Abstract: A turbine rotor blade includes a hollow airfoil joined to a platform and dovetail. A main flow channel extends longitudinally in span through the blade and is bound chordally by opposite partitions transversely bridging opposite sidewalls of the airfoil. A damper rib and transversely opposite damper pad are arranged together in a plurality of pairs spaced longitudinally apart in the airfoil and chordally positioned intermediate to the partitions to provide unobstructed forward and aft portions of the flow channel for channeling a coolant therethrough. The damper ribs and pads are configured to receive a wire damper through the channel to locally dampen vibration while minimizing obstruction of the coolant flow therethrough.

Abstract: A blade for a wind turbine includes a device for causing air at an increased velocity and/or pressure to escape from exit openings in the blade, thereby to apply a force. Pressure and diameter adjustable valves may be provided at each exit opening or orifice to allow the air to exit from the blade only when the air pressure within the blade exceeds a preset level. Also disclosed is the provision of valves at the exit openings or orifices on an opposite edge of the blade that allows air to escape from the blade to slow down but not stop rotation of the blade when a dangerous wind condition is detected.

Abstract: A means (22) for structurally stiffening or reinforcing a ceramic matrix composite (CMC) gas turbine component, such as an airfoil-shaped component, is provided. This structural stiffening or reinforcing of the airfoil allows for reducing bending stress that may be produced from internal or external pressurization of the airfoil without incurring any substantial thermal stress. The stiffener is disposed on a CMC wall and generally extends along a chord length of the airfoil.

Abstract: A blade assembly for a horizontal-axis wind turbine includes a hub, and a plurality of detachable blade units each mounted on a periphery of the hub and each including a plurality of blades detachably connected with each other by a plurality of connecting devices. Thus, the blade assembly comprises a hub and a plurality of detachable blade units so that all parts of the blade assembly are detachable to reduce the whole volume of the blade assembly before assembly, thereby facilitating packaging, storage and transportation of the blade assembly.

Abstract: A wind turbine blade (1, 3, 5) comprising at least one central spar longitudinal section (7, 7?, 7?) composed of two cap prefabricated panels (15, 17) and two web prefabricated panels (19, 21) placed side by side in a box shape and at least two shell longitudinal sections (5, 5?, 5?, 5??, 5??; 9, 9?, 9?, 9??) forming, respectively, the leading edge and the trailing edge of the corresponding blade section that are placed adjacently to a central spar section (7, 7?, 7?) and are composed of a single prefabricated panel (31, 33) or of two prefabricated panels (11, 23; 13, 25), the aerodynamic profile of the blade being defined by said cap panels (15, 17) and said single shell panels (31; 33) or said two shell panels (11, 23; 13, 25).

Abstract: An airfoil for a gas turbine engine includes: opposed pressure and suction sidewalls extending between spaced-apart leading and trailing edges, a root, and a tip; and a hollow plenum disposed between the sidewalls. The plenum has a chordwise first dimension at a radially outer end thereof, and a chordwise second dimension at a radially inner end thereof. The second dimension is substantially less than the first dimension.

Abstract: A rotor blade for use in a wind energy system is provided with the rotor blade having a rotor blade body comprising a rotor blade body tip; a rotor blade body cavity; and a lightning protection unit. The lightning protection unit includes a conductive part that is connected to a down conductor and positioned within the rotor blade body cavity close or at the rotor blade body tip, the conductive part having an extension protruding out of the rotor blade body for receiving a receptor. According to another aspect of the invention, a wind energy system is provided with at least one of such a rotor blade.

Abstract: A cooling system for an airfoil portion of a turbine engine component is provided. The cooling system includes a first cavity dedicated to cooling a trailing edge portion of an airfoil portion and a second cavity dedicated to cooling an aft portion of a pressure side wall.

Abstract: The present invention is directed to a hollow turbine airfoil having a cooling system designed to provide enhanced cooling to the fillet of a turbine airfoil. The turbine airfoil may include at least one fillet cooling channel, passing proximate to the fillet. A portion of the fillet cooling channel may be positioned proximate to the fillet outer surface without breaching an outer surface of the turbine airfoil. The turbine airfoil may include a vortex plate positioned adjacent to the end wall inner surface proximate to the fillet and an opening of the fillet cooling channel may be in fluid communication with the vortex chamber. The turbine airfoil may also include at least one end wall film cooling channel that may extend obliquely through the end wall and may be in fluid communication with the vortex chamber.

