Abstract: A rotor blade for a gas turbine, includes a blade extending in a radial direction, with a blade body having a peripheral wall with pressure-side and suction-side wall sections, a plate-shaped crown base connected to the peripheral wall in the region of the blade tip, and a sweep edge extending along the peripheral wall, the peripheral wall and the crown base defining a cavity in the blade body, the sweep edge being aligned on the outer side with the peripheral wall and protruding radially over the crown base, and cooling channels are embodied in the blade body, extending from the cavity to cooling fluid outlets provided in the sweep edge. At least one recess being formed in the front surface of the sweep edge, into which at least some of the cooling channels flow such that the cooling fluid outlets are entirely arranged in a bottom region of the recess.

Abstract: A wall of a hot gas part, having a first surface subjectable to a cooling fluid, a second surface located opposite of the first surface and subjectable to a hot gas and, at least one film cooling hole extending from an inlet area located within the first surface to an outlet area located within the second surface for leading the cooling fluid from the first surface to the second surface. The respective film cooling hole has a diffusor section located upstream of the outlet area, the diffusor section is bordered at least by a diffusor bottom and two opposing diffusor side walls, wherein the diffusor section has a delta wedge element for dividing the cooling fluid flow into two sub-flows and subsequent formation of a pair of delta vortices. The respective delta wedge element protrudes in a step-wise manner from the diffusor bottom and is, in a top view, triangular-shaped.

Abstract: A turbine blade or vane for a gas turbine has successively along a radial direction of the gas turbine, a root for attaching the turbine blade or vane to a carrier, a platform, an aerodynamically shaped hollow airfoil with a suction side wall and a pressure side wall extending with respect to the direction of a hot gas flow from a common leading edge to common a trailing edge and extending transversely thereof from the platform to an airfoil tip. The airfoil has at least one cooling cavity extending in a cooling fluid flow direction from a platform level to the airfoil tip, the cooling cavity in fluid connection with a number of cooling fluid outlets distributed along the trailing edge through an array of impingement cooling features located therebetween. The array extends into a region which is located radially outside the airfoil within the platform having impingement cooling features.

Abstract: A rotor blade for a gas turbine, includes a blade extending in a radial direction, with a blade body having a peripheral wall with pressure-side and suction-side wall sections, a plate-shaped crown base connected to the peripheral wall in the region of the blade tip, and a sweep edge extending along the peripheral wall, the peripheral wall and the crown base defining a cavity in the blade body, the sweep edge being aligned on the outer side with the peripheral wall and protruding radially over the crown base, and cooling channels are embodied in the blade body, extending from the cavity to cooling fluid outlets provided in the sweep edge. At least one recess being formed in the front surface of the sweep edge, into which at least some of the cooling channels flow such that the cooling fluid outlets are entirely arranged in a bottom region of the recess.

Abstract: A turbine blade or vane for a gas turbine has successively along a radial direction of the gas turbine, a root for attaching the turbine blade or vane to a carrier, a platform, an aerodynamically shaped hollow airfoil with a suction side wall and a pressure side wall extending with respect to the direction of a hot gas flow from a common leading edge to common a trailing edge and extending transversely thereof from the platform to an airfoil tip. The airfoil has at least one cooling cavity extending in a cooling fluid flow direction from a platform level to the airfoil tip, the cooling cavity in fluid connection with a number of cooling fluid outlets distributed along the trailing edge through an array of impingement cooling features located therebetween. The array extends into a region which is located radially outside the airfoil within the platform having impingement cooling features.

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 wall of a hot gas part, having a first surface subjectable to a cooling fluid, a second surface located opposite of the first surface and subjectable to a hot gas and, at least one film cooling hole extending from an inlet area located within the first surface to an outlet area located within the second surface for leading the cooling fluid from the first surface to the second surface. The respective film cooling hole has a diffusor section located upstream of the outlet area, the diffusor section is bordered at least by a diffusor bottom and two opposing diffusor side walls, wherein the diffusor section has a delta wedge element for dividing the cooling fluid flow into two sub-flows and subsequent formation of a pair of delta vortices. The respective delta wedge element protrudes in a step-wise manner from the diffusor bottom and is, in a top view, triangular-shaped.

Abstract: A film-cooled gas turbine component for a gas turbine has a surface exposed to a hot gas and a number of film-cooling openings open out, which film-cooling openings combined to form at least one row transverse to a flow direction of the hot gas. Each of the film-cooling openings has a duct section and a diffuser section having an upstream diffuser edge, two diffuser longitudinal edges and a downstream diffuser edge. At least two immediately adjacent film-cooling openings, of the respective row have their duct axes of the respective duct sections laterally inclined relative to the local flow direction of the hot gas and their diffuser sections are formed asymmetrically with respect to a projection of the duct axis, such that immediately adjacent corner regions of the respective film-cooling openings are in alignment without the respective diffuser sections making contact with one another.

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 turbine rotor blade with a blade root, the blade root has, with respect to a direction of flow of a work medium circulating around the turbine rotor blade, a front end face and a rear end face. At least one of the front end face and the rear end face has at least one recess located at a distance to an edge of the respective end face.

