Abstract: An annulus filler for mounting to a rotor disc of a gas turbine engine bridges the gap between adjacent disc blades. The annulus filler has a lid defining part of an airflow surface for air drawn through the engine; a separate support connectable to the lid and to the rotor disc with an engagement portion of the support extending radially past a substantially adjacent region of the lid; and a retainer configured to interconnect the lid and the support by engaging each engagement portion of the support and adjacent regions of the lid, the retainer defining another part of the airflow surface. The annulus filler can be configured to allow mounting to the rotor disc, where the support is connected to the rotor disc without the lid, and then the lid is mounted to the support such that each engagement portion remains visible from the lid's radially outermost side.

Abstract: The disclosure relates to adjacently mounted circumferentially distributed turbo machine airfoils with a vibration damping system. Each adjacent pair of airfoils includes a fixing and receiving portion, extending between the paired adjacent airfoils, each with a face that are proximal (e.g., in contact with) each other. Vibration can be suppressed by the fixing and receiving portions each having a received magnet fixingly installed therein and a non-magnetic conducting plate therebetween. Each magnet has a pole that faces the pole of the other magnet in between which the non-magnetic conducting plate is located and in which eddy currents can be induced by the relative movement of the magnets due to vibration.

Abstract: A tip shroud that includes a plurality of damping fins, each damping fin including a substantially non-radially-aligned surface that is configured to make contact with a tip shroud of a neighboring rotor blade. At least one damping fin may include a leading edge damping fin and at least one damping fin may include a trailing edge damping fin. The leading edge damping fin may be configured to correspond to the trailing edge damping fin.

Abstract: A turbine blade damper arrangement in which a damper is positioned against the undersides of the platforms of adjacent turbine blades. In operation, the damper is centrifugally urged into engagement with the blade platforms to provide damping of relative movement between the blades. The damper and platform surfaces that it engages are of part-cylindrical configuration in order to minimize gas leakage paths between the damper and blade platforms.

Abstract: An airflow assembly includes a fan, a shroud, a plurality of ribs, and a fan support. The fan has a number of fan blades. The shroud includes (i) a plenum defining a plenum opening located adjacent to the number of fan blades, and (ii) a barrel extending from the plenum so as to surround the plenum opening. The plenum further defines at least one airflow opening spaced apart from the plenum opening. Each of the plurality of ribs extends inwardly from the barrel. The fan support is attached to the plurality of ribs and is configured to support the fan. The at least one airflow opening is not an attachment structure or a guiding structure, is not configured to receive a fastening member, and does not function as a water drain.

Abstract: A steam turbine rotor blade has a profile and a cover integrally formed on and at an end of the profile, the leading edge of the cover formed on the profile and the trailing edge of a cover formed on an adjacent preceding profile being in contact and connected with each other by the torsional return force produced during rotation. The cover formed on the profile is provided with a radially-formed stepped portion at the trailing edge, the stepped portion having a height large than the thickness of the cover.

Abstract: The invention relates to a compressor, particularly a high pressure compressor, of a gas turbine, particularly of a gas turbine aircraft engine, having at least one rotating blade ring on the rotor side, and having at least one guide blade ring on the stator side, wherein the, or each, guide blade ring is formed of a plurality of guide blade segments, and wherein each guide blade segment is formed of a plurality of individual blades. According to the invention, adjacent individual blades (11, 12; 12, 13; 13, 14) are permanently connected to each other within each guide blade segment (10) of at least one guide blade ring on opposing surfaces positioned radially outward, whereas the same are not connected to each other on opposing surfaces positioned radially inward.

Abstract: A vibration damper (28) is disclosed for use in a turbomachine, the turbomachine comprising at least one turbine rotor (19) having a plurality of radially extending blades (16, 17). Each blade has an aerofoil (22), a platform (21) and a stem (20). The vibration damper (28) has a seal-region (29) which comprises a pair of sealing surfaces (24, 25) configured for engagement with respective contact surfaces (24, 25) provided on adjacent blade platforms (21). The vibration damper (28) also has a mass-region (30) which is configured to extend radially inwardly from the seal-region (29) and to terminate at a position located between adjacent blade stems (20) (FIG. 4).

