Abstract: A turbine blade for use in a gas turbine engine, where the turbine blade is made from a spar and shell construction in which a thin walled shell is held in place between the blade tip and the platform by only a mechanical fastener without using any bonding, welding or brazing. The spar is connected to an attachment by a tie bolt, and a separate platform is secured, over the attachment t in which the shell is held against. This provides for a thermally free platform and shell connection. With this arrangement, the shell is held under compression within the 2% yield stress range such that the turbine blade shell can have an infinite life. Also, the turbine blade is much lighter than prior art blades. As a result, the spar and shell blade produces less stress on the rotor disk during engine operation.

Abstract: A reinforced airfoil includes an airfoil body including opposed walls that define a hollow interior space and a reinforcement member provided on at least one of the walls within the interior space, the reinforcement member increasing the thickness of the at least one wall so as to resist deformation of the at least one wall but not extending from one wall to the other.

Abstract: A rotor blade for a wind turbine includes a plurality of individual blade segments, with each blade segment defining an internal passage extending between the longitudinal ends of the blade segment. A rigid spar member extends longitudinally through the internal passages of the individual blade segments such that the blade segments are aligned and connected end-to-end on said spar member to define a complete rotor blade from a root to a blade tip. The spar member has a cross-sectional profile that is keyed to the cross-sectional profile of the internal passage in the blade segments. The spar member includes oppositely facing spar caps that engage against the inside surfaces of the blade segments within the internal passages.

Abstract: A hollow turbine blade is provided including an airfoil profile which is formed by a suction-side profile wall and a pressure-side profile wall and around which a hot gas can flow and which has a profile height, directed along a blade axis, from a platform up to a profile tip, having at least one supporting rib which is provided in the interior of the turbine blade and connects the pressure-side profile wall to the suction-side profile wall in a respective connecting region, and having at least one slot provided in the profile wall on the hot-gas side and extending along the blade axis. The slot of the turbine blade, on the hot-gas side in the profile wall, is opposite the connecting region formed by the supporting rib and the profile wall.

Abstract: A blade of a wind turbine is provided having an improved spar cap. The spar cap includes at least one trench extending in a substantially span-wise direction in at least a portion of the spar cap. At least one shear web is connected to the spar cap. At least a portion of the shear web is positioned within the trench of the spar cap.

Abstract: A turbine blade with an aerodynamically profiled blade leaf which is produced from two assembled shell elements is provided. The shell elements forming the suction-side and the pressure-side blade leaf wall have been assembled by a high-temperature and high-pressure bonding process. In the region of the blade leaf tip, a cramp is provided which hooks the two blade walls together with one another with a form fit.

Abstract: A turbine airfoil usable in a turbine engine with a cooling system and a compliant dual wall configuration configured to enable thermal expansion between inner and outer layers while eliminating stress formation in the outer layer is disclosed. The compliant dual wall configuration may be formed a dual wall formed from inner and outer layers separated by a support structure. The outer layer may be a compliant layer configured such that the outer layer may thermally expand and thereby reduce the stress within the outer layer. The outer layer may be formed from a nonplanar surface configured to thermally expand. In another embodiment, the outer layer may be planar and include a plurality of slots enabling unrestricted thermal expansion in a direction aligned with the outer layer.

Abstract: A hybrid fan blade for a gas turbine engine is provided that includes an airfoil and a composite panel. The airfoil has a first side and a second side orientated opposite the first side. The first and second sides extend between a tip, a base, a leading edge and a trailing edge. The airfoil includes a plurality of cavities disposed in the first side of the airfoil, which cavities extend inwardly toward the second side. The cavities collectively form an opening. At least one rib is disposed between the cavities. A shelf is disposed around the opening. The composite panel is attached to the shelf first mounting surface and to the rib, and is sized to enclose the opening. The first composite panel is a load bearing structure operable to transfer loads to the airfoil and receive loads from the airfoil.

