Abstract: A turbine blade is provided for use in a gas turbine engine. The turbine blade has a platform, an airfoil radially extending from the platform, and an attachment portion comprising an asymmetric root neck portion having a higher stress side and a lower stress side. The turbine blade may further have additional material and a compound fillet for dispersing strain in a region where the airfoil overhangs the neck portion.

Abstract: A turbine bucket includes a shank portion, an adjacent and radially inner dovetail mounting portion, an adjacent and radial outer airfoil portion, with a platform at a radially inner end of said airfoil portion adjacent the shank portion, and an integral cover at a radially outer tip of the airfoil portion, wherein the shank portion is shaped to provide a torque zone establishing a desired torque characteristic for the bucket to create, upon insertion into a dovetail groove on a turbine wheel, a desired contact pressure between the integral cover and adjacent covers in a row of similar buckets.

Abstract: An apparatus for retaining blade to a rotor disc of turbine engines includes a positioning member integrally for stopping the root of the blade when sliding into an attachment slot of the rotor disc, and a locking apparatus for retaining the root of the blade in the attachment slots.

Abstract: An arrangement is disclosed for the admission of cooling air to the internal walls of a component rotating about a rotation axis, such as a moving blade in a rotary machine. A component root can be fastened to a rotor unit in a rotationally fixed manner and adjoining in a radially extending manner is a one piece component airfoil in which at least one radially extending cooling passage region (K1) is provided which, in the region of the component root, opens out via an opening into a cooling-air supply passage passing at least partly through the component root longitudinally relative to the rotation axis. A distribution plate forms a fluid-tight connection with an opening margin, surrounding the opening of the cooling passage region (K1), at least during the rotation of the component about the rotation axis.

Abstract: A gas turbine engine blade having a dovetail with an enhanced surface for durability is disclosed. The blade includes an airfoil section and a dovetail section disposed opposite the airfoil section. The dovetail section has an outer surface, at least a portion of which has residual compressive stresses resulting from a plurality of peening operations. The dovetail section also includes a copper-nickel-indium wear coating layer disposed over the outer surface having residual compressive stresses and a lubricant coating layer disposed over the copper-nickel-indium wear coating layer, wherein the lubricant coating layer is lead-free.

Abstract: A rotor (20) for a compressor, in particular in a gas turbine, has a number of rotor blades (25) which are arranged around the rotation axis of the rotor (20) in the form of a rim and are each held in a circumferential recess (21) on the rotor (20) by a blade root (26), with the blade root (26) having a widening lower part (27) which engages behind two shoulders (24) that are formed on the side walls of the recess (21). In such a rotor, the life is lengthened in that the recess depth (T) of the recess (21) is substantially greater than a minimum recess depth (Tmin) which results in the rotor (20) having sufficient strength in the area of the blade attachment for starting, based on the predetermined material characteristics of the rotor (20) and the operating conditions of the compressor.

Abstract: The invention relates to a wind turbine blade comprising one or more turbulence generating strips, where the strips are placed on a surface of the blade. The blade is characterized in that at least one joint area of the turbulence generating strips and the surface of the blade are completely or partially covered by sealing means. The invention further relates to a pitch controlled wind turbine comprising at least two pitch controlled wind turbine blades and pitch controlling means for pitching the blades. The pitch controlled wind turbine is characterized in that the blades comprises one or more turbulence generating strips, wherein at least one joint area of the turbulence generating strips and a surface of the blades are completely or partially covered by sealing means.

Abstract: An identification feature is used to unmistakably identify internal features present in different generations of turbine blade designs. The identification feature is located on a root portion of the turbine blade and protrudes to provide a visually identifiable feature that is also readable by a coordinate measuring machine, but does not interfere with installation or operation of the turbine blade. The weight of the identification feature is in a specific proportion to the weight of the turbine blade in order to prevent interfere with operation of the turbine blade during high-speed rotation in a gas turbine engine.

