217results about "Rotors" patented technology

A wind energy plant with a nacelle, a rotor, which features at least one rotor blade adjustable in its blade pitch angle, a central control device for controlling the wind energy plant and a control unit disposed in the rotor for controlling the blade pitch angle of the at least one rotor blade, wherein the central control device and the control unit in the rotor can exchange data with each other via a data link, which comprises at least one first sending and receiving device at the nacelle side and at least one second sending and receiving device at the rotor side, wherein a wireless network connection with a safety-oriented communication protocol is provided between the at least one first sending and receiving device and the at least one second sending and receiving device, and that one monitoring- and communication device at a time is associated to the first and / or the second sending and receiving device, which can monitor the function of the sending and receiving device and initiate a predetermined action in the case of an error.

An assembly for an aircraft propulsion system includes an open propulsor rotor and an open guide vane structure. The open propulsor rotor is rotatable about a rotational axis. The open guide vane structure is axially next to and downstream of the open propulsor rotor. The open guide vane structure includes a plurality of open guide vanes arranged circumferentially about the rotational axis. The open guide vanes include a first open guide vane. The first open guide vane includes a camber line, a leading edge, a trailing edge, a leading edge section and a trailing edge section. The first guide vane extends longitudinally along the camber line from the leading edge to the trailing edge. The leading edge section is configured to pivot about a pivot axis. The trailing edge section is configured to move relative to the leading edge section to change a camber of the first open guide vane.

An engine system for an aircraft includes a first turbofan engine and a second turbofan engine. The first turbofan engine includes a first low-pressure shaft, a first fan having a first fan shaft, and a counterclockwise gearbox assembly. The first fan shaft is drivingly coupled to the first low-pressure shaft through the counterclockwise gearbox assembly. The first low-pressure shaft rotates in a counterclockwise direction. The first fan shaft rotates in the counterclockwise direction such that the first fan rotates in the counterclockwise direction. The second turbofan engine includes a second low-pressure shaft, a second fan having a second fan shaft, and a clockwise gearbox assembly. The second fan shaft is drivingly coupled to the second low-pressure shaft through the clockwise gearbox assembly. The second low-pressure shaft rotates in the counterclockwise direction. The second fan shaft rotates in a clockwise direction such that the second fan rotates in the clockwise direction.

A gas turbine engine defining a centerline and a circumferential direction, the gas turbine engine comprising: a turbomachine comprising, in serial flow order, a compressor section, a combustion section, and a turbine section, the turbomachine defining a working gas flow path and a fan duct flow path; a primary fan driven by the turbomachine, the primary fan defining a primary fan tip radius R1 and a primary fan hub radius R2; a secondary fan downstream of the primary fan and driven by the turbomachine, at least a portion of airflow from the primary fan being configured to bypass the secondary fan, the secondary fan defining a secondary fan tip radius R3 and a secondary fan hub radius R4, wherein the secondary fan is configured to provide, during operation, a fan duct airflow through the fan duct flow path to generate a fan duct thrust, wherein the fan duct thrust is equal to a %Fn of a total engine thrust during operation of the gas turbine engine at rated speed during standard day operating conditions 3S ; wherein a ratio of R1 to R3 is equal to wherein EFP is between 1.5 and 11, wherein RqR Prim.‑Fan is a ratio of R2 to R1, and wherein RqR Sec.‑Fan is a ratio of R4 to R3.

Variable pitch fan blade (8a) for an unfaired turbomachine fan, the blade (8a) comprising an aerodynamic blade (9) and a base (10), the blade (9) comprising a recess opening simultaneously onto the upper and lower surfaces (14) and a leading edge (15) of the blade (9), the blade (8a) further comprising a measuring device (26) housed in the recess and clipped onto the blade (9), the measuring device (26) comprising a streamlined body (28) whose outer surface (29) is complementary to the outer surface (30) of the blade (9) to ensure aerodynamic continuity, the body (28) carrying at least one measuring instrument (31) configured to measure at least one parameter of an airflow upstream of the blade (9). Figure for the abstract: 2

An actuation system (1) for adjusting the pitch of the fan blades (20) of an aircraft engine, comprising a hydraulic power supply (3), a hydraulic cylinder (5), and a hydraulic circuit including a hydraulic channel equipped with a check valve (43) and a distribution valve (41) adapted to selectively adopt two configurations, such that when the distribution valve (41) is in the first configuration, the hydraulic cylinder (5) is configured to be position-controlled by the hydraulic power supply (3), and when the distribution valve (41) is in the second configuration, the control fluid can only flow from a retraction chamber (5f) to an extension chamber (5c) so that the piston of the hydraulic cylinder (5) moves in the first direction of travel, the fan blades orienting themselves towards a safety position known as the "feathered position." Figure for the abstract: Fig. 5

Electric motor comprising a squirrel-cage rotor in the form of a cage, which electrically conductively carries a plurality of rotor end rings formed as cast parts made of a casting material, wherein at least one reinforcing element made of a different material is cast into each of the cast parts, and wherein the casting material substantially forms the outermost free surface of the rotor end ring.

