Abstract: The invention concerns a shaft component, which can be connected or is connected to the input or output side of a gear box in a gas turbine engine, in particular an aircraft engine, wherein the shaft component has partially a region comprising fiber reinforced plastic, the fibers in this region being arranged only in an angular range of +/?40° to 50°, in particular of +/?42° to 48°, most particularly +/?45°, in relation to the main axis of rotation of the shaft component. The invention also concerns a method for producing a shaft component and a gas turbine engine.

Abstract: A drive shaft component, which can be connected or is connected to the input side of a gear box in a gas turbine engine, in particular an aircraft engine, characterized in that the drive shaft component has partially a region including fiber reinforced plastic, the fibers in this region being arranged only in an angular range of +/?40° to 50°, in particular of +/?42° to 48°, most particularly +/?45°, in relation to the main axis of rotation of the drive shaft component. The invention also concerns a method for producing a drive shaft component and a gas turbine engine.

Abstract: A gear-box casing in a gas turbine engine, in particular an aircraft engine, characterized in that the gear-box casing has partially a region including fiber reinforced plastic, the fibers in this region being arranged only in an angular range of +/?40° to 50°, in particular of +/?42° to 48°, most particularly +/?45°, in relation to the main axis of rotation of the gear-box casing. The invention also concerns a method for producing a gear-box casing and a gas turbine engine.

Abstract: A bearing element for at least one adjustable guide vane of a turbomachine, in particular an aircraft engine, is provided. The bearing element comprising a porous matrix made of carbon and/or graphite and at least one metallic phase or a metal salt that is at least partially arranged inside the pores of the matrix, so that what results is a metal-infiltrated or a metal-salt-infiltrated material.

Abstract: A gas turbine having at least one disk, wherein turbine blade elements are connected via connection means to the at least one disk, wherein the connection means are arranged in the interior of the turbine blade elements in a region radially above the disk in the radial direction of the turbine blade elements, in particular in a region which is in the driving airflow during the operation of the gas turbine, and the turbine blade elements have at least two zones composed of different materials, wherein the at least two zones adjoin one another in particular in the radial direction, and in that a zone with a material suited to compressive stress, in particular a ceramic, in particular an yttrium-stabilized zirconium oxide, is arranged radially below the connection means, and a zone with a material suited to tensile stress, in particular CMSX 4, is arranged radially above the connection means.

Abstract: An abradable coating structure with at least partially closed cells made of a metallic material with a mean cell wall porosity between 7 and 50%, in particular between 20 and 40%, a mean cell wall thickness between 50 and 200 ?m, in particular between 80 and 150 ?m, a cell size with a mean free diameter between 200 ?m and 15 mm, in particular 2 mm or in particular between 8 and 12 mm and a mean carbon concentration in the material of between 0 and 5% by weight, in particular between 0.05 and 2% by weight. The invention also relates to a component with a polyhedral cell structure and to a method for manufacturing the polyhedral cell structure.

Abstract: An aircraft gas turbine with a fan disk on which are fastened fan blades spread over the circumference and forming an intermediate space with one another, with a sealing disk arranged at the rear of the fan disk and with an inlet cone mounted at the front of the fan disk, as well as with filler elements arranged in the intermediate spaces, where the sealing disk has an annular groove and where the filler elements are at the rear inserted into the annular groove and at the front held underneath a rim area of the inlet cone, characterized in that the filler elements are designed as bending beams and that on the radially inner side of the filler element at least one rib-like reinforcing area is provided, extending in the axial direction and longitudinally to the filler element.

Abstract: The present invention relates to a compressor blade of a gas turbine having an airfoil and a blade root, said blade root and airfoil being made in one piece from a fiber-composite material, as well as a leading edge protector made from a sheet metal material, characterized in that the airfoil is provided with a radial groove extending substantially over the entire length of the airfoil and that the leading edge protector is detachably clamped onto the airfoil and is fixed with a sheet metal flange in the radial groove of the airfoil.

Abstract: A method for manufacturing a metallic component especially configured and designed for a turbomachine includes a) for a precision-casting process, a wax model with a wax structure is produced, subsequently b) the tip of the wax structure is thermally and/or mechanically treated such that a region with an undercut is formed on the wax structure, subsequently c) the metallic component is manufactured from the wax model in the precision-casting process, with a component structural element with an undercut forming on a surface of the component, and d) the component structural element is provided at least partially with a ceramic coating, a plastic-containing coating, in particular a fiber composite layer and/or a plastic component.

Abstract: A bearing element for at least one adjustable guide vane of a turbomachine, in particular an aircraft engine, is provided. The bearing element comprising a porous matrix made of carbon and/or graphite and at least one metallic phase or a metal salt that is at least partially arranged inside the pores of the matrix, so that what results is a metal-infiltrated or a metal-salt-infiltrated material.

