Abstract: The invention relates to a gas turbine assembly which substantially includes at least one compressor, at least one first burner, at least one second burner that is connected downstream of the first burner, and at least one turbine that is connected downstream of the second burner. At least the first and second burner form a component of a tubular or quasi-tubular combustion chamber element in the flow direction of the combustion path of the burners. The combustion chamber element being closed or quasi-closed and extending between the compressor and the turbine. The combustion chamber elements are arranged around the rotor of the gas turbine assembly in the shape of a ring.

Abstract: A gas turbine having a compressor, a combustor downstream from the compressor in a gas flow direction, and a turbine downstream from the combustor in the gas flow direction is described herein. The turbine includes a rotating part and a stationary part arranged around the rotating part. A gap between the rotating part and the stationary part, extends in a substantially radial direction relative to the rotation axis of the rotating part. A cooling fluid flows from the compressor to the gap, wherein at least a part of the cooling path extends in the stationary part, and wherein a pre-swirl nozzle is arranged adjacent to the gap and within the cooling path in the stationary part.

Abstract: The disclosure pertains to a cooled wall for separating a hot gas flow path of a gas turbine from a cooling flow including at least one turbulator rib extending from the wall into the cooling flow, and having a height, a width for providing heat transfer enhancement for the cooled wall. The turbulator rib has filets at its root with a filet radius. In order to increase the heat transfer enhancement of the turbulator rib, the filet at the downstream side of turbulator rib is extending into the cooled wall with a penetration depth. Further, the disclosure relates to specific embodiments in which the cooled wall with turbulator ribs is configured as the sidewall of an airfoil, a combustor wall or a heat shield.

Abstract: A gas turbine having a compressor, a combustor downstream from the compressor in a gas flow direction, and a turbine downstream from the combustor in the gas flow direction is described herein. The turbine includes a rotating part and a stationary part arranged around the rotating part. A gap between the rotating part and the stationary part, extends in a substantially radial direction relative to the rotation axis of the rotating part. A cooling fluid flows from the compressor to the gap, wherein at least a part of the cooling path extends in the stationary part, and wherein a pre-swirl nozzle is arranged adjacent to the gap and within the cooling path in the stationary part.

Abstract: The invention relates to a mounting and sealing arrangement for a guide vane of a gas turbine, which includes a vane carrier for carrying said vanes in a ring-like arrangement, whereby the vanes each comprise an airfoil and an outer diameter platform. The thermal stress is reduced and lifetime improved by providing a separate intermediate mounting element between the vane carrier and the outer diameter platform of each of the vanes, and mounting the intermediate mounting element with an outer side on the vane carrier, whereby the intermediate mounting element abuts with an inner side on the outer diameter platform in a sealing fashion.

Abstract: The invention refers to an arrangement for sealing an open cavity against hot gas entrainment. The open cavity is arranged between two axially adjacent stationary components limiting radially a hot gas path of a rotary flow machine, of which at least the downstream component carries at least one airfoil extending radially into said hot gas path. A multitude of air jets is directed into the open cavity such that hot gas flowing over the open cavity is prevented from penetrating into the open cavity completely. The invention has an arrangement of supply conduits for air jets which are directed into the open cavity with a jet momentum such that hot gas is induced not to enter the open cavity and forming at least one hot gas vortex close to the hot gas path within the open cavity and preventing hot gas from penetrating into the open cavity beyond the extension of the hot gas vortex.

Abstract: A blade is provided and includes a platform and a root configured to be connected to a blade carrier. Airfoil portions extend from opposite sides of the platform. Each airfoil portion defines an operating surface being the surface facing the other airfoil portion. An operating surface of one of the airfoil portions defines a suction side and the other operating surface of the other airfoil portion defines a pressure side.

Abstract: A gas turbine for a gas and steam turbine plant and method for operating a gas turbine is provided. The gas turbine includes a cooling system for cooling inlet air taken from compressor extractions or compressor discharge and introduced to the turbine(s) and/or combustor(s) of the gas turbine which is adapted for flexible operation in a combined cycle application as well as a simple cycle application. The cooling system includes at least a water/steam cooled cooling air cooler for use in a normal combined cycle application. The gas turbine is provided with at least one second cooling air system arranged in a bypass which is operable instead of or parallel to the water/steam cooled cooling air cooler for use in particular in a simple cycle application operation of the gas turbine.

Abstract: A gas turbine (1) includes a combustion chamber (2) followed by a stator airfoil row (4) defining a plurality of guide vanes and separated by the combustion chamber (2) by a first gap (5), and a rotor airfoil row (6) separated by the stator airfoil row (4) by a second gap (7). The stator airfoils (15) of the stator airfoil row (4) are connected to guide vane boxes (17) collecting a cooling fluid (A) and injecting it through nozzles (20) in the second gap (7) to make it to enter rotor airfoil inlets (23). The guide vane boxes (23) are provided with passages (30) connecting a zone (31) upstream of the guide vane boxes (17) to a zone (32) of the second gap (7) downstream of the guide vane boxes (32). Moreover, the mouths (3) of the passages (30) facing the rotor airfoil row (6) are closer to a hot gases path than the nozzles (20).

Abstract: The disclosure pertains to a cooled wall for separating a hot gas flow path of a gas turbine from a cooling flow including at least one turbulator rib extending from the wall into the cooling flow, and having a height, a width for providing heat transfer enhancement for the cooled wall. The turbulator rib has filets at its root with a filet radius. In order to increase the heat transfer enhancement of the turbulator rib, the filet at the downstream side of turbulator rib is extending into the cooled wall with a penetration depth. Further, the disclosure relates to specific embodiments in which the cooled wall with turbulator ribs is configured as the sidewall of an airfoil, a combustor wall or a heat shield.

