Abstract: A ceramic matrix composite blade may include an airfoil and a root. A protective metallic layer may be coupled to the root. The protective metallic layer may be plasma sprayed on the root. A portion of the protective metallic layer may be removed.

Abstract: A method for masking cooling passages of a turbine component having an external surface, an internal cavity for receiving cooling air, and cooling passages extending therebetween. The location and angle of cooling passages are determined using a robotic arm and a location system. A masking device is placed in the cooling passages located during the locating step. The masking device includes a head portion having a gripping feature for gripping by a robotic arm, and a locating feature for orientation of the masking device by the robotic arm. A retaining portion extending from the head portion is arranged and disposed to retain the masking device in a cooling passage. The retaining portion is narrower proximate a distal end than proximate the head portion. The component and head portion of the masking devices are coated. The masking devices may be removed using the robotic arm and locating system.

Abstract: A composite bonded repair method for an aircraft is provided. The method includes generating a doubling patch of a portion of the aircraft from topographic data, filling damage at the portion of the aircraft with filler material, applying adhesive to the portion of the aircraft and the filler material and adhering the doubling patch to the adhesive.

Abstract: The disclosure provides a turbine blade. A turbine blade includes a shroud at a tip of the blade and an airfoil having a leading edge extending between the shroud and a root of the blade; a cutout portion extending at least along the leading edge of the airfoil nearest the shroud; an insert welded to the cutout portion at a weld region; and a weld-free region located between the insert and the cutout portion nearest the shroud.

Abstract: A turbine blade according to an example of the present disclosure includes, among other things, a platform extending from a root section, an airfoil section extending radially from the platform to an airfoil tip, a plurality of cooling passages defined in an external wall of the airfoil tip, the plurality of cooling passages extending radially between the airfoil tip and a cavity in the airfoil section bounded by the external wall, and each of the plurality of cooling passages defining an inlet port along the cavity and an exit port adjacent the airfoil tip, and at least one internal feature within each of the plurality of cooling passages that meter flow to the respective exit port.

Abstract: This method can be used for refurbishing a facility for converting hydraulic energy into electrical or mechanical energy, and vice versa. The facility includes a Francis type pump-turbine. The pump-turbine includes a runner movable about an axis, a pre-distributor, including stay vanes defining between each pair of two adjacent stay vanes a first water passage channel and a distributor, including guide vanes arranged downstream of the stay vanes in the direction of water flow feeding the pump-turbine operating in turbine mode. The guide vanes define between each pair of two adjacent guide vanes a second water passage channel. This method includes steps of reducing the height, taken parallel to the axis of rotation of the runner, of the first water passage channels and reducing the height, taken parallel to the axis of rotation of the runner, of the second water passage channels.

Abstract: A magnetic impeller for a blood pump such as a magnetically driven, rotary ventricular assist device for pumping blood of a patient, the impeller comprising a magnetic alloy including platinum, cobalt, and boron.

Abstract: The present disclosure is directed a method for repairing a rotor blade of a wind turbine. More specifically, in certain embodiments, the rotor blade may be constructed, at least in part, of a thermoplastic material reinforced with at least one fiber material. Thus, the method includes identifying at least one defect on the rotor blade. For example, in certain embodiments, the defect(s) as described herein may include a crack, creep, void, hole, distortion, deformation, scratch, or any other blade defect. The method also includes applying at least one of heat, pressure, and/or one or more chemicals to the defect(s) for a predetermined time period until the defect is repaired.

Abstract: A turbine airfoil segment includes inner and outer platforms that are joined by at least one airfoil. The airfoil includes leading and trailing edges that are joined by spaced apart first and second sides to provide an exterior airfoil surface. The airfoil includes film cooling holes that have external breakout points that are located in substantial conformance with the Cartesian coordinates set forth in one of Tables 1 and 2. The Cartesian coordinates are provided by an axial coordinate, a circumferential coordinate, and a radial coordinate, relative to a zero-coordinate, and the film cooling holes have a diametrical surface tolerance relative to the specified coordinates of 0.20 inches (5.0 mm).

