Abstract: A turbine blade/vane includes a profiled blade/vane aerofoil, which extends along a blade/vane center line. A platform is formed on an end region of the blade/vane aerofoil, extending transverse to the blade/vane center line. This is to be designed for a particularly high capability for carrying thermal and mechanical loading. In addition, the arrangement permits reliable cooling with a comparatively small coolant requirement. For this purpose, the platform includes an outer rim which is thickened in comparison with the platform floor. The side wall of the outer rim facing toward the blade/vane center line is slanted relative to the latter. A casting system suitable for manufacturing the turbine blade/vane includes a first shell element which can be positioned in a casting mold, an essentially flat configuration second shell element being displaceable, while being guided in a direction tilted by an angle of more than 10° and of less than 80°, in the first shell element.

Abstract: 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.

Abstract: A turbine blade/vane, having a blade/vane aerofoil which extends along a blade/vane axis and through which cooling medium flows, mainly in the longitudinal direction of the turbine blade/vane, is configured comparatively simply for reliable and effective closed cooling. In particular, it is configured with the use of cooling air as the cooling medium. An incident flow duct and an efflux duct for cooling medium are routed within the blade/vane aerofoil, essentially over its complete length. The incident flow duct and the efflux duct are connected together on the cooling medium side in such a way that cooling medium passing from the incident flow duct into the efflux duct is conducted in a transverse direction along a wall inner surface, which has to be cooled, of the blade/vane aerofoil.

Abstract: The invention relates to a method for producing a monocrystalline component, having a complex moulded structure with different structure parts, from a molten metal. According to the method, said molten metal is located in a negative mould, corresponding to the moulding structure and said negative mould moves with the formation of a solidification front on a temperature drop, which is adapted to the crystallization speed of the molten metal and which includes the melting point. Said method is characterized in that at least one of the structure parts of the moulded structure experiences an individual temperature drop. If the solidification front has to grow through a structure part, which is inappropriately oriented in relation to the solidification front, said structure part or zone can experience an individual temperature drop. This enables monocrystalline components, having a complex moulded structure, to be produced in a reliable and economic manner.

Abstract: A turbine blade/vane includes a profiled blade/vane aerofoil, which extends along a blade/vane center line. A platform is formed on an end region of the blade/vane aerofoil, extending transverse to the blade/vane center line. This is to be designed for a particularly high capability for carrying thermal and mechanical loading. In addition, the arrangement permits reliable cooling with a comparatively small coolant requirement. For this purpose, the platform includes an outer rim which is thickened in comparison with the platform floor. The side wall of the outer rim facing toward the blade/vane center line is slanted relative to the latter. A casting system suitable for manufacturing the turbine blade/vane includes a first shell element which can be positioned in a casting mold, an essentially flat configuration second shell element being displaceable, while being guided in a direction tilted by an angle of more than 10° and of less than 80°, in the first shell element.

Abstract: 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.

Abstract: The present invention relates to a combustion chamber (1) for combusting a combustible fluid mixture having a burner (2) disposed on the combustion chamber (1). The invention further relates to a method for cooling a combustion chamber (1) according to the invention. The object of the invention is to provide a fixing device for liner elements (5) of a combustion chamber (1) by means of which a simplified assembly from the inside of the liner (4) can be achieved. To achieve this object there is proposed according to the present invention a combustion chamber of the generic type (1) having a burner (2) disposed on the combustion chamber (1), a liner (4) disposed in the combustion chamber (1), and an outlet opening (3), wherein the liner (4) comprises liner elements (5) which are elastically fixable to a combustion chamber casing (7) by means of rail elements (6), the rail elements (6) being disposed on the combustion chamber side and projecting outward between two adjacently disposed liner elements (5).

Abstract: This invention relates to a gas turbine having a combustion chamber arrangement and a turbine chamber connected downstream of said combustion chamber arrangement, wherein the combustion chamber arrangement includes a plurality of individual combustion chambers formed by input areas and transition areas converging in an annular gap leading to the turbine chamber and wherein the longitudinal axes of the individual combustion chambers are placed at an angle relative to an engine axis that is defined by the axial extension of the turbine chamber. In order to improve said gas turbine by visibly reducing the thermal and mechanical loads of the individual combustion chambers in the transition area so that cooling requirements can be lowered in said area, the transition area of at least one individual combustion chamber into the turbine chamber is deflected in the direction of the engine axis and substantially in the input area of the turbine chamber.

