Abstract: A containment system for an engine includes an engine case having an inner perimeter. The containment system includes a containment ring nested within the inner perimeter of the engine case and integrally formed with the engine case along a first interface and a second interface. The containment ring includes a first leg opposite a second leg, and the first interface is defined between the first leg and the engine case. The containment system includes a first plurality of perforations defined at the first interface, and the first leg of the containment ring is frangible along the first plurality of perforations to at least partially release the containment ring to protect the engine case during a containment event.

Abstract: A gas turbine engine component includes a tubular body section including a plurality of fiber wraps encompassed within a matrix composition and one or more integrally-formed stiffeners extending from an outer surface of the body section and in a component circumferential direction around the body section. The stiffener includes one or more fiber wraps extending radially outwardly from the body section over a form and to the body section from the form.

Abstract: The invention relates to a method for producing a shielding cartridge (1) of a turbomachine element (10), said method comprising a step of winding a carbon wire (3) comprising a plurality of fibres about a longitudinal axis and a step of trapping and preserving the outside surroundings of the wound-up wire (3) in a housing (2).

Abstract: Gas turbine engine for aircraft including: engine core including turbine, compressor, and core shaft connecting turbine to compressor; fan located upstream of engine core, including plurality of fan blades; gearbox receives input from core shaft and outputs drive to fan shaft to drive fan at lower rotational speed than core shaft; and fan shaft mounting structure arranged to mount fan shaft within engine. Fan shaft mounting structure includes at least two supporting bearings connected to fan shaft. Gearbox's output is at gearbox output position and fan's input is at fan input position. First bearing separation distance is defined as axial distance between input to fan and closest bearing of at least two supporting bearings in rearward direction from fan.

Abstract: A turbomachine includes a fan shaft driven by a turbine shaft via a device for reducing a speed of rotation. The turbomachine includes an uncoupling device interposed between the reduction device and the turbine shaft. The uncoupling device is configured to uncouple the reduction device and the turbine shaft in response to the exceeding of a determined resistant torque exerted by the reduction device on the turbine shaft.

Abstract: A ring for a connection element includes a contact portion intended to cooperate with a contact surface of another ring and a fastening portion intended to be secured to a support. The contact portion is made of a first metallic material and the fastening portion is made of a second metallic material, the hardness of the first material being substantially greater than that of the second material, and the toughness of the second material being substantially greater than that of the first material, the contact portion and the fastening portion being of one piece construction.

Abstract: A replenishment receptacle comprised of a volumetric region for storing a consumable good during shipping to a consumer location, a processor, a transceiver, programmatic instructions stored in a non-transitory storage medium, an a sensor for determining a quantity of the consumable good in the volumetric region. The sensor periodically detects and reports changes in the quantity of the consumable good present the volumetric region. When the quantity is low (below an established threshold), the consumer good is automatically reordered from a vender, purchased, and shipped to a consumer location. In embodiments, a rate of consumption of the consumable good affects when automatic reordering occurs to ensure the consumable good does not run-out before a new quantity arrives at the consumer location. The replenishment receptacle is able to be discarded in embodiments, where in others it is reusable and/or partially recyclable.

Abstract: A method of fabricating a fan case for a gas turbine engine defines a metallic ring including an outer surface and an inner surface. A first composite material is assembled about the outer surface of the metallic ring. A second composite material is assembled about the first composite material. The first composite material and the second material are cured about the metallic ring within a tool to form a first subassembly. The first subassembly is removed from the tool. A fan case assembly for a gas turbine engine and a gas turbine engine are also disclosed.

Abstract: A fan blade containment system is arranged to surround a fan including a plurality of fan blades in a gas turbine engine for an aircraft. The fan blade containment system includes a fan case arranged to surround the fan; a debris retainer, such as a front hook, mounted on the fan case and arranged to prevent forward debris release should all or part of a fan blade become detached from the fan; and a front panel mounted on the fan case adjacent to and rearward of the front hook, and arranged to move or break when struck by a detached fan blade or fan blade part so as to allow the detached fan blade or fan blade part to engage the front hook.

