Abstract: A leaf spring inserted between a sealing member installed between various components of a gas turbine to prevent an inflow and outflow of gases and a mating sealing member, and a sealing assembly including the same are proposed. The leaf spring includes a curved body section, a locking protrusion protruding in one direction at an end of the body section so as to be fitted into the locking groove of the sealing member, and a detachment-prevention protrusion protruding in a direction opposite to the locking protrusion and configured to comes into contact with one surface of the mating sealing member and prevent the locking protrusion from being detached from the locking groove when the body section is compressed by the sealing member and the mating sealing member.

Abstract: An assembly including a propeller blade and a pitch adjustment system for setting the pitch of the blade. A root of the blade has a bulb and the pitch adjustment system for setting the pitch of the blade includes a cup having an upper opening; —a retaining ring, which extends around the bulb, is restricted at least in terms of axial movement towards the opening relative to the cup and has a bearing face to inhibit axial movement of the root towards the opening The pitch adjustment system being further includes a lower seat for the root to engage the cup, the seat and/or the retaining ring being mounted so as to be capable of axial translation relative to the cup by at least one clamping mechanism.

Abstract: A rotor arrangement for a rotor of a gas turbine includes at least one rotor disk with attachment slots for carrying rotor blades and an annular groove having an annular opening towards the outward direction; and rotor blades having an airfoil and a blade root and assembled in an attachment slot. Each of the assembled blade roots have a root extension with a root groove. The root groove faces the annular groove when the rotor blade is assembled in the attachment slot. For each assembled rotor blade a locking element is provided, the locking element engaging the annular groove of the rotor disk and the root groove of the respective rotor blade. Preferably all locking elements are embodied as a locking sheet metal strip, which has two bended tongues embracing the respective root extension.

Abstract: A blade for use in a gas turbine engine, the blade comprising a blade portion and a fir tree root portion, the blade portion and the root portion having a connected passage for allowing cooling air to flow within the blade, the fir tree root portion having an air intake on its leading edge, the air intake allowing cooling air to enter the cooling passage and wherein the fir tree root portion comprises a plurality of projections, including at least a base projection and a top projection; and wherein the air inlet located in the base projection of the fir tree root portion and wherein the air inlet comprises at least 50% of the face of the base projection of the fir tree root portion.

Abstract: A rotary machine is provided. The rotary machine includes a rotor disk including a plurality of slots, a plurality of blades mounted on an outer circumferential surface of the rotor disk, each of the blades having a root part inserted into an associated one of the slots, and a retainer restraining axial movement of the plurality of blades, wherein the retainer includes an inner fixing member inserted into a bottom of the slot, an outer fixing member inserted into a lower surface of the root part to abut against the inner fixing member, and a spacer inserted between the outer fixing member and the bottom of the slot to prevent the outer fixing member from being separated.

Abstract: A rotor assembly for a gas turbine engine includes a rotor extending circumferentially about a central axis. The rotor includes a ring that extends around the central axis and a mount that extends axially away from the ring. The ring includes an axially facing engagement surface. A blade extends radially outward from the ring of the rotor. A damper is coupled with the mount of the rotor and engaged with the engagement surface to minimize vibrations of the rotor assembly.

Abstract: A rotor and a turbo machine including the same are provided. The rotor includes a disk including a disk slot, a blade including a root member inserted into the disk slot and an airfoil disposed outside the root member in a radial direction of the disk, a locking sheet disposed inside the root member in the radial direction, and a fixing portion disposed inside the root member in the radial direction and configured to fix the locking sheet to the root member.

Abstract: A rotor and a turbo machine including the same are provided. The rotor includes a disk including a disk slot, a blade including a root member inserted into the disk slot and an airfoil disposed radially outside the root member, and configured to form a cooling cavity between an inner surface of the disk slot and the root member, a lifting part installed in the cooling cavity and configured to press the root member in a radially outward direction, and a supporting part installed in the cooling cavity, disposed in contact with the lifting part, and configured to support the lifting part.