Abstract: A radial-flow turbine wheel is provided. The radial-flow turbine wheel includes a hub having an outer radius gradually increasing from a front end to a rear end, a rear periphery of the hub being radially extended in a plane generally perpendicular to a center axis, and a plurality of turbine blades formed around the hub at constant intervals. A plurality of slots is formed by inward cuts at the rear periphery of the hub between the turbine blades of the hub. The turbine wheel restrains creation and propagation of cracks due to thermal stress, as well as improving a turbine efficiency.

Abstract: Blade load path on a gas turbine disk can be diverted to provide a significant disk fatigue life benefit. A plurality of gas turbine blades are attachable to a gas turbine disk, where each of the gas turbine blades includes a blade dovetail engageable in a correspondingly-shaped dovetail slot in the gas turbine disk. In order to reduce gas turbine disk stress, an optimal material removal area is defined according to blade and/or disk geometry to maximize a balance between stress reduction on the gas turbine disk, a useful life of the gas turbine blade, and maintaining or improving the aeromechanical behavior of the gas turbine blade. Removing material from the material removal area effects the maximized balance.

Abstract: Blade load path on a gas turbine disk can be diverted to provide a significant disk fatigue life benefit. A plurality of gas turbine blades are attachable to a gas turbine disk, where each of the gas turbine blades includes a blade dovetail engageable in a correspondingly-shaped dovetail slot in the gas turbine disk. In order to reduce gas turbine disk stress, an optimal material removal area is defined according to blade and/or disk geometry to maximize a balance between stress reduction on the gas turbine disk, a useful life of the gas turbine blade, and maintaining or improving the aeromechanical behavior of the gas turbine blade. Removing material from the material removal area effects the maximized balance.

Abstract: Blade load path on a gas turbine disk can be diverted to provide a significant disk fatigue life benefit. A plurality of gas turbine blades are attachable to a gas turbine disk, where each of the gas turbine blades includes a blade dovetail engageable in a correspondingly-shaped dovetail slot in the gas turbine disk. In order to reduce gas turbine disk stress, an optimal material removal area is defined according to blade and/or disk geometry to maximize a balance between stress reduction on the gas turbine disk, a useful life of the gas turbine blade, and maintaining or improving the aeromechanical behavior of the gas turbine blade. Removing material from the material removal area effects the maximized balance.

Abstract: A turbine blade damper-seal assembly includes a seal nested within a damper such that both the seal and damper are disposed to provide sealing between adjacent blade platforms. The seal traverses the seal slot in the damper and seals the gap between adjacent blade platforms for the full axial length of the neck cavity between adjacent blades. The damper is located in an aft most position and includes features to facilitate vibration-dampening performance. The damper also includes features that cause entrapment between blades and therefore avoids the conventionally required protrusions on the blade to retain it in the assembled position.

Abstract: The present invention provides a turbine blade having a revised under-platform structure including a unique coating combination that reduces mechanical stress factors within the turbine blade. The turbine blade includes a platform with an airfoil extending upwardly from the airfoil and a root portion extending downwardly from the platform. Two suction side tabs extend a first distance outward from a suction side of the root potion. Two pressure side tabs extend outward from a pressure side of the root portion. One of the two pressure side tabs extends outward a distance similar to the first distance, however, the other of the two pressure side tabs extends outward a distance much smaller than the first distance, which reduces stresses acting on the turbine blade. In addition, a plurality of coatings are systematically applied to the turbine blade to further reduce mechanical stress factors and improve cooling.

Abstract: A first stage turbine disk section for a gas turbine engine, including a centrally disposed disk hub having an integrally-formed, radially outwardly extending web terminating at an outer end. The disk hub has a convex, curved hub surface exposed to high pressure, high temperature compressor bleed and discharge gases during engine operation, and the convex, curved hub surface acts to reduce the axial constraint at the location of peak axial stress, thereby mitigating the formation of undesirable axial stress in the disk hub and allows separation of the peak axial and hoop stresses, thereby reducing the magnitude of the peak hub surface effective stress.

