Siemens Aktiengesellschaft

Siemens Power Generation, Inc.

Peter Tiemann

Michael Strassberger

Hans-Thomas Bolms

Michael Scheurlen

Dirk Anding

Iris Oltmanns, Legal Representative

Robert James Bland

Charles A Ellis

Mirko Milazar

Winfried Esser

Christian Cornelius

Michael Huth

Arnd Reichert

Christian Scholz

John Carl Glessner

Burkhard Bischoff-Beiermann

Thomas Schulenberg

Dirk Lieser

Michael Haendler

Micheal Strassberger

Ariel Jacala

Reinhard Mönig

The invention relates to a gas turbine, for energy generation, with a compressor, arranged coaxially to a rotor, mounted such as to rotate, for the compression of an inlet gaseous fluid, at least partly serving for combustion of a fuel in a subsequent annular combustion chamber, with generation of a hot working medium, with an annular diffuser arranged coaxially to the rotor, between the compressor and the annular combustion chamber, for distribution and deflection of the fluid, whereby a part of the fluid is diverted as cooling fluid for the turbine stages after the combustion chamber, by means of a dividing element, arranged in the fluid flow. According to the invention, a compact diffuser and an economical gas turbine with an improved flow for the diversion of cooling air may be achieved, whereby the annular dividing element, arranged coaxially to the rotor, comprises at least one opening, facing the fluid flow and the dividing element is supported on the diffuser, by means of several hollow rib-like support elements, by means of which the cooling fluid, diverted through the opening, is first directed towards the rotor.

Diffuser having distribution element for providing part-flow

The invention relates to a combustion chamber of a gas turbine, inside of which a supplied fuel is reacted with supplied combustion air in order to produce a working medium. The inside of the combustion chamber wall is provided with a lining formed from a number of heat shield elements. The or each heat shield element, together with the combustion chamber wall, forms an interior space that can be subjected to the action of a coolant. The aim of the invention is to provide a combustion chamber with a comparatively simple design that has a high system efficiency. To this end, the invention provides that a flow element is placed in the respective interior space while serving to effect a locally specific control of the flow of coolant.

Combustion chamber

A gas turbine engine having an axial flow compressor, an annular combustion chamber, a turbine, and a diffuser. The diffuser includes a flow-dividing element formed by an inner deflecting flank and an outer deflecting flank that divides a compressed gas flow into two partial flows at a branching point. The two deflecting flanks define: an angle of less than 90&#xb0; along at least a portion of the deflecting flanks, and an angle between 15&#xb0; and 90&#xb0; between the deflecting flank angle bisector and the turbine longitudinal axis. The deflector also includes a main deflecting region arranged upstream of the branching point and directed at an acute angle from the turbine longitudinal axis toward an inner combustion chamber shell.

Diffuser arranged between the compressor and the combustion chamber of a gas turbine

The invention relates to a segmented inner ring for holding guide blades. According to the invention, a lateral wall opposing the front side of the inner ring and pertaining to a shaft shoulder formed on the rotor shaft extends radially, and respectively one half of a labyrinth seal is formed on the front side of the inner ring and on the shaft shoulder. The aim of the invention is to apply an arrangement of stacked labyrinth seals, known from aeroplane turbines, to a stationary gas turbine having a separation plane. To this end, a method is used to mount an inner ring of a gas turbine. The invention also relates to a stationary gas turbine comprising a segmented inner ring.

Labyrinth seal in a stationary gas turbine

The invention relates to a segmented inner ring for holding guide blades. According to the invention, a lateral wall opposing the front side of the inner ring and pertaining to a shaft shoulder formed on the rotor shaft extends radially, and respectively one half of a labyrinth seal is formed on the front side of the inner ring and on the shaft shoulder. The aim of the invention is to apply an arrangement of stacked labyrinth seals, known from airplane turbines, to a stationary gas turbine having a separation plane. To this end, a method is used to mount an inner ring of a gas turbine. The invention also relates to a stationary gas turbine comprising a segmented inner ring.

An attachment device for coupling one or more restrained members to a frame. In at least one embodiment, the attachment device may be configured to attach one or more turbine combustor liners to a carrier. The attachment device may include a body rotatable about an axis of rotation and including one or more arms for grasping an edge of at least one restrained member. A fastener may be coupled to the body for coupling the body to a frame. The attachment device may also include a retainer for preventing the fastener, which may be a threaded bolt, from becoming disengaged from the attachment device and traveling downstream in a turbine. An alignment device may be included for facilitating rotation of the body of the attachment device and engagement the arms with edges of retrained members. The alignment device may include an alignment indicator indicating a position of the arms.