Abstract: A method for making a blade includes stacking a plurality of modular segments having formed conduits, threading and tensioning cables through the conduits of the stacked segments, and clamping the cables to hold the modular segments together.

Abstract: There is described a turbine blade of a gas turbine comprising a suction-side blade wall and a pressure-side blade wall that, in the trailing edge area, runs toward the suction-side blade wall. This suction-side blade wall has a trailing edge projection projecting over the end edge of the pressure-side blade wall at least in a partial section of the trailing edge area. A turbine blade of the aforementioned type is designed for a particularly long serviceable life while requiring little cooling air. To this end, the pressure-side surface of the trailing edge projection has a number of trough-shaped recesses that increase the transfer of heat from the trailing edge projection by a local swirling of the cooling air.

Abstract: A turbine blade includes an airfoil having first and second tip ribs extending along the opposite pressure and suction sides thereof. The tip ribs are spaced apart between the leading and trailing edges of the airfoil to include a trifurcate tip baffle defining a triforial tip cavity.

Abstract: A turbine blade includes forward and aft serpentine cooling circuits terminating in corresponding forward and aft impingement channels. Each serpentine circuit has two metered inlets for distributing primary inlet flow to the first passes thereof and supplemental inlet flow to the last passes thereof.

Abstract: A gas turbine engine includes a compressor, combustor, and high pressure (HP) turbine operatively joined together. An interstage cooling circuit is joined in flow communication from an intermediate stage of the compressor to a forward face of an HP disk supporting a row of turbine blades for channeling interstage bleed cooling air thereto.

Abstract: A gas turbine engine includes a compressor, combustor, and high pressure turbine operatively joined together. A first interstage bleed circuit is joined in flow communication between a first preultimate stage of the compressor and hollow blades in the turbine to provide thereto pressurized primary air. A second interstage bleed circuit is joined in flow communication between a second preultimate stage of the compressor and the turbine blades to provide thereto pressurized secondary air at a lower pressure than the primary air.

Abstract: A method for manufacturing a turbine engine component comprises the steps of forming a first half of an airfoil portion of the turbine engine component and forming a plurality of microcircuits having at least one passageway on an exposed internal wall of the first half of the airfoil portion. The method further comprises forming a second half of the airfoil portion of said turbine engine component, and forming at least one additional cooling microcircuit having at least one passageway on an exposed internal wall of the second half of the airfoil portion. Thereafter, the first half is placed in an abutting relationship with the second half after the cooling microcircuits have been formed and inspected. The first half and the second half are joined together to form the airfoil portion.

Abstract: A method for forming a bond between a first and a second shell of a blade is provided. The method comprises the steps of forming a cavity between said first and second blade shells and filling an adhesive into said cavity to form a bond line between said first and second blade shells.

Abstract: Hollow fan blades for turbo fan gas turbine engines are formed of two separate detail halves. Each detail half has a plurality of cavities and ribs machined out to reduce weight. The floor and opposite interior walls of each cavity are machined simultaneously. The configuration minimizes the number of cutter plunge cuts for the internal cavities, and maximizes cutter size, in order to minimize the time required to machine them. These detail halves are subsequently bonded and given an airfoil shape in the forming operation.

Abstract: With regard to hollow blades for turbine engines, it will be understood there is a problem with respect to percussive impact resulting in excessive distortion of the blade as well as potential failure as a result of blade tip bulging. By provision of ridges 107, 207, 307 which coincide and engage each other under impact, the extent of impact deformation is limited as well as a result of the narrowing between the ridges, a reduction in the possibility for fragmentary insert movement to bulge the cavity towards the tip 102, 202, 302 of a blade 100, 200, 300.

Abstract: A blade for a rotor has integral stiffeners with at least one of unidirectional caps configured to carry blade bending loads in axial tension or webs to carry transverse shear resulting from blade bending.

Abstract: A method for manufacturing a turbine engine component comprises the steps of forming a first half of an airfoil portion of the turbine engine component and forming a first cooling microcircuit having at least one passageway on an exposed internal wall of the first half of the airfoil portion. The method further comprises forming a second half of the airfoil portion of said turbine engine component, forming a second cooling microcircuit having at least one passageway on an exposed internal wall of the second half of the airfoil portion, and placing the first half in an abutting relationship with the second half after the cooling microcircuits have been formed and inspected.