Abstract: A blade root retaining system for attachment of a turbine blade to a turbine disc of a gas turbine engine comprises a blade root having at least one projection on each of opposite sides thereof, the projection extending from a leading edge to a trailing edge of the blade in an axial direction toward a longitudinal axis of the gas turbine engine. The blade root is received in an attachment slot defined through a periphery of the turbine disc. The attachment slot is configured in shape and direction for retaining the blade root.

Abstract: Blade load path on a gas turbine disk can be diverted to provide a significant disk fatigue life benefit. A plurality of gas turbine blades are attachable to a gas turbine disk, where each of the gas turbine blades includes a blade dovetail engageable in a correspondingly-shaped dovetail slot in the gas turbine disk. In order to reduce gas turbine disk stress, an optimal material removal area is defined according to blade and/or disk geometry to maximize a balance between stress reduction on the gas turbine disk, a useful life of the gas turbine blade, and maintaining or improving the aeromechanical behavior of the gas turbine blade. Removing material from the material removal area effects the maximized balance.

Abstract: A method of repairing a generator blower hub assembly (12) and rotor blade (10) removed from the generator blower hub assembly (12) where the rotor blade (10) has fatigued material in threads of a threaded blade root (14). The method may include removing (32) at least a portion of the fatigued material to form the blade root (14) to a first or new diameter. Threads may be formed (36) on the diameter of the blade root (14) using a roll forming process. A collar (20) may be secured within a threaded hole (16) formed in the generator blower hub assembly (12) where the collar (20) includes internal threads (35) for receiving the threads formed on the diameter of the blade root (14), which may be threaded into the collar (20). Collar (20) may be fabricated of steel alloy and include relief groove (40) for improving the fatigue resistance of the threaded blade root (14) when reassembled into the generator blower hub assembly (12).

Abstract: A mask is provided for shielding a radially extending edge of an impeller vane during a weld repair process, where the impeller vane extends radially and axially outwardly from a shaft. The mask comprises a hub and a flange. The hub has an outer surface and an opening extending therethrough, and the opening is configured to allow the shaft to extend at least partially therethrough. The flange extends radially outwardly from the hub outer surface and is configured to shroud substantially all of the impeller vane radially extending edge.

Abstract: Blade load path on a gas turbine disk can be diverted to provide a significant disk fatigue life benefit. A plurality of gas turbine blades are attachable to a gas turbine disk, where each of the gas turbine blades includes a blade dovetail engageable in a correspondingly-shaped dovetail slot in the gas turbine disk. In order to reduce gas turbine disk stress, an optimal material removal area is defined according to blade and/or disk geometry to maximize a balance between stress reduction on the gas turbine disk, a useful life of the gas turbine blade, and maintaining or improving the aeromechanical behavior of the gas turbine blade. Removing material from the material removal area effects the maximized balance.

Abstract: Blade load path on a gas turbine disk can be diverted to provide a significant disk fatigue life benefit. A plurality of gas turbine blades are attachable to a gas turbine disk, where each of the gas turbine blades includes a blade dovetail engageable in a correspondingly-shaped dovetail slot in the gas turbine disk. In order to reduce gas turbine disk stress, an optimal material removal area is defined according to blade and/or disk geometry to maximize a balance between stress reduction on the gas turbine disk, a useful life of the gas turbine blade, and maintaining or improving the aeromechanical behavior of the gas turbine blade. Removing material from the material removal area effects the maximized balance.

Abstract: Described herein is a novel and useful cross fluid-flow axis turbine (CFAT) that converts the energy of a flowing fluid to rotational energy. The invention includes an axle rotatably mounted on a frame. In one embodiment a plurality of arms extending outwardly from the axle. Each arm has at least one beam that supports at least one outer foil and an inner foil, the inner foil being positioned between the inner most outer foil and the axle. The upper surfaces of the foils face toward the direction of rotation of the axle so that lift generated by the foils contributes to the rotation of the axle. The foils are oriented and aligned so that fluid flowing into the turbine is redirected by the inner foils to the leading edge of the outer foils. The flow of fluid through the center of the turbine can be restricted by a baffle or by placing the inner edge of the inner foils sufficiently close to the axle.

