Abstract: A compressor blade 1 of a gas-turbine engine has a flow-exposed leading edge 2. A component 3, made of wear-resistant material and featuring a higher erosive wear resistance than the base material, is attached to at least part of a length of the leading edge 2 The component 3 forming the flow-exposed leading edge 2 is strip-shaped, has an essentially rectangular cross-section and, with reference to a sectional plane perpendicular to the radial direction, is arranged only on an inflow side of the compressor blade profile in a partial area of a sectional surface of the inflow area.

Abstract: With a hybrid fan blade, the supporting structure (1) in fiber-composite material is enclosed by an enveloping structure (2) only in an area of the airfoil, with a suction-side sheet-metal cover of the enveloping structure being pre-stressed in tension and a pressure-side sheet-metal cover of the enveloping structure being pre-stressed in compression. This reduces the hazard of delamination between the enveloping structure and the supporting structure.

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 fibre composite layer and/or a plastic component.

Abstract: With a method for the manufacture of fiber-composite components, more particularly of fan blades for aircraft engines, dry fiber mats are superimposed in a lay-up mold, correctly positioned and fixed with locating pins and subsequently sewn in the mold. Metallic indicators are provided on at least one mat. The preformed entity of fiber mats is whipped at its edges with thermoplastic material and subsequently smoothed by application of heat. Remaining fiber residues are cut off. The blank thus available is infiltrated with synthetic resin in an infiltration mold or a sheathing remaining on the component. The fiber mats are checked for correct positioning by non-destructive inspection procedures. The fiber-composite component thus formed conforms to high quality requirements.

Abstract: On an intake cone made of fiber compound material for a gas turbine engine, the conical part, the attachment part and the fairing part are manufactured as one piece from a plurality of mutually crossing and covering winding layers (1) forming an interlace (2). The winding layers include parallel, alternately adjacent glass fiber strands and carbon fiber strands of equal thickness. The one-piece intake cone is producible and mountable cost-effectively and is characterized by elasticity, stiffness and impact resistance.

Abstract: On an engine shaft of hybrid design with an externally toothed power transmission element connected at the ends of a fiber-composite plastic tube, the power transmission element is of a two-part design and includes a toothed part made of steel of a specific hardness and, frontally weldedly connected thereto, a tapering adapter part enclosed by a fiber-composite material to form a scarf-type joint and having a thermal expansion approximately equal to the expansion behavior of the fiber-composite material.

Abstract: The present invention relates to a segment component in high-temperature casting material for an annular combustion chamber of an aircraft engine, characterized by a combustion-chamber wall which in operation shields a fuel flame extending along a burner axis from the environment, with the combustion-chamber wall having a bulge which points in a direction facing away from the burner axis. The invention furthermore relates to an annular combustion chamber, an aircraft engine with an annular combustion chamber as well as a method for the manufacture of an annular combustion chamber.

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 an aircraft gas-turbine compressor blade having an airfoil with a leading edge, with at least the area of the leading edge being made from a low-modulus titanium alloy.

Abstract: The present invention relates to a compressor blade of a gas turbine having an airfoil made of a fiber-reinforced plastic and a leading-edge element connected to said airfoil, characterized in that 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, as well as to a manufacturing method.

Abstract: The intake cone for a gas-turbine engine is wound in one piece from fiber compound material with fiber layers crossing one another. It is provided with a fiber compound belt, wound in the circumferential direction of the intake cone, in the connecting area with a mounting flange, which is of segmented design and thus features reduced circumferential stiffness, of a metallic retaining ring attached to a fan rotor disk. The intake cone is easily manufacturable with almost constant wall thickness and in high quality, can reliably be subjected to non-destructive testing and, in combination with the mounting flange, which is flexible in the radial direction, ensures safe connection of the two components despite the different thermal and elastic behavior of the respective materials.

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 behaviour of the smoothing material essentially equaling the X-ray absorption behaviour of the material of the component, as well as to a smoothing material for carrying out the method in accordance with claim.

Abstract: The present invention relates to a segment component in high-temperature casting material for an annular combustion chamber of an aircraft engine, characterized by a combustion-chamber wall which in operation shields a fuel flame extending along a burner axis from the environment, with the combustion-chamber wall having a bulge which points in a direction facing away from the burner axis. The invention furthermore relates to an annular combustion chamber, an aircraft engine with an annular combustion chamber as well as a method for the manufacture of an annular combustion chamber.

Abstract: A fan blade for a gas-turbine engine which comprises a supporting structure (2) in fiber-composite material is enclosed by a sheet-metal enveloping structure (4) only in an area of the airfoil (3) such that the loads and stresses acting on the blade root (11), as well as the risk of delamination caused by the high bending load in the transition area between airfoil and blade root, are minimized.

Abstract: A low-pressure turbine shaft (11) for a gas-turbine engine includes a fiber-composite plastic tube (13) with fiber layers provided in an inner wall area for taking up and transmitting torsional forces, and, in an outer wall area, with fiber layers suitably oriented for influencing stiffness and damping characteristics. The fiber layers are embedded in a high temperature resistant plastic matrix. A load input element (14) is an attachment flange (24) made of fiber-composite plastic material integrally formed on a fiber-composite plastic tube (13). A load output element (15) is a metallic driven protrusion inseparably connected to the fiber layers transmitting the torsional forces. Alternatively, the load input element may be provided as a metallic driving protrusion, which is firmly connected to the fiber layers taking up the torsional forces.

Abstract: On a low-pressure turbine shaft made of fiber-composite material, the connection to the metallic driven protrusion or the driving protrusion (1), respectively, is made via an adapter (3) formed onto the latter, with pylons (6) extending beyond the circumferential area of the latter and on which the fibers (4) are deflected and guided in accordance with the fiber orientation in the fiber-composite material, enabling high torsional forces to be transmitted via the small connecting zone.

Abstract: A method for manufacturing integrally bladed rotors by laser brazing a rotor disk (1) made of a nickel or titanium-base material to blades (3) made of titanium aluminide. A brazing metal similar to the blade material is used, which is introduced into the laser beam (4) as brazing power jet (5) and whose heat capacity, which is controlled by the melting temperature and the volume of the molten metal bath This provides for melting of the disk material and alloying with this material, but not for melting of the blade material, to which it is connected by adhesion. Such blisk features a low blade weight, reduced centrifugal effects and enhanced service life.

Abstract: A method for manufacturing a stator vane 7 of a stator vane cascade 6 of a fan structure of an aircraft gas turbine, includes circumferentially welding two essentially flat metallic sheets 10, 11. A space between the sheets 10, 11 is formed in an injection mold by a filling pressure during a plastics injection process.

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.

Abstract: During the radiographic inspection of components by X-rays for better detection of defects, preferably automated, an uneven surface topography (2) of the component (1) is covered with a smoothening layer (8) made of a material whose volume-specific radiation absorption corresponds to that of the component material, so that a decrease of radiation absorption or an increase in radiation intensity due to the uneven surface topography is reduced to accentuate a radiation effect caused by internal material defects.