Abstract: A cooled blade of a small gas turbine engine can enhance cooling performance without increasing the amount of cooling air in a simple configuration. A cooled turbine blade is provided with a cooling passageway 12 formed inside thereof, making the cooling air flow therein. Film-cooling holes 13 penetrate from an inner wall surface 111 to an external wall surface 112 and form a cooling film. An impingement cooling member 2 has a multiple number of small holes 21 ejecting the cooling air. A gap t made by the inner wall surface 111 and the impingement cooling member 2 has a sealing portion 14 mounted thereon which divides the gap in a blade chord direction.

Abstract: A damping arrangement for a blade (20) of an axial turbine, in particular a gas turbine, includes a damping element (17) which is arranged in a recess (16) in the blade aerofoil (10) of the blade (20) and frictionally dampens the vibrations of the blade (20). In such a damping arrangement, simplified manufacture and assembly and a reliable and effective function are achieved by the recess being configured as a cavity (16) extending in the radial direction through the inside of the blade aerofoil (10), the damping element (17) being inserted in the radial direction into said cavity (16).

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 method for generating an internal pack coating having different, controlled thicknesses includes partially filling a root opening of a turbine blade having a cavity therein with a first powder and a second powder having different formulations so that the first powder contacts a first predefined portion of the surface of the cavity and the second powder contacts a second predefined portion of the surface of the cavity. The method further includes heating the object with the first powder and the second powder therein to thereby produce a coating of the internal cavity having different coating thicknesses over the first portion of the surface of the cavity and the second portion of the surface of the cavity.

Abstract: A hollow fan blade for the fan of an aircraft engine includes a blade base section (2) and a blade tip section (5) which are assembled at opposing faying surfaces (6, 6?) by a joining process. Starting from the respective faying surface, cavities (7, 9) are produced in the two blade portions (2, 5) which are dimensioned in accordance with the loads occurring in the respective blade areas. The joining weld (4) is in a low-loaded area. The blade base section and the blade tip section can be constructed of solid material with formed-in cavities. The blade tip section can also be a sheet-metal structure.

Abstract: A lightweight component in the form of an outlet guide vane 10 for a gas turbine engine. The vane 10 being filled with hollow metal ellipsoid members 12 in a solid metal outer 14.

Abstract: A hollow component for gas turbine engine, for example an outlet guide vane, is assembled from a body having at recessed pocket formed therein, and an aerodynamic cover bonded over the pocket. Both the cover and the body are constructed from materials which are not readily fusion weldable. The cover is attached to the body by a solid state bonding process, for example friction stir welding. The hollow component may also be built up from multiple individual components which are bonded to each other by a solid state bonding process such as friction stir welding.

Abstract: A compressor blade (26) comprises an aerofoil (36) and a root (38). The aerofoil (36) comprises a concave wall (40) extending from a leading edge (44) to a trailing edge (46) and a convex wall (42) extending from the leading edge (44) to the trailing edge (46). The aerofoil (36) defines at least one internal chamber (48). The root (38) is connected to the aerofoil (36) and the root (38) has a base (50) remote from the aerofoil (38) and at least one aperture (52) extends from the base (50) to the at least one internal chamber (48) in the aerofoil (36). The dimensions, shape and position of the least one aperture (52) are selected such that the root (38) is deformable. This reduces the weight of the containment region of a fan casing (28).

Abstract: Percussive impacts due to bird strikes upon the hollow fan blades 20 are a well known problem. These percussive impacts not only deform previous fan blades but also reduce their stiffness. In accordance with the present invention, hollow fan blades 20 incorporate a cavity 23 within which, a matrix 24 with embedded expandable elements 25, is located. Thus, upon a percussive impact these expandable elements 25 are released in order to create an internal pressure within the cavity 25 which acts outwardly in order to relieve deformation and also stiffen the blade 20 as a result of the over pressure within the cavity 23.

Abstract: The control stage buckets for a steam turbine have airfoils which are hollowed out to form an interior cavity within the airfoil. The cavity opens through the radial inner platform of the bucket and terminates short of the tip of the bucket. The buckets are therefore of reduced weight. This reduced weight causes reduced creep damage in the axially extending dovetails along the rotor rim.

Abstract: An article such as a hollow gas turbine blade has an internal cavity therein with an inlet and outlet. The outlet has an outlet minimum transverse dimension of from about 0.009 inch to about 0.012 inch. An aluminiding source powder has a minimum particle size of greater than about 0.0015 inch and not greater than about 0.005 inch. The aluminiding source powder is a mixture of from about 5 to about 15 percent by weight of a metallic aluminum-containing powder and from about 85 to about 95 percent by weight of a ceramic powder. The aluminiding source powder is flowed into the internal cavity through the inlet, and the article is heated with the aluminiding source powder in the internal cavity to a temperature of from about 1750° F. to about 2000° F., and for a time of from about 2 hours to about 12 hours, to deposit an aluminum-containing coating on the internal surface of the internal cavity. The aluminiding source powder is thereafter removed from the internal cavity through the inlet.