Abstract: A turbine blade/vane includes a profiled blade/vane aerofoil, which extends along a blade/vane axis to be capable of absorbing high thermal and mechanical loading, on the one hand, and to ensure a comparatively economical consumption of coolant, on the other. For this purpose, the blade/vane aerofoil includes, integrally formed on it, a hot-gas platform extending transversely to the blade/vane axis and, above it, a load platform. A mechanical connection between the load platform and the hot-gas platform occurs exclusively by way of the blade/vane aerofoil.

Abstract: 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.

Abstract: The invention relates to a blade for a turbine engine with a base body, made from a titanium alloy, with a first and a second component piece, which are connected in the connection region by means of a bonding process. A groove with a groove wall runs through the connection region, to prevent local concentration of tension, such that the above borders directly on the second component piece and forms therewith a connection angle greater than 70 degrees.

Abstract: A gas turbine blade includes a combined convective cooling effected by a meandering cooling channel and an impact cooling channel and an impact cooling effected via an impact cooling insert. The impact cooling insert is arranged inside a first partial section of the meandering cooling channel extending along the front edge of the blade pan. The impact cooling insert tapers along this first partial section.

Abstract: The invention relates to a turbine engine with a turbine shalt which has a number of disks arranged adjacent to one another, to each of which a number of moving blades are fastened in a star arrangement, wherein these moving blades can be cooled with at least one coolant which flows through coolant channels inside the moving blades. The object of this invention is to increase the efficiency at an economically justifiable cost and at the same time to lower the operating risk and to increase the service life of the elements and the reliability of the turbine engine. To this end, coolant is supplied to and discharged from the moving blades via a simple and reliable coolant feeder system which is structured in a radial multilayered fashion between two adjacent disks.

Abstract: In order to provide a method of air guidance and a design for a gas turbine combustion chamber which permits improved cooling and reduced pressure losses and which avoids the overcooling phenomena by specific reaction to hot spots, a gas turbine combustion chamber includes a casing and at least one hollow element. The at least one hollow element is in connection with air guidance regions of the combustion chamber via at least two openings. As such, a tangential through-flow exists with a subsequent axial flow to the burner.

Abstract: A turbine blade through which cooling liquid flows, comprises several flow channels that are adjacently arranged in the flowing-off direction of the working fluid. The flow channels extend between inflow openings, which are located on a radially inner blade foot, and out-flow openings, which are situated opposite the blade foot in a more radially outer manner. The flow channels extend such that the local radial flow components of the resulting cooling fluid stream are directed outward in a, for the most part, predominately radial manner. A flow channel, which is in the front when viewing in the flowing-off direction, is provided whose outflow openings are placed inside a touching edge of the turbine blade. In addition, at least one rear edge channel is provided whose associated resulting cooling fluid stream additionally comprises local cross-flow components and whose outflow openings are placed inside a rear edge of the turbine blade.

Abstract: This invention relates to a combustion chamber arrangement for a gas turbine, with a number of individual combustion chambers which open into an annular gap, leading to a turbine chamber, whereby burners are arranged before the individual combustion chambers, connected to the individual combustion chambers through a turbine housing. According to the invention, said combustion chamber arrangement may be further developed, such that the cooling efficiency of the individual combustion chambers may be significantly improved, with an embodiment of simple construction, whereby at least one collar, arranged on one side of the turbine housing, facing the turbine chamber, running radially in the direction of the turbine chamber, at least partly surrounds a section of at least one individual combustion chamber whilst leaving a gap space.

Abstract: The invention relates to a combustion chamber for a gas turbine having at least one closed-circuit-cooled burner insert which can be disposed in an inlet opening of the combustion chamber for the purpose of feeding and/or igniting a combustible gas/air mixture, and having an outlet opening.

Abstract: 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.

Abstract: A device and a method are for producing a blade including two outer walls and at least one cavity between the outer walls, for a turbine. An outer mould and several cores are used in forming the outer walls and the at least one cavity of the blade. At least one of the cavities is divided into two channels by a middle segment. One channel is located between the first outer wall and the middle segment, while the other channel is located between the middle segment and the second outer wall. Two cores which are separate from each other are used accordingly. This provides a simple and economical means of reducing the thickness of the outer wall.

Abstract: A hot gas chamber, for example a combustion chamber of a gas turbine installation, is lined with thermal shield elements. A closed-circuit cooling system is provided by configuring the thermal shield element as a hollow body into which cool air flows via a cool air supply channel. Once the cool air is discharged from the thermal shield element through at least one opening, it is collected in a tiled intermediate space and is then used for the combustion process.