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 result is a blade with good protection from a wide spectrum of external threats.

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 one embodiment, contiguous cavities are formed around freestanding ends of the ribs to reduce the number of cavities required. The freestanding end of each rib is flared such that it has a larger width than the rest of the rib. The cavity extends continuously around the free, flared end of the rib.

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: A turbine nozzle segment includes a single hollow airfoil extending radially between radially outer and inner band segments. The airfoil has an airfoil wall with pressure and suction sides extending axially between leading and trailing edges of the airfoil. The airfoil wall surrounds a bifurcated cavity and a bifurcating rib extends through the bifurcated cavity and between the pressure and suction sides of the airfoil wall dividing the bifurcated cavity into forward and aft cavities. A stiffening rib extends radially outwardly from and along a radially outer surface of the outer band segment and axially and circumferentially from a pressure side forward corner of the outer band segment to the bifurcating rib. The outer and inner band segments, the bifurcating rib, and the stiffening rib are integral and made from a unitary one-piece casting.

Abstract: A device and a method are for producing a blade including two outer walls and at least one cavity between the outer walls, for a turbine. An outer mould and several cores are used in forming the outer walls and the at least one cavity of the blade. At least one of the cavities is divided into two channels by a middle segment. One channel is located between the first outer wall and the middle segment, while the other channel is located between the middle segment and the second outer wall. Two cores which are separate from each other are used accordingly. This provides a simple and economical means of reducing the thickness of the outer wall.

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 present invention provides a cooling apparatus for cooling a turbine blade. The cooling apparatus comprises a pressure side plate comprising a plurality of pressure ribs, a suction side plate comprising a plurality of suction ribs, and a plurality of flow redirection areas. The pressure side plate is disposed over the suction side plate, where the pressure ribs are disposed at a first angle with respect to a blade span reference line and the suction ribs are disposed at a second angle with respect to the blade span reference line, to form the flow redirection areas. The present invention provides a method of fabricating the cooling apparatus for the turbine blade comprising aligning the pressure side plate and the suction side plate to form the plurality of flow redirection areas between the pressure side plate and the suction side plate.

Abstract: The invention relates to a blade (13; 14) for a turbine (10), comprising at least one channel (22) which is delimited by walls (19, 20, 21). An insert (25) which can be subjected to the action of a liquid coolant is inserted into at least one channel (22). According to the invention, at least one of the walls (19; 20) is provided with a number of horizontal ribs (24) which are located between the insert (25) and the wall (19; 20). Said insert (25) is provided with openings (27) through which the liquid coolant passes out of the insert (25) and between the horizontal ribs (24). The liquid coolant is therefore conducted along the wall (19, 20) and guided by the horizontal ribs (24) in order to provide improved convection cooling.

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 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 wing of composite material has an upper skin, a lower skin and a plurality of intermediate spars arranged between the upper and the lower skin. The intermediate spars extend from roots to tips of the upper and the lower skins. The intermediate spars are integrally formed with the upper and the lower skins or are adhesively bonded to the upper and the lower skins. The wing has a small number of parts, needs greatly reduced time for fabrication and can be fabricated at a low cost.

Abstract: A grit blasting and alkaline etch surface pretreatment that is applied to a metallic airfoil. The metallic airfoil includes pockets or cavities that have been fabricated into the airfoil to reduce the weight of the airfoil. The pretreatment includes grit blasting the surface of the pockets or cavities, followed by washing the airfoil and treating the airfoil in an alkaline etch bath. After any remaining solution from the bath is neutralized and within twenty-four hours of the end of the pretreatment, primer is applied to the surface of the pockets or cavities. After the primer is applied to the surface of the pockets, a lightweight resin is injected into the pockets and bonds to the primer forming a mechanical bond with good fracture toughness at elevated temperatures.

Abstract: A wing of composite material has an upper skin, a lower skin and a plurality of intermediate spars arranged between the upper and the lower skin. The intermediate spars extend from roots to tips of the upper and the lower skins. The intermediate spars are integrally formed with the upper and the lower skins or are adhesively bonded to the upper and the lower skins. The wing has a small number of parts, needs greatly reduced time for fabrication and can be fabricated at a low cost.

