Abstract: A sensor device determines measured values of a property of a fluid, in particular of a gas, in a cavity of a gas turbine engine having a duct for carrying the fluid from the cavity to a sensor element. A data processing device is coupled to the sensor element and processes the measured values. The data processing device has a device for detecting changes in the measured values with respect to time, and an evaluation device, by which the changes in the measured values with respect to time can be detected. If there is a deviation in the changes in the measured values with respect to time from a predefined criterion, a signal relating to an at least partial blockage of the at least one inlet duct can be output. A measurement method is also disclosed.

Abstract: A combustion chamber assembly group, and a mounting method therefor, includes a combustion chamber for an engine that includes a curved combustion chamber wall extending along two spatial directions, and a combustion chamber shingle affixed at an inner side of the combustion chamber wall and having a shingle edge defining the outer contour of the shingle. For an at least sectional abutment of the shingle edge at the combustion chamber wall with a minimum clamping force in an operational state of the engine, the shingle is mounted to the combustion chamber wall in a mounting state in which the shingle at least at one section of the shingle edge has a curvature with respect to at least one of the spatial directions that differs from the curvature of the combustion chamber wall with respect to this spatial direction.

Abstract: A sensor device determines measured values of a property of a fluid, in particular of a gas, in a cavity of a gas turbine engine having a duct for carrying the fluid from the cavity to a sensor element. A data processing device is coupled to the sensor element and processes the measured values. The data processing device has a device for detecting changes in the measured values with respect to time, and an evaluation device, by which the changes in the measured values with respect to time can be detected. If there is a deviation in the changes in the measured values with respect to time from a predefined criterion, a signal relating to an at least partial blockage of the at least one inlet duct can be output. A measurement method is also disclosed.

Abstract: A labyrinth seal system having at least two sealing fins arranged axially one behind the other on a rotor of a turbomachine and having a radially facing run-in element for the at least two sealing fins on a stator of the turbomachine, wherein in each case one radial sealing gap exists between the run-in element and the at least two sealing fins, wherein maxima of the radial sealing gaps are arranged offset with respect to one another in a circumferential direction, such that, in the region of the maxima, a passage exists in targeted fashion, for an air flow to the sealing fins that follow downstream, in the event of rubbing against the stator. This may be applied for example in a gas turbine engine.

Abstract: A control method/system for a hybrid-electric aircraft propulsion system includes a generator, a propulsor, a controller and an electric storage unit. The generator including a gas turbine with blades separated from a casing by a clearance and driving an electric generator. An electric motor drives the propulsor. The controller controls the turbine and supply of electric power between the motor, the storage unit and the generator in response to a thrust demand and cooperates with a clearance controller: receives a command for a change in thrust demand; determines an operational profile that minimizes a function comprising a measure of fuel supplied to the turbine, transferring electric power from/to the storage unit, a difference between measures of current and demanded thrust and clearance over a time period; and operates the motor, turbine and storage unit according to the operational profile over the time period.

Abstract: A labyrinth seal system having at least two sealing fins arranged axially one behind the other on a rotor of a turbomachine and having a radially facing run-in element for the at least two sealing fins on a stator of the turbomachine, wherein in each case one radial sealing gap exists between the run-in element and the at least two sealing fins, wherein maxima of the radial sealing gaps are arranged offset with respect to one another in a circumferential direction, such that, in the region of the maxima, a passage exists in targeted fashion, for an air flow to the sealing fins that follow downstream, in the event of rubbing against the stator. This may be applied for example in a gas turbine engine.

Abstract: A diffuser component for a gas turbine is provided, where a fluid flow in the direction of a combustion chamber of the gas turbine can be slowed down by the diffuser component, and a flow cross-section of the diffuser component, which is defined by a diffuser wall is widened to do so. The diffuser wall is at least locally braced, in that at least one stiffening element with a lattice-type structure is provided on the diffuser wall.

Abstract: A system for the active setting of a radial clearance size between a blade tip of at least one compressor stage and/or at least one turbine stage of an aircraft engine and a housing that surrounds the at least one compressor stage and/or the at least one turbine stage is provided. The system comprises a model-based setting device, wherein the time-dependent clearance size can be approximated by the model-based setting device, with only those influencing variables being taken into account that have a time-dependent deformation behavior which is equal to or slower than the time-dependent deformation behavior of the housing. The clearance size in the cold state serves as the set point, with the clearance size being reduced by at least one previously saved value of a clearance proportion, which is determined at maximal thrust of the aircraft engine.

Abstract: A combustion chamber assembly group, and a mounting method therefor, includes a combustion chamber for an engine that includes a curved combustion chamber wall extending along two spatial directions, and a combustion chamber shingle affixed at an inner side of the combustion chamber wall and having a shingle edge defining the outer contour of the shingle. For an at least sectional abutment of the shingle edge at the combustion chamber wall with a minimum clamping force in an operational state of the engine, the shingle is mounted to the combustion chamber wall in a mounting state in which the shingle at least at one section of the shingle edge has a curvature with respect to at least one of the spatial directions that differs from the curvature of the combustion chamber wall with respect to this spatial direction.