Abstract: A fan blade containment system is arranged to surround a fan including a plurality of fan blades in a gas turbine engine for an aircraft. The fan blade containment system includes a composite fan case arranged to surround the fan; and a metallic insert mounted on the composite fan case and including a metallic hook arranged to prevent forward debris release should all or part of a fan blade become detached from the fan.

Abstract: A property gradient coating to be applied by additive manufacture to an inner wall of a casing mounted on the periphery of moving blades of a turbomachine rotor, the coating including in superimposed layers from an outer surface of the coating to this inner wall of the casing, on the one hand, a first layer consisting of a three-dimensional scaffolding of filaments of an abradable material forming an ordered network of channels or microchannels whose pore sizes are between 50 and 250 microns and whose porosity is greater than 85%, and, on the other hand, a second layer having a function of dissipating energy from acoustic waves striking the outer surface of the coating and consisting of a three-dimensional scaffolding of filaments of a first thermosetting material forming an ordered network of channels or microchannels whose pore sizes are between 50 and 400 microns and whose porosity is greater than 60%.

Abstract: A composite blade is formed by laying up composite layers in which reinforced fibers are impregnated with resin, and has a blade root and an airfoil extending from the blade root in a longitudinal direction. The composite blade includes a first laminate of the composite layers extending along the longitudinal direction in the airfoil and extending along a first inclination direction inclined toward a direction intersecting the longitudinal direction in the blade root; a second laminate of the composite layers extending along the longitudinal direction and contacting the first laminate in the airfoil, the second laminate extending along a second inclination direction inclined toward a direction opposite to the first inclination direction in the blade root and being separated from the first laminate; and a third laminate of the composite layers provided between the first and second laminates in the blade root.

Abstract: A turbocharger, with a turbine for expanding a first medium, a compressor for compressing a second medium utilising energy extracted in the turbine during expansion of the first medium. A turbine housing and a compressor housing are connected to a bearing housing. A casing radially and axially surrounds the turbine housing, the compressor housing, and/or the bearing housing at least in sections, which is connected to the respective housing via multiple fastening devices. At least some of the fastening devices extend through a wall of the casing into the housing include metallic damping elements made of wire on opposite sides of the wall of the casing through which the same extend.

Abstract: A composite vane includes both a vane body extending in a longitudinal direction between an inner end and an outer end and also at least one fastener base connected to the inner end or to the outer end of the vane body. The vane includes metal reinforcement associated with an organic matrix. The metal reinforcement defines integrally as a single part both a longitudinal core extending between the inner and outer ends of the vane body and also the fastener base, the longitudinal core having the organic matrix overmolded thereon, the organic matrix defining the outside shape of the vane.

Abstract: A containment assembly is disclosed for a turbofan engine. The containment assembly comprises a forward fence, an aft fence, a fan track liner, an intermediate casing member, an outer casing member, and an annular nesting cavity. The fan track liner is disposed axially between the forward and aft fences. The intermediate casing member is disposed radially outward of at least a portion of the fan track liner. The outer casing member is disposed radially outward of at least a portion of the intermediate member. The annular nesting cavity is bounded at least in part by the intermediate casing member and the outer casing member.

Abstract: A turbofan engine includes: a cylindrical fan case; a fan rotatably disposed in the fan case and including a central member and multiple fan blades arranged on an outer circumference of the central member such that the fan blades are spaced apart from one another in a circumferential direction; an annular member disposed to surround the fan; and an elastic support device that supports the annular member to the fan case radially elastically such that a predetermined clearance is radially defined between the annular member and tips of the fan blades.

