The turbomachine includes a compressor, an inner annular casing, and an outer annular casing. The inner annular casing and the outer annular casing define at least one cavity therebetween. The clearance control system includes a manifold system including at least one conduit disposed within the cavities and configured to channel a flow of cooling fluid between the cavities. The clearance control system also includes an impingement system including a header and at least one plenum configured to channel the flow of cooling fluid to the inner annular casing. The conduits configured to channel the flow of cooling fluid to the impingement system. The clearance control system further includes a channel system including at least one channels configured to channel the flow of cooling fluid to the turbomachine. The channels are configured to control the flow of cooling fluid to the manifold system.

A section of a gas turbine engine includes a rotor blade designed to rotate about an axis. The section also includes a case positioned radially outward from the rotor blade and extending circumferentially about the axis. The section also includes a control ring being annular, positioned radially inward from the case and designed to move radially relative to the case. The section also includes a segmented blade outer air seal (BOAS) including a plurality of BOAS segments each being positioned radially outward from the rotor blade, movably coupled to the control ring, and designed to move circumferentially relative to each other such that a circumferential gap between each of the plurality of BOAS segments changes in size in response to a temperature change in the section of the gas turbine engine.

A system includes a stator assembly including at least one stator airfoil. The system also includes a rotor assembly including at least one rotor airfoil configured to rotate about an axis. The system also includes an actuator coupled to the stator assembly and configured to actuate the stator assembly in an axial direction relative to the rotor assembly, creating an axial movement such that a clearance between the at least one rotor airfoil and the stator assembly varies based on an axial position of the stator assembly.

A method of calibrating a rotor tip clearance arrangement that includes measuring a rotor parameter indicative of rotor efficiency and altering a position of a tip clearance control actuator by an increment to reduce the rotor tip clearance. The method further includes iterating the measuring and altering, and calculating a rate of change of the rotor parameter between pairs of iterations. The method further includes identifying a knee point where the rate of change of the rotor parameter changes and record the corresponding tip clearance control actuator position as a calibrated position. Also included is a rotor tip clearance arrangement calibrated by the method and a method of monitoring deterioration of a rotor tip clearance arrangement.

A device for guiding variable pitch stator vanes of a turbine engine is provided. The device includes a plurality of inner ring angular sectors arranged end-to-end in order to form an inner ring, each inner ring sector having chimneys passing radially through the inner ring sector, a plurality of cylindrical bushings, each being put into place in a chimney of the inner ring from the inside and each serving to receive a guide pivot of a stator vane, a plurality of reconstitution ring angular sectors arranged end-to-end in order to form a reconstitution ring and put into place radially from the inside against the inner ring, and a plurality of blocking elements passing axially through the inner and reconstitution rings in order to assemble the rings together.

A fan assembly for a work vehicle includes a fan shroud defining an opening and a fan. The fan includes a rotor and one or more blades extending outwardly along a radial direction from the rotor. The fan assembly additionally includes a wear ring attached to the fan shroud and extending inwardly along the radial direction. The wear ring is configured to interact with tips of the blades to increase an efficiency of the fan assembly.

A rotor blade tip clearance control method and a rotor blade manufactured using same. The control method includes: coinciding, along an axial direction of an aircraft engine, a center of gravity of a rotor blade with a center of gravity of a rotor wheel disk supporting the rotor blade; rotating the rotor wheel disk to measure a leading edge deformation amount of the rotor blade; measuring a trailing edge deformation amount of the rotor blade; comparing the deformation amounts; and adjusting the center of gravity of the rotor wheel disk until the leading edge deformation amount tends to be approximately equal to the trailing edge deformation amount. The method can effectively improve or even solve the problem of inconsistent radial displacements of a leading edge and a trailing edge during the operation.

An example axial pump for delivering liquid coolant to cool an electronic device comprises a conduit defining a flow path from an inlet of the conduit to an outlet of the conduit and an impeller in the conduit. The impeller is rotatable about an axis of rotation parallel to the flow path. The pump also has a motor stator configured to drive rotation of the impeller body about the axis of rotation. In addition, the pump comprises an adjustment mechanism coupling the impeller to the conduit. The adjustment mechanism is actuatable to cause translation of the impeller relative to the conduit. This may allow for adjustment of a clearance between a blade tip of a blade of the impeller and a wall of the conduit.

A pump having a cutting wheel and a pre-cutter (26) for cutting chippings that are contained in the medium pumped by the pump, the pre-cutter being driven by a shaft portion that projects axially from the cutting wheel and having a plurality of wings (28), the pre-cutter being surrounded by a crest of anvil blocks (32) at which the free ends of the wings (28) move past in a little distance when the pre-cutter rotates, wherein the anvil blocks are configured as counter-blades (32) that are held exchangeably at a blade carrier (34) that surrounds the pre-cutter (26) at least on a part of its periphery.

An aspect includes a system including a high compressor of a gas turbine engine having a ratio of a cold-rotor build clearance to a span between 0.7% and 7%. The cold-rotor build clearance is defined for a plurality of rotor blades of the high compressor with respect to an engine casing assembly interior surface of the high compressor, and the span is defined as a gap between a rotor disk of the high compressor and the engine casing assembly interior surface of the high compressor for at least a last two stages of the high compressor closest to a combustor section of the gas turbine engine. The system also includes at least two bowed rotor management systems for the gas turbine engine to prevent damage to the rotor blades for a bowed rotor condition of the high compressor under a plurality of operating conditions.