Attachment device for turbine combustor liner

The invention relates to a turbine, in particular a gas turbine, which along a swivel-mounted rotationally symmetrical rotorhas a compressor, a combustion chamberand a turbine sectionformed of a plurality of turbine stages, in which each turbine stagecontains mutually interacting bladesand vaneswhich a hot working fluidcan flow around, with a coolant provided by the compressorfor cooling the blades, which can flow in a channel along the rotorfrom the compressorto the turbine sectionand into which a liquid can be introduced for cooling. In order to specify a turbine in which less wear occurs and in which the lifetime of the components is increased, it is proposed that the channel extends outside the rotorand that the liquid can be introduced at the end of the channel which faces the compressor

Cooling air and injected liquid system for gas turbine blades

A combustion chamber () for combusting a combustible fluid mixture having a burner () disposed on the combustion chamber () and a method for cooling a combustion chamber (). A fixing device is provided for liner elements () of a combustion chamber () by which simplified assembly from the inside of the liner () can be achieved. The combustion chamber has a burner () disposed on the combustion chamber (), a liner () disposed in the combustion chamber (), and an outlet opening (), wherein the liner () includes liner elements () which are elastically fixable to a combustion chamber casing () by rail elements (), the rail elements () being disposed on the combustion chamber side and projecting outward between two adjacently disposed liner elements ().

Combustion chamber for combusting a combustible fluid mixture

A turbine and a method for dismantling a blade of a turbine. A turbine includes a rotor that extends in an axial direction and an accessible combustion chamber, which communicates with an annular hot-gas conduit, in which a plurality of blades that form a blade row is arranged. Each blade has a blade root that is fixed to the internal housing and a blade head that lies opposite the root and faces the rotor, the head being fixed to a fixing ring of the turbine that encircles the rotor. To reduce the down time of a turbine during repair, inspection and/or maintenance work, the blade root and/or the blade head can be fixed by means of a manually detachable clamping device. The clamping device is released from the blade and locked in a parking position beyond a displacement path used for removing the blade through the combustion chamber.

Turbine, fixing device for blades and working method for dismantling the blades of a turbine

A turbine vane, especially a turbine vane of the last stages, respectively includes a lower area which is radially and externally arranged, an upper area which is radially and internally arranged, and a radial cooling air channel extending between the upper area and the lower area. Cooling air can be introduced into the channel via an inlet in the lower area, and can be at least partially discharged via an outlet in the upper area. The cooling channel includes a radial inner channel through which the cooling air flows from the lower area to the upper area, and an outer channel which is adjacent to the inner channel on the circumferential side thereof. The outer channel communicates with the inner channel and includes an outlet which is arranged in the lower area. Part of the cooling air flows back in the direction of the lower area via the outer channel and emerges via the outlet.

Turbine vane system

A turbine engine with a turbine shaft which has a number of disks arranged adjacent to one another, to each of which a number of blades are fastened in a star arrangement, wherein these blades can be cooled with at least one coolant which flows through coolant channels inside the blades is provided.

Turbine engine and a method for cooling a turbine engine

The invention relates to a combustion chamber, in particular of a gas turbine, comprising an external wall structure that surrounds an internal area and an internal wall, which is formed by the surface of a housing that is located in the internal area. Said combustion chamber can be cooled essentially by convection by an air stream that flows between the external wall structure and the internal wall, whereby the air stream is conducted in a closed cooling air canal. The housing can be mounted in the internal area of the combustion chamber by means of a suspension device, which comprises a large number of fixing elements that are arranged on the periphery of the housing and are braced against the wall structure.

Combustion chamber, in particular of a gas turbine

A turbine () with a rotationally symmetrical rotor () on rotary bearings which has on its outer perimeter moving blade retaining grooves running transverse to the outer perimeter and set at the same angular distance from one another, having a moving blade () that has a vane () and a moving blade foot () which corresponds to the moving blade retaining groove () and which is inserted in one of the moving blade retaining grooves (), wherein a platform () is located between the moving blade foot () and the vane () and extends transverse to the moving blade foot (), having a baffle cooling element () which is located between two directly adjacent moving blades () between its platforms () and the rotor () and through which a cooling fluid can flow to cool the platform (). It is proposed that the baffle cooling elements () are each formed from two independently moveable halves () and that the two halves () of the baffle cooling elements () are automatically pressed tight against each other by centrifugal force during turbine operation.

Turbine, in particular a gas turbine, and a blade

The invention relates to a heat shield arrangement for a hot gas conducting structure, in particular a metal component of a gas turbine unit or combustion chamber, with the heat shield elements anchored adjacently on a support structure to cover a surface. Said heat shield elements each comprise a lateral groove, arranged in the region of the edge of the surface thereof facing the hot gas, at least two adjacent shield elements are connected by at least one seal element mounted in the groove. The seal element is embodied as a sealing flap, which may be displaced from a first position to a second position and vice versa, whereby the first position is an open position without a sealing effect and the second position is a closed position with a sealing effect. According to said method the sealing flap is moved from the first to the second position by the movement of a heat shield element.

Heat shield arrangement for a hot-gas conducting component, in particular for structural pieces of gas turbines and method for production of said arrangement

A turbine blade or vane includes an inner contour profile which is matched to the contour profile of the outer transition region, so that a substantially uniform blade or vane wall thickness is achieved.

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Witten, Germany

Peter Tiemann