Abstract: A method for predicting a blade structure failure within a coupled blade structure including a plurality of blades supported for rotation on a rotor and a shroud structure coupling the blades. The method includes the steps of determining displacements of a plurality of predetermined circumferential locations on the shroud structure during rotation of the blade row, where the displacements are provided as a function of time relative to the periodic rotation of the shroud structure for time intervals that are integer multiples of rotor rotation. A signal characteristic related to vibrational mode and a nodal diameter of the shroud structure is derived based on the displacements of the circumferential locations on the shroud structure.

Abstract: The present invention concerns a method for reducing vibration levels likely to occur in a turbomachine comprising at least a first and a second bladed wheel, where the two wheels are moving relative to one another about an axis of rotation and a gaseous fluid is passing over them, due to perturbations of an aerodynamic origin that are produced by the second bladed wheel or an obstacle on the first bladed wheel. The method includes the following steps in the design of said two bladed wheels: an initial configuration of the blades is defined; the synchronous forced response on the first bladed wheel is calculated as a function of the harmonic excitation force produced by the second bladed wheel expressed in the form of a linear function of the generalized aerodynamic force for the mode considered; a geometric tangential shift value ? is determined for the stacked cross sections of one of the two wheels so as to reduce the corresponding term to the generalized aerodynamic force.

Abstract: A rotor blade having a wing and beam construction for a tip shroud is disclosed. Various construction details are developed for providing a transition zone that extends from the suction side and pressure side of the airfoil to provide a flow path surface of the shroud. In one detailed embodiment, the transition zone over substantially all of its extent between the leading edge region and the trailing edge region is contoured to extend to the sides of the wings to provide a spanwise taper that extends to the side of the wing.

Abstract: A gas turbine engine includes a tip shroud ring rotatably supported on a last stage blade assembly in a turbine. The tip shroud ring is formed entirely of a fiber reinforced ceramic matrix composite material to provide a very light weight tip shroud ring, but also a very strong ring. The blades include pins extending from the tips, and the tip shroud ring includes holes in which the pins slide. Thus, the tip shroud ring is completely supported by the blade tip pins, and the centrifugal force developed by the rotation of the tip shroud ring does not add to the tensile force developed along the blades. Thus, the maximum rotational speed of the turbine rotor as defined by An2 can be increased. Also, because the blade tips are secured to the tip shroud ring in a circumferential direction, the tip shroud ring acts to dampen the bladed turbine by allowing only the middle of the blade to vibrate.

Abstract: A turbine blade arrangement comprises turbine blades which are secured in adjacent positions to a rotor disc with a cavity defined between root segments and platform segments. A flow deflector within the cavity is provided as an insert such that through a recessed portion coolant flow .from a coolant path is constrained to remain adjacent to a rim surface. By constraining the coolant flow to remain adjacent to the surface greater cooling efficiency is achieved. Inner surfaces of the deflector may also be coated with low emissivity materials to reduce radiant heat flux transfer. The flow deflector supports a damper member in association with the platform segments.

Abstract: A conformal seal (20) for sealing air flow between a cooling airflow path and a hot gas flow path within a combustion turbine engine. The conformal seal (20) may be fitted within cooperating side slots of adjacent vane segments (10) within the combustion turbine engine. The conformal seal (20) may include an elongated metallic substrate (22, 40) forming an upper surface and a lower surface. A conformal coating (26, 44) may be deposited over one or both surfaces of the substrate (22, 40). The conformal coating (26, 44) may be deposited to a depth so that a point contact between the conformal coating (26, 44) and respective interior walls of the side slots wears the conformal coating (26, 44) to establish surface area contact there between. The surface area contact improves a sealing function between the conformal coating (26, 44) and the respective interior walls during operation of the combustion turbine engine.