Abstract: An arrangement between blade elements in a blade row in a turbine is described. Each blade element has at least one shroud element and a blade airfoil which abuts on, and is connected to, the shroud element, and essentially extends radially with regard to a principal axis of the blade row. When installed, the shroud element sides, which extend circumferentially, abut on the respectively adjacent shroud element of the respectively adjacent blade element, each forming an essentially radial gap. At least one blade element has a projection which projects into the shroud element of the abutting blade element and extends in the circumferential direction, and at least one blade element has a recess which accommodates such a projection. In the region of the projection or recess there is a stepped region of the radial gap, and the guiding of the radial gap in this stepped region is a labyrinth seal.

Abstract: A gas turbine blade having reduced stresses around a regions where the airfoil training edge joins the platform is disclosed. The lower stresses are achieved due to an undercut region proximate the blade trailing edge being placed in the platform. The undercut region has a compound fillet radius formed from two single and tangent radii where the first radius closer to the airfoil is larger than the second radius such that a ratio of R1/R2 is approximately 2.0 to 2.5.

Abstract: A vibration damper assembly for damping non-synchronous vibration between adjacent, spaced apart components comprises a vibration damper located in both of a pair of generally confronting passages in each of the components. The assembly comprises at least two spaced apart articulation surfaces for contact between damper and component, each of the articulation surfaces is arcuate in a first direction and is characterized by having a substantially linear portion in an orthogonal and second direction. Thereby the contact area is greatly enlarged and material loss minimized. The cross-sectional shape of the damper or passage is non-circular and there is a clearance between the damper and passage sufficiently small to prevent rotation of the damper in the passage during use.

Abstract: Embodiments of the invention can provide systems, methods, and an apparatus for thermally isolating a turbine rotor wheel. According to one embodiment, a method for thermally isolating a turbine rotor wheel can be provided. The turbine rotor wheel can include a pair of blade posts, or shanks, supporting respective turbine rotor blades such that a shank cavity is formed between respective blade posts. To thermally isolate the turbine rotor wheel, at least one seal operable to partition the shank cavity into an upper region and a lower region can be provided. In addition, this seal can be inserted within the shank cavity to partition the shank cavity into an upper region and a lower region. In doing so, cooling air flow can be divided between the upper region and the lower region so that the lower region can be maintained at a lower temperature than the upper region.

Abstract: A turbine has blade assemblies disposed about a rotor. Each of the blade assemblies have an airfoil and a bucket. Pockets are defined at trailing and leading sides of the bucket, with damper pin slots at an ends thereof. The damper pin slot at the trailing side has a depth sufficient for fully receiving a damper pin. The damper pin slot at the trailing side of a first adjacent blade assembly is positioned relative to the damper pin slot at the leading side of a second adjacent blade assembly to allow the damper pin to move. At each side of the pocket at the trailing side is a seal pin slot with seal pins therein. The seal pin slots extend beyond a line that is aligned with an inner edge of the damper pin slot at the trailing side, wherein the seal pins overlap the damper pin.

Abstract: A damper pin for a turbine bucket includes an elongated main body portion having a first substantially uniform cross-sectional shape and axially-aligned, leading and trailing end portions having a second relatively smaller cross-sectional shape at opposite ends of the main body portion. A seal element is provided on one or both of the opposite leading and trailing end portions projecting radially outwardly beyond the main body portion.

Abstract: A blade arrangement, with a rotor and a plurality of blades which are arranged on the periphery of the rotor in a ring, and a plurality of damping elements are provided. Between two directly adjacent blades at least two damping elements are arranged in series in the circumferential direction of the rotor. As a result of a centrifugal force, which acts in the radial direction, during a rotation of the rotor around a rotor axis adjacent damping elements come into contact with each other, and one of the two damping elements comes into contact with one of the two blades and the other of the two damping elements comes into contact with the other of the two blades.

