Abstract: A turbine blade or vane includes passages in the interior. Further, in the region of a radial end of a partition which delimits a passage, a shoulder is provided. On the shoulder there rests an insert which closes off the at least one passage. As such, the end of a partition is produced in such a manner that it is optimized in terms of flow retained.

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: A combustion chamber arrangement for gas turbines includes a multiplicity of individual combustion chambers which open out in a common annular combustion chamber and are arranged in the shape of a circle. The arrangement is such that it can be cooled highly efficiently using convective cooling, while also as far as possible dispensing with the use of impingement cooling. In such a combustion chamber arrangement, the annular combustion chamber, in the transition region to the individual combustion chambers, is of a height which fluctuates periodically and which is minimal in the region between adjacent individual combustion chambers.

Abstract: A turbomachine, in particular a gas turbine, includes a rotor which extends along an axis of rotation. The rotor includes a circumferential face, which is defined by the outer radial boundary surface of the rotor, and a receiving structure. Additionally, it includes a first rotor blade and a second rotor blade, which each have a blade root and a blade platform. The blade platform of the first rotor blade and the blade platform of the second rotor blade adjoin one another, and a space is formed between the blade platforms and the circumferential face. A sealing system is provided on the circumferential face in the space, the sealing system including a labyrinth sealing system.

Abstract: The invention relates to a combustion-chamber arrangement for gas turbines, having an annular combustion chamber connected to at least one burner and leading into a turbine space, an annular-combustion-chamber wall which defines the annular combustion chamber being of double-shell construction, and an inner shell of the annular-combustion-chamber wall being composed of lining elements releasably arranged on an outer shell of the annular-combustion-chamber wall, and a gap space through which a cooling medium can flow being formed between the inner shell and the outer shell of the annular-combustion-chamber wall.

Abstract: A gas turbine with axially mutually displaceable guide parts that compensates for differing amounts of thermal expansion, comprising guide parts which can be displaced with respect to one another in the axial direction and enclose a funnel-like gas duct from outside. In order to optimize a rotor blade tip gap, at least one of the funnel-like guide parts under control by means of a motor can be displaced. As a result of the axial displacement, because of the funnel-like shape of the guide parts, the width of the rotor blade tip gap is changed.

Abstract: A shaped part (30) for forming a guide ring (21) for a gas turbine (1) is, on the one hand, to be produced with particularly little outlay and, on the other hand, is to be designed so that it can be cooled particularly effectively with steam as the coolant. For this purpose, according to the invention the shaped part (30) comprises a base plate (34) which, together with an associated metal guide plate (38, 44, 52), forms a flow passage (36) for a coolant which, in cross section, extends substantially over its entire width, with respect to the axial direction (x) of the guide ring (20, 21).

Abstract: In a turbine (1) having a moving-blade row which has a moving blade (12) on a turbine shaft (8), and having a guide blade (14) which has a guide-blade row, an especially high efficiency is to be achieved with simple means with the turbine blades (12, 14) being cooled in a reliable manner. To this end, the cooling-medium passage inside the guide blade (14), according to the invention, is split up into a first and a second partial-flow passage, the cooling medium flowing in the first partial-flow passage being used for cooling the guide blade (14), and the cooling medium flowing in the second partial-flow passage being passed on more or less free of losses.

Abstract: In gas turbines, compressed air is supplied via an air duct to combustion chambers and is heated there. Pressure losses in the air duct should be minimized in order to ensure good overall efficiency. This is achieved by the compressed air flowing with approximately constant velocity in the air duct from the compressor to the inlet into the combustion chamber. This is supported by the effective cross section of the air duct being almost constant over this distance.

Abstract: A turbine blade or vane includes a platform that is cooled by way of air and steam. The air cooling may form an open circuit, where the steam cooling may form a closed circuit.

Abstract: A device for a component to be cooled includes a baffle cooling plate with at least one leadthrough for the cooling agent, arranged in front of the component. The baffle cooling plate can be impinged upon with a cooling agent. A deviation channel for the cooling agent is provided next to the leadthrough for the cooling agent and between the baffle cooling plate and the component to be cooled. The cooling agent enters the deviation channel after the component has been baffle-cooled. The cooling agent is collected and deviated by way of the deviation channel and can be specifically sued for further cooling purposes. The inventive device provides more efficient baffle cooling.