Abstract: A guide vane for a gas turbine aircraft engine includes an aerodynamic shell for turning a flow of working fluid and an internal spar spaced from the aerodynamic shell by an air gap and reinforced by stiffeners.

Abstract: A damping element for a wind turbine rotor blade is provided, the damping element comprising a laminate material made of at least one viscoelastic layer and at least one stiff layer adhered to said viscoelastic layer, wherein the damping element is adapted to be attached to a body of the rotor blade so that the at least one viscoelastic layer is in contact with the body of the rotor blade.

Abstract: A hollow airfoil and a method for manufacturing a hollow airfoil is provided.

Abstract: An aerofoil blade for a gas turbine engine, the blade having a metallic core defining a number of cavities that contain a foamed material. The cavities are shaped such that the force of an impact on one surface of the blade is dissipated through the foam and transmitted to the metallic core. A fibrous internal containment device allows the blade to fragment progressively thereby spreading the load imparted by the blade to a casing should the blade fragment upon impact.

Abstract: Aspects of the invention relate to a construction system and method for components in high temperature environments, such as the hot gas path components of a turbine engine. Such a component can include a skeleton and a coating. The skeleton can be formed by a plurality of interconnected frame members, which can give the component its general shape. The frame members can be made of ceramic matrix composite. A coating can be provided around at least a portion of the skeleton. Preferably, the coating is a refractory material, such as refractory ceramic. Examples of turbine engine components that can be constructed according to aspects of the invention are airfoils with or without platforms, blade rings, combustor tiles and heat shields. A component according to aspects of the invention can be made using low cost fabrication and construction methods.

Abstract: The present invention is a vane for us in a gas turbine engine, in which the vane is made of an exotic, high temperature material that is difficult to machine or cast. The vane includes a shell made from Tungsten, and is formed from a wire electric discharge process. The shell is positioned in grooves between the outer and inner shrouds, and includes a central passageway within the spar, and forms a cooling fluid passageway between the spar and the shell. Both the spar and the shell include cooling holes to carry cooling fluid from the central passageway to an outer surface of the vane for cooling. This cooling path eliminates a serpentine pathway, and therefore requires less pressure and less amounts of cooling fluid to cool the vane.

Abstract: A multi-segment wind turbine blade comprises at least two blade segments. A first spar cap segment is attached to a first blade segment and a second spar cap segment is attached to a second blade segment. The first and second spar cap segments are configured to form a scarf joint. First and second spar cap brackets are attached in locations of the first and second spar cap segments, respectively, selected to facilitate alignment of the first and second spar cap segments at the scarf joint. At a field site, the first and second spar cap segments are bonded after fastening the first and second spar cap brackets.

Abstract: A blade for the use in the generation of power has a leading edge and a trailing edge. A first shell portion of the blade extends from the leading edge to the trailing edge. A second shell portion of the blade also extends from the leading edge to the trailing edge. A root portion of the blade is positioned proximate the wind turbine and a tip portion which extends from the root portion away from the wind turbine. A trailing edge insert is positioned between the first shell portion and the second shell portion proximate the root portion. The trailing edge insert has a surface which extends between and separates the first shell portion from the second shell portion at the trailing edge, such that the insert provides a high lift profile for increased blade efficiency.

Abstract: The invention relates to a rotor blade (22) for a wind energy installation (10), with the rotor blade (22) having at least one cavity (28). The cavity (28) is sealed to be gas-tight and is filled with gas, in particular air, in order to prestress the rotor blade (22), with the pressure of the gas being greater than or less than the average air pressure to be expected at the point of use of the rotor blade.

Abstract: The invention introduces a reinforcement of a box girder of a wind turbine blade. The reinforcement prevents the transverse shear distortion of the blade structure, when the blade is loaded during operation. The reinforcement connects the corners diagonally opposite inside the girder, and fixes them in relation to each other. The reinforcement increases the blade's resistance to overall collapse. The reinforcement comprises one or more individual element, such as rods or plates.