Abstract: A gas turbine combustion chamber with a combustion chamber wall and with at least one shingle that is fastened to the combustion chamber wall at a distance from the same, wherein the shingle has a shingle edge that abuts the combustion chamber wall, wherein the combustion chamber wall and the combustion chamber shingle respectively have at least one mixed air hole, and wherein a tubular fastening element provided with a ring flange is inserted through the mixed air hole of the combustion chamber wall, abutting the side of the combustion chamber wall that is facing away from the combustion chamber shingle with the ring flange, characterized in that, at its exterior side, the fastening element is provided with first snap-in means that are in mesh with second snap-in means that are formed at a tubular snap-in element that braces the combustion chamber shingle against the combustion chamber wall.

Abstract: A cooled flange connection of a gas-turbine engine is annular and includes a first flange of a first component, at least a second and central flange of a second component, and a third flange of a third component. At the contact area between the first and the second flange a first circumferential duct is provided that extends over at least part of the circumference. At the contact area between the second and the third flange a second circumferential duct is provided that extends over at least part of the circumference. The first and second circumferential ducts are connected to one another by axial connecting recesses. The first flange is provided with at least one inflow recess connected to the first circumferential duct. The third flange is provided with at least one outflow recess connected to the circumferential duct.

Abstract: A gas turbine combustion chamber with a combustion chamber wall and with at least one shingle that is fastened to the combustion chamber wall at a distance from the same, wherein the shingle has a shingle edge that abuts the combustion chamber wall, wherein the combustion chamber wall and the combustion chamber shingle respectively have at least one mixed air hole, and wherein a tubular fastening element provided with a ring flange is inserted through the mixed air hole of the combustion chamber wall, abutting the side of the combustion chamber wall that is facing away from the combustion chamber shingle with the ring flange, characterized in that, at its exterior side, the fastening element is provided with first snap-in means that are in mesh with second snap-in means that are formed at a tubular snap-in element that braces the combustion chamber shingle against the combustion chamber wall.

Abstract: A gas turbine combustion chamber with a combustion chamber wall and with at least one shingle that is fastened to the combustion chamber wall at a distance from the same, wherein the shingle has a shingle edge that abuts the combustion chamber wall, wherein the combustion chamber wall and the combustion chamber shingle respectively have at least one mixed air hole, and wherein a tubular fastening element provided with a ring flange is inserted through the mixed air hole of the combustion chamber wall, abutting the side of the combustion chamber wall that is facing away from the combustion chamber shingle with the ring flange, characterized in that, at its exterior side, the fastening element is provided with first snap-in means that are in mesh with second snap-in means that are formed at a tubular snap-in element that braces the combustion chamber shingle against the combustion chamber wall.

Abstract: An assembly group for a combustion chamber of a gas turbine includes a combustion chamber shingle with a mounting element and a bolt formed as a separate structural component for supporting the combustion chamber shingle at a wall of the combustion chamber. The mounting element has a reception area into which a bolt head of the bolt is inserted transversely with respect to a longitudinal axis of the bolt and anchored therein in a form-fit manner. At the reception area, the mounting element forms an edge area that protrudes substantially radially with respect to the longitudinal axis of the bolt and that at least partially surrounds an edge of the bolt head. The edge of the bolt head does not extend exclusively in a plane that is perpendicular to the longitudinal axis of the bolt.

Abstract: A system for the active setting of a radial clearance size between a blade tip of at least one compressor stage and/or at least one turbine stage of an aircraft engine and a housing that surrounds the at least one compressor stage and/or the at least one turbine stage is provided. The system comprises a model-based setting device, wherein the time-dependent clearance size can be approximated by the model-based setting device, with only those influencing variables being taken into account that have a time-dependent deformation behavior which is equal to or slower than the time-dependent deformation behavior of the housing. The clearance size in the cold state serves as the set point, with the clearance size being reduced by at least one previously saved value of a clearance proportion, which is determined at maximal thrust of the aircraft engine.

Abstract: A diffuser component for a gas turbine is provided, where a fluid flow in the direction of a combustion chamber of the gas turbine can be slowed down by the diffuser component, and a flow cross-section of the diffuser component, which is defined by a diffuser wall is widened to do so. The diffuser wall is at least locally braced, in that at least one stiffening element with a lattice-type structure is provided on the diffuser wall.

Abstract: A cooled flange connection of a gas-turbine engine is annular and includes a first flange of a first component, at least a second and central flange of a second component, and a third flange of a third component. At the contact area between the first and the second flange a first circumferential duct is provided that extends over at least part of the circumference. At the contact area between the second and the third flange a second circumferential duct is provided that extends over at least part of the circumference. The first and second circumferential ducts are connected to one another by axial connecting recesses. The first flange is provided with at least one inflow recess connected to the first circumferential duct. The third flange is provided with at least one outflow recess connected to the circumferential duct.