A gas turbine engine including: a turbomachine having a compressor section, a combustion section, and a turbine section arranged in serial flow order; a fan defining a fan axis and comprising a plurality of fan blades rotatable about the fan axis; and a pitch change mechanism operable with the plurality of fan blades, the pitch change mechanism including a plurality of linkages, the plurality of linkages including a first linkage coupled to a first fan blade of the plurality of fan blades and a second linkage coupled to a second fan blade of the plurality of fan blades; and a non-uniform blade actuator system operable with one or more of the plurality of linkages to control a pitch of the first fan blade relative to a pitch of the second fan blade.

This description concerns a collective and cyclic blade pitch control system (2) for a turbomachine propeller (1), comprising: - a plate assembly (10) movable in translation along the propeller axis (R) so as to modify the collective pitch of the blades (2), the plate assembly (10) also being tiltable with respect to a plane normal to the propeller axis (R) so as to modify the cyclic pitch of the blades (2), - a jack (20) configured to move the plate assembly (10) in translation along the propeller axis (R), and - a tilting system (30) which cooperates mechanically with the plate assembly (10) and is configured to tilt the plate assembly (10) with respect to the plane normal to the propeller axis (R), the actuation of the jack (20) mechanically triggering the tilt of the plate assembly (10) by means of the tilting system (30). Figure for the abbreviation: Fig. 4

Turbomachine module, this module having a longitudinal axis (X) and comprising: a propeller (3) comprising a plurality of variable-pitch blades (30), a speed reducer (34), a pitch change system (50) for the blades (30), a fluid transfer device (94) configured to supply fluid to the hydraulic actuator (52) of the pitch change system (50), this transfer device (94) comprising a stator part (96) engaged in a rotor part (97), and a member (99) for mechanically connecting the stator part (96) of the transfer device (94) to the planet carrier (38) of the speed reducer (34), and for fluidically connecting a fluid inlet port (98a) of the transfer device (94) to fluid pipes (95) configured to be connected to a fluid reservoir (91). Figure for abstract: Figure 4

The invention relates to an assembly comprising a propeller blade (10) and a pitch adjustment system (34) for setting the pitch of the blade (10), wherein the root (14) of the blade (10) has a bulb (32) and the pitch adjustment system (34) for setting the pitch of the blade (10) comprises: - a cup (58) having an upper opening (58c); - a retaining ring (82) extending around the bulb (32), the retaining ring being at least limited in axial movement relative to the cup (58) towards the opening (58c) and having a bearing face (84) to hinder axial movement of the root (14) towards the opening (58c); the pitch adjustment system (34) being characterized in that it further comprises a lower seat (88) by which the root (14) is engaged with the cup (58), the seat (88) and / or the retaining ring (82) being mounted so as to be axially translatable relative to the cup (58) by at least one clamping mechanism (90).

The invention relates to a turboprop (10) comprising a propeller (12), a propeller shaft (13) carrying the propeller (12), the propeller being a variable-pitch propeller having a propeller pitch, a rotating electric machine (19) having at least a first configuration in which it is mechanically coupled to the propeller shaft (13) and at least one oil pump (21, 21B) configured to supply a hydraulic circuit for adjusting the pitch of the propeller (12). The oil pump is configured to be electrically operated. The invention further relates to an aircraft comprising such a turboprop (10) and to the methods for controlling such a turboprop (10) and such an aircraft.

A variable pitch fan for a gas turbine engine includes a hub, a fan blade supported by the hub, the fan blade including a blade root, and a retention system including a bearing and a retainer, the retainer securing the bearing to the blade root. The hub is rotatably supported by the bearing.

A system (34) for controlling the pitch of a propeller blade (10) of an aircraft turbine engine is disclosed, characterized in that the system comprises: - a cup-shaped portion (58) having an annular wall (58a) extending about an axis (A) intended to set the pitch of the blade, the annular wall (58a) having a lower axial end enclosed by a bottom wall (58b) and an open upper axial end configured such that the root (14) of the blade (10) can be mounted within the cup-shaped portion (58); and - a retaining ring (152) extending about the axis (A) and configured to be mounted around the root (14), the retaining ring (152) being a double-claw clutch ring comprising two annular rows of external claw teeth (154, 156), the two annular rows of external claw teeth constituting an active claw tooth and a safety claw tooth, respectively.