Abstract: A reinforced composite article includes a first ply, a second ply, and a third ply. A first interface material is between the first ply and the second ply. A second interface material is between the second ply and the third ply. A designated pattern of material property variation, geometric structure variation, spatial variation, or combinations thereof is distributed to selected locations identified in the first interface material or distributed between selected locations identified in the first interface material compared to the second interface material. The pattern is sufficient to measurably vary adhesion, toughness, strength, modulus, or combinations thereof. The article allows distribution of a load across the pattern when the first, second, and third plies receive a force from kinetic energy above a separation threshold by partially delaminating the first ply from the second ply, the second ply from the third ply, or both.

Abstract: A gas turbine engine for an aircraft includes: a flow path boundary, which delimits the flow path through the engine radially on the outside, and a lining, which lines the flow path boundary on the inside, at least along an axial section. Here, the lining includes a plurality of panels, which, in the circumferential direction of the flow path boundary, adjoin each other and which together line a circumferential area of 360°, wherein each panel has two end faces, which each adjoin an end face of an adjacent panel. The panels are of beveled design at their end faces, such that two mutually adjoining panels form a V-shaped gap between them, the minimum clearance of which is realized at the inside of the panels. The panels can be sound-absorbing panels. Also disclosed is a panel for a gas turbine engine.

Abstract: A turbine shroud assembly includes a blade track, a multi-piece carrier and a mount assembly. The carrier includes a fore body and an aft body. Attachment pin of the mount assembly extends between fore body and aft body and through an eyelet in an attachment post of the blade track.

Abstract: A bearing arrangement rotatably supports a shaft of an aircraft engine. The bearing arrangement comprises a bearing having rolling elements disposed between inner and outer races. The inner race is affixed to the shaft. A decoupler normally structurally couples the outer race of the bearing to a stator structure of the engine. The decoupler is configured to release the bearing from the stator structure when subject to a predetermined critical load. A bumper is mounted to the stator structure and encircles the bearing. The bumper has a radially inwardly facing surface disposed in close proximity to a radially outer surface of the outer race of the bearing and defines therewith a radial gap to accommodate and constrain an orbiting motion of the rotor about the central axis of the engine after decoupling at the bearing.

Abstract: A gas turbine engine, system, and method of severing at least one blade of a row of blades spaced about a rotor during an over speed condition are provided. The gas turbine engine includes a shaft rotatable about an axis of the shaft, a rotor coupled to the shaft and having a radially outer surface, and one or more rows of blades spaced axially along the outer portion. Each row of blades includes a plurality of blades spaced circumferentially around the radially outer surface. A gas turbine engine casing at least partially surrounds the one or more rows of blades. One or more rows of stationary vanes are spaced axially along the casing. Each row of stationary vanes includes a plurality of stationary vanes extending radially between the shaft and the gas turbine engine casing. The stationary vanes are coupled to the gas turbine engine casing using a fuse element.

Abstract: A method and a turbine rotor system for reducing over-speed potential of a turbine of a gas turbine engine involve mechanically connecting the turbine to at least two mechanical loads via first and second mechanical drives extending in opposite directions from the turbine.

Abstract: Embodiments of the disclosure relate to a device for limiting overspeeding of a turbine shaft of a turbomachine, comprising structure for destroying the rotor blades of at least one turbine stage. The structure includes a casing and at least one projectile configured for projecting into the path of the rotor blades of the stage in order for the stage in question to be destroyed.

Abstract: Flow path assemblies and methods for forming such flow path assemblies for gas turbine engines are provided. For example, a method for assembling an airfoil with a boundary structure to form a flow path assembly is provided. The method includes machining an opening into the boundary structure. The opening is sized to receive an airfoil or other component. The method also includes machining a cutout into the boundary structure proximate the opening. A locking feature is inserted into the cutout. When the airfoil is inserted into the opening, the locking feature interlocks the airfoil with the boundary structure. To seal the airfoil with the boundary structure, the airfoil is pressed against or into the boundary structure. When the airfoil is pressed, the locking feature is compressed such that a seal is formed between the airfoil and the boundary structure to seal the flow path assembly.

Abstract: A containment assembly for a machine having a rotor radially bounded at least in part by the containment assembly. The containment assembly may comprise and inner casing member formed from a composite material, a containment liner bonded to a radially inner surface of the inner casing member, an outer casing member formed from a composite material and positioned radially outward of the inner casing member, and a containment core embedded between the inner and outer casing members.

