Abstract: A wind turbine blade root having fully bonded insert bushings with internal threads for mounting bolts for releasable attachment to a hub of a wind turbine is manufactured by arrangement of a plurality of the insert bushings on a first layer of fibre mat arranged on a holder having a plurality of spaced recesses for accommodation of the insert bushings such that the bushings extend largely in a longitudinal direction of the blade accommodated in the spaced recesses. Subsequently, a number of additional layers of fibre mat are arranged on top of the bushings, and the fibre mat layers are consolidated.

Abstract: A method of manufacturing a fibre reinforced metal matrix composite article, the method comprising placing metal coated (18) fibers (14) between a first metal ring (30) and a second metal ring (36). Each of the metal-coated (18) fibres (14) having a glue (22) to hold the metal-coated (18) fibers (14) in position. A solvent is supplied to the glue (22) on the metal-coated (18) fibers (14) to soften the glue (22) and pressure is applied to allow the metal-coated (18) ceramic fibers (14) to become more closely packed. Thereafter the glue (22) is removed and the metal coated (18) fibers (14) and first and second metal rings (30, 36) are consolidated and diffusion bonded together.

Abstract: A composite blading member includes an airfoil and a base of composite material layers and a platform of composite material layers interfused with the airfoil and base. The platform comprises a platform shelf and a plurality of spaced-apart platform supports integral with the platform shelf and interfused with surfaces of the base. Such blading member was made by providing a partially cured airfoil-base preform and a platform preform, including an airfoil shaped opening therethrough and preforms of the platform supports. The airfoil-base preform was inserted through the airfoil shaped opening whereby cooperating surfaces of the platform preform and airfoil-base preform were in juxtaposition. The preforms were heated to partially cure them together. Then they were interfused with a binder compatible with material of the preforms under conditions to substantially fully cure the structures.

Abstract: A method of manufacturing a composite airfoil includes the step of providing a core made of a metal or ceramic material. A plastic airfoil portion is molded to envelope at least a portion of the core.

Abstract: A method for surface treating a titanium gas turbine engine component. The method includes providing a gas turbine engine component having a titanium-containing surface. The component is heated to a temperature sufficient to diffuse carbon into the titanium and below 1000° F. The surface is contacted with a carbon-containing gas to diffuse carbon into the surface to form carbides. Thereafter, the carbide-containing surface is coated with a lubricant comprising a binder and a friction modifier. The binder preferably including titanium oxide and the friction modifier preferably including tungsten disulfide. The coefficient of friction between the surface and another titanium-containing surface is less than about 0.6 in high altitude atmospheres.

Abstract: A method of assembling a turbine bucket is provided. The bucket includes a dovetail portion, an airfoil portion, and a root that extends between the airfoil portion and the dovetail portion. The turbine bucket includes a pressure side and a suction side that are connected together at a leading edge and a trailing edge. The method includes forming at least one pocket within the turbine bucket, such that the pocket is formed within the pressure side. The method also includes forming at least one pin within the at least one pocket.

Abstract: Apparatus and methods related to wind turbine blades and manufacturing of the blades.

Abstract: The invention relates to a wind turbine rotor blade comprising a blade tip and a lightning protection system. The rotor blade includes at least one lightning receptor at the surface of the blade in an external distance (Lex) from the distal end of the blade tip, and a lightning receptor base inside the rotor blade arranged at an first internal distance (Li1) from the distal end of the blade tip. The rotor blade further includes means for changing at least one electric property of the rotor blade at the lightning receptor base, as compared to the electric properties of the ambient air by increasing the electric field supported between the lightning receptor base and the inner surface of the rotor blade. The invention further relates to a method for manufacturing a wind turbine rotor blade.

Abstract: A component (10) and method for manufacturing a component. The component comprises first and second panels (16, 18) spaced apart from each other, the first panel (16) having at least one protrusion (20) extending therefrom towards the second panel (18). The protrusion (20) extends partially across the space (19) between the first and second panels so as to define a free end of the protrusion. The free end is surrounded by damping material (24).