Abstract: A rotor for turbomachines with shrouded blades comprising a row of a plurality of blades covered by an external ring, coaxial to said row of blades; the external ring being composed of a plurality of sectors, commonly referred to as shrouds, which are adjacent to one another and transversely and circumferentially attached to the end of respective blades; a damping member being arranged so as to straddle a respective pair of adjacent sectors and being pushed, in use, against radially internal surfaces of the sectors of the pair of adjacent sectors.

Abstract: A turbine blade is disclosed. The turbine blade may have an airfoil extending from a first surface of a turbine platform. The turbine blade may further have a first side pocket of the turbine platform that is configured to substantially entirely house a first moveable seal between a forward wall of the first side pocket and an aft wall of the first side pocket. The first side pocket may have a convex surface, extending between the forward wall and the aft wall, and a concave surface. The turbine blade may also have a second side pocket of the turbine platform configured to receive a portion of a second moveable seal.

Abstract: A vibration damper for a rotor blade of a gas turbine engine, includes a plate for placing between an inner bearing surface and an outer bearing surface of the rotor, the bearing surfaces extending in directions that are substantially axial and being radially spaced apart from each other, the plate being folded so as to form two plate portions that are inclined relative to each other, the mass of the plate being distributed in such a manner that its center of gravity is situated on the side of one of these plate portions, the free end of the other plate portion being designed to bear radially against the inner bearing surface, and the junction zone between the two plate portions being designed to bear radially against the outer bearing surface under the effect of centrifugal force when the rotor is rotating so as to provide axial sealing against gas.

Abstract: A fan blade assembly of a turbojet with fan and a fan blade damper is disclosed. The fan blade includes a root and a platform. The fan blade damper is suitable for being attached in a housing formed in the lower face of the platform. The housing includes an upstream side perpendicular to the root of the blade. The damper includes a front edge of which a first portion is parallel to the upstream side of the housing and of which a second portion is inclined relative to said upstream side. A fan blade damper and a method for calibrating the damper are also disclosed.

Abstract: The present invention relates to a rotor for gas or steam turbines having a disk (1), an outer cover (2) and a plurality of blades (3). The disk (1) defines the portion of less diameter of the rotor and has an inner surface configured to be coupled on a shaft (100) of the turbine or connected to an axially adjacent disk (1?, 1?); an outer surface configured for housing a plurality of blades (3). The outer cover (2) defines the portion of greater diameter of the rotor. The blades (3) have a root (31) configured to be coupled on the disk (1) and a tip (32) opposite to the outer cover (2). Between these blades there is a plurality of blades (3) welded together at the tip (32) to form groups of blades (3a) and individual blades (3b) intercalated between groups of blades (3a).

Abstract: A section of a gas turbine engine includes a damper which extends from a disk between a first blade and a second blade to provide a dissipating vibratory motion energy in the form of friction heating.

Abstract: Composite containment casing including a body having an interior, an abradable system integrally joined to the interior of the body of the containment casing wherein the abradable system comprises, a sandwich structure including a first facesheet and a second facesheet position about at least one core layer, and at least one abradable layer applied to the sandwich structure wherein the at least one core layer comprises any of a cell configuration, a columnar configuration, or a truss configuration and wherein the sandwich structure is strong radially and weak circumferentially.

Abstract: A device on a shroud which is provided on a turbine rotor blade tip has a cutting structure designed like a line of ribs, locally projecting over the shroud radially to the rotational axis around which the turbine rotor blade is rotatably arranged. The sealing structure has a longitudinal extent (S) oriented in the direction of rotation (U) of the turbine rotor blade, tapers with increasing radial distance from the shroud, and terminates with a flat end face (St) which radially faces away from the turbine rotor blade, and can be divided into five interrelated surfaces.

