A compressor for a heating, ventilating, air conditioning, and refrigeration (HVAC&R) unit includes an impeller, a shaft configured to rotate the impeller, and a fastener. The impeller includes an opening and does not include keys, splines, pins, or any combination thereof. The fastener is coupled to an end of the shaft and extends through the opening of the impeller, and the fastener is configured to stretch in an axial direction relative to the shaft via a tensioner during assembly of the compressor.

A turbine rotor blade according to at least one embodiment to be connected to a rotational shaft so as to be rotatable around an axis includes a hub having a hub surface inclined with respect to the axis in a cross-section along the axis; and a plurality of rotor blades disposed on the hub surface. In a throat portion where a blade-to-blade distance between two adjacent rotor blades is smallest, a value (Lt/r) obtained by dividing the blade-to-blade Lt at a given radial position by a distance r from the axis to the radial position is maximum at a position where a dimensionless span length is in a range of 0.2 to 0.65, assuming that the dimensionless span length is 0 at a position of a root end portion on a hub side and is 1 at a position of a tip end portion opposite to the hub side.

A rotor assembly adapted for a gas turbine engine includes a shaft, a wheel, and a retaining nut. The shaft extends along an axis and includes a first tapered surface. The wheel is arranged circumferentially around the shaft and includes a second tapered surface. The retaining nut is fastened to the shaft and applies an axial force to the wheel to couple the wheel with the shaft.

A turbine section for a gas turbine engine according to an example of the present disclosure includes, among other things, a first turbine rotor coupled to a first turbine shaft. The first turbine shaft is rotatable about a longitudinal axis. A second turbine rotor is coupled to a second turbine shaft. The second turbine shaft is rotatable about the longitudinal axis, and the second turbine rotor is axially aft of the first turbine rotor relative to the longitudinal axis. An aft bearing assembly rotatably supports the second turbine shaft. The second turbine rotor includes a disk assembly that carries at least one row of turbine blades. The disk assembly is mechanically attached to the second turbine shaft at an attachment point. The attachment point is axially aft of the aft bearing assembly such that an aft portion of the second turbine shaft is cantilevered from the aft bearing system with respect to the longitudinal axis. The disk assembly includes a metallic material. Each of the turbine blades comprises a ceramic matrix composite (CMC) material.

An impeller attach mechanism for a turbocharger including a stud extending from a central bore of a compressor impeller toward a turbine wheel, the stud having a first threaded region and a second threaded region; a shaft coupled to the turbine wheel and extending toward the compressor impeller, the shaft having a leading portion, the leading portion having a threaded interior configured to engage the second threaded region of the stud; and an insert having an internal portion and an external portion, the internal portion having a threaded external surface to engage the compressor impeller, the internal portion having a threaded internal surface to engage the first threaded region of the stud, the external portion configured to surround the leading portion of the shaft.

An impeller attach mechanism for a turbocharger including a stud extending from a central bore of a compressor impeller toward a turbine wheel, the stud having a first threaded region and a second threaded region; a shaft coupled to the turbine wheel and extending toward the compressor impeller, the shaft having a leading portion, the leading portion having a threaded interior configured to engage the second threaded region of the stud; and an insert having an internal portion and an external portion, the internal portion having a threaded external surface to engage the compressor impeller, the internal portion having a threaded internal surface to engage the first threaded region of the stud, the external portion configured to surround the leading portion of the shaft.

This invention relates to a multi-stage rotor (10). More specifically, the invention relates to a multistage rotor (10) for the compressor stage of a machine that, through a concentric configuration of its innermost (12), outermost (24) and intermediary (16) blade sets co-operative with a reverse flow convoluting ducting arrangement, provides an axially compact, lighter and more easily maintainable compressor rotor for such machine. The multi-stage rotor (10) includes innermost (30), outermost (34) and intermediary (32) duct ports comprising a radial duct spans, as measured between respective diametrically inner and outer duct walls of the duct port, being greater than respective innermost (48), outermost (54) and intermediary (50, 52) radial blade spans of the respective blade sets rotatable at least partially within such duct port. In this manner, a gap is defined between: (i) the at least one diametrical ends of the radial rotating blades ending radially short of the respective radial duct span to form free ends of the blades; and (ii) a stationary part of the respective duct the free ends of the blades sweep neared to; for generating a friction wash between such free ends of the blades and the stationary part of the respective duct.

A sealing element and a rotor of a gas turbine having at least one rotor disc and having an annular rotor component arranged adjacently to the rotor disc and having a plurality of sealing elements arranged distributed around the circumference. The sealing elements are fastened to the rotor disc at least in the axial direction. An inner edge portion of each of the sealing elements is adjacent to a sealing portion of the rotor component. In order to provide a seal between the sealing element and rotor component whilst at the same time enabling a relative axial displacement, a ring seal is arranged in a receiving space formed by the sealing element and rotor component.

The present invention thus proposes an inlet cone for an aircraft turbine engine, comprising a frustoconical body and a tip made from elastically deformable material fixed to an end of smaller diameter of said body, the tip comprising a top configured to be situated on an axis of rotation of the cone and a fastening base for attachment on said end of said body. Said base extends in a connecting plane P. Said connecting plane P is inclined relative to said axis of rotation. Said base has a generally circular or oval shape. According to the invention, said connecting plane P is inclined relative to a transverse plane T perpendicular to said axis of rotation.

An assembly for a turbine engine turbine includes a turbine rotor disc centered on a longitudinal axis and a turbine shaft centered on the longitudinal axis and driven in rotation by the rotor disc. Torque from the rotor disc is transmitted to the shaft, wherein the rotor disc is locked in translation relative to the shaft in the direction of the longitudinal axis by a screwed member on the shaft. Torque from the rotor disc is transmitted from the rotor disc to the screwed member when the torque ceases being transmitted from the rotor disc to the shaft. The screwed member has an unscrewing direction identical to the direction of rotation of the rotor disc in operation.