Abstract: This centrifugal blower fan wheel has extruded-aluminum airfoil blades. The preferred blades have two hollow regions, separated by an angled mid-load support beam, inside the blade. The nose section of the blade (at the leading edge) and the tail section of the blade (at the trailing edge) have open semi-circular grooves which face the adjacent hollow sections and are designed to accept thread cutting or thread forming screws for attachment of the blade to a ring-shroud and back-plate.

Abstract: A tuning rib is added preferably in the aft cavity of a cored turbine bucket to alter the bucket's natural frequencies. The tuning rib may be a solid rib or a segmented rib and is particularly suited for altering high order frequency modes such as 2T, 4F and 1-3S. As such, detrimental crossings of natural bucket frequencies and gas turbine stimuli can be avoided to thereby improve the reliability of a gas turbine without impacting other features of the bucket that are important to the performance of the gas turbine.

Abstract: A welded hollow partition formed from sheet metal and having a concave side and a convex side, includes end caps to close off the partition cavity. The end caps are recessed into the hollow partition to provide wall strengthening within the hollow partition. Epoxy or other water resistant material can be used adjacent to the end cap and within the partition. A gasket can also be fitted adjacent to the end cap and within the partition to provide a water resistant seal. The hollow partition can be filled with temperature resistant material to further seal the partition against moisture.

Abstract: The present invention relates to a lightweight shrouded turbine blade. The turbine blade comprises an airfoil section and a hollow blade tip shroud joined to the airfoil section. The hollow tip shroud is preferably a cast, compartmentalized structure and has a plurality of ribs acting as load bearing structures and defining a plurality of shroud core sections. Each of the shroud core sections communicates with a supply of cooling fluid and has a plurality of apertures for supplying cooling fluid to exterior portions of the shroud.

Abstract: A thin film that forms an interface between metal surfaces and polymeric inserts of aircraft engine fan blades. The thin film is formed from a chemical comprised of carbon black, at least one evaporable solvent, a combination of phenolic-like resins dissolved in the solvent and the balance filler and inert ingredients. The thin film, upon drying of the evaporable solvent, forms a phenolic that bonds with the metallic portion of the fan blade. After the polymeric material is applied to recesses in the fan blade and cured, a chemical bond is formed between the thin film and the elastomeric inserts. The bonded structure has improved FWT peak stresses and improved fracture toughness.

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: The present invention relates to an improved attachment air inlet configuration for a turbine blade. It defines several modifications that reduce attachment rib concentrated stresses in highly loaded, single crystal turbine blades that cause cracking in conventional blade inlet configurations. It desensitizes rib concentrated stresses to variations in secondary crystal orientation, permitting highly loaded single crystal blades to have random secondary crystal orientation for lower cost, or seed with any secondary orientation needed to solve other blade or manufacturing problems.

Abstract: A tuning rib is added preferably in the aft cavity of a cored turbine bucket to alter the bucket's natural frequencies. The tuning rib may be a solid rib or a segmented rib and is particularly suited for altering high order frequency modes such as 2T, 4F and 1-3S. As such, detrimental crossings of natural bucket frequencies and gas turbine stimuli can be avoided to thereby improve the reliability of a gas turbine without impacting other features of the bucket that are important to the performance of the gas turbine.

Abstract: A blade for a ceiling fan includes at least one raised portion formed on the surface of the blade with a chamber in the raised portion of the blade. The raised portion will reduce the drag when the air passes over the blade because the blade has a smooth surface due to the raised portion. The recess in the raised portion provides flexibility to the blade of the present invention to lengthen the useful life of the blade for a ceiling fan.

Abstract: A nozzle vane or airfoil structure is provided in which the nozzle ribs are connected to the side walls of the vane or airfoil in such a way that the ribs provide the requisite mechanical support between the concave side and convex side of the airfoil but are not locked in the radial direction of the assembly, longitudinally of the airfoil. The ribs may be bi-cast onto a preformed airfoil side wall structure or fastened to the airfoil by an interlocking slide connection and/or welding. By attaching the nozzle ribs to the nozzle airfoil metal in such a way that allows play longitudinally of the airfoil, the temperature difference induced radial thermal stresses at the nozzle airfoil/rib joint area are reduced while maintaining proper mechanical support of the nozzle side walls.

Abstract: A fan blade includes a metal airfoil having a pocket disposed in a first side thereof, with the pocket including a filler bonded thereto. The pocket includes a plurality of cells separated by corresponding ribs which are submerged in the filler.