Abstract: A film-cooled gas turbine component for a gas turbine has a surface exposed to a hot gas and a number of film-cooling openings open out, which film-cooling openings combined to form at least one row transverse to a flow direction of the hot gas. Each of the film-cooling openings has a duct section and a diffuser section having an upstream diffuser edge, two diffuser longitudinal edges and a downstream diffuser edge. At least two immediately adjacent film-cooling openings, of the respective row have their duct axes of the respective duct sections laterally inclined relative to the local flow direction of the hot gas and their diffuser sections are formed asymmetrically with respect to a projection of the duct axis, such that immediately adjacent corner regions of the respective film-cooling openings are in alignment without the respective diffuser sections making contact with one another.

Abstract: An internally coolable component for a gas turbine with at least one cooling duct, on the inner surface of which swirl elements are arranged in the form of turbulators which extend transversely with respect to the main flow direction of a coolant is provided. In order to reduce pressure losses for the coolant in the cooling duct, pins with different heights are set up between the turbulators.

Abstract: A turbine blade includes a first cooling air duct and a second cooling air duct which is separated from the first cooling air duct by a wall and which has a main direction, wherein the first and the second cooling-air duct are connected to one another by a first opening in the wall, wherein the wall has a second opening, to permit an improved cooling air action and thus higher operating temperatures and higher efficiency of the turbine. The second opening is adjoined by a diverting duct, the main direction of which, in the region in which the diverting duct issues into the second cooling air duct, is oriented substantially parallel to the main direction of the second cooling air duct.

Abstract: A wall structure (32, 42, 68, 70, 80) with layers (A, B, C, D, E) of non-random voids (26A, 26B, 28B, 30B) that interconnect to form discretely defined tortuous passages between an interior (21) and an exterior surface (23) of the wall for transpiration cooling of the wall. A coolant flow (38) through the wall may be metered by restrictions in coolant outlets (31) and/or within the passages to minimize the coolant requirement. Pockets (44) may be formed on the exterior surface of the wall for thermal Insulation (46). The layers may be formed by lamination, additive manufacturing, or casting. Layer geometries include alternating layers (A, B, C) with different overlapping void patterns (42), 3-D lattice structures (70), and offset waffle structures (80).

Abstract: A turbine blade for a turbomachine has an outer wall which delimits an inner cavity. Cooling fluid flows in the inner cavity. A through duct is arranged in the outer wall through which the cooling fluid flows from the inner cavity to an outside of the turbine blade. The through duct is inclined with respect to a trailing edge of the turbine blade, wherein a marginal portion of an entrance of the through duct is designed on an upstream side to be sharp-edged in relation to other marginal portions of the entrance such that a separation zone of a cooling fluid flow is formed in the through duct. A pair of swirls builds up in the through duct, wherein velocity vectors of the cooling fluid flow between the swirl centers point toward a downstream side of the through duct.

Abstract: A turbine blade for a turbomachine has an outer wall which delimits an inner cavity. Cooling fluid flows in the inner cavity. A through duct is arranged in the outer wall through which the cooling fluid flows from the inner cavity to an outside of the turbine blade. The through duct is inclined with respect to a trailing edge of the turbine blade, wherein a marginal portion of an entrance of the through duct is designed on an upstream side to be sharp-edged in relation to other marginal portions of the entrance such that a separation zone of a cooling fluid flow is formed in the through duct. A pair of swirls builds up in the through duct, wherein velocity vectors of the cooling fluid flow between the swirl centers point toward a downstream side of the through duct.

Abstract: An internally coolable component for a gas turbine with at least one cooling duct, on the inner surface of which swirl elements are arranged in the form of turbulators which extend transversely with respect to the main flow direction of a coolant is provided. In order to reduce pressure losses for the coolant in the cooling duct, pins with different heights are set up between the turbulators.

Abstract: A wall structure (32, 42, 68, 70, 80) with layers (A, B, C, D, E) of non-random voids (26A, 26B, 28B, 30B) that interconnect to form discretely defined tortuous passages between an interior (21) and an exterior surface (23) of the wall for transpiration cooling of the wall. A coolant flow (38) through the wall may be metered by restrictions in coolant outlets (31) and/or within the passages to minimize the coolant requirement. Pockets (44) may be formed on the exterior surface of the wall for thermal Insulation (46). The layers may be formed by lamination, additive manufacturing, or casting. Layer geometries include alternating layers (A, B, C) with different overlapping void patterns (42), 3-D lattice structures (70), and offset waffle structures (80).

Abstract: There is described a vane wheel of a turbine comprising at least one vane, the footing thereof being held on a wheel disk. At least one cooling channel is arranged between the wheel disk and the vane footing. The vane wheel having a plurality of turbulators is embodied on at least one of the walls of the cooling channel, the turbulators being configured in such a way that the turbulence and thus the heat transfer of a cooling fluid flowing through the cooling channel are increased.

Abstract: Layered systems in prior art are inefficient at cooling an external hot gas. The inventive layered system comprises an external porous layer, in which the pore walls of the pores have differing thicknesses. This improves the cooling action by preventing too much heat from entering the layered system.