The invention relates to an aircraft turbomachine assembly comprising a propeller constituting a non-fan, the turbomachine comprising a cyclic pitch control device (50) for the propeller blades, the cyclic pitch control device (50) comprising at least two actuators (52), each actuator being fixed to a stator (70), the stator (70) being rotationally fixed relative to a bearing support supporting the rotating shaft, the stator (70) comprising a rim (70C) connected to a hub (70D) by a set of spokes (70E), the rim (70C) being fixed to the bearing support, the hub being rotationally mounted about the rotating shaft, at least two sets of two adjacent spokes (70E) being configured for the attachment of an actuator (52) of the variable pitch control device, each actuator (52) being at least partially inserted between the two corresponding adjacent spokes (70E). Figure 3.

Assembly for an aircraft turbomachine propeller (10), this assembly comprising a variable-pitch blade (14), a blade pitch control system (76) comprising a sleeve (80) engaged in a metal body (102) forming the blade root, and a tie rod (120) passing through the sleeve (80) and the metal body (102) to connect the blade to the control system (76) and pre-stress the metal body (102). Figure for the abbreviation: Figure 4

A propeller for an aircraft turbine engine having a hub, vanes, bearings for guiding the roots of the vanes into orifices in the hub, and systems for retaining the roots of the vanes in the orifices in the hub along the pitch axis, this system for retaining the root including a ring configured to be fitted in a groove in the root, and a screw-nut system mounted on the root between the ring and one of the bearings.

An assembly (1) of variable-pitch blades for an unfaired aircraft turbomachine, the assembly comprising a rotor (3) movable about a radial axis (X) of the turbomachine, a stator (4) fixed about said radial axis (X), and at least one lightning arrester (5). The lightning arrester (5) comprises a pin (6), at least one face of said pin (6) being in contact by friction with the rotor (3), a screw (7) radially extending from said pin (6), and a means (8) for connecting the pin (6) and the screw (7), said means (8) comprising a spring interposed between the pin (6) and the screw (7). Figure for the abbreviation: Fig 1

This disclosure relates to a rocker arm assembly for adjusting the angle of engine blades and an engine blade assembly. The rocker arm assembly includes: a rocker arm with a groove at one end of its bottom for receiving a boss of a blade; the groove has a longitudinal first through hole and engages with the boss to rotate the blade; a bushing with a second through hole coaxial with the first through hole, the bushing passing through the first through hole and having a clearance fit with the inner wall of the first through hole; a stud with a first thread at its top and a second thread at its bottom, passing through the second through hole, the second thread being used for threaded connection with a threaded hole in the blade extending longitudinally from the top of the boss; and a nut connected to the first thread of the stud, its bottom surface abutting against the top surface of the bushing. The rocker arm assembly of this disclosure includes a bushing, which improves the accuracy of installation and positioning between components, prevents loosening of fasteners, avoids safety accidents, and thus improves the reliability of torque transmission by the rocker arm.

A fan assembly for a gas turbine engine includes a fan actuation system, a fan blade hub, a plurality of fan blades, and a counterweight system arranged disconnected from the plurality of fan blades. The counterweight system includes a counterweight hub, and a plurality of counterweight levers each having a counterweight trunnion rotationally connected to the counterweight hub, a cantilever arm with a counterweight connected thereto. The fan blades are rotationally connected to the fan blade hub and to the fan actuation system, and the counterweight levers are rotationally connected to the counterweight hub and to the fan actuation system. The fan actuation system is arranged to correspondingly rotate each of the plurality of fan blades about a respective fan blade pitch change axis, and each of the plurality of counterweight levers about a counterweight lever rotational axis in unison.

The invention relates to a device (10) for transferring a plurality of fluid pathways, which device comprises: a central portion (20); a peripheral portion (22) rotating about the central portion; and transfer chambers (24A-24C) between the central portion and the peripheral portion. The central portion defines first fluid paths (FP1A-FP1C) connecting fluid inlets (14A, 14C) to the transfer chambers. The peripheral portion (22) defines second fluid paths (FP2A-FP2C) connecting fluid outlets (16A-16C) to the transfer chambers. The central portion comprises an outer body (210) defining a recess (215), and an inner body (212) comprising a central portion (212A) extending into the recess (215) and a connection flange (224) bearing the fluid inlets (14A to 14C). The outer body and the inner body are prevented from rotating relative to each other by the engagement of radial teeth (226) of the inner body (212) with axial teeth (218) of the outer body (210).

A method for reducing the clearance between multiple rotor blades and a shroud assembly of an engine includes: determining, using a flight control system, that the aircraft is in a first flight condition. The method further includes adjusting the clearance to a first clearance distance associated with the first flight condition. Additionally, the method includes receiving a demand for a second flight condition using the flight control system. During the second flight condition, the method includes adjusting at least two independently controllable parameters, the at least two independently controllable parameters including at least a first parameter for optimizing the clearance and a second parameter for meeting the engine's thrust requirements, the first parameter having a first effect on the clearance and the second parameter having a second effect on the clearance, the first effect being greater than the second effect.