Abstract: The present invention proposes a lightning protection device of a windmill blade, comprising a tip receptor assembly having receptors which are mounted, in particular to, the tip end part of the blade where the blade profile becomes smaller, and having a simple structure so as to be readily assembled, and an intermediate receptor assembly having a receptor which is set at the outer surface of the blade so as to exhibit an effective countermeasure against increasing of the air pressure in the blade upon a lightning striking upon the blade, a lightening current being discharged to the ground from the receptors through connection equipment such as lead wires laid through the inside of the blade the windmill body and bypass brushes, characterized in that a base plate made of an electrically conductive material and connected thereto with the lead wires is embedded in the blade, and the receptors are secured and supported to the base plate, being exposed at the outer surface of the blade.

Abstract: An integrally bladed rotor is constructed from a plurality of layers of organic matrix composite material wound together in a spiral fashion to form the disc portion of the rotor, and at each rotor blade position at least the outermost one of the layers is turned substantially radially outwards from the periphery of the rotor disc to form a blade. Each blade is finished with further pieces of organic matrix composite material bonded into position on the. An encircling blade tip shroud may be formed by further layers of material wound around the tips of the blades in conjunction with closed loop inserts in the spaces between blades.

Abstract: A butt connection of divided hollow profile members, which is suitable in particular for rotor blades of wind power installations, comprises a multiplicity of straps which are arranged along the joint and which bridge over same and which are respectively fixed with their ends to one of the profile members to be connected. In this respect the arrangement is preferably such that one of the two bolts fixing the strap at the ends thereof has a wedge-shaped flattening, by means of which a tensile prestressing can be imparted to the strap.

Abstract: A method of manufacturing a component such as a gas turbine fan blade. The method comprises: disposing a plurality of webs between a plurality of panels; and deforming the panels and the webs by applying internal pressure between the panels, thereby forming a series of internal cavities (15) partitioned by the webs (20-22). Weakened regions (30a,30b,30c) are formed in the webs which open to form holes (30) in the webs (20-22) during the deformation step. A damping material (17) is introduced into at least one of the cavities (15) whereby it flows from the cavity into the other cavities via the holes in the webs.

Abstract: The present invention relates to a rotorcraft blade (1) provided with a distributed spar (10) on a leading edge (4), a suction side (2), and a pressure side (3) of the blade, and the invention also relates to a method of fabricating such a spar. The blade is provided with a fastener insert (20) integrated in the blade root for the purpose of fastening the blade to a rotor, the fastener insert (20) comprising a horizontal shaft (21) perpendicular to the span of the blade and to an axis of rotation of said rotor, with the distributed spar (10) of the blade being wound in part about the horizontal shaft.

Abstract: The present invention provides near-surface cooled airfoils that can be made with near-surface cooling passages that are completely free of any leachable or otherwise sacrificial material in the recessed portion of the outer surface of the core. The turbine airfoil comprises a metallic core or substrate having an outer surface and one or a plurality of recessed portions of the outer surface; an intermediate metallic skin or foil having a back surface and a top surface, the back surface of the intermediate skin being bonded to the outer surface of the core such that the recessed portion(s) is sufficiently enclosed so as to form at least one or more near-surface cooling passages or pathways; and at least one or more metallic coatings of a high temperature-resistant metallic material deposited on a top surface of the intermediate skin.

Abstract: A rotor blade assembly includes a shank, an airfoil that is formed integrally with the shank, and a removable platform coupled between said shank and said airfoil via a friction fit. A method of assembling a gas turbine engine rotor blade assembly that includes a removable platform, and a gas turbine engine rotor assembly including the removable platform are also described herein.

Abstract: A method for producing a blade tip armor plating on a blade of a turbomachine, in particular on a high-pressure compressor rotary blade of a gas turbine, having the following steps: a) provision of a blade made of a nickel base material; b) application of a solder to a blade tip of the blade, the solder being a nickel base solder; c) application of hard material particles of cubic boron nitride to the solder, the hard material particles of cubic boron nitride being coated with titanium or with a titanium base material; d) fusion of the solder in a vacuum in order to form a matrix that at least partially surrounds the hard material particles; and e) fixing of the blade tip armor plating through cooling.

Abstract: A cooled turbine engine component is made by providing first and second pieces respectively having first and second surfaces. At least one circuit is formed in at least one of the first and second surfaces. A first plurality of apertures is provided in the first piece to form inlets to the at least one circuit. A second plurality of apertures is provided in the second piece to form outlets to the at least one circuit. A combination of the first and second pieces is assembled and integrated.