Abstract: A gas turbine engine component may include a coating adapted to protect the component during use. The coating may be applied by sintering metallic particles to form a metallic matrix fused to the component.

Abstract: A gas turbine comprises a compressor module, with a lower pressure compressor section including a plurality of stages, with at least one of the plurality of stages being an integrally bladed rotor. A higher pressure compressor section includes a plurality of stages with at least one of the plurality of stages being an integrally bladed rotor. A fan drive turbine shaft drives a fan rotor through a gear reduction. The fan rotor delivers a portion of air into a bypass duct, and a portion of air into the compressor module. A bypass ratio defined by the volume of air delivered into the bypass duct compared to the volume of air delivered into the compressor module is greater than or equal to about 6.0.

Abstract: A method for repairing or servicing a wye ring of a generator includes the steps of, dismantling the generator to gain access to the wye ring, determining a fault location in the wye ring; and attaching a patch to the wye ring in an area of the fault location. The patch provides an electrical path around the fault location so that the generator is repaired.

Abstract: A process for rejuvenating a turbine disk having a plurality of slots includes the steps of determining a depth of a damaged layer containing M23C6 carbide dissolution; and removing the damaged layer from the slots in accordance with the determined depth.

Abstract: A method for producing a component, in particular a rotor wheel, includes positioning a main body, a fusion-weldable attachment layer, and a barrier layer. The main body has a first alloy in an attachment region. The fusion-weldable attachment layer is positioned in the attachment region of the main body and has a second alloy which differs from the first alloy. The barrier layer is positioned between the main body and the fusion-weldable attachment layer. The barrier layer is configured to prevent a reaction of the first alloy of the main body with the second alloy of the fusion-weldable attachment layer during a thermal treatment. The method further includes heating the main body, the barrier layer, and the fusion-weldable attachment layer to connect the main body, the barrier layer, and the fusion-weldable attachment layer to one another.

Abstract: A layer system including a substrate and a layer is provided. The layer system has a hole at least in the layer, and wherein at least one recess is present in a surface of the layer in a vicinity around the hole or directly adjoining a boundary line of the hole.

Abstract: A method for manufacturing a turbine assembly, which includes at least two aerofoils arranged adjacent towards each other and embodied as a twin vane segment, is provided. The method includes: processing the turbine assembly with a first protection technique providing a first protection for the at least two adjacent aerofoils, wherein at least one region of one aerofoil remains unprocessed; pre-processing of the at least one region that will remain unprocessed or post-processing of the at least one region that remained unprocessed with a second protection technique to provide a second protection for the at least one region of the one aerofoil, wherein the first and the second protection techniques differ from one another.

Abstract: A reservoir groove for storing condensate water is formed on a lower part side of a link chamber inside a turbine housing. The reservoir groove is formed such that, assuming that water vapor contained in an exhaust gas retained inside the link chamber is completely condensed and liquefied after an engine stops its operation, a height position of a reference water surface of the condensate water stored in a reservoir area, which includes a space inside the reservoir groove in the link chamber, is lower than a height position of a lowermost part of the link mechanism.

Abstract: A process is provided for repairing a damaged portion of a gas turbine blade comprising: providing a preform comprising a low-melt alloy material and a base alloy material; locating the preform on the gas turbine engine blade damaged portion; and heat treating the preform and the blade such that the preform is brazed to the blade.

Abstract: A hydro turbine assembly includes a hub configured to rotate about a center axis and configured to be mounted in a water passage. The hub includes an upstream end, a downstream end and an outer surface between the upstream and downstream ends. The hub includes at least three mounting recesses arranged in the outer surface wherein each mounting recess includes a first hub mounting surface and a second hub mounting surface, and the second hub mounting surface is downstream and radially inward of the first hub mounting surface. The assembly includes at least three runner blades each including a base configured to seat in a respective one of the mounting recesses, wherein the base includes a first blade mounting surface arranged to abut the first hub mounting surface and a second blade mounting surface arranged to abut the second hub mounting surface.