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: A fan rotor includes a hub (10), a plurality of blades (20) radially disposed around the hub, and a magnet (30) located in an inner side of the hub. The hub includes a top wall (12) and a sidewall (14) surrounding the top wall. A plurality of spaced first ribs (17) is disposed between the top wall of the hub and a top wall of the magnet. A plurality of spaces (171) is formed between adjacent first ribs for accommodating equilibrium clay (50) therein.

Abstract: A propulsor apparatus for a mobile platform, for example an aircraft is provided. The propulsor apparatus includes an aerodynamically shaped propeller duct that houses at least a portion of a turboprop engine, and fully houses a propeller driven by the engine. Inside the propeller duct is a circumferential ring that closely surrounds the propeller such that only a small clearance is provided between the outermost tips of the propeller blades and an inner surface of the circumferential ring. The circumferential ring includes sound deadening material that attenuates noise generated by the tip vortices created at the outermost tips of the propeller blades. The propeller duct and circumferential ring are supported by a plurality of structural rods that couple to structure of the mobile platform, for example to the spars within a wing of an aircraft. The apparatus and method significantly reduces the noise associated with a turboprop engine.

Abstract: A weight adjustment portion is formed at an upper end of a hub to project radially inward. An approximately U-shaped balance clip is attached to the weight adjustment portion. A gap to be constantly ensured is formed axially between the weight adjustment portion and a rotor holder. For ensuring the gap, a plurality of seats are provided on the lower surface of the weight adjustment portion at regular circumferential intervals. The rotor holder is press-fitted to the hub until the rotor holder comes into contact with the seats. The upper surface of the weight adjustment portion is positioned lower than the uppermost portion of the hub, thereby preventing the balance clip projecting from the uppermost portion of the hub when the balance clip is attached to the weight adjustment portion.

Abstract: A wheel for a rotating flow machine, comprising a rotor (1) on which a plurality of blades (2) are arranged projecting radially from the rotor (1), with each blade (2) having a root (3) attached to the rotor (1) and a blade top (4). In addition, the wheel comprises a ring (5) arranged concentrically in relation to the rotor (1) outside the blade tops (4) in such a way that the ring (5) and the blades (2) overlap each other when the ring and the blades are viewed in a radial direction, which ring (5) is arranged to be caused to rotate by the blades (2) when the rotor (1) rotates. The extent (9) of the blades (2) in a radial direction and the internal dimension (10) of the ring (5) are matched to each other in such a way that there is a clearance (?) between each blade top (4) and the ring (5) in radial direction when the rotor (1) is not rotating.

Abstract: Vibration damping is important with regard to such components as hollow turbine blades in gas turbine engines. Traditionally damping has occurred through damping elements secured at the root or tip of such blades. Such damping is not optimised and results in potential problems with wear in operational life. By providing a tube of deformable material which can be located within a hollow cavity it is possible to provide an element which through friction engagement can absorb vibration energy and therefore damp such vibration. The tube incorporates a number of cuts and/or grooves in an appropriate pattern in order to define a deformation profile once the tube is expanded in location. The tube is secured in position internally upon an expandable element which is typically an inflatable device. Once in position the tube is retained in its expanded deformable profile and the engagement between the tube and the hollow cavity wall surface results in energy absorption through vibration episodes.

Abstract: A vibration damper assembly is provided between two shrouds of adjacent turbine blades. A pair of confronting passages are provided, one in each of the shrouds and a vibration damper is located in both of the confronting passages. The vibration damper comprises a structure which when heated expands outwardly so as to engage with the walls of the passages. The vibration damper may be a sheet metal shape memory nickel base alloy wound into an open spiral and provided in circular cross-section passages.

Abstract: A method facilitates reducing stress in a turbine bucket including a metallic base metal. The method comprises providing a turbine bucket including at least one pocket formed therein, and filling at least one pocket defined in the bucket with a polymer composite having continuous fibers oriented in a resin matrix, such that the fibers have an orientation determined in accordance with a pre-selected frequency tuning of the bucket.

