A blade wheel of a turbomachine, which blade wheel has a multiplicity of blades which are suitable and provided for extending radially in a flow path of the turbomachine, wherein the blades form a blade entry angle and a blade exit angle. Provision is made whereby the blade wheel forms N blocks of blades, where N≥2, wherein the blades of a block have in each case the same blade entry angle and the same blade exit angle, and the blades of at least two mutually adjacent blocks have a different blade entry angle and/or a different blade exit angle. According to a further aspect of the invention, partial gaps that the blades form in relation to an adjacent flow path boundary are varied in mutually adjacent blocks.

Disclosed are a compressor, which is a cantilever type that is easy to manufacture and assemble, and is capable of minimizing the vane tip clearance as the elastic member absorbs the impact and is compressed when the vane collides with the shroud segment due to expansion of the vane, and a gas turbine including the same.

Gas turbine engine (GTE) airfoils, such as rotor and turbofan blades, having multimodal thickness distributions include an airfoil tip, and an airfoil root opposite the airfoil tip in a spanwise direction. The GTE airfoil has a first, second and third locally-thickened region, with the first locally-thickened region defined at the airfoil root. A maximum thickness of each chord between the airfoil root and the airfoil tip transitions toward the leading edge between the first locally-thickened region and the second locally-thickened region, and the third locally-thickened region extends in the spanwise direction. A chord line that extends through the third locally-thickened region contains a first local thickness maxima and a second local thickness maxima interspersed with at least two local thickness minima, and the first local thickness maxima is defined by the third locally-thickened region and is greater than the second local thickness maxima.

Disclosed is a cooled airfoil having a hub end and tip, an airfoil height being defined between the hub end and the tip. The airfoil has a leading edge, trailing edge, suction side and pressure side. The airfoil has a first airfoil height section adjacent the hub end and extending towards the tip, wherein, in a meridional view, the leading edge and trailing edge are straight along the first airfoil height section. The airfoil has a second airfoil height section adjacent the tip and extending towards the hub end, wherein, in a meridional view, the airfoil is concavely shaped at the leading edge and is convexly shaped at the trailing edge along the second airfoil height section. At least one cooling channel has a length principally extending along the airfoil height, extends straight in a first cooling channel length section, and is bent in a second cooling channel length section.

The invention relates to an assembly for a turbomachine extending along an axis (X) and comprising: —a ferrule (32) designed to define a fan duct (5) of a gas stream of the turbomachine, —a fan casing (2) radially surrounding the ferrule (32) and defining with the ferrule (32) the fan duct (5), —a rectifier (6) comprising a plurality of vanes (7) comprising a first vane (7a) and a second vane (7b) adjacent to the first vane (7a), the vanes defining between them a converging flow channel (13) designed to direct and accelerate the stream by means of an inlet section (14a) included in a plane non-perpendicular to the axis of the turbomachine and an outlet section included in a plane (14b) perpendicular to the axis (X) of the turbomachine, the first vane (7a) and the second vane (7b) each having an unducted downstream portion which forms a trailing edge.

A compressor aerofoil rotor blade for a turbine engine includes: a root portion spaced apart from a tip portion by a main body portion. The main body portion is defined by a suction surface wall having a suction surface and a pressure surface wall having a pressure surface. The suction surface wall and the pressure surface wall meet at a leading edge and a trailing edge. The tip portion comprises a shoulder provided on the pressure surface wall. A tip wall extends from the aerofoil leading edge to the aerofoil trailing edge. A transition region of the pressure surface wall tapers from the shoulder in a direction towards the tip wall. The tip wall includes a squealer defined by a first tip wall region which extends from the trailing edge to a winglet.

A flow-straightener vane of a bypass turbomachine includes a plurality of vane sections stacked radially with respect to a longitudinal axis (X) along a stacking line (L) between a root end and a tip end. Each vane section has a pressure-face surface and a suction-face surface extending axially between an upstream leading edge and a downstream trailing edge. Between the leading and trailing edges of each vane section there is formed a profile chord (CA) the length of which is substantially constant between the tip end and the root end, and the stacking line (L) exhibits a curvature in a plane passing more or less through the axis (X) and through the stacking line (L), situated in the vicinity of the tip end and oriented from downstream towards upstream.

A blade wheel of a turbomachine, which blade wheel has a multiplicity of blades which are suitable and provided for extending radially in a flow path of the turbomachine, wherein the blades form a blade entry angle and a blade exit angle. Provision is made whereby the blade wheel forms N blocks of blades, where N≥2, wherein the blades of a block have in each case the same blade entry angle and the same blade exit angle, and the blades of at least two mutually adjacent blocks have a different blade entry angle and/or a different blade exit angle. According to a further aspect of the invention, partial gaps that the blades form in relation to an adjacent flow path boundary are varied in mutually adjacent blocks.

Fairing installations disclosed herein may include a damper for mitigating vibration of a cantilevered fairing disposed in a bypass duct of a gas turbine engine. The bypass duct may include a first shroud radially spaced apart from a second shroud to define a bypass passage between the first and second shrouds. The fairing may be disposed in the bypass passage and cantilevered from the first shroud. The fairing may have a secured end secured to the first shroud and a free end proximate the second shroud. The damper may be engaged with the free end of the fairing to damp movement of the free end of the fairing.

An air turbine starter comprising a housing defining an inlet, an outlet, and a flow path, a turbine having a rotor with circumferentially spaced blades extending into the flow path, a drive shaft operably coupled to and rotating with the rotor, and at least one vane located within the flow path, upstream of the blades. The at least one blade being defined by an acute axial angle and an acute tangential angle.