Abstract: A fan case assembly for a gas turbine engine includes a honeycomb structure. The fan case assembly also includes a septum operatively coupled to the honeycomb structure and located radially inward of the honeycomb structure. The fan case assembly further includes a rubstrip in contact with the septum and located radially inward of the septum.

Abstract: A metal leading edge includes a nose positioned along the leading edge of a fan blade airfoil body. The metal leading edge also includes a first edge extending axially aftward from the nose along a pressure side of the fan blade airfoil body. The metal leading edge further includes a second edge extending axially aftward from the nose along a suction side of the fan blade airfoil body. The first edge and the second edge forming a notch at the conjunction of the first edge, the second edge, and the nose. The metal leading edge also includes a nose length extending from a nose tip to the notch. The nose length at a first radial location is different from the nose length at a second radial location.

Abstract: A gas turbine engine with a geared turbofan arrangement with a gearbox in a drive shaft assembly driven by a turbine is provided. A driving side of the gearbox being driveably connected with at least one propulsive fan, with at least one mechanical fuse in the drive shaft assembly enabling a controlled disengagement of at least one engine part from the drive shaft assembly in case of a mechanical failure of the gas turbine engine or a part thereof and at least one load stop for bearing a load, in particular an axial or radial load in case of the mechanical failure of the gas turbine or a part thereof. A first mechanical fuse is positioned in a torque carrying shaft or a torque carrying part of a shaft, in particular in a torque bearing coupling between the shaft and the gearbox.

Abstract: In combination a wear indicator and a component of a gas turbine engine is provided. The wear indicator is secured to a surface of the component of the gas turbine engine. The wear indicator comprising: a first side delaminates when impacted by a blade of the gas turbine engine; a second side parallel to the first side, the second side being secured to a surface of the component of the gas turbine engine; a mid-section interposed between the first side and the second side; and a reference dimension remains constant when the first side delaminates.

Abstract: A gas turbine engine component includes a rotor blade having a squealer tip comprising a projecting lip, the rotor blade further having a raised rim running along both edges of each projecting kip of the squealer tip, and an abrasive coating formed of hard particles embedded in a retaining matrix covering a tip region of the rotor blade within an area bounded by the raised rim. The raised rim has a height of between 50% and 75% of a mean diameter of the hard particles.

Abstract: A backsheet for use in a turbine engine fan case includes a first portion including a first portion first end and an opposing first portion second end. The first portion is tapered between the first portion first end and the first portion second end. The backsheet also includes a second portion coupled to the first portion. The second portion includes a second portion first end and an opposing second portion second end. The second portion defines a constant thickness between the second portion first end and the second portion second end.

Abstract: A fitted platform for positioning between two adjacent blades of an aviation turbine engine fan includes a flow passage wall made of composite material having a central portion, and first and second margins each extending in a longitudinal direction of the wall. Each margin extends over a determined distance from the central portion in a transverse direction of the wall. The flow passage wall includes a fiber reinforcement densified by a matrix. The fiber reinforcement present in the central portion presents three-dimensional weaving, and the fiber reinforcement present in the first and second margins presents two-dimensional weaving, at least in part.

Abstract: The invention relates to a mechanical fuse intended to be rigidly mounted between a drive unit (8, 9), and a receiver unit (10), each rotating about the same axis (7) of rotation, said fuse comprising a body (13) extending in a longitudinal direction parallel to said axis (7) of rotation, once the fuse is mounted between said drive unit (8, 9) and receiver unit (10). The invention is characterized in that said body (13) comprises a plurality of longitudinal bars (14), each bar (14) being deformable by bending, such as to form a twist-breakable mechanical fuse.

Abstract: A blade track for a gas turbine engine includes a plurality of blade track segments. The blade track segments are arranged circumferentially around a central axis to form the blade track.