Abstract: A turbine blade tip shroud. The tip shroud may include a platform, a number of Z-notched positioned within the platform, a number of seal rails positioned on the platform, and a number of cavities positioned within the platform and adjacent to seal rails and the Z-notches.

Abstract: A turbine airfoil of a turbine engine having cooling channels positioned on side surface edges of a platform of the airfoil. The platform may include at least one angled side surface that may be aligned with a side surface of a platform of an adjacent turbine blade. The airfoil may include a suction side edge formed from a first surface at an obtuse angle relative to an upper surface of the platform and a second surface at an obtuse angle relative to a bottom surface of the platform. One or more film cooling slots may be positioned in the first surface and may include a diffusion portion. A damper may also be positioned in a groove between the second surface and an adjacent platform. The damper may include cooling slots on a side surface proximate to the second surface of the suction side edge.

Abstract: A damping system for a turbine bucket. The damping system includes a damper pocket with a variable tangential depth and a damper pin positional within the damper pocket.

Abstract: A structure of a balancing fan is disclosed. The fan comprises a hub and a rotation balancing unit. The hub comprises a top portion and a ring portion disposed around the top portion. A plurality of blades extends from the hub. The rotation balancing unit is disposed among the outer side of the top portion and the inner side of the ring portion and comprises a guiding portion and at least one movable weight member. The movable weight member is positioned on the inner side of the guiding portion such that the movable weight member can move along the guiding portion when the hub rotates to provide the dynamic-balance effect and thereby stabilize the rotation of the fan.

Abstract: A vibration damper assembly for damping non-synchronous vibration between adjacent, spaced apart components comprises a vibration damper located in both of a pair of generally confronting passages in each of the components. The assembly comprises at least two spaced apart articulation surfaces for contact between damper and component, each of the articulation surfaces is arcuate in a first direction and is characterised by having a substantially linear portion in an orthogonal and second direction. Thereby the contact area is greatly enlarged and material loss minimised. The cross-sectional shape of the damper or passage is non-circular and there is a clearance between the damper and passage sufficiently small to prevent rotation of the damper in the passage during use.

Abstract: A damper pin for a turbine bucket includes an elongated main body portion of substantially uniform cross-sectional shape having opposite ends, one only of said opposite ends coated with a corrosion and oxidation-resistant coating.

Abstract: The present invention provides a centrifugal fan by which air velocity distribution in the height direction of impellers at a fan blow-out port is made uniform and running noise is reduced, and an air conditioner whose running noise is reduced by applying the centrifugal fan thereto. The centrifugal fan is provided with a hub for fixing the rotary shaft of a motor, a main plate formed on the outer circumference of the hub, a shroud disposed to be opposed to the main plate, which forms an air passage, a plurality of impellers disposed between the main plate and the shroud, and a bellmouth disposed at the suction side of the shroud. Projections or recesses are formed on the surface of the shroud facing the bellmouth, which projections or recesses form air streams flowing from the center of the shroud toward the outer circumference thereof along the surface when running a fan. The air conditioner has a thus constructed centrifugal fan mounted therein.

Abstract: A method enables a stator assembly for a turbine engine to be assembled. The method comprises forming a recess within a portion of each base, and coupling the stator vanes within the turbine engine in a circumferentially-spaced arrangement such that the recessed portion of each base facilitates reducing excitation responses of each of the plurality of stator vanes during engine operation.

Abstract: A composite turbine blade and fan blade assembly for use in a turbo machine such as a gas turbine engine. The turbine blade assembly includes a plurality of metallic turbine blades secured to a rotor disk and forming an outer shroud assembly. A fiber reinforced composite fan blade assembly includes an annular ring with a plurality of fan blades extending there from. The annular ring has a conical shape to allow for easy installation over the outer shroud assembly, and is secured to the outer shroud assembly to force the fan blade assembly to rotate with the turbine blades, yet allow for a slight radial movement between the fan blades and the turbine blades. The turbine blades are formed of a high temperature super alloy to allow for use in a gas turbine engine, and the fan blade assembly is made from a fiber reinforced composite to allow for higher rotational speeds and smaller diameters for smaller gas turbine engines without exceeding allowable stress limits.

