Abstract: A gas turbine component, for forming part of a stage of the turbine, operable to change cooling scheme, includes an airfoil profiled section, a cooling passageway, film holes and interchangeable connectors. The profiled section includes pressure and suction sides joined together at chordally opposite leading and trailing edges. The cooling passageway extends between the pressure and suction sides along the leading edge to enabling cooling fluid to flow therefrom. The film holes are configured on the cooling passageway to enable the flow of a portion of the cooling fluid to a portion of the profiled section. The interchangeable connectors configured to the cooling passageway, one at a time, to change the cooling scheme. An insert may also be provided to close and open the film holes.

Abstract: A blade assembly having a modular structure, wherein the blade elements include at least a rotor blade airfoil, a footboard mounting part and a heat shield. The elements each have at its one ending an interchangeable connection for connection among each other, wherein the connection of the airfoil with respect to the other elements is based on a fixation in radial or quasi-radial extension compared to the rotor axis of the turbomachine. The assembling of the blade airfoil in connection with the footboard mounting part is based on a force-fit or form-fit fixation, or on use of a metallic and/or ceramic surface for the purpose of a friction-locked bonding actuated by adherence interconnecting, or on friction-locking with a detachable, permanent or semi-permanent fixation.

Abstract: A turboengine component is disclosed, having at least one first, receiving, member and at least one second, received, member, the receiving member having at least one receiver opening, the received member including a body and at least one fixation post extending from the body. A retainer cavity is provided with a first retainer groove at an inner surface of the receiver opening. A second retainer groove is provided on a surface of the fixation post. A retainer member has a cross section and a longitudinal extent aligned with the lengthwise extent of the retainer cavity, the retainer member being displaceable within the retainer cavity along the lengthwise extent. The retainer member may be removed and the turboengine component may be disassembled without the damaging any of the received member and the receiving member.

Abstract: A blade assembly of a power plant having a modular structure, wherein blade elements include at least one blade airfoil, and at least one footboard mounting part. Blade elements can each have at its one ending a configuration for an interchangeable connection among each other. The connection of the airfoil with respect to other elements can be based on a fixation in radially or quasi-radially extension relative gas turbine axis, wherein the assembling of the blade airfoil in connection with the footboard mounting part is based on a friction-locked bonding actuated by adherence interconnecting, or on use of a metallic and/or ceramic surface fixing blade elements to each other, or on closure configuration with a detachable, permanent or semi-permanent fixation.

Abstract: The disclosure refers to a method of assembling or disassembling a rotor blade or guide vane assembly, wherein the rotor blade or guide vane includes an airfoil and additional structured peripheral members forming at least an inner and/or an outer platform of the rotor blade or guide vane. The airfoil includes at least one airfoil sub-structure designed for anchoring at least one superposed component for the purpose of a thermal protection. The connection between the airfoil sub-structure and the superposed component is supported on friction-locked device, wherein the airfoil sub-structure is formed by a spar and the superposed component includes at least one flow-charged outer shell. Connection between the spar or airfoil understructure and flow-charged outer shell is formed by a force-fit and/or a form-fit fixation or a shrinking joint, wherein the airfoil and additional structured peripheral members is formed by friction-locked device with a detachable, permanent or semi-permanent fixation.

Abstract: A gas turbine component, for forming part of a stage of the turbine, operable to change cooling scheme, includes an airfoil profiled section, a cooling passageway, film holes and interchangeable connectors. The profiled section includes pressure and suction sides joined together at chordally opposite leading and trailing edges. The cooling passageway extends between the pressure and suction sides along the leading edge to enabling cooling fluid to flow therefrom. The film holes are configured on the cooling passageway to enable the flow of a portion of the cooling fluid to a portion of the profiled section. The interchangeable connectors configured to the cooling passageway, one at a time, to change the cooling scheme. An insert may also be provided to close and open the film holes.