Abstract: A gas turbine rotor blade capable of effectively reducing creep damage by forming a cooling through hole to cool a target area in which significant creep damage of a shroud cover is predicted based on analysis of stress and temperature acting on the gas turbine rotor blade. A gas turbine using the rotor blade, and a power plant using the gas turbine are also provided. The gas turbine rotor blade includes a blade section provided with a shroud cover at an outer peripheral end thereof, and a platform, a shank and a dovetail which are formed in integral structure to successively continue from the blade section. An inner cooling hole is formed to penetrate through the gas turbine rotor blade from the dovetail to the shroud cover. The shroud cover has a cooling through hole formed to open in an outer surface of the shroud cover and extend in communication with the inner cooling hole.

Abstract: A damping arrangement for guide vanes, in particular for guide vanes of a gas turbine or an aircraft engine, is disclosed. Referring to this damping arrangement, the radially external ends of the guide vanes of a guide vane grid or a guide vane ring are mounted to a housing. The radially internal ends of the guide vanes form an inner shroud. At least one seal bearing is mounted to the inner shroud of the guide vanes. At least one spring element is installed between the inner shroud of the guide vanes and the, or each, seal bearing. The, or each, spring element is configured as a leaf spring.

Abstract: An enhanced damping system for a turbine bucket. The damping system includes a damper pocket and a damper pin with an offset center of gravity positioned within the damper pocket.

Abstract: A rotatable member is adapted for dynamic balancing by the addition of balancing weights in the form of high speed projectiles fired at the member during rotation in timed relationship with the rotation of the member so as to impact the member at specific circumaxial locations required for balancing the member. The member has an annular groove for receiving the projectiles and securing the same at the desired locations. The groove has a cross sectional configuration slightly smaller than that of the projectile in at least one direction of measurement so as to engage and capture the projectile in retention in the groove.

Abstract: A geothermal turbine comprises a turbine casing, a turbine rotor shaft, a plurality of nozzle diaphragm outer rings fixed to the turbine casing, a plurality of nozzle diaphragm inner rings located radially inside the nozzle diaphragm outer rings, a plurality of nozzles placed between the nozzle diaphragm outer rings and the nozzle diaphragm inner rings, which form a steam passage, a plurality of rotor blades mounted on the turbine rotor shaft, which face the nozzles and form a blade cascade in a circumferential direction, a shroud arranged on tip of the rotor blade, an overhang attached to the nozzle diaphragm outer ring, which extends downstream of the nozzles and has an inner surface facing radially outside the shroud, and a plurality of sealing fins which protrude radially outward from the shroud and face the overhang.

Abstract: A rotor blade having a wing and beam construction for a tip shroud is disclosed. Various construction details are developed for providing the shroud with a depression in the shroud and a transition zone that extends from the suction side and pressure side of the airfoil to provide a flow path surface of the shroud. In one detailed embodiment, the tip shroud has a depression generally outwardly of the airfoil and generally following the curve of the pressure and suction sides of the airfoil from the leading edge region to the trailing edge region.

Abstract: A turbine blade damper seal assembly includes a seal and a damper that both abut a radially outermost non-gas path surface. The seal is fabricated from a plastically deformable material and nests within a recess of the damper. The damper is fabricated from a rigid material that absorbs vibrational energy generated during operation. The recess within the damper provides for both the damper and the seal to be positioned at the radially outermost non-gas path surface.

Abstract: The segments of the vanes of a stator and the blades of a rotor for a gas turbine engine are dampened by utilizing braided wire cable that is affixed to adjacent segments such that each of the wire strands forming the cable is allowed to move relative to each other to dissipate the vibratory energy inflicted on the stator or rotor. The ends of the cable are welded or heated and then tapered to facilitate fitting into a sleeve or collar which in turn is affixed to the adjacent segments by braze, weld or the like.

Abstract: A rotor of a steam or gas turbine is equipped with rotor blades, which are held in the rotor (6) in a plurality of radial rows and comprise a blade foot 1 installed in the rotor (6), a blade leaf (2) and a cover plate (3). An open pocket (5) is prepared in the sloping surfaces of the cover plates (3) of a row of rotor blades, which the sloping surfaces are located opposite each other. The pockets (5) of two adjacent cover plates (3) form together an essentially closed cavity, which expands in the radial direction of the rotor (6). A pin, whose largest cross section is smaller than the largest cross section of the cavity but larger than the smallest cross section of the cavity, is placed freely movably into each cavity.