Abstract: The present invention is a vane for us in a gas turbine engine, in which the vane is made of an exotic, high temperature material that is difficult to machine or cast. The vane includes a shell made from either Molybdenum, Niobium, alloys of Molybdenum or Niobium (Columbium), Oxide Ceramic Matrix Composite (CMC), or SiC—SiC ceramic matrix composite, and is formed from a wire electric discharge process. The shell is positioned in grooves between the outer and inner shrouds, and includes a central passageway within the spar, and forms a cooling fluid passageway between the spar and the shell. Both the spar and the shell include cooling holes to carry cooling fluid from the central passageway to an outer surface of the vane for cooling. This cooling path eliminates a serpentine pathway, and therefore requires less pressure and less amounts of cooling fluid to cool the vane.

Abstract: The present invention relates to a reinforced blade for a wind turbine, and in particular to a wind turbine blade comprising a shell having a section with an aerodynamic profile, and at least one internal reinforcing floor connected inside the shell and extending substantially along the profile chord in order to increase the strength of the blade and to prevent or reduce deformations of the surface of the blade caused by edgewise and flapwise loading of the blade structure.

Abstract: A turbine blade includes an airfoil tip with first and second ribs extending along the opposite pressure and suction sides. The ribs extend outwardly from a tip floor and are joined together at opposite leading and trailing edges. A cascade tip baffle transversely bridges the two ribs above the tip floor forward of the maximum width of the tip to partition the tip chordally into corresponding tip pockets on opposite sides of the baffle.

Abstract: A variable geometry inlet guide vane for a gas turbine aircraft engine includes an aerodynamic shell for turning inlet flow to a turbine or compressor and an internal spar spaced from the airfoil shell by an air gap. A number of actuation mechanisms grounded to the spar and connected to the inner surface of the aerodynamic shell adjust the shape of the shell in response to varying operating conditions of the engine, imbalanced aerodynamic loading of the shell or vibration or other transient loads on the shell.

Abstract: A noise control cassette for a gas turbine engine includes a perforated face sheet configured for exposure to an airflow, a non-perforated backing sheet, a core arranged between the face sheet and the backing sheet and defining a cavity between the face sheet and the backing sheet having an effective length tuned so as to provide acoustic reactance control, and an attachment face for attaching the cassette to an airfoil-shaped structure.

Abstract: A CMC airfoil (20) formed with CMC stitches (37) interconnected between opposed walls (26, 28) of the airfoil to restrain outward flexing of the walls resulting from pressurized cooling air within the airfoil. The airfoil may be formed of a ceramic fabric infused with a ceramic matrix and dried, and may be partially to fully cured. Then holes (32, 34) are formed in the opposed walls of the airfoil, and a ceramic stitching element such as ceramic fibers (36) or a ceramic tube (44) is threaded through the holes. The stitching element is infused with a wet ceramic matrix before or after threading, and is flared (38) or otherwise anchored to the walls (26, 28) to form a stitch (37) there between. The airfoil and stitch are then cured. If the airfoil is cured before stitching, a pre-tension is formed in the stitch due to relative curing shrinkage.

Abstract: A method for making a gas turbine engine turbine blade comprising an airfoil section a platform section, an under platform section, and a dovetail section, the exterior surface of the dovetail section comprising a shank exterior surface and a serrated exterior surface. The blade further includes a silicon-modified diffusion aluminide layer a surface of a turbine blade section selected from the group consisting of the exterior surface of the under platform section, the exterior surface of the dovetail section, and combinations thereof, the aluminide layer having a concentration of silicon at a surface of the aluminide layer in the range of about 1 weight percent to about 10 weight percent and a concentration of aluminum at the surface of the aluminide layer in the range of about 5 weight percent to about 25 weight percent.

