Abstract: A combustor arrangement for a gas turbine includes a first burner, a first combustion chamber, a mixer for admixing a dilution gas to the gases leaving the first combustion chamber during operation, a second burner, and a second combustion chamber arranged sequentially in a fluid flow connection. These elements of the combustor arrangement are arranged in a row to form a flow path extending between the first combustion chamber and the second burner. The arrangement includes a central lance body in the flow path, extending from the first burner into the second burner, which lance body includes at least one air duct for providing air for the mixer, wherein the air is injected into the combustor through air supply elements.

Abstract: The invention refers to a sequential combustor arrangement including a first burner, a first combustion chamber, a mixer for admixing a dilution gas to the hot gases leaving the first combustion chamber during operation, a second burner, and a second combustion chamber arranged sequentially in a fluid flow connection. The mixer includes a plurality of injection pipes pointing inwards from the side walls of the mixer for admixing the dilution gas to cool the hot flue gases leaving the first combustion chamber with a low pressure drop. The disclosure further refers to a method for operating a gas turbine with such a combustor arrangement.

Abstract: The invention refers to a combustor arrangement of a gas turbine engine or power plant, having at least one combustion chamber, at least one mixer for admixing a dilution medium or air to the hot gas flow leaving the combustion chamber. The mixer is configured to guide combustion gases in a hot gas flow path extending downstream of the combustion chamber, wherein the mixer includes a plurality of injection pipes pointing inwards from the side walls of the mixer for admixing the dilution medium or air to cool the hot gas flow leaving combustion chamber. The mixer includes at least one dilution air plenum having at least one pressure-controlled compartment which is directly or indirectly connected to at least one injection pipe.

Abstract: A blade includes a leading edge, a trailing edge, a pressure surface, a suction surface, a root end, a tip end, a tip shroud attached to the tip end, the tip shroud comprising a platform and a fin, wherein the fin having a leading edge side facing towards the leading edge of the blade, a trailing edge side facing towards the trailing edge of the blade, a back end and a front end, the leading edge side and the trailing edge side extending between the back end and the front end, the fin extending across the tip end of the blade at an angle to the chord of the blade at the tip end of the blade. A first platform portion extends from the leading edge side of the fin to the suction surface. A second platform portion extends from the trailing edge side of the fin to the pressure surface at the tip end of the blade.

Abstract: The present invention generally relates to a diverting system to be used in a combined-cycle power plant. According to the invention, the diverting system is integrated within the by-pass stack and carries out, in a sole component, the functions of by-passing the hot exhaust flow in a single-cycle configuration or guiding it through the HRSG, acting as a diffusor, in an open combined-cycle configuration.

Abstract: The invention concerns a gas turbine fuel pipe, having a fuel line, the fuel line having a fuel line volume, a fuel line outer wall and an opening in the fuel line outer wall, a damper having a damper volume and a damper outer wall and attached in fluid communication with the fuel line, wherein the damper covers the opening in the fuel line outer wall, and a perforated lining extending across at least part of the opening in the fuel line outer wall.

Abstract: The application describes a method of operating a gas turbine during a cool-down phase. The gas turbine provides a compressor, a combustor downstream of the compressor, and a turbine downstream of the combustor, with the turbine providing a turbine vane carrier. The method includes feeding a flow of cooling air from the compressor to the turbine vane carrier, measuring a temperature of the flow of cooling air and measuring a temperature of the turbine vane carrier. In the method, the flow of cooling air is fed at a first flow rate when the temperature of the turbine vane carrier is lower than the temperature of the cooling air, and the flow of cooling air is fed at a second flow rate when the temperature of the turbine vane carrier is higher than the temperature of the cooling air, wherein the first flow rate is higher than the second flow rate.

Abstract: A locking pin arrangement with a basic element for receiving an insert and a fall-proof locking pin for fixing the insert in the basic element. The locking pin is at least partially inserted into a borehole of the basic element and is pushed towards an opening of the borehole by a spring. The locking pin has a guiding surface for axial guidance in the borehole, an outer thread and a constriction positioned between the outer thread and the guiding surface, wherein the diameter of the outer thread is smaller than the diameter of the guiding surface and the diameter of the locking pin at the constriction is smaller than the diameter of the outer thread. Further, the arrangement comprises an inner thread going from the wall of the borehole into the space formed by the constriction.

Abstract: A gas turbine blade includes a blade root and a blade aerofoil, a cooling fluid plenum extending inside the gas turbine blade through the blade root, the blade aerofoil and the blade tip, a blade root impingement plate in the cooling fluid plenum inside the blade root and a blade tip impingement plate in the cooling fluid plenum inside the blade tip, the blade tip impingement plate having at least one cooling fluid hole configured and arranged to enable a cooling fluid to flow from the blade tip into the blade aerofoil via the cooling fluid hole or holes, and a pipe extending in the cooling fluid plenum from the blade root impingement plate to the blade tip impingement plate. The blade root impingement plate can direct the cooling fluid from the blade root to the pipe.

Abstract: An airfoil for a gas turbine engine, the airfoil comprising a wall having a first surface, a second surface, and a passageway extending through the wall from a first opening in the first surface to a second opening in the second surface, the passageway having one or more sections between the first opening and the second opening, the one or more sections in fluid communication with each other, the plurality of sections comprising a first diffuser section providing a first change in cross-sectional area within the passageway, a second diffuser section providing a second change in cross-sectional area within the passageway, a flow conditioning section, and an edge section having two surfaces set opposite each other across the passageway, the two surfaces extending along the passageway substantially in parallel to one another, the edge section being located adjacent to the second opening.