Abstract: A wheel for a rotating flow machine, comprising a rotor (1) on which a plurality of blades (2) are arranged projecting radially from the rotor (1), with each blade (2) having a root (3) attached to the rotor (1) and a blade top (4). In addition, the wheel comprises a ring (5) arranged concentrically in relation to the rotor (1) outside the blade tops (4) in such a way that the ring (5) and the blades (2) overlap each other when the ring and the blades are viewed in a radial direction, which ring (5) is arranged to be caused to rotate by the blades (2) when the rotor (1) rotates. The extent (9) of the blades (2) in a radial direction and the internal dimension (10) of the ring (5) are matched to each other in such a way that there is a clearance (?) between each blade top (4) and the ring (5) in radial direction when the rotor (1) is not rotating.

Abstract: A fan assembly is provided. The fan assembly includes a fan, and a shroud having a motor mounting portion on which a fan motor of the fan is mounted. One or more coupling units are provided at a lower side of the shroud, and on a planar surface of an apparatus, and one or more vibration preventing members are interposed between the coupling units. Because the fan and the motor mounted thereto cause vibration and noise to the fan assembly in the apparatus, the vibration preventing members are interposed between the coupling units of the shroud and the apparatus to reduce the amount of vibration and noise transmitted from the fan and the motor to the apparatus.

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: A tip shrouded turbine blade that may include a target pad disposed on an outer radial surface of the tip shroud, the target pad including a raised surface that protrudes radially outward from the outer face of the tip shroud. The target pad may be substantially cylindrical in shape such that an outer radial face of the target pad is substantially circular in shape. The size of the surface area of the outer radial face of the target pad may be configured to be substantially the same size as an area of measurement for a conventional proximity sensor.

Abstract: A retaining ring is mounted in frictional engagement with a static gas turbine engine part, such as a compressor shroud, in order to provide frictional damping.

Abstract: The aft stiffening frame of an air intake structure for an aircraft engine nacelle is made with an inflection area such that the part between the inflection area and one of its ends could be subjected to a large deformation when a constraint occurs in the direction of the frame part defined by the inflection area and its other end, usually normal to the casing. This particular shape of the aft stiffening frame makes it possible to improve its flexibility so that the air intake structure has better resistance to severe deformed shapes that could occur, for example, when a fan breaks on a wide body aircraft.

Abstract: A durable blade, a method of manufacturing it and a method of repairing blades are described and claimed. Durable blade includes an airfoil having a pressure side and a suction side, a midspan shroud located on the airfoil, at least one recess on the midspan shroud and a wear pad attached to the recess of the midspan shroud.

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: A gas turbine or compressor blade, includes a convex outer surface or suction side and a concave inner surface or pressure side, these sides being joined at their upstream ends by a leading edge and at their downstream ends by a trailing edge for the gases and terminating on a platform, one part of which includes a downstream fairing that includes at least one through-notch that extends substantially between a transverse line passing through the trailing edge and the downstream end of the fairing.

Abstract: A turbine moving blade has a blade portion extending from the basal portion to the tip of the moving blade. The blade root portion is formed at the basal portion of the blade portion and engaged with a corresponding disk groove of a turbine rotor on a one-by-one basis, and the integral cover is formed at the tip of the blade portion integrally with the blade portion. The integral cover includes at least a pair of pressure and suction sloped surfaces inclined relative to the direction of the rotational axis of a turbine so as to restrain the elastic restoring force of the moving blade torsionally deformed at the time of installation of the moving blade by bringing the integral covers of mutually adjacent blades into contact.