Abstract: A balance system for a rotational airfoil assembly includes a bulkhead, an elongated member, a support member and at least one balance weight. The elongated member extends through the bulkhead along a bore toward an airfoil tip segment of the airfoil assembly. The balance weights are mounted to a tip segment of the elongated member positioned within the bore outboard of an airfoil mid-span segment of the airfoil assembly. A reduced quantity of weight is required to balance the airfoil assembly because of the location of the balance weights outside of the mid-span segment of the airfoil assembly.

Abstract: A lightweight wind turbine blade formed with a truss support structure assembly of composite truss joints including composite spar and cross members attached to and supporting in spaced relation a spine of lightweight rib panels. The rib panels are oriented in parallel spaced relation from one another and individually molded with perimeters defining individual areas of curvature for the finished blade assembly. The truss support structure is covered with a lightweight fiberglass or hardened fabric skin attached to and fitted on respective rib panel edges forming an airfoil structure.

Abstract: A turbine airfoil usable in a turbine engine and having at least one cooling system. At least a portion of the cooling system may be positioned in an outer wall of the turbine airfoil and be formed from at least one suction side serpentine cooling chamber and at least one pressure side serpentine cooling chamber. Each of the suction and pressure side serpentine cooling channels may receive cooling fluids from a cooling fluid supply source first before being passed through other components of the cooling system. The cooling fluids may then be passed into a mid-chord cooling chamber to cool internal aspects of the turbine airfoil, yet prevent creation of a large temperature gradient between outer surfaces of the turbine airfoil and inner aspects.

Abstract: An apparatus and method for impingement cooling of a turbine vane, in which the vane includes an insert having a plurality of impingement cavities formed therein and in series such that cooling air flows from a first impingement cavity onto the wall for impingement cooling, and is then directed into the second impingement cavity and redirected for impingement cooling on another part of the wall. Supports for the insert form seals that direct cooling air from one impingement cavity into the next impingement cavity in the series. A trailing edge impingement cavity directs cooling air through holes to provide impingement cooling to the trailing edge region, the cooling air passing through a trailing edge discharge passage to cool the trailing edge. The insert is formed as a single piece, and has from 3 to 5 impingement cavities separated by ribs.

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: Hollow fan blades for turbo fan gas turbine engines are formed of two separate detail halves. Each detail half has a plurality of cavities machined out to reduce weight and form ribs. The ribs extend continuously and without intersecting one another in order to facilitate machining. The ribs extend in different directions in different regions of the detail half. In at least one region, the ribs extend in a serpentine pattern, such that concatenated rib segments are substantially perpendicular to one another, thereby providing stiffness in multiple directions.

Abstract: The invention refers to a coated turbine blade for a gas turbine, having blade walls divided into sections with locally adapted material temperatures. The cooler material temperatures of the blade walls are at the points where the support ribs merge into the blade walls. The regions with higher material temperatures of the blade walls are at the positions where cavities are arranged inside the blade walls. This is achieved via a ceramic thermal barrier coating having different layer thicknesses that allow different material temperatures. The region of the surface of the blade wall which faces the working medium and lies opposite an internal rib has a thicker thermal barrier coating than a region of the surface of the blade wall which is cooled via a cooling medium that flows in the cavities. This results in homogenization of the material temperature in the connecting regions, which results in prolonged blade life.

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 fluid turbine, such as a wind or hydro turbine, for generating electrical power, the turbine having a rotor assembly comprising a hub, a rim and a plurality of tensioned cable members joining them together, the cable members being grouped into sets to define a plurality of supports for individual skin members to create a plurality of lightweight blade members.

Abstract: A variable geometry inlet guide vane for a gas turbine aircraft engine includes an aerodynamic shell for turning inlet flow to a turbine or compressor and an internal spar spaced from the airfoil shell by an air gap. A number of actuation mechanisms grounded to the spar and connected to the inner surface of the aerodynamic shell adjust the shape of the shell in response to varying operating conditions of the engine, imbalanced aerodynamic loading of the shell or vibration or other transient loads on the shell.