Abstract: A rotor blade having a wing and beam construction for a tip shroud is disclosed. Various construction details are developed for providing a transition zone that extends from the suction side and pressure side of the airfoil to provide a flow path surface of the shroud. In one detailed embodiment, the transition zone is contoured under the beam differently than a portion under the rear portions of the wings.

Abstract: A bucket tip shroud in a gas turbine includes a seal rail having a cutter tooth at one end. Bucket tip shroud creep rate can be reduced by removing the cutter tooth from the bucket tip shroud seal rail after use or with a pre-groove honeycomb shroud. Preferably, the remaining geometry matches a geometry of the seal rail within a predetermined tolerance.

Abstract: A one-piece blade for a turbine section of a gas turbine engine, the blade comprising a root, an airfoil and a shroud. The shroud extends generally perpendicularly from a tip of the airfoil and is defined by a pair of opposed bearing faces and a pair of opposed non-bearing faces. The bearing faces each have a contact portion adapted to contact a shroud of an adjacent blade. The shroud has a substantially constant nominal thickness and the bearing faces have a substantially constant face thickness across the contact portion, the face thickness being greater than the nominal thickness. The transition between the face thickness and the nominal thickness is substantially discontinuous.

Abstract: This arrangement relates to a moving blade (1), cast in one piece, for a turbomachine, in particular a turbine or compressor, comprising an aerodynamically shaped profile body (2) which has a shroud band (5). The shroud band (5) projects beyond the profile body (2) in the circumferential direction (6) and has a fin (12). The fin (12) possesses a base portion (14), a transitional portion (15) and a sealing portion (16). An axially measured wall thickness decreases in a transitional portion (15) from the base portion (14) to the sealing portion (16). To avoid a porous structure in the base portion (14), two depressions (17), which reduce the wall thickness of the base portion (14), are integrally molded on the outside of the base portion (14).

Abstract: An assembly for preventing rotation of a damper in a stator system includes a slot in the damper and a block for engaging the slot to prevent rotation of the damper. The block is positioned within a groove located at a mid span portion of an inner air seal.

Abstract: The division wall is made up of a plurality of division wall sections forming a passage wall of high temperature gas which are connected in the direction of arrangement of blades to form a wall surface having a roughly circular cross section as a whole, a gas flow restricting structure for preventing high temperature gas from passing through a gap formed at a connecting portion between the division wall sections in the flow direction of the high temperature gas from the opening on the upstream side of the high temperature gas in the gap, or a gas flow restricting structure for preventing the high temperature gas from being embraced in the gap, for example, a sealing member formed into a prism having a T-shape cross section as a whole composed of a plane portion as a sealing portion and a projected portion for filling the gap is provided.

Abstract: A damper pin is disposed between adjacent buckets of a turbine rotor. A first bucket has circumferentially extending supports defining a pair of axially spaced surfaces on which a damper pin rests in a cold condition of the turbine. The adjacent bucket is undercut adjacent its platform to provide an angled surface overlying a generally correspondingly angled surface of the damper pin. The damper pin fits slightly loose within the recess and, upon turbine rotation at speed, the angled surfaces of the damper pin and recess cooperate to bias the damper pin against the first bucket whereby the damper pin engages both buckets and dissipates vibratory action.

Abstract: The present invention relates to a turbine bucket to be provided at the low pressure last stage of a steam turbine and an object of the present invention is to provide a turbine bucket in which the adjacent blades are connected without using a connecting member at a blade intermediate portion. In order to achieve the object of the present invention, a turbine bucket of the present invention is formed in such a manner that the blade sectional configuration is twisted from a blade root portion to a blade tip side, and when assuming two axial directions in a blade section of the bucket on horizontal plane and taking one axial direction as X axis and the other axial direction perpendicular to X axis as Y axis, the blade sections at predetermined heights from the blade root portion of the turbine bucket are formed in a range of ±0.3 mm from respective points defining blade section configurations as shown respectively in chart 1, chart 4, chart 7, chart 10, chart 13, chart 16 and chart 18.

