Abstract: Combustor device having a twin-shell tubular casing including: an inner tubular member which extends roughly coaxial a longitudinal axis of the combustor device, delimits the combustion chamber and surrounds the burner; and an outer tubular housing which extends roughly coaxial outside of the inner tubular member. An intermediate supporting structure includes an outer annular supporting member, an inner annular supporting member, and a series of three or more oblong connecting beams angularly staggered about the longitudinal axis of combustor device in cantilever manner to stably connect the inner and outer annular supporting members.

Abstract: A gas turbine unit having a combustor having a liner, a turbine, arranged downstream of the liner along a main flow gas direction and including a plurality of first stage vanes, a rotor cover support located inwardly of the vanes, and a sealing arrangement at a combustor to turbine interface, wherein the sealing arrangement includes a first dogbone seal extending between the rotor cover support and an inner downstream end of the liner or between the rotor cover support and a bulkhead located at the inner downstream end of the liner.

Abstract: A combustion device (1) for a gas turbine includes portions (12) having an inner and an outer wall (13, 14) with an interposed noise absorption plate (15) having a plurality of holes (16). The combustion device (1) further has first passages (17) connecting zones between the inner wall (13) and the plate (15) to the inside of the combustion device (1) and second passages (21) for cooling the inner wall (13). The portions (12) also have an inner layer (22) between the inner wall (13) and the plate (15) defining inner chambers (23), each connected to at least a first passage (17), and an outer layer (24) between the outer wall (14) and the plate (15) defining outer chambers (25) connected to the inner chambers (23) via the holes (16) of the plate (15).

Abstract: The invention relates to a burner arrangement for using in a single combustion chamber or in a can-combustor comprising a center body burner located upstream of a combustion zone, an annular duct with a cross section area, intermediate lobes which are arranged in circumferential direction and in longitudinal direction of the center body. The lobes being actively connected to the cross section area of the annular duct, wherein a cooling air is guided through a number of pipes within the lobes to the center body and cools beforehand at least the front section of the center body based on impingement cooling. Subsequently, the impingement cooling air cools the middle and back face of the center body based on convective and/or effusion cooling. At least the back face of the center body includes on the inside at least one damper.

Abstract: A turboengine as disclosed includes an outer wall structure and an inner wall structure, wherein the inner wall structure is provided at a radially inner position with respect to the outer wall structure, and each of the wall structures has a surface, the surfaces being arranged facing each other in the radial direction. At least one guide vane member includes at least one airfoil, a radially inner end and a radially outer end. The inner wall structure and the outer wall structure are jointly provided as a vane carrier unit, wherein the inner wall structure and the outer wall structure are fixedly connected to each other by at least one bridging member extending between the inner wall structure and the outer wall structure.

Abstract: A burner for a combustion chamber of a gas turbine with a mixing and injection device, which includes a limiting wall that defines a gas-flow channel and at least two streamlined bodies. Each streamlined body extends in a first transverse direction into the gas-flow channel, and has two lateral surfaces that are arranged essentially parallel to the main-flow direction. The lateral surfaces are joined to one another at their upstream and downstream sides to form leading and trailing edges of the body, respectively. At least one of the streamlined bodies includes a mixing structure and at least one fuel nozzle at its trailing edge for introducing at least one fuel essentially parallel to the main-flow direction into the flow channel. At least two of the streamlined bodies have different lengths along the first transverse direction such that they may be used for a can combustor.

Abstract: A burner for a combustion chamber of a gas turbine with a mixing and injection device, which includes a limiting wall that defines a gas-flow channel and at least two streamlined bodies. Each streamlined body extends in a first transverse direction into the gas-flow channel, and has two lateral surfaces that are arranged essentially parallel to the main-flow direction. The lateral surfaces are joined to one another at their upstream and downstream sides to form leading and trailing edges of the body, respectively. At least one of the streamlined bodies includes a mixing structure and at least one fuel nozzle at its trailing edge for introducing at least one fuel essentially parallel to the main-flow direction into the flow channel. At least two of the streamlined bodies have different lengths along the first transverse direction such that they may be used for a can combustor.

Abstract: The invention relates to a damper for reducing the pulsations in a combustion chamber of a gas turbine. The damper includes a resonator cavity with an inlet and a neck tube in flow communication with the interior of the combustion chamber and resonator cavity, and a compensation assembly pivotably connected with the neck tube. The compensation assembly is inserted between the resonator cavity and the combustion chamber to permit relative rotation between the combustion chamber and the resonator cavity. With the damper according to the present invention, by way of providing the compensation assembly, it is assured the relative rotation between the combustion chamber and the resonator cavity is compensated, hence operation life is elongated.

Abstract: The invention relates to a damping device for a gas turbine combustor with significantly reduced cooling air mass flow requirements. The damping device includes a wall with a first inner wall and a second outer wall, arranged in a distance to each other. The inner wall is subjected to high temperatures on a side with a hot gas flow. A plurality of cooling channels extend essentially parallel between the first inner wall and the second outer wall, and at least one damping volume bordered by cooling channels. Furthermore, the damping device includes a first passage for supplying a cooling medium from a cooling channel into the damping volume and a second passage for connecting the damping volume to the combustion chamber. An end plate, fixed to the inner wall, separates the damping volume from the combustion chamber.