Abstract: Low-noise fan exit guide vanes are disclosed. According to the present invention a fan exit guide vane has an outer shell substantially shaped as an airfoil and defining an interior cavity. A porous portion of the outer shell allows communication between the fluctuations in the air passing over the guide vane and the interior cavity. At least one acoustically resonant chamber is located within the interior cavity. The resonant chamber is in communication with the porous portion of the outer perimeter. The resonant chamber is configured to reduce the noise generated at a predetermined frequency. In various preferred embodiments, there is a plurality of acoustically resonant chambers located within the interior cavity. The resonant chambers can be separated by one or more partitions within the interior cavity. In these embodiments, the resonant chambers can be configured to reduce the noise generated over a range of predetermined frequencies.

Abstract: A CMC airfoil (20) formed with CMC stitches (37) interconnected between opposed walls (26, 28) of the airfoil to restrain outward flexing of the walls resulting from pressurized cooling air within the airfoil. The airfoil may be formed of a ceramic fabric infused with a ceramic matrix and dried, and may be partially to fully cured. Then holes (32, 34) are formed in the opposed walls of the airfoil, and a ceramic stitching element such as ceramic fibers (36) or a ceramic tube (44) is threaded through the holes. The stitching element is infused with a wet ceramic matrix before or after threading, and is flared (38) or otherwise anchored to the walls (26, 28) to form a stitch (37) there between. The airfoil and stitch are then cured. If the airfoil is cured before stitching, a pre-tension is formed in the stitch due to relative curing shrinkage.

Abstract: A turbine airfoil usable in a turbine engine and having at least one cooling system. The cooling system may be positioned in an outer wall of the turbine airfoil, and the airfoil may include a hot gas receiving cavity positioned in a mid-chord region of the airfoil. The hot gas receiving cavity may have an opening in a tip of the airfoil to enable hot gases to circulate into the hot gas receiving cavity. In at least one embodiment, the cooling system in the outer wall and the hot gas receiving cavity may include a plurality of ribs. Cooling fluids may be passed through the cooling system in the outer wall, and hot combustion gases may be passed into the hot gas receiving cavity to moderate the temperature of the inner portions of the outer wall to reduce the temperature gradient in the outer wall.

Abstract: The present disclosure is directed to a power system component for a work machine. The power system component may include a wheel having one or more free-form welded members. The wheel may include a front end, a back end, and a central, axial spine member extending from the front end to the back end. The wheel may also include an outer shell adjacent the spine member at the front end, extending toward the back end, and flaring radially away from the spine member near the back end. The outer shell may include a front surface, a back surface, and one or more protrusions extending from the front surface. The wheel may further include at least one void defined between the spine member and the back surface of the outer shell and at least one reinforcing structure connecting the outer shell with a portion of the spine member between the front end and the back end.

Abstract: An internally-cooled airfoil for a gas turbine engine, wherein the airfoil comprises at least one substantially chordwise-extending and internally-projecting stiffener located on an internal surface of one of the sidewalls and adjacent to the insert.

Abstract: An exit guide vane array for a turbine engine includes a set of guide vanes 28 having a solidity and defining fluid flow passages 74 with a chordwisely converging forward portion 80. The high solidity and convergent passage portion 80 resist fluid separation. The vanes may also cooperate with each other to restrict an observer's line of sight to planes upstream of the vane array.

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: The invention relates to a blade for use on a wind turbine and a method of assembling laminated profiles (3, 5) for a blade for a wind turbine. The development towards increasingly larger blades requires, that the technology of manufacture is reviewed and novel methods applied, in particular with regard to aspects regarding strength and weight. It is an object of the invention to provide a blade that can be manufactured more accurately and with large uniformity with regard to strength, from blade to blade, and which is lighter than prior art blades.