Abstract: A gas turbine engine rotor having an axis, comprising: a body about the axis having an inner surface, a seat having an outer seat edge at a first radial location surrounded by the inner surface at a second radial location, a lip along the inner surface having an inner lip edge spaced axially away from the seat to define a gap, the lip at a rated axial location between the inner lip edge and the seat facing toward the seat at a normal angle; a damper in the gap having first and second damper surfaces, the first damper surface adjacent the seat; and a split ring in the gap adjacent the second damper surface, having an outer ring edge spaced from the inner surface, engaging the lip at the rated axial location and resiliently expandable against the lip deflects the split ring to load the damper against the seat.

Abstract: Methods, apparatus, systems, and articles of manufacture are disclosed for a slot to accommodate a blade of an engine. An example apparatus includes: a socket to receive an end of a blade; and a retaining device to interact with the socket and the blade for retention of the end of the blade in the socket. The example retaining device includes: a spacer to be positioned in the socket with the end of the blade; a wedge positioned inside the spacer; and a spring to tighten to pull the wedge to expand the spacer.

Abstract: A flow inducer assembly and a method for cooling turbine blades of a gas turbine engine are presented. The gas turbine engine includes a rotor disk having circumferentially distributed disk grooves and turbine blades. Each turbine blade includes a blade root inserted into blade mounting section of the disk groove. Seal plates are attached to an aft side circumference of the rotor disk. The flow inducer assembly is integrated to each seal plate at a side facing away from the rotor disk. The flow inducer assembly is configured to function as a paddle due to rotation of the rotor disk and the seal plate therewith during operation of the gas turbine engine to drive ambient air as a cooling fluid into the disk cavity and enter inside of the turbine blade from the blade root for cooling the turbine blade.

Abstract: A compressor rotor disk for a gas turbine includes a circumferential surface; and a plurality of dovetail slots formed along the circumferential surface, each dovetail slot comprising a cutout portion formed by removing a portion of a radially outer convex surface of the dovetail slot. The compressor rotor disk can reduce stresses generated in the compressor rotor disk by centrifugal force, by reducing the centrifugal force of the compressor rotor disk through the formation of the cutout portion at the radially outer portion of the compressor rotor disk, thereby enhancing the durability and reliability of the compressor rotor disk.

Abstract: A blade retainer adapted for use in a gas turbine engine is configured to block axial movement of a blade. The blade retainer includes a first brace, a second brace, and a web that extends between the first brace and the second brace. The blade retainer is configured to block axial movement of a root of the blade out of a blade receiver slot.

Abstract: A gas turbine engine is provided. The gas turbine engine defines a radial direction. The engine includes: an airfoil positioned within an airflow and extending between a root end and a tip along the radial direction; and a mount coupled to or formed integrally with the root end of the airfoil for mounting the airfoil to the engine, the mount including an outer surface along the radial direction exposed to the airflow and defining an air-cooling channel extending between an inlet and an outlet, the inlet positioned on the outer surface of the mount.

Abstract: A rotor disk assembly for a gas turbine engine includes a rotor disk that defines an axis; and a damper pivot support that extends from a rim of the rotor disk, the damper pivot support comprising an aperture that defines a pivot axis parallel to the axis. A method to dampen a rotor blade of a gas turbine engine, includes independently pivoting a first damper and a second damper of a damper assembly about the pivot axis to respectively contact a first blade and a second blade in response to centrifugal force below a platform of the respective first blade and second blade.

Abstract: A blade is provided including a structure made of composite material including a fibrous reinforcement obtained by three-dimensional weaving and a matrix in which the fibrous reinforcement is embedded, the structure made of composite material including an aerodynamically profiled blade portion and a blade root portion, the blade root portion including two sections each connected to the blade portion, a blade root fastening part including a wall delimiting a cavity and an opening formed in the wall, the structure made of composite material extending through the opening so that the blade portion is located outside the fastening part and the blade root portion is located inside the cavity, and a blocking part disposed in the cavity, between the two sections of the blade root portion, in order to keep the two sections apart from each other so as to oppose withdrawal of the blade root portion from the cavity via the opening.