Abstract: An abradable coating system (2-8) for a turbine or a compressor is proposed has at least one layer of woven or non-woven tissue (15) or foam made of: ceramic, glass, glass-ceramic, ceramic-metal composite and combinations thereof. A free-standing preformed element (18) has at least one layer of such a woven or non-woven tissue or foam is filled with fillers and/or binder. A method prepares such a system. Finally, a method for the repair of an abradable coating portion of a turbine or the compressor component (9) includes in a first optional step that the damaged portion is at least partly removed and/or cleaned and/or surface treated, and in a second step at least one free-standing preformed abradable coating element (18) or a complete abradable coating system is attached to the component (9), preferably by using a matrix, wherein for example only a center part (2) or a fraction of a center part (2) is replaced.

Abstract: An abradable coating system (2-8) for a turbine or a compressor is proposed has at least one layer of woven or non-woven tissue (15) or foam made of: ceramic, glass, glass-ceramic, ceramic-metal composite and combinations thereof. A free-standing preformed element (18) has at least one layer of such a woven or non-woven tissue or foam is filled with fillers and/or binder. A method prepares such a system. Finally, a method for the repair of an abradable coating portion of a turbine or the compressor component (9) includes in a first optional step that the damaged portion is at least partly removed and/or cleaned and/or surface treated, and in a second step at least one free-standing preformed abradable coating element (18) or a complete abradable coating system is attached to the component (9), preferably by using a matrix, wherein for example only a center part (2) or a fraction of a center part (2) is replaced.

Abstract: A turbomachine (1), in particular a turbine or compressor, includes a rotor (2) which has at least one moving blade row (4) with a plurality of moving blades (5), and a stator (3) which has at least one guide vane row (6) with a plurality of guide vanes (7), at least in one of the blade or vane rows (4, 6) all the blades or vanes (5, 7) each having a shroud plate (10). In order to simplify the maintenance of the turbo-machine (1), the shroud plates (10) are separate components with respect to the respective blade or vane (5, 7) and are mounted on the respective blade or vane (5, 7), shroud plates (10) adjacent in the circumferential direction (9) being coupled to one another for the transmission of tensile forces in the circumferential direction (9).

Abstract: Reducing the tip leakage for an existing compressor or turbine blade (10), including internally cooled blades, can be achieved by retrofitted the blade with a tip leakage-reducing device in the form of a winglet (12). The winglet (12) can be fixed onto the tip portion by bonding the winglet onto the face or faces of the blade by a key and slot arrangement.

Abstract: A turbo machine (1, 101), in particular turbine or compressor, includes a rotor (2, 102) which has at least one moving blade row (4, 104) with a plurality of moving blades (5, 105), and a stator (3) which has at least one guide vane row (6, 106) with a plurality of guide vanes (7, 107), at least one moving blade row (4, 104) having a shroud (8, 108). The shroud (8, 108) is designed to be self-supporting in such a way that it can at least partially absorb the centrifugal forces arising during operation and discharge them in the circumferential direction.

Abstract: A turbomachine (1), in particular a turbine or compressor, includes a rotor (2) which has at least one moving blade row (4) with a plurality of moving blades (5), and a stator (3) which has at least one guide vane row (6) with a plurality of guide vanes (7), at least in one of the blade or vane rows (4, 6) all the blades or vanes (5, 7) each having a shroud plate (10). In order to simplify the maintenance of the turbo-machine (1), the shroud plates (10) are separate components with respect to the respective blade or vane (5, 7) and are mounted on the respective blade or vane (5, 7), shroud plates (10) adjacent in the circumferential direction (9) being coupled to one another for the transmission of tensile forces in the circumferential direction (9).

Abstract: A thermally highly loaded machine component (10), which is protected from overheating by film cooling, is provided with conical cooling passages (30) of circular cross section. The cooling passages are designed so as to be divergent from the cold-gas side (13) toward the hot-gas side. The conical cooling passages, as compared with cylindrical passages with regard to the mass flow of cooling medium (35) fed through the cooling passages, are substantially less dependent on the pressure ratio between the cold-gas side (13) and hot-gas side (14) of the component. The straight conical cooling passages of round cross section throughout may be produced in a very simple manner by laser drilling, with a convergent cutting beam being used for the machining.