Abstract: In accordance with certain embodiments, a wind turbine blade includes an outer skin and an internal support structure. The skin has a closed airfoil shape and includes an outer surface, an inner surface, and opposite sides along a length of the blade. The internal support structure is disposed within the closed airfoil shape. The internal support structure includes at least one shear load bearing member disposed across the opposite sides and oriented lengthwise along the length of the blade. The internal support structure further includes a plurality of bending load bearing members disposed along the inner surface in a lengthwise direction along the length of the blade. The internal support structure is configured to position a shear center and a center of aerodynamic pressure loading of the blade at different locations relative to one another.

Abstract: A propeller or propeller blade is manufactured using lattice block material to provide a structure which is generally hollow but for a three-dimensional lattice of support spars. The propeller or propeller blade, being predominately hollow, is substantially lighter than a solid cast propeller or propeller blade while retaining the desired strength due to the three-dimensional lattice of support spars.

Abstract: A vane assembly 10 suitable for a turbine engine features a refractory vane 12 with an internal cavity 20 and a pair of flexible metallic baffles 26 extending into the cavity from spanwisely opposite ends of the vane. A rigid fastener 48, such as a nut and bolt assembly applies a tensile load to the baffles. The tensile load is reacted out as a compressive load applied to the vane. In another embodiment, the baffle is relatively rigid but the fastener is flexible. The compressive loading exerted on the vane counteracts the brittleness customarily exhibited by refractory materials and imparts damage tolerance to the vane. The arrangement also allows the use of a metal baffle that can be easily secured to the vane and dispenses with any need for a potentially troublesome seal between the baffles and the spanwise extremities of the vane.

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. These detail halves are subsequently bonded and given an airfoil shape in the forming operation. In the present invention, the ribs are oriented biased to provide stiffness as needed in different sections of the fan blade with smooth transitions between regions. The ribs and cavities do not run solely radially or chordwise, but curve and change direction. The path of each cavity is curved sufficiently to eliminate any regions with long, straight cavities, which would have low inertia in cross sections through the cavity parallel to the rib. Preferably, cavities do not continue straight in any direction for lengths greater than half the blade chord (e.g. leading edge to trailing edge) before curvature causes the rib to intersect cross section.

Abstract: A hollow turbine blade cooled with compressor air is divided into a cooling air chamber (6) and into impingement air cooling chambers (8, 9) by inner, supporting partitions (3, 4). The cooling air is conveyed from the cooling air chamber into the impingement air cooling chamber via impingement air channels (7) provided in the partitions. The impingement air channels are concave with regard to the adjacent outer wall (2) of the blade airfoil (1) and arranged completely in the hot area near the outer wall and, in addition, have an oblong or elliptical cross-section whose longitudinal axis agrees with the radial orientation of the turbine blade. By reduced stress concentration in the area of the impingement air channels, the fatigue and creep characteristics are improved and life is increased.

Abstract: In accordance with the present invention, an aircraft engine is provided with a component which helps reduce the noise generated by the engine. The component has a first aerodynamic surface, a second aerodynamic surface, and a system for reducing noise without altering a pressure differential between the first aerodynamic surface and the second aerodynamic surface.

Abstract: A gas turbine engine turbine blade includes a hollow airfoil extending radially from a blade root to a radially outer airfoil tip. The airfoil includes an airfoil outer wall having transversely spaced apart pressure and suction side walls meeting along chordally spaced apart leading and trailing edges of the airfoil. A radially extending cooling air supply channel within the airfoil includes a bank of pins integral with and extending transversely between the pressure and suction side walls. The bank of the pins is tuned such that a natural frequency of the blade associated with an engine forced driving mode of the blade is sufficiently away from a steady state engine operating frequency to substantially avoid natural frequency resonance of the blade during steady state engine operation. The bank of the pins is tuned by locations of the pins within the cooling air supply channel.