Abstract: A method of manufacturing a root portion of a wind turbine blade includes, in an exemplary embodiment, providing an outer layer of reinforcing fibers including at least two woven mats of reinforcing fibers, providing an inner layer of reinforcing fibers including at least two woven mats of reinforcing fibers, and positioning at least two bands of reinforcing fibers between the inner and outer layers, with each band of reinforcing fibers including at least two woven mats of reinforcing fibers. The method further includes positioning a mat of randomly arranged reinforcing fibers between each pair of adjacent bands of reinforcing fibers, introducing a polymeric resin into the root potion of the wind turbine blade, infusing the resin through the outer layer, the inner layer, each band of reinforcing fibers, and each mat of random reinforcing fibers, and curing the resin to form the root portion of the wind turbine blade.

Abstract: A vane wheel for torque converter with a plurality of blades formed integrally with a shell in a casting process is disclosed. The blade tilts relative to an axial direction of the vane wheel and has an inner surface located on a positive pressure side thereof and an outer surface located on a negative pressure side thereof. A root thickness in a middle section between an outer periphery side and an inner periphery side of the shell is thicker than a root thickness on the inner periphery side relative to a blade cross section along a circumferential direction from an axis center of the vane wheel.

Abstract: Aspects of the invention are directed to a turbine blade assembly and an associated method of manufacture. The turbine blade assembly can include: (a) a unitary structure having an airfoil and a primary root portion; (b) a separately-formed platform; and (c) a secondary root portion that cooperates with the primary root portion to define a root. The platform can be secured in place between the structure and the secondary root portion. In a preferred embodiment, the structure can be formed by casting, and the secondary root portion can be formed by powder metallurgy. Ideally, the structure is configured so as to avoid large changes in section size or to at least make such changes gradual, particularly in transitioning from the airfoil to the primary root portion. The turbine blade assembly can reduce the propensity for defects to form during the manufacturing process. Thus, higher casting yields can be realized.

Abstract: A turbomachine rotor including at least one disk having a ring constituted, for example, by a winding of metal-coated silicon carbide or alumina yarn. According to the invention, the composite ring is received in a closed annular cavity, e.g. defining by machining a groove in an enlarged portion of the disk and closing the groove by means of an annular plate after the ring has been put into place.

Abstract: A turbine blade locking device for axial retention of a turbine blade (1) having a blade root (3) inserted in an axially extending profiled lobe slot (5) of a turbine disk (4), wherein, a sealing fin of a blade platform (10) includes, over part of its axial length, a retaining groove (6) featuring a semicircular cross-section, a disk lobe (7) of the turbine disk (4) is provided with a semi-spherical recess (8), and a ball (2) is fitted into the recess (8) and the retaining groove (6).

Abstract: A rotor blade, a rotor disc (24) or load bearing assembly having at least one bearing surface (22) for contact with a corresponding bearing surface (28). At least one selected area of the bearing surface (30,32) which, in operation, is an area of alternating stress greater than about 50 MPa (peak to peak), is configured to have a co-efficient of friction lower than the remainder of the bearing surface (34). The selected area (30,32) having a co-efficient of friction lower than the remainder of the bearing surface (34) may be provided by the application of a dry film lubricant, with the remaining area(s) (34) of said bearing surfaces (22) being substantially free of the said coating.

Abstract: An improvement for a turbomachinery blade having an airfoil portion, a neck portion, and a root portion, the neck portion extending from the root portion, and the airfoil portion extending from the neck portion, and the root portion being tear-drop shaped, includes a jacket attached to the root portion and extending along a portion of the neck portion. Additionally, a process of forming a turbomachinery blade includes steps of providing a laminate of a material; providing a blade insert; wrapping the laminate around to insert to form a blade having a root portion, a neck portion extending from the root portion, and an airfoil portion extending from the neck portion; and providing for a jacket secured around the root portion and a portion of the neck portion extending from the root portion, the jacket having such shape as to prevent delamination of the laminates at a critical point due to centrifugal force acting on the blade.

