Abstract: A wall for a hot gas channel in a gas turbine is described, the wall having a back side and a front side and an impingement sheet having impingement cooling holes, the wall being for exposure to a hot fluid at the front side, and the wall having an array of pins attached to the back side and extending between the back side and the impingement sheet, the wall additionally having a plurality of ribs attached to the back side, each rib extending between two pins to delineate an array of cells on the back side, and/or at least one compartment wall attached to the back side to delineate compartments on the back side. Embodiments include an impingement sheet with impingement cooling holes and cooling exit holes. A gas turbine including the wall is also described.

Abstract: The invention concerns a turbine blade comprising a surface, a recess within the surface, and a damping inlay within the recess. The damping inlay comprises a chamber with a damping material, for example particles. The damping inlay should substantially maintain the aerodynamic profile of the blade to enable normal operation. A further embodiment of the invention describes the method of manufacture of a turbine blade with a damping inlay. The method comprises the steps of manufacturing a turbine blade having a surface and a recess in the surface, and providing one or more damping inlays within the recess such that the damping inlay substantially maintains the aerodynamic profile of the blade, the damping inlay comprising a chamber and a damping material disposed within the chamber.

Abstract: A turboengine blading member having a platform and an airfoil. The airfoil includes a pressure side, a suction side, a leading edge and a trailing edge. An upstream region of the airfoil extends from the leading edge in a direction towards the trailing edge and a downstream region of the airfoil extends from the trailing edge in a direction towards the leading edge. The airfoil is connected to the platform in the upstream region, and is disconnected from the platform in the downstream region, such that the downstream region cantilevers from the upstream region, whereby a gap is formed between a cross sectional face of the airfoil in the downstream region and an opposed surface of the platform facing the cross sectional face. A sealing member is provided inside airfoil and platform pockets and bridges the gap.

Abstract: The invention relates to a turbine blade, including a blade having a front edge and a rear edge, which blade transitions by means of a shaft into a blade root designed for fastening the turbine blade, and including a platform, which is arranged at the lower end of the blade in order to bound a flow channel. The platform is designed as a separate component and can be connected to the blade in a form-fit manner. Flexible application is achieved in that the platform is composed of several individual platform elements, which enclose the blade in the assembled state.

Abstract: A rotor blade arrangement (20), especially for a gas turbine, which can be fastened on a blade carrier (19) and includes in each case a blade aerofoil element (10) and a platform element (14), wherein the platform elements (14) of a blade row form a continuous inner shroud. With such a blade arrangement, a mechanical decoupling, which extends the service life, is achieved by the blade aerofoil element (10) and the platform element (14) being formed as separate elements and by being able to be fastened in each case separately on the blade carrier (19).

Abstract: Disclosed is a turboengine blading member, having at least one platform member and at least one airfoil member. A receiver through opening extends through the platform member. The connector post is received within the receiver through opening. Each of the connector post and the receiver through opening have a retainer flute provided on the circumferential wall and extending along at least a part of the circumferential extent. The retainer flutes are arranged juxtaposed each other with the open sides facing each other such as to jointly form a joint retainer cavity. First and second retainer members are bonded to each other to provide a common retainer member extending into both corresponding retainer flutes, thereby retaining the connector post within the receiver through opening and interlocking the airfoil member and the platform member.

Abstract: An internally cooled airfoil for a rotary machine, for example, a gas turbine engine includes a suction and pressure side wall each extending in an axial direction, i.e. from a leading to a trailing edge of the airfoil. A suction wall sided cooling channel and a pressure wall sided cooling channel extend in the axial direction. A feed chamber is defined between a first and second inner wall for feeding the suction wall and pressure wall sided cooling channel each by at least one through hole inside of the first and second inner wall. The suction wall sided cooling channel and the pressure wall sided cooling channel extend into the trailing edge region separately. The suction wall sided cooling channel and the pressure wall sided cooling channel join before discharging at the trailing edge.

Abstract: A cooled wall segment in the hot gas path of a gas turbine. The wall segment includes a first surface, exposed to a medium of relatively high temperature, a second surface, exposed to a medium of relatively low temperature, and side surfaces connecting the first and second surface and defining a height of the wall segment. At least one cooling channel for a flow-through of a fluid cooling medium extends through the wall segment. Each cooling channel is provided with an inlet and an outlet for the cooling medium. The at least one cooling channel includes at least two heat transfer sections, a first heat transfer section extending essentially parallel to the first surface at a first distance from the first surface and a second heat transfer section extending essentially parallel to the first surface at a second distance, whereby the second distance is less than the first distance.

Abstract: The invention relates to a component for a thermal machine, in particular a gas turbine, which includes a corner and/or edge subjected to a thermally high load. The cooling of the component is improved in a manner such that at least one cooling channel is countersunk into the surface of the component in the immediate vicinity of the corner and/or edge in order to cool the corner and/or edge.

