Abstract: A plug for brazing in a through-hole and for sealing a through-hole in a substrate of a gas turbine component. The plug has at least one elongate part and a head, wherein the head is made wider in cross section than the elongate part. The elongate part has a cross section which becomes smaller towards its end face along its longitudinal axis, wherein the elongate part has a peripheral region on its end face, wherein the peripheral region of the elongate part has a brazing filler metal. The head has a peripheral region on its end face, wherein the peripheral region of the head has a brazing filler metal, and the remaining and largest part of the elongate part and of the head has a mixture of brazing filler metal and a substrate material.

Abstract: An assembly is provided for a gas turbine engine. This assembly includes a rotating structure of the gas turbine engine, a stationary structure of the gas turbine engine and a volume formed by and extending between the rotating structure and the stationary structure. The rotating structure is rotatable about an axis. The rotating structure includes a seal land and a fluid nozzle. The stationary structure includes a seal element. The seal element is arranged with the seal land to form a seal assembly that seals an annular gap between the stationary structure and the rotating structure. The volume is adjacent the seal element and the seal land. The fluid nozzle is configured to direct a fluid jet out from the rotating structure and into the volume.

Abstract: A fan casing for an aircraft turbomachine includes an annular body extending around an axis and equipped with an annular attachment flange at each of its axial ends. The casing further includes an annular coating made from abradable material, wherein the body is made from a composite material, and further includes an annular stiffener that is arranged inside the body and that carries the coating. The stiffener includes an annular ring, a radially outer face of which is separated radially from the body and a radially inner face of which receives the coating. The stiffener further includes annular attachment tabs configured to the wall to the body.

Abstract: Proposed are an apparatus using root cause analysis (RCA) for predicting blade damage and a method therefor. The method for predicting damage includes generating a reduced-order model that derives damage information indicating damage to a blade according to operation conditions of a gas turbine apparatus and a displacement of the blade, measuring the displacement of the blade through a displacement measurement sensor when an operation of the gas turbine apparatus begins under the operation conditions according to a user's input, and deriving the damage information corresponding to the operation conditions and the displacement of the blade through the reduced-order model.

Abstract: A single-piece annular vane array for a gas turbine engine. The single-piece annular vane array comprises at least one mounting feature, a ring of stator vanes, and at least one baffle. The at least one baffle is attached at a first radial location to the ring of stator vanes, and is attached at a second radial location to the at least one mounting feature. The at least one baffle is deformable.

Abstract: A composite airfoil assembly for a turbine engine. The composite airfoil assembly includes a woven core having a core exterior. The woven core includes a first region, a second region, and a transition region. A skin is applied to at least a portion of the woven core at the core exterior.

Abstract: The invention concerns an aircraft engine with an intake cone and a fan stage coupled thereto with a plurality of fan blades, wherein the fan blades are each connected to a drive shaft via a connecting means, wherein a sealing device for blocking air is arranged between a blocking air space and the interior of the intake cone, and during operation of the aircraft engine a leakage air stream flows out of the blocking air space via the sealing device, characterized by at least one outflow means for the leakage air stream in a component of the aircraft engine to allow a flow of the leakage air stream into the interior of the intake cone.

Abstract: A gas turbine engine for an aircraft including: an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan including a plurality of fan blades; a gearbox that can receive an input from the core shaft, and can output drive to a fan shaft via an output of the gearbox so as to drive the fan at a lower rotational speed than the core shaft; and a fan shaft mounting structure arranged to mount the fan shaft within the engine, the fan shaft mounting structure including at least two supporting bearings connected to the fan shaft.

Abstract: A high-pressure gas turbine for a turbomachine includes a nozzle guide vane assembly, an annular array of movable blades mounted downstream of the nozzle guide vane assembly, a first recirculation cavity, a second recirculation cavity, and a bleed cavity. An upstream sealing element is mounted on the nozzle guide vane assembly and a downstream sealing element is mounted on the annular array of movable blades.

Abstract: The present disclosure provides an electric propulsor. The electric propulsor has a propeller having a hub and blades extending from the hub. The blades are movable along respective flapping axes between a deployed position and a stowed position. The electric propulsor also has an actuator and an electric motor coupled with the hub and arranged to rotatably drive the propeller. Further, the electric propulsor has one or more processors configured to: receive an input indicating an azimuthal position of the blades; control the electric motor to align the blades with a target azimuthal position based at least in part on the input; and with the blades of the propeller arranged at the target azimuthal position, cause the actuator to translate a control rod coupled thereto and with the hub to actively fold or unfold the blades along the respective flapping axes of the blades.

Abstract: A turbomachine includes a housing, a rotating group that is housed within the housing, and a fluid bearing system with a thrust member. The turbomachine further includes a compressor stage configured to receive a fluid from an area within the housing. Also, the turbomachine includes a seal plate of the housing. The seal plate includes a first face and a second face. The first face has a flow surface that at least partly defines a fluid flow path toward the compressor wheel. The second face has a thrust member support surface configured to support the thrust member during rotation of the rotating group. The seal plate includes an aperture that fluidly connects the area to the fluid flow path.