Abstract: A boot provides for removable attachment to a tail rotor assembly to flexibly protect the interior of the tail rotor blade assemblies. The boot includes a first fastener strip mounted adjacent a first boot edge and a second fastener strip mounted adjacent a second boot edge. The first and second fastener strips are “hook and loop” material to removably attach the edges which permits relatively uncomplicated installation and removal of the boot.

Abstract: A rotating blade body is provided that can restrain vibrations of rotating blades effectively. The rotating blade body comprises a rotor disc, a plurality of rotating blades being assembled so as to extend from the outer circumference of the rotor disc in a radial pattern, and sealing pins extending along the direction of the rotating shaft in the gaps between the platforms of the rotating blades being adjacent in a circumferential direction. The sealing pins have a through-hole made therein, penetrating axially from one end surface to the other end surface.

Abstract: A tuning notch is defined preferably in the back of a blade root to tune the blade natural frequency in a gas turbine engine.

Abstract: The present invention provides an apparatus for use in automated welding repairs, particularly those repairs intended for impellers, blisks, and similar structures. There is provided a welding shield which is shaped to fit securely within valleys of an impeller. An individual shield is sized to match with a particular impeller design. When in place, the shield covers exposed areas of the valley and thus protects that area from welding damage. The shield may include welding resistant materials on exposed surfaces. A plurality of shields can protect as many impeller valleys as needed. The shield fits securely so that it does not dislodge during automated welding operations. If desired shields can be moved and repositioned so as to expose previously protected areas thus allowing welding to take place in the newly exposed areas.

Abstract: A gas turbine engine rotor blade includes an airfoil, platform, shank, and dovetail integrally joined together. A dust escape hole extends through the platform adjacent a fillet bridging the platform and shank to bleed dust therefrom during operation.

Abstract: A compressor forward stub shaft comprising a plurality of axially spaced, annular rows of dovetail grooves, at least a first and a second of the plurality of rows having dovetail slots coated in part with an anti-wear coating.

Abstract: A cooling arrangement comprises a support (20) for a plurality of blades (22). The support (20) comprises a plurality of blade mounting members (28) provided between adjacent blades (22). The blades (22) are mounted upon the blade mounting members (28). The cooling arrangement defines a pathway (36) for a cooling fluid. The cooling arrangement further includes a fluid directing formation (40, 44, 44A) to direct the cooling fluid across the blade mounting member (28).

Abstract: A method for laser shock peening a gas turbine engine blade includes laser shock peening a thin airfoil of the blade, forming a laser shock induced twist in the airfoil, and altering a root of the blade to counter the laser shock induced twist in the airfoil. The altering may be done after the laser shock peening. The altering may be done before the laser shock peening during casting or forging of the blade or during a machining or broaching procedure which cuts a shape of the root. One embodiment of the altering includes forming the root with an altered root centerline having an altered centerline angle with respect to a predetermined root centerline designed for a non-laser shock peened airfoil of the blade.

Abstract: A deflector arrangement is provided for improving turbine efficiency by imparting added tangential velocity to a leakage flow entering the working fluid flowpath of a gas turbine engine.

Abstract: A blade of an axial compressor comprising: an airfoil is disclosed that has a leading edge and a root; a platform attached to the root of the airfoil; a dovetail attached to a side of the platform opposite to the airfoil; a neck of the dovetail adjacent the platform, and a slot in the neck and generally parallel to the platform, and the slot extends from a front of the neck to position in the neck beyond a line formed by the leading edge of the blade.

Abstract: An interblade platform for a blade support disc on a turbojet blower includes a deflecting part with a lower face provided with a first fixing lug provided with a first orifice for the passage of a first fixing stud and a second fixing lug provided with second and third orifices for the passage of second and third fixing studs. The fixing studs are intended to fix the two fixing lugs to the support disc between two adjacent blades.