Abstract: A compressor blade includes an airfoil portion having a leading edge, a radially inner attachment portion, and a platform between the airfoil portion and the attachment portion, wherein material is removed from the attachment portion to form an undercut at a front face thereof to thereby provide an overhang radially inward of the platform and leading edge of the airfoil portion, the undercut defined by a narrow entry slot opening into a rearward transverse groove.

Abstract: A steam turbine that passes more turbine driving steam by off-setting the turbine moving blade throat•pitch ratio before operation and, when a blade untwist is generated during operation, causing more turbine driving steam to flow by maintaining an appropriate value, and, at the same time causing the turbine moving blade throat•pitch ratio to swell by giving the blade untwisting angle to the blade cross-sections in regions where the aerodynamic loss is small. The steam turbine is one in which the throat•pitch ratio (S/T) distribution of a turbine moving blade is offset by forming a curve providing at least one minimal value and maximal value by giving blade twist angle to the blade cross-sections in the blade height direction from blade root to blade tip and, at the same time, the distribution of throat•pitch ratio (S/T) taking into consideration blade untwist generated during operation.

Abstract: A turbine blade arrangement includes rotating blades respectively provided with a leg which can be respectively inserted into a groove on the outer circumference of a turbine disk in a radial, positive fit and which are respectively provided with a profiled section which has a lateral platform located in an end area on the side of the disk. At least one part of the platform is joined to the turbine disk by way of a retaining element which is independent from the leg of the blades in order to extend the profiled sections of the rotating blades.

Abstract: A vibration damper 46 for a gas turbine engine has convergent friction surfaces 48a and 48b and is located radially inward of the platforms 20a and 20b of two adjacent turbine blades. The angle subtended by the friction surfaces 48a and 48b is smaller than that subtended by the angled faces 22a and 22b associated with the platforms 20a and 20b. The center of mass of the damper 46 lies in a plane bisecting the angle subtended by the friction surfaces 48a and 48b. In use the damper 46 is urged radially outwards by centrifugal force so that at least one of the friction surfaces 48a and 48b makes planar contact with at least one of the angled faces 22a and 22b. Vibrational energy is dissipated by the resultant sliding movement between the friction surfaces 48a and 48b and the angled faces 22a and 22b. A secondary vibration damping mechanism arises from the oscillation of the damper 46 between the platforms 20a and 20b.

Abstract: The moving blades of a rotor included in a steam turbine are formed so that the width of a gap between opposite end surfaces of first connecting members of adjacent blades along the direction of rotation of the rotor and the width of a gap between opposite end surfaces of second connecting members of the adjacent blades along the direction of rotation of the rotor are determined so as to make the second connecting members start coming into contact with each other at a rotating speed of the rotor higher than a rotating speed at which the first connecting members of the adjacent blades start to come into contact with each other.

Abstract: The invention relates to a turbo-machine (1) comprising a rotor (25) that extends along a rotational axis (15). Said rotor (25) has a peripheral surface (31) which is defined by the outer radial delimitation surface of the rotor (25) and has a receiving structure (33) as well as a first moving blade (13A) and a second moving blade (13B). Each moving blade comprises a blade footing (43A, 43B) and a blade platform (17A, 17B). The blade platform (17A) of the first moving blade (13A) and the blade platform (17B) of the second moving blade (13B) border one another, and a gap (49) is formed between the blade platforms (17A, 17B) and the peripheral surface (31). A sealing system (51) is provided in the gap (49) on the peripheral surface (31).

Abstract: An article that is configured to be rotatably mounted includes an embedded hollow cavity into which at least one dampening member is inserted. The dampening member or members frictionally engage the hollow cavity such that relative movement between the dampening member or members and the hollow cavity operates to attenuate vibration that is generated when the article is rotated.