Abstract: A gas turbine engine fan blade (26) comprises a root portion (36) and an aerofoil portion (32). The aerofoil portion (32) has a convex surface (44), a concave surface (42), a leading edge (38) and a trailing edge (40). The leading edge (38) of the aerofoil portion (32) is formed from a harder material than the regions of the concave surface (42) and convex surface (44) immediately adjacent the leading edge (38) such that the leading edge (38) of the aerofoil portion (32) remains pointed. This improves the efficiency of the fan blade (26) and hence the gas turbine engine (10) and reduces flutter of the fan blade (10).

Abstract: A nozzle vane or airfoil structure is provided in which the nozzle ribs are connected to the side walls of the vane or airfoil in such a way that the ribs provide the requisite mechanical support between the concave side and convex side of the airfoil but are not locked in the radial direction of the assembly, longitudinally of the airfoil. The ribs may be bi-cast onto a preformed airfoil side wall structure or fastened to the airfoil by an interlocking slide connection and/or welding. By attaching the nozzle ribs to the nozzle airfoil metal in such a way that allows play longitudinally of the airfoil, the temperature difference induced radial thermal stresses at the nozzle airfoil/rib joint area are reduced while maintaining proper mechanical support of the nozzle side walls.

Abstract: An axial flow mixer impeller which is high in efficiency in terms of flow is made more effective, when used in a sparging system for dispersion and for mass transfer of a liquid phase or a gaseous phase into a liquid which is being mixed or agitated, by forming a shear field which breaks the phase being dispersed into fine bubbles which are dispersed by the impeller. This is accomplished without a major effect on the high flow efficiency of the axial flow impeller. To this end a structure which forms a Venturi is located on the side or sides of the blade where the phase (bubbles of fluid of different density or viscosity which are to be mixed or dispersed by the impeller) occurs, in the high velocity region near the tip of the blade. This Venturi creates a high shear field which breaks up the bubbles and disperses them as fine bubbles as they leave the impeller. The structure may be provided by a pair of proplets in the vicinity of the tip end of the blade which form a wedge shaped flow path therebetween.

Abstract: A formulation used in the construction of lightweight aircraft engine fan blades. The formulation comprises a polyurethane elastomer composition, which is formed from a prepolymer, a curative and an antioxidant and molded into the blades. Optionally, a hindered amine light stabilizer and/or an ultraviolet absorber may be added to the formulation. The aircraft engine fan blade is formed from a metal, such as titanium alloy. Pockets are machined into the fan blade, so as to lessen the amount of metal used to construct the blade, thereby lowering the total weight of the blade. The formulation of the present invention is positioned in the pockets of the fan blade, so as to maintain the structural integrity of the blade against bird strikes and the like, while at the same time providing a fan blade which is significantly lighter than all-metal fan blades.

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 method for preparing a Ni base superalloy inner wall surface of a body open end, such as an end of a turbomachinery blade, and an end plate, such as a blade tip cap, for brazing together at a rim of the end plate includes electrochemically removing oxides from the inner wall surface. The end plate is prepared, at least at its rim, by first removing surface and subsurface oxides, for example by mechanical abrading or a combination of such abrading and acid cleaning. Then at least the rim is electroplated with Ni which is heated to diffuse the Ni into the rim substrate. This provides an improved combination of surfaces for brazing, for example with a Ni base brazing alloy. After brazing the rim to the inner wall, there is provided an article with an improved relatively low oxide brazed joint, including less than about 20 volume % oxides, and a plate rim of substantially Ni along with elements diffused from the brazing alloy and the rim.

Abstract: A hollow airfoil (34) for a gas turbine engine having a leading edge (50), a trailing edge (52), a pressure side (54), and a suction side (56) includes a solid region of enhanced chordal thickness (78), and buttresses (82, 84, 86, 88). The present invention airfoil is preferentially strengthened so as to reduce the risk of airfoil fracture due to impact of airfoil with successive airfoils during an airfoil loss condition.

Abstract: A fan blade includes a metal airfoil having first and second opposite sides extending radially between a root and a tip, and axially between a leading edge and a trailing edge. The airfoil further includes a pocket disposed in the first side having an elastomeric filler bonded therein. A panel is bonded to the filler along the pocket for allowing differential movement therebetween for damping vibratory response of the airfoil.