Abstract: A turbine airfoil system for forming a turbine airfoil that is usable in a turbine engine. The airfoil may be formed from a porous material shaped into an outer airfoil shape. The porous material may include an inner central spar capable of supporting the turbine airfoil an outer porous region and an outer coating. The porous material facilitates efficient cooling of the turbine airfoil.

Abstract: The present invention includes a method for fabricating elongated wind turbine blades and wind turbine blades formed by the method. The method includes providing a first shell reinforcing fiber structure. At least one shear web reinforcing fiber structure is positioned adjacent to the first shell reinforcing fiber structure. The method includes infusing the first shell reinforcing fiber structure and shear web reinforcing fiber structure with a matrix material and curing the matrix material to form a unitary composite first shell component. Thereafter a composite second shell component is attached to the composite first shell component to form an elongated composite airfoil suitable for use as a wind turbine blade. The wind turbine blades formed include unitary components providing reduced number of adhesive joints.

Abstract: A blade or a vane preferably used in a gas turbine engine, the blade or vane being formed of a metallic core or insert with a ceramic airfoil surface bonded to the metallic insert. The insert includes a series of concave and convex portions arranged along the outer surface, in which a plurality of hoop fibers are secured within a resin matrix to firmly secure the metallic insert to the ceramic airfoil body. Axial loads between the insert and the ceramic body are transferred to the hoop fibers. The hoop fibers are strong enough to resist extending in length that would permit the hoop fibers from sliding over the widest portion of the insert.

Abstract: An article includes a blade casting core combination. The combination includes a ceramic feedcore and a metallic core. The ceramic feedcore has: a root end; a tip end; a leading end; a trailing end; a first side; a second side; and a plurality of legs extending between the root and tip ends and arrayed between the leading and trailing ends. The metallic core has: a first face; a second face; a first portion extending from the feedcore trailing end; and a second portion extending from the tip end. The article may be a pattern where the core is embedded in a wax or may be a shell formed from such a pattern. The article may be used in a method for forming the resultant blade.

Abstract: A method for fabricating an outlet guide vane includes the steps of fabricating a structural spar from a first material, fabricating a fairing from a second material, and installing the fairing onto the spar so as to at least partially surround the spar and form an airfoil. An outlet guide vane includes an airfoil including a leading edge and a trailing edge, a structural spar formed from a first material located within the airfoil, and a fairing formed from a second material at least partially surrounding the spar. A gas turbine engine assembly includes a core gas turbine engine, a fan assembly including a plurality of fan blades coupled to the core gas turbine engine, and a plurality of outlet guide vanes coupled downstream from the fan blades, at least one of the outlet guide vanes including an airfoil having a leading edge and a trailing edge, a structural spar formed from a first material located within the airfoil, and a fairing formed from a second material at least partially surrounding the spar.

Abstract: A process for forming an airfoil portion of a turbine engine component, such as a turbine blade, is described. The process comprises the steps of placing a ceramic core having a configuration of a passageway to be formed in the airfoil portion within a mold, attaching a refractory metal core element to the ceramic core to stabilize a tip region of the ceramic core, and casting the airfoil portion.

Abstract: A turbine blade for turbo-engines as well as a method for manufacturing such a turbine blade is disclosed. According to the task set, the turbine blades should be capable of withstanding high thermal stress and able to maintain an adequate mechanical strength even at raised operating temperatures. The turbine blades are so designed that on the surface of a core element a heat-insulating layer of a metallic open-cell foam is integrally connected to said core element by sintering. The outer contour of the turbine blade is formed with at least one shell element. The shell element comprises a nickel-base alloy, which is also integrally connected by sintering to the open-cell foam which forms the heat-insulating layer.

Abstract: A method of manufacturing a fibre reinforced metal matrix composite article, the method comprising placing at least one metal coated (18) fibre (14) between a first metal ring (30) and a second metal ring (36). Heating to a first temperature and applying a first pressure to partially consolidate the metal (18) on the at least one metal coated (18) fibre (14) and heating to a second temperature and applying a second pressure to consolidate the metal of the first and second metal rings (30, 36), wherein the first temperature is different to the second temperature and the first pressure is different to the second pressure.