Abstract: Disclosed is a gas turbine blade, in which a ceramic covering, which is mechanically fastened to a metal platform, is arranged in a manner that the metal platform is protected against a hot gas in a hot gas duct of a gas turbine.

Abstract: A component can be subjected to hot gas. At least one duct is provided which can be subjected to a cooling fluid. The duct is bounded by two first walls opposite to one another. The walls include turbulators with the same direction of inclination. In order to avoid constrictions, the turbulators of the first wall have a different angle of inclination relative to a flow direction of the cooling fluid to the turbulators of the second wall.

Abstract: A method is for producing a turbine blade, in particular, a gas turbine blade, comprising a head, a foot, and a blade section, in addition to an internal canalization system, including individual channels through which coolant gas can pass along a flow path within the turbine blade. The turbine blade also includes a throttle device which influences the passage of the coolant gas without impairing the flow of the coolant gas in the intake area. The throttle device is located in the rear section of the flow path, and is positioned upstream of the exit openings.

Abstract: The invention relates to a guide vane (17) for a turbomachine (1), in particular a gas turbine guide vane, in which the platform (48) has a separating region (50), which is embodied as a separate component. This has, in particular, advantages with respect to the simplification of cast blade/vane (17) in terms of manufacturing technology, with respect to the variability of a material selection, the quality of a protective coating to be applied and efficient cooling.

Abstract: The invention relates to a turbine vane (1), especially a turbine vane of the last stages, respectively comprising a lower area (2) which is radially and externally arranged, an upper area (3) which is radially and internally arranged, and a radial cooling air channel (4) extending between the upper area and the lower area. Cooling air (23) can be introduced into said channel via an inlet (36) in the lower area, and can be at least partially discharged via an outlet (35) in the upper area. The cooling air channel comprises a radial inner channel through which the cooling air flows from the lower area to the upper area, and an outer channel (9) which is adjacent to the inner channel and which at least partially surrounds the inner channel on the circumferential side thereof. Said outer channel communicates with the inner channel and comprises an outlet (12) which is arranged in the lower area.

Abstract: The present invention relates to a process for producing a turbine blade or vane (13; 14), which has at least one chamber (22; 23, 24, 25) and an inlet (30; 31) for applying a cooling medium to the chamber (22; 23, 24, 25), at least one inlet (30) running at an angle with respect to a longitudinal axis (37) of the turbine blade or vane (13; 14). According to the invention, to form the inlet (30) a core (35) with a projection (33) is used, which projection is arranged at a distance from a mold (40). Therefore, after removal from the mold the inlet (30) of the turbine blade or vane (13; 14) is closed, and is opened up by machining. The invention also relates to a turbine blade or vane, in particular for a gas turbine (10), which has at least one chamber (22; 23, 24, 25) and at least one inlet (30; 31) for applying a cooling medium to the chamber (22; 23, 24, 25).

Abstract: A turbomachine includes a coolable arrangement, having a first wail element and having a second wall element. Coolant is used to cool the arrangement in a turbomachine. After the impingement cooling of the first wall element, the coolant is further used for cooling the second wall element.

Abstract: The invention relates to a turbine blade (1) through which cooling liquid flows and which comprises several flow channels (4, 5, 6) that are adjacently arranged in the flowing-off direction (2) of the working fluid (3). Said flow channels (4, 5, 6) extend between inflow openings (7, 8, 9), which are located on a radially inner blade foot (10), and outflow openings (11, 12, 13), which are situated opposite the blade foot in a more radially outer manner. The flow channels (4,5,6) extend such that the local radial flow components (15) of the resulting cooling fluid stream (14) are directed outward in a, for the most part, predominantly radial manner. A flow channel (4), which is in the front when viewing in the flowing-off direction (2), is provided whose outflow openings (11) are placed inside a touching edge (16) of the turbine blade (1).

Abstract: A covering element is for the protection of components in a machine subjected to high thermal load, in particular of components in a gas turbine. The covering element includes a wall with a hot side capable of being exposed to a hot medium, and with a cool side located opposite the hot side. The cooling side includes a cooling surface capable of being acted upon by a coolant. Holding elements are provided on the cool side.