Abstract: A stator component of a turbomachine includes at least one axially extending outer ring which serves as a frame of an inner ring composed of partial segments. The partial segments are arranged on one another such that, on the rotor side, to form a coherent circular circumferential surface in relation to the rotational movement of rotor blades. The individual partial segment is composed of a material of uniform construction or, at least in a radial direction, of multiple partial bodies constructed from different materials, such as for example ceramic, wherein a partial segment thus formed exhibits predetermined stress and/or expansion behavior as a function of the load ranges of the turbomachine.

Abstract: The present disclosure generally relates to a guide vane for a gas turbine, and provides for example an innovative guide vane with improved flexibility leading to a reduction of stresses at the interface between the vane platform and the vane carrier. Exemplary embodiments provide only circumferential line contact or point contact between the guide vane and the guide vane carrier.

Abstract: A turboengine blading member includes at least one airfoil and at least one platform provided at least one of a base and a tip of the airfoil. The airfoil has a profile body, a leading edge provided at a first side of the profile body, and a trailing edge section extending from a second side of the profile body and opposite the leading edge. The profile body is connected to the at least one platform. The trailing edge section cantilevers from the profile body and is provided without connection to the platform.

Abstract: A turboengine component is disclosed, having at least one first, receiving, member and at least one second, received, member, the receiving member having at least one receiver opening, the received member including a body and at least one fixation post extending from the body. A retainer cavity is provided with a first retainer groove at an inner surface of the receiver opening. A second retainer groove is provided on a surface of the fixation post. A retainer member has a cross section and a longitudinal extent aligned with the lengthwise extent of the retainer cavity, the retainer member being displaceable within the retainer cavity along the lengthwise extent. The retainer member may be removed and the turboengine component may be disassembled without the damaging any of the received member and the receiving member.

Abstract: Disclosed is a turbo-engine component and a method for cooling a turbo-engine component. The method includes guiding a working fluid flow along a hot gas side surface of a wall of the component and in a main working fluid flow direction, discharging a coolant discharge flow at the hot gas side surface from a coolant discharge duct provided in the wall, and supplying a coolant supply flow to the coolant discharge duct and through a coolant supply path. The method also includes discharging the coolant supply flow into the coolant discharge duct as a free jet oriented across a cross section of the coolant discharge duct, and directing the free jet onto an inner surface section of the coolant discharge duct, thus effecting impingement cooling of the inner surface section.

Abstract: Disclosed is a turbo-engine component comprising a wall, the wall comprising a hot gas side surface and a coolant side surface. At least one coolant discharge duct is provided in said wall and opening out onto the hot gas side surface at a coolant discharge opening. A coolant flow direction is defined from the interior of the coolant discharge duct towards the discharge opening, the coolant discharge duct further being delimited by a delimiting surface thereof provided inside the wall. The coolant discharge duct has a first cross sectional direction and a second cross sectional direction. The coolant discharge duct is a blind cavity and is closed towards the coolant side surface, and further a dimension of the coolant discharge duct measured in the first cross sectional direction decreases in the coolant flow direction.

Abstract: Disclosed is a method for manufacturing a blading member assembly for a fluid flow machine, wherein the blading member assembly includes at least one airfoil member and one platform member. The method includes manufacturing a platform member blank, manufacturing an airfoil member blank, and processing at least one of said blanks thus obtaining the at least one airfoil member and at least one platform member, and subsequently joining the at least one platform member and the at least one airfoil member, thus providing the blading member assembly. The method, for instance, provides larger flexibility in arranging cooling openings.

Abstract: The invention refers to a stationary gas turbine arrangement with at least one turbine stage that includes at least a first row of vanes being mounted at a stationary component arranged radially outside of the first row of vanes and extending radially into an annular entrance opening of the turbine stage facing a downstream end of a combustor. Further a method for performing maintenance work on a stationary gas turbine is described. The invention is characterized in that the stationary component provides for each vane a radially orientated through-hole designed and arranged for a radial insertion and removal of the vane, and each of said vanes comprises an airfoil having at its one end directed radially outwards a contour being adapted to close the through-hole airtight by a detachable fixation means.

Abstract: The invention concerns a turbine for a gas turbine comprising a blade, a vane and an abradable lip attached to the blade or to the vane, wherein the blade and the vane are separated by a gap and the abradable lip extends part of the distance across the gap. Embodiments include the addition of an abrasive layer attached on the other side of the gap from the abradable lip. A method of manufacturing is also described.

Abstract: A multi-airfoil guide vane unit includes a first platform member, a second platform member, and at least two airfoil members, the airfoil members having a longitudinal extent along an airfoil span width and two longitudinal ends, wherein all airfoil members are rigidly affixed to a platform member at a first longitudinal end and at least one airfoil member is rigidly affixed to both platform members at both longitudinal ends. At least one of the airfoil members is floatingly connected to one of the platform members at a second longitudinal end. A longitudinally floating joint arrangement is thus provided between the second longitudinal end of the airfoil member and the platform member, such as to allow displacement of the airfoil member relative to said platform member along the longitudinal extent of said airfoil member.

Abstract: A method for manufacturing a part by means of an additive manufacturing technique and by post additive manufacturing process steps.