Abstract: A method of manufacturing a fibre reinforced metal matrix composite article (10) comprises forming an annular groove (32) in the first metal component (30), forming a second metal component (36) and forming a number of fibre preforms (20). The fibre preforms (20) are placed in an annular channel (80) in a cassette (70). The cassette (70) and the first metal component (30) are arranged such that the annular channel (80) in the cassette (70) is coaxial with and faces the annular groove (32) in the first metal component (30). The fibre preforms (20) are moved from the annular channel (80) in the cassette (70) to the annular groove (32) in the first metal component (30). The second metal component (36) is placed on the first metal component (30) such that the fibre preforms (20) are arranged between the first metal component (30) and the second metal component (36). The second metal component (36) is sealed to the first metal component (30).

Abstract: A blade (10, FIG. 1) for a gas turbine engine includes an aerofoil portion (12, FIG. 1) and a root portion 14 defined by concave and convex walls 16, 18 having opposing inner surfaces 16a, 18a. A reinforcing member 20 is located between the concave and convex walls 16, 18 and is bonded to the inner surfaces 16a, 18a thereof, and the root portion 14 includes an unbonded region 24a, 24b in which the reinforcing member 20 contacts an inner surface 16a, 18a of one of the concave and convex walls 16, 18 but is not bonded thereto.

Abstract: A method of manufacturing an article by diffusion bonding at least two metal components (30,36) comprises assembling the metal components (30,36) into a stack relative to each other so that the surfaces (34,40) are in mating contact. The edges of the metal components (30,36) are sealed together, except for an aperture (49) where a pipe (50) is to be connected. A connector (51) has a first end (52) and a second end (53). The first end (52) has a smaller diameter than the second end (53) and a bore (54) extends through the connector (51) from the first end (52) to the second end (53). The second end (53) of the connector (51) is joined (57) to the stack. The pipe (50) is joined (58) to the first end (52) of the connector (51). The interior of the sealed assembly is evacuated via the pipe (50) and the pipe is sealed. Heat and pressure are applied to diffusion bond the metal components (30,36) together to form an integral structure.

Abstract: A welded component includes at least one high temperature segment of a high alloy Cr steel with high creep strength and a low temperature segment of a low alloy steel with high toughness and/or a high yield strength which are connected materially to one another via a weld joint. In one such component a gradual transition of chemical, physical and mechanical properties in the joining area is achieved in that between the weld joint and the high temperature segment there are at least two successive clad layers of at least two lower alloy weld metals with a total content of elements which increase the creep strength, such as for example Cr, Mo, W and V, which total content decreases toward the weld joint, and/or an increasing total content of elements which increase the toughness and/or yield strength, such as for example Ni and Mn, which total content increases toward the weld joint.

Abstract: A disbond resistant multi-laminate composite product is provided having, a first member; a second member of composite material bonded to the first member; a third member of composite material bonded to at least the second member and being so positioned and arranged such that when the second and third members are stressed at least a portion of the third member will have higher stress than the adjacent second member. The third member preferably has a terminal portion adjacent the second member in which terminal portion the strain energy release rate is higher than in the adjacent second member when the second and third members are stressed. The second and third members are preferably so positioned and arranged such that the third member initiates delamination from the second member before the second member initiates delamination from the first member when the first, second and third members are under stress.

Abstract: Fabricating an IBR, in particular a two-part IBR. A metal container is defined, made up of a plurality of parts that define between them at least one annular cavity, an insert made of composite material is positioned in the or each cavity, the assembly is subjected to hot isostatic compacting, and a rotor disk is machined.

Abstract: A repaired gas turbine blade includes a turbine blade body having a monocrystalline airfoil made of a first nickel-base superalloy, and a repair squealer tip welded to a tip of the airfoil. The repair squealer tip is made of a second nickel-base superalloy different from the first nickel-base superalloy and having less than about 0.15 weight percent total of carbon, boron, silicon, zirconium, and hafnium.