Abstract: A gas turbine engine (20) and method for containing a fan inside an engine after a fan thrust bearing assembly failure. The engine (20) may comprise a fan (42), a housing (100) including a compartment (102), a fan shaft (104) inside the compartment (102) and comprising a bowl (108), a support structure (110) inside the compartment (102), a speed sensor pickup (114) mounted on the outer surface (120) of the bowl (108), a speed sensor (112) mounted on the support structure (110), and a fan thrust bearing assembly (41) disposed forward of the bowl (108). The fan thrust bearing assembly (41) including a bearing (126). The speed sensor (112) and the sensor pickup (114) define a defining a sensor gap (116). The bearing (126) and the outer surface (120) defining a fan thrust bearing gap (130), wherein the sensor gap (116) is less than the fan thrust bearing gap.

Abstract: A method for repairing a fan casing in which the inner surface is damaged, includes attaching a reinforcement element to the fan casing, the reinforcement element being attached to the outer surface of the fan casing opposite the damage.

Abstract: A bearing arrangement rotatably supports a shaft of an aircraft engine. The bearing arrangement comprises a bearing having rolling elements disposed between inner and outer races. The inner race is affixed to the shaft. A decoupler normally structurally couples the outer race of the bearing to a stator structure of the engine. The decoupler is configured to release the bearing from the stator structure when subject to a predetermined critical load. A bumper is mounted to the stator structure and encircles the bearing. The bumper has a radially inwardly facing surface disposed in close proximity to a radially outer surface of the outer race of the bearing and defines therewith a radial gap to accommodate and constrain an orbiting motion of the rotor about the central axis of the engine after decoupling at the bearing.

Abstract: A composite annulus filler adapted to bridge the gap between two adjacent rotor blades attached to a rim of a rotor disc of a gas turbine engine. The annulus filler includes a unitary body portion having a platform which extends between the adjacent rotor blades and defines an airflow surface for air being drawn through the engine and a support body extending beneath the platform and terminating in a root which, in use, extends along a groove provided in the rim of the rotor disc.

Abstract: A gas turbine engine impact liner is disclosed. The impact liner may include a base sheet, a plurality of stanchions extending from the base sheet, and a plurality of supports, each being operatively associated with one of the plurality of stanchions.

Abstract: A fan housing assembly of a turbofan engine and method of manufacture includes a fan housing that defines a flowpath radially inward, and a groove structure attached to and located aft of the housing for connection to a stationary portion of an outer shell of a thrust reverser. The groove structure includes a base portion concentrically located about an engine axis, and aft and leading legs spaced axially apart from one another and projecting radially outward from the base portion. The fan housing, the base portion, and the aft and leading legs are made of a one-piece composite material.

Abstract: An improved composite structure and method for making same has been provided. The provided improved composite structure has locally strengthened areas within a reinforcement region. The locally strengthened areas within the reinforcement region have load distribution devices to redistribute load in order to (i) locally strengthen an area around damage induced by an initial momentary and direct transmitted load, and (ii) limit growth and propagation of damage induced by an initial momentary and direct transmitted load during a subsequent unbalance load. The improved composite structure reduces the impact of the fan blade out phenomenon in a weight efficient manner.

Abstract: A turbine component for a gas turbine engine includes a multiple of self-opening cooling passages each of which defines a self-opening cooling passage axis that extends through a gas path surface.

Abstract: A composite external casing for a compressor of an axial-flow turbomachine, the casing comprising a generally circular wall having a matrix and a woven fibrous reinforcement. The latter is formed by a stack of fibrous plies having fibers woven according to a weft positioned in the axis of the circular wall, and a warp positioned around the circumference of the wall. The stack of plies comprising, depending on thickness, two external layers and one central layer. The layers include differences in the proportions between the weft fibers and the warp fibers. In the external layers the warp fibers make up the majority. In the central layer the weft fibers make up the majority. The stack exhibits interlayers inserted between the central layer and each external layer. The plies corresponding to the interlayers exhibit fibers woven at +45° and ?45° in relation to the axis of rotation of the turbomachine.