Abstract: The present invention relates to a gas turbine cooled moving blade in order to cool the gas turbine using steam alone. In the moving blade, two cavities are installed in the lower part of the moving blade root section of the moving blade, a steam inlet is located in one cavity, and a steam outlet is installed in the other cavity. Steam is charged from the steam inlet and passes through a serpentine cooling passage consisting of a plurality of steam passages, and discharged to the steam outlet and collected. Bypasses are located near the base of the moving blade in the middle of each steam passage, so that cold steam on the leading edge side can be led into the bypass, mixed with the hot steam which comes from the tip section side and goes into the next downstream section of the steam passage, and flows to an upstream portion of the steam passage, thereby equalizing the steam temperature and cooling the steam over the entire range from the upstream to downstream of the steam flow.

Abstract: A fan blade includes a metal airfoil having first and second opposite sides extending radially between a root and tip, and axially between a leading edge and a trailing edge. The airfoil further includes a plurality of pockets disposed in the first side and separated by corresponding ribs. A filler is bonded in the pocket, and is coextensive with the airfoil first side. Radial and diagonal ribs respectively intersect solely each other for selectively increasing torsional and bending stiffness to increase frequency margin between adjacent torsional and bending resonant modes of vibration.

Abstract: A fan blade includes a metal airfoil having first and second opposite sides extending radially between a root and tip, and axially between a leading edge and a trailing edge. The airfoil further includes a plurality of pockets disposed in the first side and separated by corresponding ribs. A filler is bonded in the pocket, and is coextensive with the airfoil first side. Radial and diagonal ribs respectively intersect solely each other for selectively increasing torsional and bending stiffness to increase frequency margin between adjacent torsional and bending resonant modes of vibration.

Abstract: A composite tip cap assembly for a helicopter main rotor blade includes a composite main fairing, a unitary composite rib detail internally integrated in combination with the composite main fairing, a composite closure fairing bonded in combination with the composite main fairing, and an abrasion strip bonded in combination with the composite main fairing and the composite closure fairing.

Abstract: A blade for a rotor, especially a turborotor, is made of two hollow blade foot sections having matching seam surfaces to form a hollow blade with a thickened foot that is butt-welded with its blade foot sole (7) to an integral welding stub of a rotor hub or disk. The two hollow blade foot sections are soldered to each other along a radially outer seam section (5) between the seam surfaces and welded to each other along a radially inner seam section (6) between the seam surface. The radially inner seam section extends from said blade foot sole. The welded seam normally ends in the thickened blade foot section (6) and is thus shorter than the soldered seam section (5).

Abstract: The structure and process for making a hollow article in which the core structure is formed from two cell sheets diffusion bonded to a rib pattern sheet having a preselected pattern of ribs having passageways. The cell sheets are diffusion bonded to the rib pattern sheet using tooling pattern sheets having a geometry matching that of the rib pattern sheet thereby creating contiguous cells, interconnected by the passageways, between the cell sheets at each rib and the periphery of the structure. Inert gas is then applied to the cells during a superplastic forming process to create expanded envelopes of each cell of the core. Face sheets of the sandwich structure may then be formed from the cell sheets themselves by further expanding the cell sheets out against a mold surface. Alternatively, the envelopes of the core may be expanded and simultaneously or subsequently diffusion bonded to separate face sheets.

Abstract: A hollow airfoil (10) for a gas turbine engine having a leading edge (12), a trailing edge (14), a pressure side (20), and a suction side (22) includes a plurality of internal spanwise stiffening ribs (31-35) that are arranged in a logarithmic pattern. The particular arrangement of internal fibs (31-35) optimizes stiffness of the airfoil (10) without significantly increasing the weight thereof.

Abstract: An aerofoil (10) for a turbocharger stator for a large diesel engine has a convex, upper wall (12) and a concave, lower wall (14) with a leading edge (16) and a trailing edge (18). A wiggle strip (22) extends the length of the aerofoil and is brazed to the walls to provide added strength. The inner surface of the wall (12) is reinforced by a plate (20) extending from the trailing edge to a point in the region of the maximum spacing between the walls (12, 14). The plate is a shaped reinforcement piece and gives added strength to the area of the wall (12) which is impacted by cleaning medium. It provides a localized reinforcement to the area which is prone to damage by cleaning.