Abstract: A compressor blade of a gas turbine includes an airfoil made of a fiber-reinforced plastic and a leading-edge element connected to said airfoil. The leading-edge element includes two partial elements, which are fabricated as separate elements and connected to one another, with the leading-edge element being connected to the airfoil essentially using the clamping effect of the two partial elements.

Abstract: An aircraft gas turbine with a fan disc at which fan blades, distributed around the circumference and forming an intermediate space in between each other, are attached, with a sealing disc that is arranged on the back side of the fan disc, and with an inlet cone that is mounted at the front side of the fan disc, as well as with filling elements that are arranged in the intermediate spaces, wherein the inlet cone includes strip-shaped lugs which are configured in one piece with the same and which can be inserted into the intermediate spaces, with their free ends being inserted into a ring groove that is formed at the sealing disc, and in that a filling element is arranged on each strip-shaped lug.

Abstract: The present invention relates to a compressor blade of a gas turbine having an airfoil made of a fiber-reinforced plastic, which is fastened by means of a blade root to a disk, as well as a metallic leading-edge element, which is arranged on the leading-edge side of the airfoil and partially encompasses the latter, with the leading-edge element itself being fastened to the disk.

Abstract: The present invention relates to a compressor blade of a gas turbine having an airfoil and a blade root, said blade root and airfoil being made in one piece from a fiber-composite material, as well as a leading edge protector made from a sheet metal material, characterized in that the airfoil is provided with a radial groove extending substantially over the entire length of the airfoil and that the leading edge protector is detachably clamped onto the airfoil and is fixed with a sheet metal flange in the radial groove of the airfoil.

Abstract: The present invention relates to a method for radiographically inspecting a component by means of X-rays, where at least one component surface to be radiographed is provided with a surface structure, with at least the surface provided with the surface structure being smoothed by means of a smoothing material to level out the surface structure, with at least one organic compound and at least one metal powder being used as smoothing material, with the X-ray absorption behavior of the smoothing material essentially equaling the X-ray absorption behavior of the material of the component, as well as to a smoothing material for carrying out the method in accordance with Claim.

Abstract: An aircraft gas turbine with a fan disk on which are fastened fan blades spread over the circumference and forming an intermediate space with one another, with a sealing disk arranged at the rear of the fan disk and with an inlet cone mounted at the front of the fan disk, as well as with filler elements arranged in the intermediate spaces, where the sealing disk has an annular groove and where the filler elements are at the rear inserted into the annular groove and at the front held underneath a rim area of the inlet cone, characterized in that the filler elements are designed as bending beams and that on the radially inner side of the filler element at least one rib-like reinforcing area is provided, extending in the axial direction and longitudinally to the filler element.

Abstract: An aircraft gas turbine with a fan disc at which fan blades 31, distributed around the circumference and forming an intermediate space in between each other, are attached, with a sealing disc that is arranged on the back side of the fan disc, and with an inlet cone that is mounted at the front side of the fan disc, as well as with filling elements that are arranged in the intermediate spaces, characterized in that the inlet cone includes strip-shaped lugs which are configured in one piece with the same and which can be inserted into the intermediate spaces, with their free ends being inserted into a ring groove that is formed at the sealing disc, and in that a filling element is arranged on each strip-shaped

Abstract: A method for manufacturing a workpiece 1, 2 uses a forming tool 4, 5 having a cavity 3. An embossing element 6 profiled on at least one surface is arranged on at least one surface of the forming tool 4, 5 for producing a surface structure, and with the profile of the embossing tool 6 is imparted to the surface of the workpiece 1, 2 during a forming process.

Abstract: When manufacturing hybrid components, especially the fan blades or stator vanes of an aircraft gas turbine including a metallic enveloping structure and a supporting structure in fiber-composite material, the mating faces of the metallic structure are cleaned and roughened with plasma prior to fitting the fiber-composite material and the metallic molecules activated so that high attractive forces take effect and an intimate, extremely strong adhesive connection is produced between the metal and the fiber-composite material. Fan blades or stator vanes of the fan structure of a gas-turbine engine which are manufactured according to this method are capable of transmitting high loads and feature a long service life.

Abstract: With filmless radiographic inspection of components by means of digital X-ray technology, an uneven surface geometry of the component is smoothened by defining a digital virtual smoothening layer for better, preferably automated, recognition of defects, where the digital radiation signals generated by an X-ray detector are overlaid with digitized surface measurement signals, so that a change in absorption and intensity of radiation due to the surface topography of the component, i.e. due to an uneven surface, is compensated for and only a density caused by internal material defects is represented in the X-ray image.