A propeller blade pitch control mechanism configured to adjust and control the blade pitch of the blades of a multi-blade propeller simultaneously. The pitch control mechanism may have a pitch plate which is raised or lowered along a liner path along an axial direction using a linear drive, which may be a power screw. The pitch plate is coupled to the blade roots with kinematic links which transfer the linear axial motion of the pitch plate along the propeller spin axis into rotary motion of the propeller blade roots, thereby controlling the pitch of the propeller blades. The propeller pitch control mechanism may include springs or gas springs adapted to place the propeller blades in a low drag position in the event of failure of the blade pitch control mechanism.

The invention relates to a fluid pumping assembly (40) intended to be arranged in a compartment (12) of a turbomachine (2), the fluid pumping assembly comprising:-a fluid reservoir (42); a pump (44) fluidly coupled to the fluid reservoir by means of a connecting device; characterised in that the connection means comprise:-a normally closed valve configured to remain in an open position when the pump is coupled and to close the reservoir when the pump is separated from said reservoir.

Electric motor with a squirrel-cage rotor in the form of a cage, which electrically carries several short-circuit end rings formed as castings made of a casting material, wherein at least one reinforcement element made of another material is cast into each of the castings, the casting material essentially forming the outermost free surface of the short-circuit end ring.

This pitch change mechanism (70) comprises a frame (72), a movable part (102) movable in translation along a longitudinal axis (X), and a locking device (160) for immobilizing the movable part (102) relative to the frame (72). The locking device (160) comprises a movable member (210) movable in translation relative to a locking member (164) between a retracted position in which it leaves the locking member (164) free to be in an unlocking configuration away from a surface (162) and a deployed position in which it forces the locking member (164) into a locking configuration engaged with the surface (162). It also comprises a return member (220) urging the movable member (210) towards its deployed position, and a holding device (222) for holding the movable member (210) in its retracted position. Figure for abstract: Fig. 4

The invention relates to an aircraft turbomachine module, in particular for a test machine, comprising: - an annular hub (10) comprising at least one blade root housing (12), - at least one blade (11) provided with a blade root (13) intended to be arranged inside the blade root housing (12), and - a system (15) for locking the blade (11) in rotation, comprising: - a split insert (16) having an external face (17) of conical shape and an opening (18) in which the blade root (13) is arranged, - a platform (21) provided with a through hole (22) in which the split insert (16) is inserted, said through hole (22) being delimited by an internal face having a conical shape of a shape complementary to the external face (17) of the split insert (16), and - at least one fixing member (25) capable of fixing the platform (21) on the annular hub (10),so that tightening of the fixing member (25) ensures compression of the split insert (16) between the platform (21) and the blade root (13) so as to lock the blade (11) in a desired angular position. Figure 4,

The application provides a dolphin head profile imitated biomimetic leading edge slot wing, a blade and a preparation method. The biomimetic leading edge slot wing is made of a leading edge slot wing profile based on a dolphin head profile. The leading edge slot wing profile is obtained by tangent transition connection of a biomimetic leading edge profile line and an inner profile line of the leading edge slot wing profile through a smooth transition line. The dolphin head profile is obtained by equal ratio scaling according to the size of a basic blade airfoil. The corresponding part of the basic blade airfoil leading edge and the biomimetic leading edge profile line is the inner profile line of the leading edge slot wing profile. The application is based on a passive control technology. The biomimetic leading edge slot wing is designed according to the characteristics of a dolphin head. The biomimetic leading edge slot wing and the basic blade are directly connected through an adjusting device and a slot wing connecting piece. The aerodynamic performance of the blade is improved without changing the overall shape of the basic blade, so as to inhibit the negative influence of flow separation and dynamic stall on the blade.

Assembly for a turbomachine (1) comprising: - a variable-pitch fan blade (8) of an unfaired fan (2) including an aerodynamic blade and a base for mounting in a recess of a pivot (12); - a measuring device (25) comprising an aerodynamic body adjacent to and upstream of the blade, the body carrying at least one measuring instrument configured to measure at least one parameter of an airflow (F) upstream of the blade, the measuring device (25) being configured to be integral with the blade (8) and / or the pivot (12), such that the measuring device (25) moves in conjunction with the blade (8), so as not to disturb the aerodynamic and mechanical behavior of the blade (8). Figure for the abstract: 1

Variable pitch turbine blade (114) for an aircraft turbomachine, this blade (114) comprising: - a blade (16), - a foot (18) defining a pitch axis (A), and - an eccentric (78) attached to and fixed to the foot (18), characterized in that the eccentric (78) is coupled to the foot (18) via a curvic® coupling (100). Figure for the abbreviation: Figure 4