Abstract: A method for analyzing and designing structures with structural aliasing. The concept of structural aliasing applies in one embodiment of the present invention to structural systems which are cyclically symmetric. In analyzing and designing structural systems such as wheel assemblies of gas turbine engines, it is possible to group the discrete components so as to aliasingly couple particular ordered excitations and harmonic families. With structural aliasing, it is possible to couple the ordered excitations and harmonic families such that resonant response of the wheel assembly does not occur within the operating range of the engine, and also to have the resonant vibratory mode excited in such a manner that the drive coupling of the wheel dampens the wheel assembly.

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

Abstract: A rotor, such as a rotor of an aircraft gas turbine engine, uses damper-ring devices for damping unwanted rotor vibrations. One device has inner and outer damper rings with the outer damper ring providing ring-rotor frictional damping only at large vibrational amplitudes and with the inner and outer damper rings providing inter-ring frictional damping at both small and large vibrational amplitudes. Other devices include a damper ring and a viscoelastic layer, a hollow damper ring containing particulate matter, and a damper ring in the form of a cable made of twisted single-wire strands.

Abstract: An improved gas turbine blade is disclosed that utilizes a blade tip shroud with tapered shroud pockets to remove excess weight from the shroud while not compromising shroud bending stresses. Excess shroud weight removed by the tapered pockets reduces overall blade pull while maintaining material necessary to increase shroud stiffness. The resulting rib created between the tapered pockets provides a surface, away from possible areas of stress concentration, for drilling radial cooling holes.

Abstract: The present invention relates to a blade arrangement with damping elements and to a method of damping vibrations of a blade arrangement. The blade arrangement comprises a rotor and blades arranged on the circumference of the rotor. Damping elements (2, 3) are arranged between the blades and are in contact with the blades due to a centrifugal force, acting in the radial direction, during rotation of the rotor. In this case, at least two damping elements (2, 3) are arranged one behind the other in the circumferential direction of the rotor between adjacent blades, which damping elements (2, 3), during rotation of the rotor, come into contact with one another via a contact surface (6), and of which a first damping element (2) comes into contact with a first friction surface (4) of one of the blades and a second damping element (3) comes into contact with a second friction surface (5) of the other blade. A multiplicity of different vibration states can be effectively damped with the blade arrangement.

Abstract: A vibration damper 46 for a gas turbine engine has convergent friction surfaces 48a and 48b and is located radially inward of the platforms 20a and 20b of two adjacent turbine blades. The angle subtended by the friction surfaces 48a and 48b is smaller than that subtended by the angled faces 22a and 22b associated with the platforms 20a and 20b. The center of mass of the damper 46 lies in a plane bisecting the angle subtended by the friction surfaces 48a and 48b. In use the damper 46 is urged radially outwards by centrifugal force so that at least one of the friction surfaces 48a and 48b makes planar contact with at least one of the angled faces 22a and 22b. Vibrational energy is dissipated by the resultant sliding movement between the friction surfaces 48a and 48b and the angled faces 22a and 22b. A secondary vibration damping mechanism arises from the oscillation of the damper 46 between the platforms 20a and 20b.

Abstract: A bucket cover to be attached to the tip of a turbine bucket has an aperture with an engageable geometry for receiving a tenon extending from the bucket and retaining the bucket cover to the bucket. The cover has an outer planar surface, an inner planar surface to abut against the bucket and an aperture extending between its surfaces. The aperture is defined by a sidewall that has multiple surfaces of varying cross-sectional size so as to structurally engage a tenon that is deformed during assembly. The sidewall of the aperture has a concave curved portion adjacent the outer planar surface leading to a narrow throat portion that corresponds generally to the size and shape of an undeformed tenon. In a method of making a turbine blade attached to a bucket cover, after a tenon is received in the aperture, the tenon is deformed by application of compressive force so that the tenon fills the aperture of the cover.

