Abstract: A turbine blade for a gas turbine, has a blade, with a blade wall and a crown base arranged at the free end of the blade. The blade wall surrounds at least a first and a second flow chamber, between which a first separating wall is arranged, wherein a free through-hole for connecting the first and the second flow chambers is present between the first separating wall and the crown base. Furthermore, at least one cooling air opening is present in the crown base. The cooling air opening is designed as a relief groove, which extends to a point over the through-hole at least in some sections.

Abstract: A gas turbine includes an insert element for fastening to a ring segment body of a turbine of the gas turbine. The ring segment body has a recess on a hot gas side. The insert element is designed to cover the recess. The insert element includes a cover plate, which has a concavely shaped front side and a back side, and at least one foot arranged on the back side for positioning on the ring segment body. A mounting method includes mounting the insert element where the insert element is fixed to a recess of a ring segment body of a turbine of a gas turbine.

Abstract: An insert element fastens to an annular segment body of a turbine of a gas turbine. The annular segment body has a recess on a hot-gas side. The insert element is designed to cover the recess and has a concave front side and a rear side having at least one shaped portion for positioning on the annular segment body. An annular segment has the insert element, a gas turbine has the insert element, and a mounting method mounts the insert element.

Abstract: A rotor blade for a gas turbine, the rotor blade having an aerofoil with a pressure side and an intake side which transition into a platform to which an attachment region is connected, the region having a fir tree spring which has a plurality of lateral toothed extensions arranged on the pressure side and the intake side, which extensions extend in an azimuthal direction. In addition, the azimuthal extent of one lateral toothed extension is reduced in an axial region of the lateral toothed extension, in which region the pressure load during operation exceeds a predefined threshold.

Abstract: A coolant bridging line for a gas turbine having inner and outer sides which are separated by a wall, wherein the coolant bridging line extends from a first component of the gas turbine to a second component of the gas turbine is provided herein. The coolant bridging line for a gas turbine provides a further increase in the service life of the components of the gas turbine possible with a degree of efficiency which is nevertheless as high as possible. To this end, the coolant bridging line has means which change the heat transfer between and/or the flow conditions on the inner and outer sides.

Abstract: A process for producing a layer system is provided wherein the layer system has at least a substrate, a ceramic layer, which is applied to a surface structured in a targeted manner, in which process the intermediate layer, in particular the metallic layer, is applied in such a way that the recesses form during the coating. By introducing recesses into a surface, the stresses in the ceramic layer on the metallic substrate are reduced in such a manner that a longer lifespan for the ceramic layer is achieved.

Abstract: A turbine guide vane having an aerodynamically bent vane airfoil with a channel system equipped with a throttle element is provided herein. The channel system includes channel sections for the guidance of coolant. In order to provide an alternative turbine guide vane by means of which both a partial coolant flow flowing in the interior and a partial coolant flow guided out of the turbine guide vane are adjustable; therefore, in an embodiment, the throttle element is designed for the removal of coolant.

Abstract: A hollow blade body for a hollow blade, has a blade wall which has a pressure side which has at least one first blade connection element on the inner side thereof, and a suction side which has at least one second blade connection element diametrically opposite to the first blade connection element on the inner side thereof, wherein the first blade connection element can be engaged with at least one first rib connection element which is arranged on a first longitudinal end of an insertion rib, and the second blade connection element can be engaged with at least one second rib connection element which is arranged on a second longitudinal end of the insertion rib facing away from the first longitudinal end in such a way that the insertion rib is fixed on the hollow blade body and effects a stiffening of the hollow blade.

Abstract: By continuously long engravings, the spallation behavior of the thermal barrier coating is reduced and the lifetime of the TBC is prolonged.

Abstract: A turbine blade having an internally cooled turbine blade airfoil in which a hollow space is divided by rib elements into at least one cooling duct carrying a coolant, wherein in at least one of the rib elements a separating tear initiating device for producing a separating tear is disposed, which extends at least partially in the longitudinal direction of the at least one rib element.

Abstract: A turbine rotor blade is provided with a blade root, platform adjoining it, and turbine blade on that side of the platform which faces away from the blade root, with at least one opening for feeding coolant into the turbine rotor blade interior on an underside of the blade root, which opening merges into a coolant duct. An axial rotor section for a rotor is provided, having an outer circumferential surface adjoining two end-side first side surfaces with rotor blade holding grooves distributed over the circumference and extending along an axial direction, wherein a turbine rotor blade is arranged in every holding groove, wherein a multiplicity of sealing elements are at the side of a side surface of the rotor section, and lie opposite the end sides of blade roots to form a gap. Multiple outlet holes for impingement cooling of the sealing elements are provided in the end surface.