Abstract: A turbine, a gas turbine including the same, and a method of assembling and disassembling the same, capable of ensuring stable sealing performance as well as facilitating assembly and disassembly are provided. The turbine may include a rotor disk having a lower hook, a plurality of turbine blades installed on an outer peripheral surface of the rotor disk, each of the turbine blades having a blade hook located above the lower hook, and a retainer sealing a cooling passage defined between the turbine blade and the rotor disk and supported by the blade hook and the lower hook.

Abstract: A turbine blade includes a blade having a first end, and a second end opposite the first end. A tip is located at an outer radial portion of the blade. A blade dovetail complements a corresponding dovetail slot in a turbine rotor. The dovetail has a body, and a plurality of projections extending from the body in opposing directions for complementing a plurality of recesses in the corresponding dovetail slot. A tapered groove extends through the body from the first end to the second end. The tapered groove has a tapered profile such that a first depth of the tapered groove near the first end is greater than a second depth of the tapered groove near the second end. The tapered profile gradually transitions from the first depth to the second depth, so that the tapered groove is open at a bottom surface and sized to engage a shim.

Abstract: The gas turbine engine rotor assembly includes a rotor disc with a plurality of sealing tabs projecting radially out from a peripheral surface of the rotor disc in the rear end portion thereof. A said sealing tab is disposed at a tip portion of a fixing member of the disc, formed between pairs of blade root slots, adjacent a trailing edge of the rotor disc. The sealing tabs help to reduce the leakage of secondary air out the back of a blade pocket defined between adjacent blades mounted on the rotor disc.

Abstract: A turbine includes a stator and a movable rotor, rotating around an axis, in relation to the stator. The rotor comprises at least one disc, the radially outer periphery of which comprises cavities. The rotor further includes blades, each having a blade root axially engaged in a cavity of the disc so as to radially hold the blade on the disc. At least one of the blade roots has an axially flaring shape, with the corresponding cavity having a matching shape so as to axially lock the blade root in the cavity in a first axially oriented direction. Locking means lock the blade root in the cavity in a second axially oriented direction, opposite the first direction.

Abstract: A sealing system for a rotary machine is provided. The sealing system includes a pair of circumferentially-adjacent rotary components and an axial seal. Each of the rotary components includes a platform including a first side channel and an opposite second side channel, a shank extending radially inwardly from the platform, and a dovetail region extending radially inwardly from the shank. The axial seal is sized and shaped to be received in the first side channel of a first of the rotary components and the second side channel of a second of the rotary components, such that the axial seal sealingly interfaces with the first and second channels.

Abstract: A turbine blade has a first end, and an opposing second end. A base includes a dovetail for complementing a corresponding dovetail slot in a rotor. The dovetail has a body and a plurality of projections extend from the body in opposing directions. A tapered groove extends through the body from the first end to the second end, and has a tapered profile such that a first depth of the tapered groove near the first end is greater than a second depth of the tapered groove near the second end. The tapered profile gradually transitions from the first depth to the second depth. The tapered groove is open at a bottom surface of the body and is sized to engage a shim. The tapered groove includes a flat section near the first end or the second end, and the flat section has a constant depth.

Abstract: A composite wear pad for being coupled to a slider block of a convergent nozzle of a gas turbine engine includes a high heat capacity composite having a resin and a plurality of carbon fibers bonded together by the resin. The composite wear pad also includes a first rod coupled to the high heat capacity composite at a first axial end of the composite wear pad such that a first end thickness. The composite wear pad also includes a second rod coupled to the high heat capacity composite at a second axial end of the composite wear pad such that the first axial end and the second axial end of the composite wear pad each have an end thickness that is greater than a middle thickness of the composite wear pad.

Abstract: Fixing set for securely assembling and locking a blade shaped element, for instance a turbine blade, on a support, for instance a turbine rotor. The blade shaped element is adapted to be inserted into a groove of the external surface of the support and the fixing set includes first, second and third locking elements adapted to be inserted into the groove, between the lower end surface of the root portion of the blade shaped element and the groove base surface. The first, second and third locking elements are adapted to securely lock the blade shaped element on the support and are further adapted to allow easy and simple installation by an operator with no need of additional working and extra machining.