Abstract: The present invention relates to a rotating blade (5), in particular for a compressor or turbine stage of a gas turbine, having a radially outer rib arrangement with at least one rib (2), onto which a coating (3) is disposed, whereby, in a meridian section, the coating (3) has an outer contour (3.1), which extends axially outwardly in the radial direction.

Abstract: The invention refers to a method for manufacturing a hybrid component including the following steps of manufacturing a preform as a first part of the hybrid component, then successively building up on that preform a second part of the component from a metallic powder material by means of an additive manufacturing process by scanning with an energy beam, thereby establishing a controlled grain orientation in primary and in secondary direction of at least a part of the second part of the component. The controlled secondary grain orientation is realized by applying a specific scanning pattern of the energy beam, which is aligned to the cross section profile of the component or to the local load conditions for the component.

Abstract: A rotary blade is provided that includes a root, an airfoil, and a tip insert. The airfoil is attached to the root, and has a suction side surface, a pressure side surface, a leading edge, a trailing edge, a tip, and a slot disposed within the tip. The slot extends in a chordwise direction between the leading edge and trailing edge. The tip insert has a base end and a rub end. The base end of the tip insert is disposed within the slot. The rub end of the tip insert extends radially outward from the airfoil tip.

Abstract: A process of fabricating a shield, a process of preparing a component, and an erosion shield are disclosed. The process of fabricating the shield includes forming a near-net shape shield. The near-net shape shield includes a nickel-based layer and an erosion-resistant alloy layer. The nickel-based layer is configured to facilitate secure attachment of the near-net shaped to a component. The process of preparing the component includes securing a near-net shape shield to a substrate of a component.

Abstract: The gas turbine engine (S1) includes: turbine blades (7b); and a cooling air supply unit (11) to supply cooling air to the turbine blades (7b). A flow path surface (31) is formed so as to be positioned in an upstream side of the turbine blades (7b) and so as to be connected to a base surface (32) in which the turbine blades (7b) are provided. The flow path surface (31) includes: depression portions (31a) depressed relative to the base surface (32), each depression portion (31a) including at least an area overlapping with a front end (7b1) of each turbine blade (7b), when viewed from a direction of the turbine axis (L); and protrusion portions (31b) protruding relative to the base surface (32), each protrusion portion (31b) being at least part of each area positioned between front ends (7b1) of the turbine blades (7b), when viewed from the above direction.

Abstract: A method for removing a rotor blade from a wind turbine may generally include installing a blade sock around an outer perimeter of the rotor blade, coupling a support cable to the blade root, lowering the rotor blade relative to the hub using the support cable, coupling at least one pulley cable between the rotor blade and a winch using at least one pulley, moving the pulley cable relative to the pulley to lower the rotor blade relative to the hub, applying a force through the blade sock as the pulley cable is moved relative to the pulley in order to control an orientation of the rotor blade and further lowering the rotor blade to a location on or adjacent to the support surface.

Abstract: A manufacturing method includes providing a substrate with an outer surface and at least one interior space, selectively deposited a coating on a portion of the substrate to form a selectively deposited coating having one or more grooves formed therein. The method further includes processing at least a portion of the surface of the selectively deposited coating to plastically deform the selectively deposited coating in the vicinity of the top of a respective groove. An additional coating is applied over at least a portion of the surface of the selectively deposited coating. A component is disclosed and includes a substrate, a selectively deposited coating disposed on at least a portion of the substrate, and defining one or more grooves therein, and an additional coating disposed over the selectively deposited coating. The substrate, the selectively deposited coating and the additional coating defining one or more channels for cooling the component.