Abstract: A turbine rotor blade for a thermal continuous flow machine, wherein a transition region and an aerodynamically shaped turbine blade connected to same, follows a blade foot for securing the turbine rotor blade to a rotor along a notional blade longitudinal axis of the turbine rotor blade from bottom to top. The blade foot has two flat end surfaces facing one another, two contoured side surfaces facing one another and joining the two end surfaces to one another, in which side surface at least one respective carrying edge is formed by creating dovetail- or fir tree-like end surface contour. The carrying edges become free edges or extend into the transition region via concave rounded portions. A channel bordering one of the two end sides is arranged in at least one concave rounded portion, the extension of which channel along the side surface is less than that of the carrying edge.

Abstract: A turbine blade with an internally cooled turbine blade airfoil, in which a hollow space is divided by rib elements in at least one cooling channel carrying a coolant, wherein a recess of material, which is arranged next to at least one rib element on a turbine blade airfoil wall, is embodied such that tensions occurring within the turbine blade airfoil can be reduced in a region surrounding the at least one rib element.

Abstract: A component which can be subjected to hot gas for a gas turbine, the component having at least one wall, which includes a first surface as far as an edge, the first surface for delimiting a hot gas flow path of the gas turbine, and which includes a second surface, which adjoins the edge and is arranged transversely to the first surface. A groove for receiving a sealing element is arranged in the second surface and extends at least partially along the edge at a distance from the edge. The groove includes a groove base lying opposite the groove opening and two mutually facing side walls which adjoin the groove base and extend along the edge. At least one of the side walls has at least one groove-shaped recess.

Abstract: A component for a turbomachine, the component including a coating for minimizing thermal gradients within the component, the coating including a first coating on at least a portion of a surface of the component, wherein the first coating is substantially free from rhenium. Further, the component includes a second coating on the first coating. The second coating includes rhenium for better thermal resistance. Furthermore, the component includes a homogeneous high porosity (HHP) layer on the second coating. Finally, the component includes an abradable layer on the homogeneous high porosity layer.

Abstract: A rotor blade for a gas turbine has a pressure-side wall and a suction-side wall, a tip cap, a cavity which is formed by the inner surface of the pressure-side wall, the suction-side wall and the tip cap, and a squealer which extends radially from the pressure-side and suction-side walls, a half-space formed by the outer surface of the tip cap and the squealer, and a cooling channel which leads from the cavity to the outside of the squealer. The tip cap has a recess which extends from the half-space into the tip cap such that the recess divides the cooling channel into a first part that communicates with the cavity and a second part that communicates with the outside space.

Abstract: A turbine blade with aerodynamic aerofoil, which extends from a bottom to a top end and has two aerofoil walls extending between, which transversely thereto extend from a common leading to a common trailing edge, and with at least one platform, which extends transversely in relation to the aerofoil and is arranged at one of the two ends of the aerofoil. A number of openings are provided in the trailing edge between the bottom end and the top end. Either the opening in the trailing edge arranged closest to the platform surface extends into the platform surface or the platform surface facing the aerofoil is lowered locally in the region of that opening that is closest to the platform, and the opening opens out into a portion of the trailing edge that has become free as a result of the lowering.

Abstract: A turbine blade for a gas turbine has a securing region and a platform region that adjoins the securing region and has a platform on which a blade airfoil with a profiled cross-section and with a pressure-side wall and a suction-side wall is arranged, wherein each outer face of the pressure-side wall and/or the suction-side wall transitions into the platform surface via an outer rounded section. The turbine blade also has at least one cavity which is arranged in the blade airfoil and extends into the platform region and in which at least one rib is provided that connects the pressure-side wall to the suction-side wall, the rib extending along a longitudinal direction of the blade airfoil so as to divide the cavity. At least one slot is introduced into the rib in the platform region, the rib passing through the rib and being arranged in the longitudinal direction.

Abstract: A component (1) for a turbomachine and a method for construction thereof: The component includes an inner core (6) and a first thermal bond coating (11) on the inner core (6), and a second thermal bond coating (12) on the first thermal bond coating (11). The first thermal bond coating (11) is sandwiched between the inner core (6) and the second thermal bond coating (12). A second operating temperature (T2max) of the second thermal bond coating (12) is higher than a first operating temperature (T1max) of the first thermal bond coating (11).

Abstract: An arrangement of a plurality of cooling channels within a turbine blade conveys cooling fluid, wherein the cooling channels lead through the turbine blade, which has a blade root, a blade tip, a leading edge, and a trailing edge, to one or more cooling-fluid outlets, wherein the cooling channels are connected to each other at selected locations and extend separately from each other in other regions in such a way that, in the event of damage to the turbine blade in the region of one cooling channel, the cooling by the other cooling channels remains largely unimpaired, wherein at least one cooling channel begins in a region near the leading edge and near the blade root and leads as a diagonal channel through the turbine blade into a region near the trailing edge and near the blade tip.