Abstract: An airfoil (44) formed of a plurality of pre-fired structural CMC panels (46, 48, 50, 52). Each panel is formed to have an open shape having opposed ends (54) that are free to move during the drying, curing and/or firing of the CMC material in order to minimize interlaminar stresses caused by anisotropic sintering shrinkage. The panels are at least partially pre-shrunk prior to being joined together to form the desired structure, such as an airfoil (42) for a gas turbine engine. The panels may be joined together using a backing member (30), using flanged ends (54) and a clamp (56), and/or with a bond material (36), for example.

Abstract: A platform arrangement for components such as aerofoil vanes 1, 21 is provided whereby the necessary rigidity provided by the platform does not create excess problems with respect to stress and strain fatigue failure at thinner sections of the component. The leading edge 6, 26 and trailing edge 7, 27 of a vane generally have a thinner section and so when subject to greater flex deformation such flex deformation may lead to premature failure through stress and strain. Structural configurations are now provided which comprise cut out windows 24 in the platform 22 beneath the edges 26, 27 or alternatively differential material types at these locations with a less stiff nature.

Abstract: A shaft portion is processed to have a desired diametrical thickness by rolling in place of cutting so that a cutting operation requiring much time may be eliminated from a manufacturing process of the adjustable blade. A shaft elongation and a sharp edge caused by the rolling can be inhibited to thereby eliminate an additional cutting operation conventionally performed for correcting the shaft elongation and the sharp edge. In particular, a metal shaped material for an adjustable blade integrally provided with a blade portion and a shaft portion is formed to have a shape and a size close to those of the desired adjustable blade. In forming the shaft portion, a shaft portion forming section of the shaped material is rolled to have a desired diametrical thickness.

Abstract: A method of manufacturing a blade for assembly on a steam turbine rotor wheel includes forming an airfoil portion with plural pockets and filling the pockets with more than one filler material chosen as a function of required temperature capability.

Abstract: The composite turbomachine blade of the present invention comprises a preform made of yarns or fibers woven in three dimensions and a binder maintaining the relative disposition between the yarns of the preform. Said preform is made up of warp yarns and of weft yarns, the direction of the warp yarns forming the longitudinal direction of the preform. In characteristic manner, said preform comprises at least a first portion made using a first weave forming the airfoil of the blade, and a second portion made using a second weave forming the root of the blade, and the first and second portions are united by a transition zone in which the first weave is progressively modified to end up with the second weave, thereby obtaining a reduction at least in the thickness of the blade between the second portion and the first portion. The invention is applicable to making a fan blade for a turbojet.

Abstract: A method of manufacturing a composite structure uses a layer of an insulating material (22) as a mold for forming a substrate of a ceramic matrix composite (CMC) material (24). The insulating material may be formed in the shape of a cylinder (10) with the CMC material wound on an outer surface (14) of the cylinder to form a gas turbine combustor liner (20). Alternatively, the insulating material may be formed in the shape of an airfoil section (32) with the CMC material formed on an inside surface (36) of the insulating material. The airfoil section may be formed of a plurality of halves (42, 44) to facilitate the lay-up of the CMC material onto an easily accessible surface, with the halves then joined together to form the complete composite airfoil. In another embodiment, a box structure (102) defining a hot gas flow passage (98) is manufactured by forming insulating material in the shape of opposed airfoil halves (104) joined at respective opposed ends by platform members (109).

Abstract: A method of processing a rotor. The rotor is formed by casting an ingot to have first and second regions formed of different alloys that intermix during casting to define a transition zone therebetween. The ingot is forged to yield a rotor forging that contains axially-aligned first and second alloy regions and a transition zone therebetween. A three-dimensional approximation of the transition zone is generated, which can be used to predict the effects of the transition zone on the dynamic performance of a rotor machined from the forging.

Abstract: An airfoil (30) having a continuous layer of ceramic matrix composite (CMC) material (34) extending from a suction side (33) to a pressure side (35) around a trailing edge portion (31). The CMC material includes an inner wrap (36) extending around an inner trailing edge portion (38) and an outer wrap (40) extending around an outer trailing edge portion (42). A filler material (44) is disposed between the inner and outer wraps to substantially eliminate voids in the trailing edge portion. The filler material may be pre-processed to an intermediate stage and used as a mandrel for forming the outer trailing edge portion, and then co-processed with the inner and outer wraps to a final form. The filler material may be pre-processed to include a desired mechanical feature such as a cooling passage (22) or a protrusion (48).