Abstract: An airfoil includes preformed first and second sides joinable together at a bond line extending between first and second opposite edges thereof. The first side includes a first extension having a proximal end forming a first end of the bond line, an intermediate section, and a distal end. The second side includes a second extension extending from the airfoil first edge which is fixedly joinable to the distal end of the first extension. The intermediate section of the first extension is spaced from the airfoil first edge to define therewith a working void so that pressure loading on the intermediate section is carried in part through the bond line first end for diffusion bonding thereof.

Abstract: An aerodynamic structure having an aerodynamic surface for an aircraft, has at least two spars, several ribs made up of rib elements, and two shells made up of a plurality of shell elements. The ribs extend in the direction of the profile depth of the structure and the longitudinal extension of the spars and shells corresponds approximately to the span width of the structure. The shells and ribs are subdivided in the direction of the profile depth with an approximately equal division, to form at least two structural components open along one length side and closed by a spar along the opposite length side. Each component (21, 22) has two shell elements (24, 25; 34, 35, etc.

Abstract: A pultruded composite blade for a Darrieus-type vertical axis wind turbine is disclosed, along with a method for fabricating such a blade. The blade is a composite structure with a uniform cross-section with reinforcing fibers at least some of which extend parallel to a lengthwise central axis and run continuously from end to end. The composite blade is pultruded straight and, when installed on the tower, the blade is elastically bent into a curved shape that approximates a troposkein. The composite structure includes a skin composed of multiple layers of resin-coated glass fabric that define an airfoil shape. The composite structure also includes one or more internal ribs or spars composed of multiple layers of resin-coated glass fabric.

Abstract: Horizontal axis, free yaw, self-regulated wind turbines which have strong, lightweight, fatigue resistant, fixed pitch, wood/GRE blades and exhibit superior performance a range of wind speeds. The blades are designed by employing defined inboard, midspan, and outboard airfoil profiles and interpolating the profiles between the defined profiles and from the latter to the root and the tip of the blades.

Abstract: Hollow airfoils, such as fan blades, nozzles and struts used in axial flow gas turbine engines, that have improved resistance to impact from foreign objects are disclosed. Relocation of airfoil material is used to preferentially strengthen the airfoil to respond to the stress from the impact of a foreign object. Internal spacers are redistributed toward the leading edge and the material from the skin of the concave side and the convex side is shifted from one side to the other and toward the leading edge of the airfoil where impact stress is greatest.

Abstract: An expandable spar filler block device for use in combination with a spar member of an articulated main rotor assembly of a helicopter. A preferred embodiment of the device includes first and second filler members and an expansion mechanism. Each filler member includes a spar engagement surface that is complementary to the inner mold line (IML) surfaces of the spar member and first and second tapered surfaces. The expansion mechanism for the preferred embodiment of the expandable spar filler block device is a pull shim that includes a first segment and a constant tapered segment having tapered surfaces that are complementary to the first tapered surfaces of the filler members.

Abstract: A hollow blade for a turbomachine includes a unitary body having a multiplicity of transverse cavities in at least the aerofoil shaped portion of the blade, and plugs disposed within the cavities for restoring the surface continuity of the intrados and extrados faces of the aerofoil shaped portion, the plugs being rigidly secured to the unitary body, such as by welding.

Abstract: A hollow blade for a turbomachine comprises a unitary body having a multiplicity of transverse cavities in at least the airfoil shaped portion of the blade, and plugs disposed within the cavities and restoring the surface continuity of the intrados and extrados faces of the aerofoil shaped portion, the plugs being rigidly secured to the unitary body, such as by welding.

Abstract: A coolable airfoil for use in gas turbine engine component such as a turbine blade or vane is provided with an integrally formed double shell outer wall surrounding at least one radially extending cavity. The inner and the outer shells are integrally formed of the same material together with tying elements which space apart the shells and mechanically and thermally tie the shells together. The present invention contemplates tying elements including pedestals, rods, and/or continuous or intermittent ribs. Impingement cooling means for the outer shell, in the form of impingement cooling holes, is provided on the inner shell to direct the coolant in impingement jet arrays against the outer shell, thereby, cooling the outer shell.

Abstract: A hollow rotor blade of a gas turbine includes a shaped core accommodated within a hollow space formed between two skins and joined to the skins. The fatigue strength at a portion where the skins are joined to the core is enhanced. More specifically, the portions of the skins joined with the core are in the form of island-shaped protrusions, and the core is joined to these protrusions by solid-phase diffusion bonding, diffusion-brazing, liquid-phase diffusion bonding, and the like.