Abstract: An embedding element (11) for embedment in the root of a wind turbine rotor blade (15) of a fibre composite material, said embedding element being elongated and having a first end portion (1) and a second end portion (2) and provided with fastening means, eg a threaded hole, a threaded rod or the like in its first end portion (1). Between its two end portions (1, 2) the embedding element (11) is provided with a first longitudinal lateral face (14) extending substantially concavely in a cross-sectional view perpendicular to the longitudinal axis of the embedding element, and with a second longitudinal lateral face (16) facing opposite the first lateral face (14) and extending substantially correspondingly convexly in a cross-sectional view perpendicular to the longitudinal axis.

Abstract: A blade fixing and blade mounting slot arrangement for a gas turbine engine has a mismatch fit along a portion of the length of the blade fixing and slot where contact stress would otherwise be maximal.

Abstract: A rotor blade (1) for a flow machine includes an aerodynamic blade (2) and a foot (3). The foot (3) has an anchoring section (5) on a side facing away from the blade (2). A rotor (4) of the flow machine has an axial anchoring groove (6) which is complementary to the anchoring section (5). The anchoring section (5) can be inserted axially into the anchoring groove (6) in order to radially attach the rotor blade (1) to the rotor (4). At least one pulling-off contour (7) is formed on the foot (3), via which an axial pulling-off force can be introduced into the rotor blade (1), which is attached to the rotor (4) with the aid of a pulling-off tool (15) which is complementary to this pulling-off contour (7).

Abstract: A hovering wind turbine (1) whose blades (2) lie on the surface of an imaginary horizontal cylinder and have their pitch angle changing as a function of the rotational angle, allowing them to gather wind energy mainly in the upwind and downwind sectors of the cylindrical path and to use a fraction of the gathered energy to create lift by deflecting air downwards, mainly in the upper and lower sectors of the cylindrical path. The remainder of the gathered wind energy is used to drive a pair of on board electrical power generators (9), and anchoring tethers (13) carry electrical current to a land of offshore site. The turbine (1) is tied to a blimp (11) that is purposely located in its wind shade and keeps it airborne during periods of light or no wind.

Abstract: A blade of an turbomachine having an airfoil with a leading edge and a root; a base attached to the root of the airfoil; a dovetail portion of the base engageable with disk; a slot in the base generally parallel to a face of the base extending between opposite sides of the base, and a vibration adsorbing insert snuggly fitted into the slot.

Abstract: On a high-pressure turbine rotor wheel, where a blade neck interspace (14) exists between two adjacent turbine blades (4) which is defined by the disk lobe (2) and the opposite blade necks (5) and blade platforms (6), part of the air supplied by the high-pressure cooling channel (9) is passed via a diverter cooling channel (10) into the blade neck interspace (14) to effectively cool the blades and the disk rim (1) in this area and to simultaneously supply the low-pressure channels (12) with cooler air and displace the hot sealing air. Thus, the life of the turbine rotor wheel is increased with simple means.

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: Blades of increased blade angle may be retrofitted onto existing compressor rotor wheels as replacement blades whereby higher performance and efficiency without replacing the rotor wheel may be achieved. The base attachments of replacement blades are provided with lateral extensions or flanges and the rim of the compressor wheel is provided with recesses straddling the grooves to accommodate the flanges. The leading edge of the replacement blade is positioned circumferentially for support only by the platform extension thereby limiting the mean and vibratory stresses in the leading edge of the airfoil and maximizing damage tolerance.

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: A fastening device for fastening moving blades to a disk of a turbomachine rotor includes moving-blade roots which have a mirror-symmetrical cross section with projections, axial disk slots holding the moving-blade roots in a positive-locking manner, and one riveted joint per moving blade for axially fixing the latter. Each blade root and the disk, in the region of each moving-blade root, have axial extensions. The extensions of moving blade and disk are opposite one another in pairs, in each case a hole lying in a radial plane relative to the disk axis is formed half in the extension of the moving blade and half in the extension of the disk. A rivet is inserted into each hole and fixed by plastic deformation.