Abstract: A method of protecting a part of a surface of an aircraft against contamination with insect residues and/or against icing is described. The method includes the following steps: a) coating the part of the surface to be protected with a coating that can, within a period of ?3 hours from take-off of the aircraft and in normal flying conditions, reduce the adhesion to the part of the surface to be protected until detachment occurs on the basis of (i) temperature change and/or (ii) UV radiation and/or (iii) addition of a decomposing component and/or (iv) difference in air humidity, and b) exposing the part of the surface to be protected to (i) a temperature change and/or (ii) UV irradiation (iii) and/or addition of a decomposing component and/or (iv) a difference in air humidity, so that the coating together with any contaminants and/or icing thereon detaches.

Abstract: A method of repairing a flange of a casing, or a casing made of aluminum, for a turbine engine or an airplane turboprop or turbojet, the flange including at least one through hole for passing a bolt for fastening equipment. The method includes: forming a spot face in the flange around the hole for passing the bolt; anodizing the bottom surface of the spot face; placing a washer on the spot face; and fastening the washer to the flange using adhesive.

Abstract: An abradable coating for application to a gas turbine engine part is formed from a titanium aluminide alloy, a filler material, and porosity. The coating may be applied to a part such as a casing made from a titanium alloy.

Abstract: A method for repair of composite materials includes applying a formulation to a ceramic matrix composite substrate. The formulation comprises a liquid carrier, a ceramic filler dispersed within the carrier, and a polymeric binder disposed in the carrier. The method further includes removing the carrier from the formulation to form a green composition; pyrolyzing the green composition to form a porous composition; disposing a liquid metal or metalloid within the pores of the porous composition to form an intermediate composite composition; and converting the liquid metal or metalloid to solid state to form a solid composite composition.

Abstract: A method of anchoring an implant in hard tissue, and/or hard tissue replacement material, includes the steps of providing an initial opening in the hard tissue, providing a thermoplastic augmentation element, a tool and a counter element, compressing the augmentation element between the tool and the counter element while energy is coupled into the tool and while a periphery of a liquefaction interface of the tool and the augmentation element and/or of a liquefaction interface of the augmentation element and the counter element is in the opening, thereby liquefying material of the augmentation element at the liquefaction interface(s) to yield liquefied material, causing portions of the liquefied material to penetrate into structures of the hard tissue, allowing the liquefied material to harden and to thereby become augmentation material, removing the tool and the counter element, and anchoring the implant in the opening including at least some of the augmentation material.

Abstract: A turbine component is provided. The turbine component includes an airfoil having a first surface and a second surface. A thermal barrier coating is coupled to the second surface, wherein the thermal barrier coating includes a first portion, a second portion and a trench defined between the first and second portions. A channel is coupled in flow communication to the first surface and the trench, wherein the channel includes a first sidewall and a second sidewall opposite of the first sidewall. The first and second sidewalls extend from the first surface and toward the trench at an angle. The turbine component includes a cover coupled to the second surface, wherein the cover includes a first end coupled to the first portion and a second end extending into the trench and spaced from the second portion.

Abstract: A method of removing a bucket from a turbomachine rotor wheel includes exposing a base portion of the bucket, positioning a pulling device radially outwardly of the base portion, connecting the base portion of the bucket to the pulling device through a linking rod, exerting an axially outwardly directed force on the linking rod through the pulling device, and removing the base portion from the rotor wheel.

Abstract: A fuel nozzle assembly for a gas turbine engine includes a flange and a pre-mix tube. The flange includes a first end that is configured to couple to an end cover of the combustor, and a second end that is opposite said first end. The pre-mix tube is coupled at a first end to the flange second end. The flange and the pre-mix tube are fabricated to operate at a natural frequency that is different from an operating frequency of the gas turbine engine.

Abstract: It is an object to provide a repairing method for a wall member with flow passages capable of reducing the time and the cost required for a partial repairing work on the wall member with flow passages. The repairing method for the wall member with flow passages includes a removal stage of removing a portion of the wall member with flow passages (1) provided with plural flow passages (4) between the base member (2) and the plate member (3), the base member welding stage of welding the repairing base member (6) to the portion removed in the removal stage, a flow passage forming stage of providing plural overlay welds on the repairing base member (6) and forming the flow passages (4) between the overlay welded beads (8), and a plate member welding stage of welding a repairing plate member (9) onto surfaces of the plural overlay welded beads (8).