Abstract: Described is a shaft support system for a gas turbine engine comprising: a rotatable fan shaft; first and second support structures extending in parallel from the shaft to a load bearing structure to provide radial location of the shaft within an engine casing, wherein the first support and second support structures include first and second respective mechanical fusible joints; wherein the first fusible joint is a two-stage fuse which partially fails within a first predetermined load range, the second fusible joint fails within a second predetermined load range which is different to the first load range, and the first fusible joint fully fails only when the second fusible joint has failed.

Abstract: A fan containment system arranged to be fitted around an array of radially extending fan blades mounted on a hub in an axial gas turbine engine. Each fan blade has a respective tip. The system includes: a cylindrical fan case including a hook projecting in a radially inward direction and positioned axially forward of the radial array of fan blades; a fan track liner disposed on the radially inner surface of the fan case; and a damaging tool which projects radially inwards from the fan case towards the tips. The damaging tool has a tip radially outward of the fan blade tips. The damaging tool is configured that in the event that one of the fan blades is released from the hub, the tip of the damaging tool damages the fan blade tip of the released fan blade to promote penetration of the fan blade into the fan track liner.

Abstract: A sealing device of an axial turbomachine, comprising a circular strip with an outer surface and an inner surface; an annular layer of abradable material on one of the outer and inner surfaces of the strip, designed to provide a sealing with an annular row of rotor blades. The surface of the strip covered by the abradable layer comprises at least two perforated circular zones, each with a series of distributed perforations. The surface of the strip covered by the abradable layer further comprises a smooth circular zone disposed axially between the perforated areas. The smooth zone is centered on the outer ends of the rotor blades. The smooth circular zone and the two perforated circular zones improve the adhesion of the abradable layer on the metal strip, and avoid a premature breakdown of the abradable layer in contact with the rotor blades, despite the presence of perforations.

Abstract: A turbomachine casing assembly includes a first casing element locatable radially outward of one or more rotating aerofoil elements of a turbomachine, a second casing element located radially distal to the first casing element and one or more arcuate fence elements positioned between the first and second casing elements. Following impact of a detached fan blade, a radially outward movement of the first casing element causes the or each fence element to project into the air flow passing through the turbomachine.

Abstract: Embodiments of a gas turbine engine rotor including stress relief features are provided, as are embodiments of method for producing a gas turbine engine rotor. In one embodiment, the method includes producing a hub preform in which a plurality of elongated sacrificial cores are embedded. Blades are attached to an outer circumference of the hub preform by, for example, bonding a blade ring to the outer circumference of the preform. The blades are spaced about the rotational axis of the gas turbine engine rotor and circumferentially interspersed with the plurality of elongated sacrificial cores. The plurality of elongated sacrificial cores are then removed from the hub preform to yield a plurality of stress distribution tunnels extending in the hub preform.

Abstract: A turbine blade for a gas turbine engine comprises an airfoil having a pressure side, a suction side, a span direction and a chord-wise direction. The airfoil has an airfoil span on a pressure line being a projection of the stacking line onto the pressure side. The airfoil has a plurality of chords extending between a leading edge and a trailing edge of the airfoil. A generally round dimple is disposed on the pressure side. The dimple is contained in an area extending on the stacking line between 0% and 23% of the airfoil span from the inner end, and in the chord-wise direction between 0% of a first chord and 82% of a second chord from the leading edge. The dimple is configured to initiate fracture of the blade at a predetermined speed of rotation. A method of preventing rupture of a disk of a turbine rotor is also presented.

Abstract: Various methods and systems are provided for an axial turbine including a containment shroud. In one example, an axial turbine for use in an engine system comprises a turbine disc/blisk and a shroud housing the turbine disc/blisk, the shroud including a first region, a second region, and a third region. A thickness of the second region is dependent on a length between the first region and the third region, and the third region has a burst strength that under a burst condition retains one or more fragments of the turbine disc/blisk.