Abstract: A fan for gas turbine engine that includes hub and annular array of fan blades mounted to hub and extending radially outwardly from hub, each fan blade includes blade root. The hub includes a plurality of hub grooves having shape complimentary to shape of blade root and each hub groove receives and radially retains one blade root. A slot is provided on the suction side and on the pressure side of each blade root. Plurality of keys is provided, each key engages with slots in one blade root and with corresponding hub groove so as to axially retain fan blade. The key includes two arms and joining member, two arms are received in slots provided on blade root and joining member extends between the arms. Interface is defined between key arms and slots in blade root, and greater or equal percentage of interface is curved than is planar.

Abstract: Damping systems are provided for a rotor blade platform. The damping system may include a blade platform defining a damper pocket and a CMC wedge damper positioned within the damper pocket. The CMC wedge damper has at least one damper angled surface parallel to a longitudinal axis. The damper pocket comprises a pocket angled surface positioned about the at least one damper angled surface.

Abstract: A wheel assembly for a gas turbine engine is disclosed. The wheel assembly includes a disk arranged for rotation about a central axis and formed to include a plurality of slots. The wheel assembly also includes a plurality of blades sized to be received in the plurality of slots so that the blades are coupled to the disk for common rotation about the central axis. The wheel assembly further includes a plurality of blade biasers positioned in the slots between the disk and the blades so that the blade biasers are engaged with the disk and the blades to preload the blades away from the central axis.

Abstract: A blade-disk assembly of a turbomachine is provided, the blade-disk assembly having a plurality of adjacent rotor blades and a closure blade which are tilted into an anchoring groove, and at least one circumferential retention element which interlockingly cooperates with at least one blade, as well as a plurality of tilt-out prevention elements which are disposed between the groove base and the root portions and which, in the rest state, space the blades from the groove base when in the upper position. The blade-disk assembly further has a locking element, a portion of which is located between the groove base and the root portions and which, in the rest state, spaces the closure blade from the groove base when in the upper position. A method for assembling such a blade-disk assembly, as well as a turbomachine, is also provided.

Abstract: In order to prevent deterioration in performance of an oil-free screw compressor and scuffing caused by rust, surfaces of both male and female rotors are coated with heat-resistance coatings containing a solid lubricant. A coating contains Polyimide resin to which Molybdenum disulfide, as a solid lubricant, and Aluminum oxide and Titanium oxide, as additives, are added. Accordingly, it is possible to realize a coating that is higher in heat resistance and longer in lifetime than a conventional one.

Abstract: A gas turbine rotor and blade include a root portion, a platform and airfoil portion arranged along a span direction of the rotor blade, the platform located between the root and airfoil portion. The platform has an upstream and downstream side, side faces which extend from upstream to downstream side, an axial groove in each side face extends perpendicular to the span direction with a minor component of extension in span direction. A radial groove in each side face extends towards the axial groove with a component of extension in span direction and a component of extension perpendicular to the span direction. The radial groove has a first end that shows away from the axial groove and a second end that shows towards the axial groove. The second end is located a distance from the axial groove forming a groove free section between the second end and axial groove.

Abstract: A gas turbine engine component has an axial seal pin assembly for sealing between adjacent platform regions where the seal pin assembly design provides a seal pin damper slot extending to a forward axial stop in a forward buttress region. The forward axial stop includes a relief slot to interrupt the loadpath from the airfoil into the blade neck and attachment regions to reduce the stress concentrations in these regions.

Abstract: A securement system for securing an attachment portion of at least one rotor blade or rotor blade segment in a mounting site of a rotor base. The securement system has at least one holder element and a second holder element, each of which has a holding portion and a connecting portion. The connecting portions can be introduced into the mounting site so as to establish a clip connection and can be locked together in the mounting site. The holding portions are configured so as to fasten the attachment portion of the at least one rotor blade or of the rotor blade segment in the mounting site in the locked state of the connecting portions. Further disclosed is a method for securing an attachment portion of at least one rotor blade or rotor blade segment in a mounting site of a rotor base and a rotor having a securement system.