Abstract: A rotor blade includes a blade main body having a proximal end connected to the blade root, and a tip shroud connected to a tip end of the blade main body so as to constitute a part of an annular shroud, a tip end side of a leading edge of the blade main body is made continuous to an outside surface of the tip shroud via a curved portion curved toward a trailing edge, fillets are provided between the tip end of the blade main body and an inside surface of the tip shroud, and the curved portion is located closer to the leading edge than the fillets, whereby the turbine rotor blade and the turbo machine can realize weight reduction and suppress performance degradation.

Abstract: In accordance with an exemplary embodiment, a method for repairing a damaged metallic component using additive manufacturing techniques includes separating a damaged portion of the damaged metallic component from an undamaged portion of the damaged metallic component, measuring the undamaged portion to determine the dimensions of the removed damaged portion, fabricating a replacement portion using additive manufacturing techniques in accordance with the determined dimensions of the removed damaged portion, and joining the replacement portion with the undamaged portion of the damaged metallic component to form a repaired metallic component.

Abstract: Tenons can be installed at the ends of turbine blades by applying a solid state bonding procedure such as a friction welding technique to join the tenons to the turbine blades. The tenons can be employed to assemble the turbine blades to one or more covers for the turbine blades. The tenons can be installed in connection with the restoration of tenons for turbine buckets in steam turbines and in connection with the renovation of rotor assemblies that include turbine blades secured to a common disk, tenons located at one end of respective turbine blades and one or more covers secured to the turbine blades by means of the tenons. In instances in which tenon remnants are present at the ends of the turbine blades, tenon enlargement elements can be joined to the tenon remnants by the solid state bonding procedure. The tenon installations can be carried out in situ.

Abstract: A masking system for an airfoil comprises a locator for positioning along the trailing edge of the airfoil, and first and second mask sections. The first mask section has a trailing edge portion with locator pins for cooperating with a positioning tab on the locator to position the first mask section along the trailing edge of the airfoil, and a leading edge portion for extending around the leading edge of the airfoil to locate the first mask section along the convex surface. The second mask section has a retainer portion for cooperating with a retainer process on the first mask section to retain the second mask along the leading edge of the airfoil, and a flange portion for extending between the locator pins to locate the second mask along the concave surface.

Abstract: A composite material turbine nozzle blade including an airfoil adapted to have a cooling fluid flow through it and extending between a shroud and a root is provided. The shroud is shaped to be attached to one or more turbine casings of a turbomachine and the root is shaped to provide a junction with a turbine internal casing. The root is produced with a loosened texture and includes an external upstream side loosened texture lug and an external downstream side loosened texture lug, the ends of the two external side lugs extending axially relative to the rotation axis of the turbomachine to form a device for supporting and centering the internal casing.

Abstract: A method of servicing an airfoil for use in a gas turbine engine. The airfoil assembly is defined by a base material and includes an airfoil and a platform from which the airfoil extends. A predetermined amount of the base material is removed from the airfoil assembly proximate to a fillet area of the airfoil assembly via water jet material removal. The fillet area comprises a junction between the airfoil and the platform and is located at an intersection between the airfoil and the platform. A remainder of the base material comprising base material of the airfoil assembly other than proximate to the fillet area is left intact.

Abstract: The invention relates to a stirrer organ in composite construction comprised of a metallic part and a hybrid casting, with the hybrid casting being fixed to the metallic part by means of at least one anchoring element which forms at least one back-cutting in the hybrid casting. Furthermore, the present invention relates to a method for rehabilitating a damaged metallic stirrer organ which when applied provides a rehabilitated stirrer organ in composite construction. The invention tackles the task of creating a stirrer organ in hybrid construction which features a largest possible metallic part because higher strength can hereby be achieved. This task is inventively solved in that the anchoring element is provided with a perforated plate which is spot-wise fastened to the metallic part and which is arranged at least partly at a small distance to the surface of the metallic part so as to configure the at least one back-cutting.