Abstract: The present invention relation to a burner for a combustion chamber of a gas turbine with a mixing and injection device. The mixing and injection device includes a limiting wall that defines a gas-flow channel and at least two streamlined bodies, each extending in a first transverse direction into the gas-flow channel. Each streamlined body has two lateral surfaces that are arranged essentially parallel to the main-flow direction, the lateral surfaces being joined to one another at their upstream side to form a leading edge of the body and joined at their downstream side to form a trailing edge of the body. Each streamlined body has a cross-section perpendicular to the first transverse direction that is shaped as a streamlined profile.

Abstract: A combustor arrangement for a gas turbine includes a first burner, a first combustion chamber, a mixer for admixing a dilution gas to the gases leaving the first combustion chamber during operation, a second burner, and a second combustion chamber arranged sequentially in a fluid flow connection. These elements of the combustor arrangement are arranged in a row to form a flow path extending between the first combustion chamber and the second burner. The combustor arrangement includes acentral lance body arranged inside the flow path and extending from the first burner through the first combustion chamber into the mixer and into the second burner, wherein the lance body includes a fuel duct for providing fuel for the first burner and/or for the second burner.

Abstract: A coated substrate is described including a substrate material, which is coated at least in part with an oxidation-resistant coating, wherein the coating consists of a wear-resistant abrasive coating layer, which contains or consists of coated abrasive particles embedded in an oxidation-resistant matrix material, wherein at least some of the abrasive particles consist of ?-Al2O3 and the abrasive particles are coated with a first particle coating layer disposed on the abrasive particles and an optional second particle coating layer disposed on the first particle coating layer, wherein the matrix material contains or consists of the compound MCrAlY, wherein M is at least one element selected from the group consisting of Ni, Co and Fe. A method for manufacturing such a coated substrate is also disclosed.

Abstract: The invention concerns a sequential liner for a gas turbine combustor, having a sequential liner outer wall spaced apart from a sequential liner inner wall to define a sequential liner cooling channel between the sequential liner outer wall and the sequential liner inner wall. The sequential liner outer wall includes a first face, a first adjacent face and a second adjacent face, the first and second adjacent faces each being adjacent to the first face, the first face of the sequential liner outer wall having a first convective cooling hole adjacent to the first adjacent face and a second convective cooling hole adjacent to the second adjacent face, each convective cooling hole being arranged to direct a convective cooling flow into the sequential liner cooling channel adjacent to each adjacent face. The invention also concerns a method of cooling using the sequential liner and a method of retrofitting a gas turbine.

Abstract: A method is disclosed for operating an axially staged mixer arrangement with dilution air or dilution gas injection in connection with a combustion arrangement of a gas turbine engine or turbo machinery, having at least one combustion chamber, at least one mixer arrangement for admixing air or gas portions to the hot gas flow leaving the combustion chamber. The mixer arrangement includes a plurality of injectors for admixing air portions to cool at least the hot gas flow leaving combustion chamber. The spacing between the last injection location of acting injectors and at least one subsequently arranged dilution air injection, inside of the mixer arrangement in the hot gas flow, where there is no node between the injection locations, corresponds to a distance equal or approximating to half of convective wave length.

Abstract: A first stage vane arrangement having an array of first stage vanes and an array of frame segments and method for cooling frame segments of the first vane arrangement of a gas turbine are disclosed. The frame segments are designed for axially receiving aft ends of a combustor transition pieces. The first stage vanes include extended vanes, each vane having a leading section, a trailing edge, and an airfoil extending between an outer platform and an inner platform. The frame segments having an I-beam with an upper horizontal element, a lower horizontal element, and a vertical web. The vertical web having a downstream face facing towards a first stage of a turbine when installed in a gas turbine. The downstream face of the vertical web of at least one of the frame segments overlaps, at least partially, the leading section of at least one of the extended vanes.

Abstract: The invention refers to a blade of a rotary flow machine including an airfoil having a suction surface and a pressure surface joining each other along a trailing and a leading edge. A radially outward directed airfoil tip and a radially inward directed end joining an inner platform connect the airfoil to a shank at a radial end of the airfoil and providing, at least one shank pocket radially encircled by an axially extending portion of the platform. At least one radially extending rim extends from the trailing edge side of the shank and has an essentially radially orientated first slot for receiving a seal. A mount extends radially inwardly from said shank pocket. The first slot has a first aperture on a shank surface orientated in an axial direction.

Abstract: A turbine blade having an airfoil and a fir tree root is disclosed. The fir tree root has a lengthwise direction, a crosswise direction extending between two lateral sides of the fir tree root, and a span direction extending from a root base towards an airfoil tip. The fir tree root includes a web interposed between each pair of at least two neighboring channels. For each web, a web-to-channel ratio with each of the two neighboring channels is chosen to be larger than or equal to 0.5 and is smaller than or equal to 0.85 at least at a position where the root width is a minimum load bearing root width in a load bearing section of the fir tree root. In another aspect, overall web-to-channel ratio is defined as a ratio between the sum of all web lengths and the sum of all channel lengths.

Abstract: The present invention generally relates to a gas turbine and more in particular it is related to a damper assembly for a combustion chamber of a gas turbine. According to preferred embodiments, the present solution provides a damper assembly including protrusions on a wall of the neck. These protrusions result in a side wall reactance to the acoustic field that has the effect of decreasing the effective speed of sound in the neck. The decrease of the effective speed of sound in the neck is equivalent to an increase of the effective neck length.

Abstract: The present invention generally relates to a vane for a gas turbine, and more in particular it provides an innovative vane with improved flexibility leading to a reduction of stresses at the transition from the vane trailing edge to the vane platform, without interfering into the cooling scheme of such component. The present invention can increase flexibility of the vane platform by introducing on the vane platform a material cutback confined in the proximity of the trailing edge portion of the vane airfoil.