Abstract: The present invention relates to a blade assembly with damping elements, which comprises a rotor (5) and blades (6) which are installed on the circumference of the rotor and have a blade platform (7), shank (14) and root (8). A damping element (9) is arranged between respectively adjacent blades (6), said damping element (9) being frictionally connected on rotation of the rotor (5), to at least a first region (1, 2) of a first of the respectively adjacent blades (6), and to a second region (3) of a second of the respectively adjacent blades (6). The blade assembly is characterized by the fact that the damping element (9) is configured and arranged between the first and second blades in such a way that the first region (1, 2) and the second region (3) are located at positions which are significantly spaced apart from one another in the radial direction. This blade assembly provides efficient oscillation damping even in the case of small relative movements between adjacent blades.

Abstract: An article that is configured to be rotatably mounted includes an embedded hollow cavity into which at least one dampening member is inserted. The dampening member or members frictionally engage the hollow cavity such that relative movement between the dampening member or members and the hollow cavity operates to attenuate vibration that is generated when the article is rotated.

Abstract: A bucket cover to be attached to the tip of a turbine bucket has an aperture with an engageable geometry for receiving a tenon extending from the bucket and retaining the bucket cover to the bucket. The cover has an outer planar surface, an inner planar surface to abut against the bucket and an aperture extending between its surfaces. The aperture is defined by a sidewall that has multiple surfaces of varying cross-sectional size so as to structurally engage a tenon that is deformed during assembly. The sidewall of the aperture has a concave curved portion adjacent the outer planar surface leading to a narrow throat portion that corresponds generally to the size and shape of an undeformed tenon. In a method of making a turbine blade attached to a bucket cover, after a tenon is received in the aperture, the tenon is deformed by application of compressive force so that the tenon fills the aperture of the cover.

Abstract: A moving turbine blade apparatus has a plurality of moving blades (5) adapted to be mounted on a rotor shaft, a plurality of snubber covers (6) formed on outer ends of the moving blades (5), respectively, so as to be arranged successively in a circle having its center on the axis of the rotor shaft, and a plurality of ribs (10) projecting from outer surfaces of the snubber covers (6), respectively, so as to extend in a circle having its center on the axis of the rotor shaft. At least one of the opposite end portions of the rib (10) has a thickness measured in a direction of the axis of the rotor shaft greater than that of a middle portion of the same rib (10). Steam loss attributable to the leakage of steam through gaps around the outer ends of the moving blades (5) is reduced.

Abstract: A damper for reducing vibrations in an integrally bladed turbine disk is provided. The damper includes an annular member and a plurality of fingers. The annular member is configured so that it is coupled to a face of the integrally bladed turbine disk. The plurality of fingers are circumferentially spaced around the annular member. Each of the fingers includes a base portion which is coupled to the annular member and extends radially therefrom. Each of the fingers is tangentially movable relative to the annular member when the turbine disk vibrates in a diametral mode shape such that the plurality of fingers contacts a surface of the turbine disk to absorb vibrations.

Abstract: A steam turbine that passes more turbine driving steam by off-setting the turbine moving blade throat-pitch ratio before operation and, when a blade untwist is generated during operation, causing more turbine driving steam to flow by maintaining an appropriate value, and, at the same time causing the turbine moving blade throat-pitch ratio to swell by giving the blade untwisting angle to the blade cross-sections in regions where the aerodynamic loss is small. The steam turbine is one in which the throat-pitch ratio (S/T) distribution of a turbine moving blade is offset by forming a curve providing at least one minimal value and maximal value by giving blade twist angle to the blade cross-sections in the blade height direction from blade root to blade tip and, at the same time, the distribution of throat-pitch ratio (S/T) taking into consideration blade untwist generated during operation.

