116results 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.

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.

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.

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.

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.

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.

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

A turbofan engine for an aircraft includes a fan, one or more sensors for providing sensor data indicative of an excitation load, and a fan actuation system. The fan has a plurality of fan blades coupled to a fan shaft having one or more fan bearings. The fan actuation system is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The fan actuation system is characterized by a fan actuation system length envelope in a range from 8.5 to 24 and given byNFB×DFTLAXIAL×(RTBNFB).NFB is a number of the fan blades, DFT is a fan tip diameter of the fan blades, RTB is a thrust bearing radius of the radial thrust bearings, and LAXIAL is an axial length from a fan hub tip to the fan bearings.

A rotor of a rotary dynamoelectric machine incudes a magnetically conductive body, having substantially axially running slots distributed around the circumference. A squirrel cage includes electrical conductors which are arranged in the slots. The electrical conductors are electrically contacted at the two end faces of the rotor by short-circuit rings. The magnetically conductive body includes a base body and at least two further additional bodies, which axially adjoin the base body. A first one of the at least two additional bodies directly axially adjoins the end face of the base body, and a second one of the at least two additional bodies and optionally any further additional body axially adjoin the first additional body. The slots have radially exposed slot portions in the axial end regions of the rotor such that the conductors can be moved radially outward.

The invention relates to a turboprop (10) comprising: - a turbine (12) mechanically connected to an input (21) of a reducer (20); - a tail shaft (30) bearing a variable-pitch propeller (31); - a rotary electric machine (46); and - an electric oil pump (40) configured to supply a hydraulic circuit (33) for adjusting the pitch of the propeller (31), wherein the electric oil pump (40) is configured to supply a lubrication circuit (53) for lubricating the bearings (25) of the reducer (20).

A fan hub (19) comprising a central body (17) extending axially between an upstream side (25) and a downstream side (26), the central body (17) having a plurality of radial ports (20) each configured to hold fan blades in place. The fan hub (19) comprises an upstream annular housing (18) fixed to the upstream side (25) via a bolted joint.

A gas turbine engine includes a turbomachine comprising 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 is rotatable about the fan axis. A pitch change mechanism is operable with the plurality of fan blades to control a pitch of the plurality of fan blades. A counterweight is connected to each respective fan blade. A spring mechanism is operably associated with each fan blade of the plurality of fan blades to create a spring twisting moment on the respective fan blade that is counter to a centrifugal twisting moment of the respective fan blade.

The rotor control system includes a main control valve (25) having an inlet (102) for receiving liquid and an outlet (104) for issuing liquid to a rotor pitch change actuator (24). The rotor control unit is configured to control flow of liquid from the inlet (102) to the outlet (104) to modify pitch angle of rotor blades (17). A feathering system (100) with a first conduit (106) connected in fluid communication with the outlet (104) of the main control valve (25), a second conduit (108) for fluid communication with the rotor pitch change actuator (24), and a drain conduit (110). The feathering system (100) has a normal operation mode for supplying liquid from the main control valve (25) to the rotor pitch change actuator (24) by allowing flow through the feathering system (100) from the first conduit (106) to the second conduit (108). The feathering system (100) has a feathering mode for controlling drainage from the rotor pitch change actuator (24) to the drain conduit (110).

An actuation system (1) for adjusting the pitch of aircraft engine fan blades (200) comprising a hydraulic cylinder (5), including an isolation valve (71) adapted to selectively adopt a first and a second configuration, such that: - when the isolation valve (71) is in the first configuration, the hydraulic cylinder (5) is configured to be position-controlled by a control fluid from a main hydraulic circuit (4a); and - when the isolation valve (71) is in the second configuration, the piston of the hydraulic cylinder (5) can move in a first direction of travel by the action of a control fluid from an auxiliary hydraulic circuit (4b), the fan blade orienting itself towards a large pitch position, and the auxiliary hydraulic circuit (4b) not being hydraulically connected with the main hydraulic circuit (4a). Figure for the abstract: Fig. 5

The invention relates to a pitch-change mechanism (70) which comprises a frame (72), a moving part (102) which can be translated along a longitudinal axis (X), and a locking device (160). The locking device (160) comprises a support member (162) which is translatable along the longitudinal axis (X) between an operating position and a locked position, and a screw-nut system (164). The screw-nut system (164) is formed by a system of satellite roller screws (195) comprising a screw (176) which is translatably secured to the support member (162) and rotatably mounted about the longitudinal axis (X), and a nut (178) which is secured to the moving part (102). The screw (176) has an abutment surface (186) which is separated from the frame (72) when the support member (162) is in the operating position and bears against the frame (72) when the support member (162) is in the locked position.

Propulsion assembly (54) for an aircraft, this propulsion assembly (54) comprising: - a gas generator, - a fan (24) located upstream of the gas generator, - a fan casing (26) extending around the fan (24), - a nacelle (58) extending around the fan casing (26) and at least a part of the gas generator, characterized in that the fan casing (26) comprises two axial sections, the upstream axial section comprising two half-shells (46a, 46b) whose circumferential ends are joined, each of these half-shells comprising a part of the external cylindrical wall (28) and a part of the abradable coating (30), a first of these half-shells being integral with the downstream axial section, and a second of these half-shells being removable and fixed at its circumferential ends to the circumferential ends of the first half-shell. Figure for the abbreviation: Figure 4