Abstract: Provided is a wind turbine blade, with a generally hollow blade body including half shells and webs, with each web including flanges with the first and second webs being supported via respective reinforcement structures relative to the respective half shell, which reinforcement structures are arranged between an outer and an inner layer of the upper and lower half shell of the shells and extend in the lengthwise direction of the blade, wherein the first and second reinforcement structures each include at least one stack composed of several pultruded composite strips including carbon fibers with the strips being fixed in the resin, wherein at least one stiffening element extending parallel to the reinforcement structures over at least a part of their length comprising at least one stack composed of several glass fiber layers infused with resin is arranged between the reinforcement structures.

Abstract: A module including a device for changing the pitch of the vanes about the pitch axes thereof, the device including: a hydraulic actuator capable of rotating about the longitudinal axis (X) and configured to rotate the vanes about their pitch axes (C), a supply pump for supplying fluid to the hydraulic actuator, the supply pump including a drive shaft capable of rotating about the longitudinal axis (X) and a shell that is rotationally fixed relative to the longitudinal axis (X), and an electric machine configured to operate the supply pump, the electric machine having a rotary member for driving the drive shaft and a fixed member that is rotationally fixed relative to the longitudinal axis (X).

Abstract: A modular wind turbine blade includes a blade root, an intermediate portion and a blade tip. The intermediate portion includes a plurality of modular blades, two adjacent modular blades being provided at edge thereof with a first connecting portion and a second connecting portion that cooperate with each other, and wherein the plurality of modular blades includes a trailing edge shell, a leading edge shell and a main beam, respectively, the first connecting portion at the edge of the trailing edge shell and the leading edge shell being fixedly connected to the second connecting portion at the edge of the main beam. The wind turbine blade according to the present invention adopts segmented modular structural design, which effectively reduces the mold occupancy time of the blade, shortening the production cycle and improving the molding efficiency.

Abstract: The invention relates to a bypass turbomachine (2) for an aircraft, comprising a gas generator (5) and a ducted fan (4) comprising variable pitch blades (18) configured to take a reverse thrust position driving a reverse flow (24) of air within a secondary duct (16), the gas generator (5) being connected to a fan casing (3) by a stator blade assembly (40) that passes through the secondary duct, first openings (28) for letting in air from the reverse flow being located on an outer casing (17) at least partially internally delimiting the outer duct, and second openings (29) for letting said air out being located on an inner casing (14) at least partially externally delimiting an inner duct (12). The first openings are located within a plane that is perpendicular to a longitudinal axis (C) of the turbomachine and passes substantially through the middle of the blade assembly.

Abstract: A passive auxiliary lubrication system for supplying lubricant to a gearbox assembly of an engine. The passive auxiliary lubrication system includes an auxiliary tank for storing lubricant. The auxiliary tank is passively pressurized. The passive auxiliary lubrication system also includes an auxiliary tank feed line fluidly coupled with the auxiliary tank and a primary lubrication system. The lubricant is supplied to the auxiliary tank from the primary lubrication system through the auxiliary tank feed line and the lubricant is passively pressurized in the auxiliary tank. The passive auxiliary lubrication system further includes an auxiliary supply line fluidly coupled with the auxiliary tank and the gearbox assembly. The lubricant in the auxiliary tank is supplied to the gearbox assembly through the auxiliary supply line.

Abstract: A gearbox assembly for a gas turbine engine includes a hollow gear. The hollow gear has one or more reservoirs configured to store a lubricant and one or more passages configured to supply the lubricant to one or more gear meshes in the gearbox assembly. The hollow gear provides a dampening system configured to absorb vibrations and loading experienced in the gearbox assembly and configured to provide impact resistance for the gearbox assembly.

Abstract: A gas turbine engine for an aircraft includes an engine core with a turbine, a compressor, and a core shaft connecting the turbine and compressor; a fan upstream of the engine core including a plurality of fan blades; and a gearbox that receives an input from a gearbox input shaft portion of the core shaft and outputs drive to a fan shaft so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox including a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon, and wherein the sun gear receives input from the core shaft. At cruise conditions the product of the torque on the core shaft and radial bending stiffness or tilt stiffness of the gearbox input shaft portion is within a specified range.

Abstract: Provided is a method for repairing a root of a rotor blade of a wind turbine, the root including several bores for receiving a bolt and extending form the root front surface into the blade, wherein at least a part of the front surface blade material surrounding the bore is removed using a removing tool and that at least one shim plate is attached to the processed front surface surrounding the bore.

Abstract: A seal assembly at a rotor-stator interface includes at least one non-contacting seal interface and at least one rub detection feature. The rub detection feature(s) is configured to generate a signal upon the rotor and the stator making contact at the rotor-stator interface and causing wear above a certain threshold at the rotor-stator interface. The seal assembly also includes at least one sensor arranged at the rotor-stator interface. The sensor is configured to sense the signal. The seal assembly further includes a controller communicatively coupled with the sensor(s). The controller is configured to receive the signal and estimate at least one of an amount and a location of the wear at the rotor-stator interface based on the signal.

Abstract: Disclosed is a blade lock assembly, which may include a drag brace coupled to a fold crank assembly and to a rotary wing blade and may further include an actuator assembly configured to be coupled to a rotary wing cuff. The actuator assembly may include an actuator pin housing, an actuation pin at least partially sheathed by the actuator pin housing, and an actuator configured to extend and retract the actuation pin away from and toward the actuator pin housing.