Abstract: The invention concerns a gas turbine having a picture frame, a first vane, and a sealing arrangement to seal a gap between the picture frame and the first vane, the sealing arrangement including two seals arranged in series between the picture frame and the first vane. In exemplary embodiments, one of the seals is a honeycomb seal, a dogbone seal, a hula seal or a piston seal and the other seal is a honeycomb seal, a dogbone seal, a hula seal or a piston seal. A method of supplying cooling fluid to the gap between the picture frame and the first vane is also disclosed.

Abstract: The invention concerns a conical hula seal wherein an inner part of the conical hula seal is closer to a conical hula seal longitudinal axis than an outer part of the conical hula seal, and wherein the conical hula seal extends in a circumferential direction relative to the conical hula seal longitudinal axis. A gas turbine which includes the conical hula seal is also disclosed, along with a method of operating a gas turbine which includes a conical hula seal that seals a gap between a first vane and a picture frame, the method including purging the gap with a cooling fluid.

Abstract: A centering arrangement includes a ring-like inner part and a ring-like outer part, whereby the outer part surrounds the inner part in a concentric arrangement and with an interspace established between the outer part and the inner part, and whereby the outer part and the inner part are subject to a differential radial expansion. A centering contact element, procures a centering mechanical contact between the outer part and the inner part at a plurality of circumferentially distributed contact points, and is able to deform in order to compensate for differential radial expansion.

Abstract: The present invention relates to dampers for gas turbines and, for example, to a compensation assembly for a damper of a gas turbine for reducing the pulsations occurring in the combustion chamber. The damper can include a resonator cavity with a neck tube in flow communication with the interior of the combustion chamber, wherein the compensation assembly includes a spherical joint associated to the neck tube and configured to allow relative rotation between the combustion chamber and the resonator cavity, and having a bulb portion disposed around the neck tube and a spherical socket configured to internally host the bulb portion, wherein the spherical socket can have a top collar portion and a bottom collar portion connected to each other.

Abstract: A sequential burner for an axial gas turbine comprises: a burner body, which is designed as an axially extending hot gas channel and further comprises a fuel injection device, which extends into said burner body perpendicular to the axial direction. The manufacturing of the burner body is simplified and the fuel injection is stabilized by designing said fuel injection device as a mechanically stiff component, and fixing said fuel injection device to said burner body in order to keep it aligned with said burner body and to stiffen said burner body against creep.

Abstract: The disclosure relates to a vane arrangement with a vane carrier, an array of rocking first stage vanes, and an array of frame segments for axially receiving aft ends of a combustor transition pieces. The frame segments comprise an I-beam with an upper horizontal element, a lower horizontal element, a vertical web, a radially outer fixation to the vane carrier, and an arm extending from the lower horizontal element in axial direction below the inner rim segment for supporting the inner platform of the vane and for sealing a gap between the inner platform and the lower horizontal element. Besides the vane arrangement the disclosure relates to a method for assembly of such an arrangement as well as to a gas turbine comprising such a vane arrangement.

Abstract: A fuel lance is disclosed for injecting a gaseous and/or liquid fuel mixed with air into an axial hot gas flow flowing through a sequential combustor of a gas turbine, the fuel lance having at least one finger extending in a longitudinal direction into the axial hot gas flow of the gas turbine essentially perpendicular to the hot gas flow. The finger is configured as a streamlined body which has a streamlined cross-sectional profile. The body has two lateral surfaces essentially parallel to the axial hot gas flow. The body includes an enclosing outer wall having a longitudinally extending air plenum for the distributed introduction of air into the at least one finger. The air plenum is provided with a plurality of distributed impingement cooling holes, such that air exiting through the impingement cooling holes impinges on the inner side of the leading edge region of the body.

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 damping device for reducing pressure oscillations in a combustion system includes at least a portion provided with a first, outer wall, a second, inner wall, an intermediate plate interposed between the first wall and the second wall. This intermediate plate forms a spacer grid to define at least one chamber between said first wall and said second wall, first passages connecting each of said at least one chamber to the inner of the combustion system, and second passages connecting each of said at least one chamber to the outer of the combustion system. The second passages open at the same side of said chambers as the first passages, the second passages have a portion extending parallel to the inner wall. This parallel portion of said second passages is equipped with heat transfer enhancing means.

Abstract: An impingement cooled wall arrangement includes a flow diverter arranged in the cooling flow path between the cooled wall and a sleeve to divert a cross flow away from a second aperture. The flow diverter extends in downstream direction of the cross flow beyond the second aperture with a first leg extending along one side of the second aperture in downstream direction of the cross flow and a second leg extending along the other side of the second aperture. No impingement cooling aperture is arranged in a first convective cooling section of the wall between the upstream end and downstream end of the flow diverter outside the section shielded by the diverter.

Abstract: The invention relates to an annular combustion chamber of a gas turbine having a machine axis. The combustion chamber includes at least two zones. A first zone receives the fuel/air mixture of a plurality of burners. A second zone guides the hot gases being produced by the burners from the first zone to an entrance of a turbine section of said gas turbine. An annular transition liner includes a plurality of liner segments located at the transition between said first zone and second zone. Each of the liner segments includes with respect to the axial hot gas flow a leading edge, a trailing edge, and two sidewalls, and is provided with axial mounting means at the leading and trailing edges, such that the liner segment can be installed in axial direction and is axially fixed by respective segments of the neighboring first zone.