraft engine, is disclosed. Referring to this damping arrangement, the radially external ends of the guide vanes of a guide vane grid or a guide vane ring are mounted to a housing. The radially internal ends of the guide vanes form an inner shroud. At least one seal bearing is mounted to the inner shroud of the guide vanes. At least one spring element is installed between the inner shroud of the guide vanes and the, or each, seal bearing. The, or each, spring element is configured as a leaf spring.

Abstract: For precision balancing of the rotor (3) that, in fully assembled condition of a gas turbine engine, is encompassed by an inner and an outer case, inserts (7) with a through hole (11) and a female threaded section are positioned at an even spacing in at least one circumferential groove (6) of the rotor disks (5) that receives the roots of the rotor blades (4). A temporarily installed guide tube (15) that runs through the openings (13, 14) connects to the through hole that in a certain rotor position is in line with the openings (13, 14) in the inner and outer cases (1, 2) for inserting and screwing in a balancing screw (16).