Abstract: A bladed rotor wheel for a turbine engine, the rotor wheel including a disk including, formed in its outer periphery, at least one slot for mounting the roots of blades, and a spacer mounted between each blade root and a bottom of the slot. The spacer includes at least one member of type that is bistable in position and that can occupy a first stable position for assembly or disassembly in which the member does not exert force on the blade root, and a second stable position in which the bistable member exerts a radial force on the blade root. The spacer also includes at least one resilient damper member mounted between the bottom of the slot and the blade root.

Abstract: Fan assemblies, and in particular fan wheels and stator assemblies for fan assemblies, are disclosed. In one embodiment, the fan wheel includes a wheel back having an outer surface forming one of a curved dome-shape and a truncated cone-shape. The fan wheel may also include a plurality of fan blades radially spaced about and mounted to the outer surface of the wheel back. In one embodiment, each of the fan blades is formed from a segment of an airfoil-shaped aluminum extrusion defining at least one internal cavity. The fan blade first ends can be provided with a compound cut profile with at least one curved cut such that the first end of the blade is mounted flush to the wheel back outer surface. The stator assembly can also be provided with a plurality of stator blades formed from airfoil-shaped aluminum extrusion segments and provided with compound cut profiles.

Abstract: A rotor assembly includes a rotor disc and a plurality of circumferentially-spaced, radially outwardly extending rotor blades. The rotor disc has a rim and a plurality of circumferentially spaced slots provided in the rim of the rotor disc, each rotor blade having a root, the root of each rotor blade being arranged in a corresponding one of the slots in the rim of the rotor disc, and at least one of the slots having a chocking device accommodated within a space between a radially outer surface of the slot and a radially inner surface of the root of the rotor blade. The chocking device includes at least a first spring and a second spring, the first and second springs each abutting a radially inner surface of the root of the corresponding rotor blade and arranged transversely relative to a longitudinal axis of the slot.

Abstract: A turbine rotor disk includes a row of buckets secured about a radially outer periphery of the rotor disk, each bucket having an airfoil, a platform, a shank and a mounting portion, the mounting portion received in a radial slot formed in the rotor disk such that adjacent buckets in adjacent radial slots are separated by a rotor disk post located between adjacent mounting portions and a shank cavity between adjacent shanks, radially outward of the rotor disk post and radially inward of adjacent platforms. The shank cavity is substantially filled with at least one discrete thermal plug.

Abstract: A beam for suspending a turboshaft engine from an aircraft structure, including a first attachment mechanism configured to be secured to the aircraft structure and at least one second attachment mechanism configured to be secured to the engine. The beam is at least partially made from a metal-matrix composite material including reinforcing fibers. In one embodiment, the beam takes a form of a circle arc.

Abstract: A blade for a gas turbine engine includes a shank interconnecting a root and a platform, and an airfoil extending radially from the shank. The shank includes a pocket with the platform overhanging the pocket. A rail extends axially along a lateral edge of the platform and extends radially inward from the platform in a direction opposite the airfoil. A gusset extends from an underside of the platform facing the pocket and in a circumferential direction between the rail and the shank.

Abstract: A support rib for securing a rotor pole to a flange in a gearless drive is described. The support rib is undivided and passes through an opening of the flange. The support rib can be joined with the flange and/or the rotor pole using an adhesive. The rib preferably has at least one notch that engages an edge of the flange opening to establish mechanical connection between the rib and flange. An insert with two ends that slideably engages a rib edge and rotor pole channel can be used to couple the rib to the rotor pole.

Abstract: An adjustable blade root spring device for turbine blade fixation in turbomachinery. The device is designed to be placed in a space in a rotor disk cavity adjacent to a tip of a blade root fir tree, where the device applies a radial outward force on the turbine blade to fix the blade position in the rotor disk. The device includes an accordion-shaped spring which is compressed by a bolt and a coil spring. When the accordion spring is compressed in length, it increases in height and makes contact with the rotor disk and the turbine blade. The force of the accordion spring on the turbine blade can be adjusted via the bolt, and the coil spring provides an increased compliance range. The device can be inserted into the space without scraping against the blade root or the rotor disk, and expanded once it is in position.