Abstract: Single suction centrifugal pumps, pump casings, pump systems and vehicles for using pumps which include a single piece casing having a first cut-water wall and a second cut-water wall, corresponding joint-water paths communicating with respective discharge nozzles, the second cut-water wall spanning less than 180 degrees from a leading edge of the second cut-water wall to a trailing edge of the second cute-water wall, and in aspects, the discharge nozzles are offset from a horizontal centerline of a center opening of the casing, the cut-water walls define a diameter which is greater than 1.07 times the impeller diameter, and the diameter is constant for a portion of the walls and transitions to a varying diameter at trailing ends of the walls.

Abstract: An aircraft includes a gas turbine engine and an optically-based measurement system. The gas turbine engine is configured to ingest a first mass flow and to exhaust a second mass flow. The optically-based measurement system is configured to determine the first and second mass flows in response to performing an imaging process on the gas turbine engine.

Abstract: An insertion tool is provided for an engine defining an access opening and including a component defining at least in part a cavity. The insertion tool includes: an insertion tool arm having a plurality of segments, the insertion tool arm configured for insertion through the access opening into the cavity and the plurality of segments configured to be in a fixed position relative to one another within the cavity; and a base coupled to the insertion tool arm and configured to be positioned outside the cavity and to move the insertion tool arm along at least two degrees of freedom.

Abstract: Counter-rotating turbine of a turbomachine extending about an axis of rotation and comprising an inner rotor rotating about the axis of rotation, and comprising at least one inner movable blade rotatably supported by a first shaft, an outer rotor rotating about the axis of rotation in a direction opposite to the inner rotor, and comprising at least one outer movable blade rotatably supported by a second shaft coaxial with the first shaft, the first and second shafts extending axially from upstream to downstream of the turbine, wherein the first shaft is guided in rotation by a first bearing disposed between the first shaft and an upstream casing of the turbine, and the second shaft is guided in rotation by a second bearing disposed between the second shaft and said upstream casing of the turbine.

Abstract: An aircraft engine, has: an upstream stator having upstream stator vanes distributed about a central axis; and a downstream stator having downstream stator vanes distributed about the central axis, the downstream stator located downstream of the upstream stator, a number of the upstream stator vanes different than a number of the downstream stator vanes, the downstream stator vanes including: a first vane, a major portion of a leading edge of the first vane circumferentially overlapped by one of the upstream stator vanes; and a second vane differing from the first vane by a geometric parameter, the geometric parameter causing the second vane to have one or more of: a stiffness greater than that of the first vane, and a major portion of a leading edge of the second vane circumferentially overlapped by another one of the upstream stator vanes.

Abstract: A bellcrank for a variable adjustment device for an aircraft turbomachine, is articulated respectively to a control housing, to a synchronization bar, to a piston stem of an actuator and to a turnbuckle. The bellcrank includes a part for articulation to the control housing which includes, on the one hand, a tab and, on the other hand, a tubular portion spaced apart from the tab, the tab and the tubular portion including orifices that are aligned and are configured to receive a common axis of articulation of the control housing.

Abstract: A counter rotating propeller system comprises a structure configured to control the front propeller outer slipstream and prop wash. Specifically disclosed is a structure configured to vary the effective hub-tip radius ratio of the rear propeller.

Abstract: An aircraft engine, has: an upstream stator having upstream stator vanes circumferentially distributed about a central axis; and a downstream stator having downstream stator vanes circumferentially distributed about the central axis, the downstream stator located downstream of the upstream stator relative to an airflow flowing within a core gaspath of the aircraft engine, a number of the upstream stator vanes being different than a number of the downstream stator vanes, major portions of leading edges of the downstream stator vanes circumferentially overlapped by the upstream stator vanes, the downstream stator vanes including: a first pair of circumferentially adjacent vanes of the downstream stator vanes spaced apart by a first pitch, and a second pair of circumferentially adjacent vanes of the downstream stator vanes spaced apart by a second pitch different than the first pitch.

Abstract: The invention relates to a blade pivot with adjustable orientation for a turbomachine fan hub, comprising a stud having means for retaining a fan blade root and coupling means for the transmission of a torsional torque, a rolling bearing for absorbing the centrifugal forces having an inner bushing mounted in transverse abutment against an external radial portion of the stud, a ring for radially retaining the stud relative to the rolling bearing housed in an annular groove formed in the stud, a plurality of clamping parts each mounted in transverse abutment against the retaining ring and in radial abutment against the inner bushing of the rolling bearing, and a clamping nut screwed onto an outer thread of the stud to come into conical abutment against the clamping parts such that said clamping parts ensure a clamping of the inner bushing of the rolling bearing on the stud and a clamping.

Abstract: A component (1, 2) for supporting at least one bearing (3) for a turbine engine (10) comprising: two coaxial walls, internal (4) and external (5) walls respectively, defining a gas flow vein (6) between them and interconnected by a row of arms (7); an external ferrule (50) comprising an internal peripheral edge (51) connected to the external wall (5) and an external peripheral edge (52) connected to an external mounting flange (53); an internal ferrule (40) comprising an external peripheral edge (41) connected to the internal wall (4) and an internal peripheral edge (42) comprising an internal mounting flange (43); at least one of the ferrules (4, 5), which at the peripheral edge (41, 51) thereof is connected to the corresponding wall (4, 5), having a general shape which is corrugated about an axis (X-X) of the component (1, 2).