Abstract: Gas turbine engines include a compressor section configured to compress a core flow and a combustor section having a combustor that is arranged downstream from the compressor section along a path of the core flow. A turbine section is arranged downstream from the combustor along the path of the core flow, and includes a plurality of first vanes arranged at an outlet of the combustor, with at least one first vane of the plurality of first vanes having an internal vane path. A compressor diffuser is arranged to direct the core flow from the compressor section through the internal vane path. The path of the core flow, in a flow direction, passes through the compressor section, the compressor diffuser, the internal vane path, the combustor, and then between the first vanes of the plurality of first vanes to enter the turbine section.

Abstract: A seal assembly includes a plurality of seal segments for sealing between a rotor and a stator, wherein, when arranged together, the rotor, the stator, and the plurality of seal segments define a high pressure region and a low pressure region. The seal assembly further includes a biasing member engaged with the plurality of seal segments and at least one feature formed in the stator. As such pressurized fluid engaging with the at least one feature provides an aft-to-forward axial force that reduces friction between interfacing aft surfaces of the plurality of seal segments and the stator.

Abstract: A system for controlling the pitch setting of a propeller blade for an aircraft turbine engine, includes a blade having a vane connected to a root. A hub accommodates the root of the blade, and a ring is mounted around the root and in the hub. Stops are also mounted around the root and in the hub. A nut is threadedly engaged with a thread of the ring and is configured to bear axially against the hub to secure the assembly.

Abstract: A downhole power generation system may include a downhole tubular having a generator recess. The generator recess is radially offset from a central bore of the downhole tubular, and the generator recess is connected to the central bore via an opening. The system may further include a generator disposed at least partially within the generator recess. Additionally, the system may include at least one turbine blade configured to drive rotation of the rotor in response to fluid flow through the central bore with the at least one turbine blade in an extended position. The at least one turbine blade is disposed at least partially within the central bore in the extended position, and the at least one turbine blade is configured to move at least partially into the generator recess toward a retracted position in response to contact with a downhole tool moving along the central bore.

Abstract: A stator vane assembly for a compressor of an aircraft turbine engine includes an inner shroud, an outer shroud, and stator vanes. The stator vanes are attached only to the inner shroud and are in non-immobilizing mechanical contact with the outer shroud.

Abstract: An air inlet of a nacelle of an aircraft propulsion assembly assembly comprising an inner wall and an outer wall which are connected upstream by an air inlet lip. The air inlet lip comprising thick portions and thin portions distributed alternately over the circumference of the air inlet lip. Each thick portion comprising an upstream leading edge to separate an upstream air flow into an outer air flow guided by the outer wall and an inner air flow guided by the inner wall during a thrust phase. Each thin portion comprising a downstream leading edge situated longitudinally downstream of each upstream leading edge so as to detach a reverse air flow at the air inlet lip during a thrust reversal phase.

Abstract: A propulsion system includes a gas turbine engine and a nacelle. The nacelle has inner and outer nacelle portions, an upper bifurcation, and a fluid drain system. The inner nacelle portion is disposed radially outside of the gas turbine engine. The outer nacelle portion is disposed radially outside of the inner nacelle portion. An annular bypass duct is disposed between the inner and outer nacelle portions. The upper bifurcation extends through the bypass duct between the inner and outer nacelle portions. The fluid drain system includes a catch tray engaged with a drain mast. The catch tray is disposed radially inside of an inner nacelle wall. The drain mast extends through the bypass duct between the catch tray and a drain mast outer radial end. The outer radial end of the drain mast is engaged with the outer nacelle portion. The drain mast includes first and second fluid passages.

Abstract: A variable guide vane assembly for a gas turbine engine stator is provided. The variable guide vane assembly includes a plurality of vanes and a plurality of RT mechanisms. The vanes extend between a shroud and hub. The vanes are circumferentially disposed and spaced apart from one another. Each vane includes inner and outer radial ends, and inner and outer radial posts. Each vane is pivotally mounted to rotate about its rotational axis. Each RT mechanism is in communication with the inner or outer radial post of a respective vane. The RT mechanism includes a pin connected to the vane that is disposed in a ramp slot non-rotational relative to the pivotable vane. The ramp slot extends between first and second lengthwise ends. Rotation of the vane relative to the ramp slot causes the pin to travel within the ramp slot and the vane to translate linearly.

Abstract: A vane assembly includes an outer case, a plurality of vanes, and an inner body assembly. The outer case defines an outer boundary of a flow path of the vane assembly. The plurality of vanes extend radially inward from the outer case relative to a central axis. The inner body assembly is configured to secure the plurality of vanes radially relative to the central axis. The inner body assembly includes an inner case arranged circumferentially about the central axis, a band coupled with the inner case, and a retainer ring that extends into the band and into each of the plurality of vanes.