Abstract: The present invention relates to a device for separating dirt and debris in flowing media by making use of centrifugal forces, said device comprising a separation chamber (1), and a dirt collection chamber (3) connected with said separation chamber (1) by at least one dirt discharge opening (2). Arranged in the region of dirt discharge opening (2) within dirt collection chamber (3) are one or more flow guide elements (4, 5, 6), which make the backflow of dirt and debris into separation chamber (1) difficult. In the present device the danger of backflow of dirt and debris particles from the dirt collection chamber into the separation chamber is reduced.

Abstract: The present invention relates to a device for separating dirt and debris in flowing media by making use of centrifugal forces, said device comprising a separation chamber (1), and a dirt collection chamber (3) connected with said separation chamber (1) by at least one dirt discharge opening (2). Arranged in the region of dirt discharge opening (2) within dirt collection chamber (3) are one or more flow guide elements (4, 5, 6), which make the backflow of dirt and debris into separation chamber (1) difficult.

Abstract: In a gas-turbine guide element (30) around which a hot air flow (23) flows and which, at least in a trailing edge region (21), in which the air flow (23) separates from the guide element (30), comprises at least two walls (10, 11) arranged essentially in parallel and connected to one another by ribs (16, 17, 20) in such a way as to form internal cooling passages (18, 19, 25, 26, 27), and which is cooled on the inside with cooling medium (28, 29) flowing through the cooling passages (18, 19, 25, 26, 27), the cooling medium discharging from the guide element (30) at the trailing edge (21) essentially parallel to and between the walls (10, 11), and in a method of producing it, easier reworking and less susceptibility to foreign particles are achieved owing to the fact that at least some of the ribs are arranged as choke ribs (24) so as to terminate essentially flush with the trailing edge (21).

Abstract: Apparatus is disclosed for providing cooling channels in the interior of a gas turbine rotor blade. The cooling channels are formed by metallic inserts which extend from adjacent the root of the blade toward the tip. The inserts are substantially flat and are secured in the interior of the airfoil section by means of rails which engage the longitudinal edges of the inserts and serve as a guide during insertion. The rails are preferable formed integrally with the blade casting.

Abstract: A component of a fluid-flow machine including a flow-deflecting device which is connected in a root region to a platform, the penetration of the flow-deflecting device with the platform being provided by a pressure-side profile line, a suction-side profile line, and a leading edge base point and a trailing edge base point, and the top view of the platform essentially exhibiting the shape of a parallelogram on which there is arranged a nose running in the circumferential direction, wherein a long side of the parallelogram is arranged as a tangent to the suction-side profile line.

Abstract: The temperature can nowadays be reduced by about 400° C. as a result of the cooling of the guide vane (15) of the first guide wheel. The many film bores leading outward from the guide vane (15) must therefore remain free of blockages or deposits caused by suspended matter in the cooling medium. Only pure gas must therefore be used as the cooling medium for the first guide wheel. This purpose is served by a dust separator which is designed advantageously as an axial cyclone (24, 24′). The latter has no moving structural parts, has a simple design, is therefore insusceptible to faults and can be exposed to high temperatures. The axial cyclone is a centrifugal deduster without gas stream reversal.

Abstract: A hot-gas stream (40) flows along the surface of a segment arrangement for shroud bands, in particular in a gas turbine. The segment arrangement comprises segments (20, 20′) arranged next to one another and in each case separated from one another by a gap (12). The hot-gas stream (40), in at least one section (70) of the gap (12), has a velocity component perpendicular to the direction of the gap from a first segment (20) to a second segment (20′). In this case, in said section (70), along that edge (26) of the first segment (20) which faces the gap (12), at least one film-cooling bore (52) connects a cooling-air chamber (50), allocated to the first segment, to the surface (22) subjected to the hot-gas stream (40), and/or, in said section (70), along that edge (26′) of the second segment (20′) which faces the gap (12), at least one edge-cooling bore (56) connects a cooling-air chamber (50′), allocated to the second segment, to the inside (28′) of the gap (12).