Abstract: The invention relates to a method for carrying out nondestructive testing of a gas turbine blade during which corrosion areas, which are located near the surface and which consist of oxidized carbides or of base material that is sulfidized near the surface, are determined by means of an eddy current measurement. As a result, the blades can be ground or sorted, in particular, before subjecting the gas turbine to a complicated cleaning and coating process.

Abstract: A part is used that constitutes a platform web and that has a shape close to the shape of the platform with an inside face and an outside face that are opposite each other. Fiber reinforcement plies are draped on a blade core and at least some of these plies extend beyond an end of the blade that is to be connected to the platform so as to form ply extensions. The ply extensions are folded down against the inside and outside faces of the platform to form flaps securing the blade to the platform, and fiber reinforcement plies are draped around the platform web and over the flaps so as to form a platform box retaining the flaps.

Abstract: A method of producing an element of composite material, including the steps of forming a first distribution of first elements defining the matrix of the element of composite material; forming a second distribution of second elements defining the reinforcing structure of the element of composite material; and compacting the first and second elements to obtain a distribution of the reinforcing structure inside the matrix; the first elements being metal wires; and the step of forming the first distribution including the step of assigning each second element an orderly distribution of metal wires.

Abstract: An article includes turbulation material bonded to a surface of a substrate via a bonding agent, such as a braze alloy. In an embodiment, the turbulation material includes a particulate phase of discrete metal alloy particles having an average particle size within a range of about 125 microns to about 4000 microns. Other embodiments include methods for applying turbulation and articles for forming turbulation.

Abstract: In a process for the manufacture of a blade/vane of a turbomachine, the blade/vane is made of a metallic enveloping structure and a supporting structure made of composite material provided inside of the enveloping structure and suitably joined to the enveloping structure. A metal-felt weave is welded or soldered/brazed to the sheet making up the enveloping structure, whereupon this sheet-weave-assembly is shaped by hydroforming or an internal high-pressure forming process to provide the enveloping structure of the blade/vane. Finally, the composite material making up the supporting structure is introduced into the cavity of the enveloping structure.

Abstract: A fiber reinforced composite article is made by providing a plurality of layers of first, substantially dry, unimpregnated reinforcing fibers, sometimes called in-plane fibers. Such layers are stacked into a preform of a stack of layers. While in the dry, unimpregnated condition, a plurality of spaced apart second reinforcing members are inserted into the preform at an angle, for example transversely, to the stack of layers. In this condition, the preform is impregnated with a matrix about the first reinforcing fibers and the second reinforcing members. The article product comprises the plurality of stacked layers of the first reinforcing fibers and the plurality of the spaced apart second reinforcing members disposed beside or adjacent the first reinforcing fibers, with a substantially solid matrix disposed about the reinforcing fibers and members.

Abstract: The present invention comprises a rotor blade and method of manufacture thereof. The rotor blade of the present invention incorporates a number of novel features in order to increase its reliability and strength while reducing its weight. In certain embodiments, the rotor blade of the present invention incorporates a flexural strap in place of traditional pitch bearings. In certain embodiments, the rotor blade of the present invention is manufactured using a unique fiber placement layup technique to reduce failure modes and maximize the fatigue strength of the part.

Abstract: A hollow liner (14) for fitting in the interior of a aerofoil vane (10) is crimped prior to fitting within the vane (10). Following fitting, the liner (14) is expanded so as to fit snugly within the vane (10). The technique permits the insertion of liners (14) into vanes (10) that are so configured as to preclude the insertion of conventional liners.

Abstract: A thermal barrier coating system for a superalloy substrate is disclosed. The superalloy is preferably of the type that is capable of forming an adherent alumina layer. A bond coat is applied to a local area of the substrate, so that a portion of the substrate remains exposed. The localized area is defined to be the area(s) at which a TBC typically fails first, e.g., the leading and trailing edges of an airfoil, or other area. An alumina layer is formed on the remaining portion of the substrate, and also on the bond coat. A ceramic layer is then applied on the alumina layer. Even if the ceramic material is removed, the localized bond coat remains, and reduces the rate at which the underlying substrate oxidizes. A coated article is also disclosed, as is a system that utilizes a conventional superalloy and aluminide coating with the localized bond coat.