Abstract: A component such as a blade for use in a gas turbine engine comprises a body 2 formed from outer panels 8, 10 defining an internal cavity 12. A damping element 14 is provided in the cavity 12 and is secured at one end between the panels 8, 10 at one end of the body 2, for example adjacent the root of the blade. Damping material fills the cavity 12 and so extends between the damping element 14 and the panels 8, 10. Vibration induced in the blade causes relative movement between the body 2 of the blade and the damping element 14, causing energy loss in the damping material which damps the vibration. The components 8, 10, 14 may be secured by a diffusion bonding process.

Abstract: A turbine blade for a turbine engine having at least one secondary flow deflector proximate to a blade tip for reducing the effective flow path between the blade tip and an adjacent outer seal. The turbine blade may be a superblade having a central opening forming a hollow turbine blade. The turbine blade may include a secondary flow deflector on upstream sides of the pressure side wall and the suction side wall. The downstream sides of the pressure and suction side walls may include chamfered corners. The secondary flow deflector reduces the effective flow path between the blade tip and an outer seal in numerous ways.

Abstract: A turbomachine blade includes a metal alloy airfoil having a cavity closed off by a cover on a hollowed side, the cover providing aerodynamic continuity of the hollowed side and being bonded via an edge thereof to the rest of the airfoil by a weld bead. The weld bead emerges on the hollowed side and penetrates into the airfoil to a depth P at least equal to the thickness EC of the edge of the cover so as to provide continuity of material between the edge of the cover and the rest of the airfoil over a depth at least equal to the thickness EC of the edge of the cover. A process for manufacturing such a blade is also disclosed.

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: A component, for example a fan blade for a gas turbine engine, comprises panels 2 and 4 which define between them a cavity containing a warren girder structure 6. Internal surfaces of the panels 2 and 4 are provided with a damping material 14, disposed between regions of contact 12 between the warren girder 6 and the panels 2, 4. The damping material damps vibrations of the component, so extending its fatigue life.

Abstract: The blade structure of a wind turbine which has enough strength for sporadic large load and increased durability. The blade can be formed into a thinner blade profile, with which blade performance is improved, increased efficiency of the wind turbine is attainable, and the noise induced by Karman vortex street is reduced. The blade is made of metal and formed into a one-piece blade or a split-type blade consisting of a main blade body made of metal and a rear end member made of metal other than that of the main blade body, the rear end member being fixed to the rear end of the main blade body, dentation being formed in the trailing edge part of the blade along the length thereof, the dentation being shaped in a serration of triangular teeth, of trapezoidal teeth, or of sawteeth, the ratio(h/&dgr;) of the tooth height h to the thickness &dgr; of the boundary layer developed on the surface of the blade is: h/&dgr;=1.0˜10.0, and the ratio (h/p) of the tooth height h to the pitch p is: h/p=0.

Abstract: A gas turbine engine (10) fan outlet guide vane (34) is manufactured by diffusion bonding titanium alloy sheet metal workpieces (52, 54) together to form an in integral structure (66). A titanium alloy is weld deposited at predetermined positions (70, 72) and in a predetermined shape to build up bosses (42, 44) at the radially outer end (40) of the hollow integral structure (68). The integral structure (66) is inflated at high temperature to form a hollow integral structure (68) of predetermined aerofoil shape. Apertures (46, 48) are drilled through the bosses (42, 44) to enable the radially outer end (40) of the fan outlet guide vane (34) to be attached to a fan casing (32) of the gas turbine engine (10).

Abstract: A rotor blade system with reduced blade-vortex interaction noise includes a plurality of tube members embedded in proximity to a tip of each rotor blade. The inlets of the tube members are arrayed at the leading edge of the blade slightly above the chord plane, while the outlets are arrayed at the blade tip face. Such a design rapidly diffuses the vorticity contained within the concentrated tip vortex because of enhanced flow mixing in the inner core, which prevents the development of a laminar core region.