Abstract: An apparatus for shaping an edge of an aerofoil, the apparatus having first and second brushes and each brush having a plurality of bristles. A first motor rotates the first brush about a first axis and a second motor rotates the second brush about a second axis. The axes are arranged substantially parallel to the bristles of the respective brush. A support structure holds the first brush such that the first axis intersects a first surface of the edge of the aerofoil and holds the second brush such that the second axis intersects a second surface of the edge of the aerofoil. There are means to produce relative movement the first and second brushes and the aerofoil such that the first and second brushes move longitudinally along the edge of the aerofoil to shape the edge of the aerofoil.

Abstract: A gas turbine engine fan stand (40) comprising a base frame (42) and a fan case frame (44) coupled together at one edge by a hinge (46). The fan case frame (44) is arranged to: rotate about the hinge (46) between abutting the base frame (42) and substantially perpendicular thereto; and tilt about a yaw axis (52) perpendicular to the hinge (46) and in the plane of the fan case frame (44). The fan case frame (44) further comprising a coupling arrangement (56) arranged to rotate about a roll axis (60) and to couple a fan case (36) to the fan case frame (44). A gas turbine engine stand assembly (68) comprising a fan stand (40) and a core stand (38). A method of splitting a gas turbine engine (10) and a method of reassembling a gas turbine engine (10).

Abstract: A method for defining an airfoil with a radiused shape edge includes joining a filler material to the airfoil such that the filler material defines excess material that extends beyond an operational condition edge location of the edge of the airfoil, placing the airfoil within a barrel, providing an abrasive finishing media in the barrel with the airfoil, rotating the barrel to produce relative motion between the abrasive finishing media and the airfoil. Contact between the abrasive finishing material and the airfoil facilitates removing the excess material that extended beyond the operational condition edge location of the edge of the airfoil.

Abstract: Method for laterally replacing a heavy component of a plant assembly, the method including disconnecting the heavy component from other components of the plant assembly and from a base plate to which the heavy component is fixed; lifting the heavy component above the base plate with a lifting system provided within the base plate; connecting at least a pair of rails to the base plate, under the lifted heavy component, such that the at least a pair of rails extends at substantially a right angle relative to a longitudinal axis of the heavy component; lowering the heavy component on crawling mechanisms disposed on the at least a pair of rails; and laterally replacing the heavy component from the base plate and the other components of the plant assembly by actuating the crawling mechanisms.

Abstract: An apparatus for removing a turbine blade, the apparatus including a main body having a front plate including at least one securing opening for receiving a securing bolt, a rear plate having a fixture groove, and first and second side plates, and a base block configured to slide along the fixture groove of the rear plate.

Abstract: In a method for repairing worn or damaged rotor blades of an aircraft engine, laser cladding for the reconstruction of the blade tips is not performed in a continuous welding process along the blade edge but during a cycle-by-cycle rotation of the rotor blades attached to a rotor disk, in only one limited section of the respective blade edge for each cycle. The method reduces time expenditure for the repair of the blades and improves the tribological properties of the blade material in the area of the repaired blade tip.

Abstract: A method of replacing a wind turbine blade includes suspending the wind turbine blade from support hub of a wind turbine, connecting one or more cable climbing members between the support hub and the wind turbine blade, and lowering the one or more cable climbing members and the wind turbine blade from the support hub.

Abstract: The invention relates to a method for producing a plating (5) of a vane tip. Said method consists of the following steps: a) a vane having a vane tip which is arranged opposite the base of the vane (2) and which comprises a surface which points radially outwards is provided, and b) a porous layer (7) is applied to at least the surface (4) of the vane tip and/or c) a bulge (8) which increases the surface of the vane tip is applied to at least one part of the flanks of the vane tip, said flanks surrounding the surface of the vane tip, and d) the plating (5) is applied to the porous layer and/or the bulge. The invention also relates to corresponding vanes or gas turbines with corresponding vanes.