Abstract: A turbine disk assembly having an integrally bladed turbine disk and a plurality of captured damper members. The integrally bladed turbine disk includes a plurality of turbine blades that terminate at an integral tip shroud. A damper aperture is formed between each pair of adjacent integral tip shrouds and includes a first slotted portion and a second slotted portion. The first slotted portion is formed concentric to the two adjacent integral tip shrouds, extending circumferentially and axially through the two adjacent integral tip shrouds. The second slotted portion extends radially outwardly between the two adjacent integral tip shrouds. One of the damper members is disposed in the first slotted portion of the an associated damper aperture and frictionally engages at least one surface of each of the two adjacent tip shrouds to dissipate energy when relative motion occurs between the two adjacent integral tip shrouds to dampen vibration.

Abstract: A blisk includes a disk having a rim from which extends a row of blades. The rim includes axially opposite overhangs one of which includes an arcuate balance land disposed eccentrically to the centerline axis of the disk for balancing the blisk.

Abstract: A turbine engine blade damper for damping vibrations of blades in a turbine engine. The blade damper includes an elongate body extending between a forward end and a rearward end opposite the forward end. The body is sized and shaped for receipt within a gap formed between adjacent platforms of the blades so the body frictionally engages the adjacent platforms to dampen vibrations of the blades and to prevent air from passing through the corresponding gap during engine operation. The damper also includes a retainer mounted on at least one of the forward and rearward ends of the body. The retainer is sized and shaped for receipt within a recess formed in at least one of the adjacent blades to hold the body between the blades. The retainer is sized and shaped to prevent air from passing between the blades and the retainer during engine operation.

Abstract: A steam turbine for use in a thermal or nuclear power plant includes moving blades of a rotor which vibrate as the rotor rotates, inducing an excessive stress in joints of connecting members for connecting the adjacent moving blades and blade portions of the moving blades and damaging the joints. The moving blades are formed so that a width of a gap between opposite end surfaces of first connecting members of adjacent blades along the direction of rotation of the rotor and a width of a gap between opposite end surfaces of second connecting members of the adjacent blades along a direction of rotation of the rotor are determined so as to make the second connecting members start coming into contact with each other at a rotating speed of the rotor higher than a rotating speed at which the first connecting members of the adjacent blades start come into contact with each other.

Abstract: The present invention relates to a blade arrangement with damping elements and to a method of damping vibrations of a blade arrangement. The blade arrangement comprises a rotor and blades arranged on the circumference of the rotor. Damping elements (2, 3) are arranged between the blades and are in contact with the blades due to a centrifugal force, acting in the radial direction, during rotation of the rotor. In this case, at least two damping elements (2, 3) are arranged one behind the other in the circumferential direction of the rotor between adjacent blades, which damping elements (2, 3), during rotation of the rotor, come into contact with one another via a contact surface (6), and of which a first damping element (2) comes into contact with a first friction surface (4) of one of the blades and a second damping element (3) comes into contact with a second friction surface (5) of the other blade.

Abstract: The present invention relates to a blade assembly with damping elements, which comprises a rotor (5) and blades (6) which are installed on the circumference of the rotor and have a blade platform (7), shank (14) and root (8). A damping element (9) is arranged between respectively adjacent blades (6), said damping element (9) being frictionally connected on rotation of the rotor (5), to at least a first region (1, 2) of a first of the respectively adjacent blades (6), and to a second region (3) of a second of the respectively adjacent blades (6). The blade assembly is characterized by the fact that the damping element (9) is configured and arranged between the first and second blades in such a way that the first region (1, 2) and the second region (3) are located at positions which are significantly spaced apart from one another in the radial direction.

Abstract: A damper for reducing vibrations in an integrally bladed turbine disk is provided. The damper includes an annular member and a plurality of fingers. The annular member is configured so that it is coupled to a face of the integrally bladed turbine disk. The plurality of fingers are circumferentially spaced around the annular member. Each of the fingers includes a base portion which is coupled to the annular member and extends radially therefrom. Each of the fingers is tangentially movable relative to the annular member when the turbine disk vibrates in a diametral mode shape such that the plurality of fingers contacts a surface of the turbine disk to absorb vibrations.