Abstract: A turbine blade comprises an external pressure side plate comprising a plurality of external pressure ribs, an internal pressure side plate comprising a plurality of internal pressure ribs, and a plurality of pressure flow redirection areas. The external pressure side plate is disposed over the internal pressure side plate to form a pressure side double wall. The external pressure ribs are disposed at a first angle with respect to a blade span reference line and the internal pressure ribs are disposed at a second angle with respect to the blade span reference line to form the pressure flow redirection areas. The pressure side double wall is configured to pass a coolant through the pressure side double wall. A fabrication method aligns the external pressure side plate and the internal pressure side plate to form the pressure flow redirection areas between the external pressure side plate and the internal pressure side plate.

Abstract: A turbomachine blade comprising a root portion and an aerofoil portion, the aerofoil portion having a leading edge, a trailing edge, a concave metal wall portion extending from the leading edge to the trailing edge and a convex metal wall portion extending from the leading edge to the trailing edge. The concave metal wall portion and the convex metal wall portion form a continuous integral metal wall. The aerofoil portion has a hollow interior defined by at least one internal surface and the hollow interior of the aerofoil portion is at least partially filled with a vibration damping and stiffening system that comprises a vibration damping material. The vibration damping material is bonded to the at least one internal surface and the vibration damping and stiffening system comprises a variation of material properties between the wall portions.

Abstract: A hollow blade for a ceiling fan is made from a selected material and has a hollow interior structure to result in a reduced weight for decreasing motor loading and saving electric power consumption, and saving material consumption for reducing costs.

Abstract: A novel blade configuration does not exceed the permitted stresses for particular loads, especially as a result of centrifugal forces and which at the same time, allows the turbomachine to function with a high degree of efficiency. To this end, a moving blade for the turbomachine contains at least partially a cellular material, especially a foamed metal. The cellular material can be provided e.g. in the hollowed-out part of the moving blade.

Abstract: A damper that provides damping for turbine engines to reduce vibrations in the blades, vanes, shrouds, ducting, liner walls and/or other components of the turbine engine. The damper includes an air cavity in a blade, vane, shroud or other engine component and a material covering at least a portion of the air cavity. Vibrations in the turbine engine and its components produce movement of the air inside the air cavity resulting in a corresponding viscous force that dampens the vibrations.

Abstract: A runner for a hydraulic turbine or pump includes a hub and a plurality of blades extending from the hub at spaced intervals therearound. Each blade includes an inner edge secured to the hub and a distal outer edge, a leading edge and an opposed trailing edge, and a curved suction surface and an opposed curved pressure surface. At least one of the blades is fabricated from a curved pressure-side member and a curved suction-side member secured together. The pressure-side member includes at least a substantial portion of the curved pressure surface and an opposed first inner surface. The suction-side member includes at least a substantial portion of the curved suction surface and an opposed second inner surface. The first and second inner surfaces face each other, and at least a portion of the first inner surface is spaced from at least a portion of the second inner surface to form a cavity within the blade. A method of making the hollow blade is also disclosed.

Abstract: A structure includes a cooled article having an open-cell solid foam of ceramic or metal cell walls with a porous interconnected intracellular volume therebetween. A source of a pressurized gas is in communication with a source region of the cooled article. The source of the pressurized gas includes a gas plenum in gaseous communication with the source region, and a compressor having a compressed gas output in gaseous communication with the gas plenum. Gas flows from the source of the pressurized gas through the porous intracellular volume, to cool the cooled article.

Abstract: A welded hollow nozzle partition formed from sheet metal and having a concave side and a convex side, includes an interior plate welded to an inner surface of the partition. The interior plate serves to strengthen the outer walls of the partition to withstand wall buckling and ballooning under certain operating conditions.

Abstract: A gas turbine engine fan blade (26) comprises a root portion (40) and an aerofoil portion (42). The aerofoil portion (42) has a leading edge (44), a trailing edge (46), a concave metal wall portion (50) extending from the leading edge (44) to the trailing edge (46) and a convex metal wall portion (52) extending from the leading edge (44) to the trailing edge (46). The aerofoil portion (42) has a hollow interior (54) and the interior (54) of the aerofoil portion (42) is at least partially filled with a vibration damping material (56). The vibration damping material (56) comprises a material having viscoelasticity for example one formed by mixing an amine terminated polymer and bisphenol a-epichlorohydrin epoxy resin.

Abstract: A turbine airfoil includes at least one spar arrangement having a length less than an associated turbine airfoil length. During operation, the turbine airfoil has an outer skin surface which operates at a substantially higher temperature than that of an internal supporting parted spar arrangement. The parted spar arrangement permits the turbine airfoil outer skin surface to thermally expand between spar arrangements, thus preventing self-constraining thermal stresses from forming within the spar arrangement or the airfoil skin surfaces.