Abstract: A rotor assembly. The rotor assembly may include a rotor, at least one axial slot positioned about the rim of the rotor having a first staking recess positioned therein, a blade positioned within each of the axial slots having a second staking recess positioned therein, a staking insert having a base portion and a projection extending therefrom with the base portion being disposed within the first staking recess and the projection being disposed within the second staking recess, and a shim positioned within the first staking recess adjacent to the base portion, opposite the projection, of the staking insert. The first staking recess retains the staking insert while the projection of the staking insert retains the blade in both the aft and forward directions.

Abstract: A safety device for a turbine blade is provided. The safety device includes a shear pin, a clamping piece and a securing element, wherein the shear pin is arranged in a corresponding bore in the turbine blade foot and the clamping piece is designed having an upper limb and a lower limb, wherein the clamping piece exerts a radial force on the turbine blade foot for radial safeguarding.

Abstract: An axial turbomachine rotor is provided. The turbomachine includes a rotor body designed in a rotation symmetrical manner around the rotor axis, a rotor blade ring including rotor blades that are each fixed to the rotor body by the blade foot thereof and a sealing disc designed in a rotation symmetrical manner around the rotor axis that is arranged having the outer edge thereof radially inside, neighboring an axially extending protrusion of the blade foot, so that a hollow space is formed between the blade foot and the sealing disc, wherein a groove exists on the outer edge leading radially to the outside, wherein a sealing ring is mounted that may slide radially to the outside in the groove during operation of the rotor by centrifugal force until the sealing ring radially contacts the inside of the protrusion and thus seals the hollow space on the blade foot.

Abstract: A retention system retains a turbine blade dovetail in a retention ring dovetail slot. The retention system includes a retention opening formed in the blade dovetail. A through hole is formed in the retention ring, wherein the through hole is positioned such that it is aligned with the retention opening when the dovetail is assembled in the dovetail slot. A connector is disposed in the through hole and engages the retention opening. In this manner, undesirable relative movement of parts can be reduced or eliminated, thereby preventing excessive wear.

Abstract: The invention relates to a device (30) for blockage of vanes in a circumferential groove (10) of a turbine engine disc, the device being intended to deploy in a radial direction to be locked in the circumferential groove open radially. According to the invention, the blocking device comprises a first and a second piece (32, 34) designed to deploy in the radial direction by relative displacement of these pieces in the circumferential direction (23).

Abstract: A gas turbine engine fan assembly includes a rotor disk, a fan blade and a spacer. The rotor disk includes a longitudinally extending slot with a first spacer contact surface. The fan blade includes a root with a second spacer contact surface arranged within the slot. The spacer includes a leaf spring backbone and a compliant member. The leaf spring backbone includes one or more root contact segments connected longitudinally to one or more slot contact segments. A first of the root contact segments contacts the second spacer contact surface, and a first of the slot contact segments contacts the first spacer contact surface. The compliant member is radially between the first spacer contact surface and the first root contact segment, and radially between the second spacer contact surface and the first slot contact segment.

Abstract: Spacers interposed between blades attached to the disk of a rotor having slots and bottoms of the slots are disclosed. Each spacer is made of an elastically deformable material and at least one longitudinal segment of the spacer presents a transverse profile of arcuate shape to define two lateral contact zones.

Abstract: A lock assembly includes a lock body with an undercut slot which receives a retaining wire of a polygon shape.

Abstract: A vibration-damping device for a turbomachine blade including a blade root configured to be inserted into a pocket of a carrier disk of a bladed wheel, the device configured to be positioned between the blade root and a retaining wall of the pocket. The device includes at least one shim including an assembly of layers of rigid materials and of viscoelastic materials, and including two portions, each forming a lateral branch which can be inserted along one of the two retaining walls, the two lateral branches being connected to one another, so as to constitute a single part, by a third portion forming a shim bottom of rigid material. At least one layer of viscoelastic material is positioned between two layers of rigid materials.

Abstract: A cover plate for a rotor disk in a gas turbine machine includes a cylindrical body having multiple outward facing snaps and multiple inward facing snaps.