Abstract: The invention refers to an arrangement for sealing an open cavity against hot gas entrainment. The open cavity is arranged between two axially adjacent stationary components limiting radially a hot gas path of a rotary flow machine, of which at least the downstream component carries at least one airfoil extending radially into said hot gas path. A multitude of air jets is directed into the open cavity such that hot gas flowing over the open cavity is prevented from penetrating into the open cavity completely. The invention has an arrangement of supply conduits for air jets which are directed into the open cavity with a jet momentum such that hot gas is induced not to enter the open cavity and forming at least one hot gas vortex close to the hot gas path within the open cavity and preventing hot gas from penetrating into the open cavity beyond the extension of the hot gas vortex.

Abstract: The invention relates to a gas turbine assembly which substantially includes at least one compressor, at least one first burner, at least one second burner that is connected downstream of the first burner, and at least one turbine that is connected downstream of the second burner. At least the first and second burner form a component of a tubular or quasi-tubular combustion chamber element in the flow direction of the combustion path of the burners. The combustion chamber element being closed or quasi-closed and extending between the compressor and the turbine. The combustion chamber elements are arranged around the rotor of the gas turbine assembly in the shape of a ring.

Abstract: A blade is provided and includes a platform and a root configured to be connected to a blade carrier. Airfoil portions extend from opposite sides of the platform. Each airfoil portion defines an operating surface being the surface facing the other airfoil portion. An operating surface of one of the airfoil portions defines a suction side and the other operating surface of the other airfoil portion defines a pressure side.

Abstract: A gas turbine (1) includes a combustion chamber (2) followed by a stator airfoil row (4) defining a plurality of guide vanes and separated by the combustion chamber (2) by a first gap (5), and a rotor airfoil row (6) separated by the stator airfoil row (4) by a second gap (7). The stator airfoils (15) of the stator airfoil row (4) are connected to guide vane boxes (17) collecting a cooling fluid (A) and injecting it through nozzles (20) in the second gap (7) to make it to enter rotor airfoil inlets (23). The guide vane boxes (23) are provided with passages (30) connecting a zone (31) upstream of the guide vane boxes (17) to a zone (32) of the second gap (7) downstream of the guide vane boxes (32). Moreover, the mouths (3) of the passages (30) facing the rotor airfoil row (6) are closer to a hot gases path than the nozzles (20).

Abstract: Suitable means for driving a flow are arranged in annular or ring-segment-shaped cavities that are formed in particular in multi-shell casings of turbomachines. Arranged inside the cavity are ejectors that are supplied via suitable means with a motive-fluid flow which in turn excites the flow, preferably a circumferential flow or a helical flow. The invention is suitable in particular for avoiding casing distortions when turbomachines are at rest.

Abstract: In order to avoid the “buckling” of multi-shell turbomachine casings during a cooling phase following the shutdown, it is proposed to provide means in order to suitably discharge warm medium which collects at a point of the casing cavity situated at the highest geodetic level. In an embodiment, it is proposed to allow the corresponding lines to open out in the stack of a power plant; the stack draft effect assists the flow even further. Regulating members are advantageously provided in the lines in order to shut off these lines during operation of the turbomachine on the one hand and in order not to arrange the flow through the casing cavities too intensively on the other hand.

Abstract: In order to avoid the “buckling” of multi-shell turbomachine casings during a cooling phase following the shutdown, it is proposed to provide means in order to suitably discharge warm medium which collects at a point of the casing cavity situated at the highest geodetic level. In an embodiment, it is proposed to allow the corresponding lines to open out in the stack of a power plant; the stack draft effect assists the flow even further. Regulating members are advantageously provided in the lines in order to shut off these lines during operation of the turbomachine on the one hand and in order not to arrange the flow through the casing cavities too intensively on the other hand.

Abstract: Suitable means for driving a flow are arranged in annular or ring-segment-shaped cavities that are formed in particular in multi-shell casings of turbomachines. Arranged inside the cavity are ejectors that are supplied via suitable means with a motive-fluid flow which in turn excites the flow, preferably a circumferential flow or a helical flow. The invention is suitable in particular for avoiding casing distortions when turbomachines are at rest.

Abstract: A rotor (1) of a turbomachine comprises a basic rotor element (2) and a multiplicity of rotor blades (3, 3′, 3″, 3′″). The rotor blades (3, 3′, 3″, 3′″) are arranged in a distributed manner in at least one row on the circumference of the basic rotor element (2). An area for a row of guide vanes is furthermore provided in front of and/or behind the row of rotor blades. A heat shield element (4) or a plurality of heat shield elements (4′, 4′, 4′″) lined up on the circumference of the rotor (1) is arranged in such a way between the basis rotor element (2) and the rotor blades (3, 3′, 3″, 3′″) of the row of rotor blades that the basic rotor element (2) is completely covered in the areas of the row of rotor blades and the row of guide vanes.

Abstract: In a gas turbine having sequential combustion and two combustion chambers and in each case associated turbines, a secondary guide row (6a-c) formed in an annular transition duct (1) is arranged on the downstream side of the first turbine before entry into the second combustion chamber. A number of blades of this secondary guide row effect an irrotational flow with regard to the mixing elements, which are located in the second combustion chamber and in the flow plane of the blades, which in each case are combined to form an assembly. Due to the irrotational incident flow in the second combustion chamber, the mixing elements produce a uniform separate swirl flow which optimizes the subsequent combustion, designed for self-ignition, with regard to efficiency and pollutant emissions.