A propeller fan includes a hub rotatable about an axis, and a plurality of blades extending radially outward from an outer peripheral surface of the hub. A ring surrounds the plurality of blades and is connected to each blade tip. A maximum camber location is on a camber line where a distance from a chord line to the camber line in an arc-shaped cross section of each blade about the axis is maximum. An axial height is from a trailing edge to the camber line in a direction along the axis. Each blade has a first portion provided inside in the radial direction of rotation and whose axial height at the maximum camber location is substantially constant, and a second portion provided outside in the radial direction of rotation and whose axial height at the maximum camber location increases toward the blade tip.

A radial foil bearing includes a base foil, a first top foil, a first corrugated foil, a second top foil and a second corrugated foil. The first corrugated foil is arranged radially on top of the first top foil and fixed to the first top foil to form a first curved segment fixed to the base foil and extending in a circumferential direction. The second corrugated foil is arranged radially on top of the second top foil and fixed to the second top foil to form a second curved segment fixed to the base foil and extending in the circumferential direction. The second curved segment is placed after the first curved segment on the base foil so that, when the base foil is rolled up, the first curved segment and the second curved segment form at least a portion of a tubular base.

According to various embodiments, disclosed is a fixed support and gutter system for a ring gear of a turbine engine gear train, the fixed support and gutter system comprising a support structure having an undulating path which forms at least a portion of a gutter structure for capturing oil shed off the gear train.

A compressor for a turbine engine includes an annular row of blades including a first blade with an extrados surface and a second blade with an intrados surface; and a connecting surface joining the intrados surface to the extrados surface, the connecting surface having a bulge of volume VB and a recess of volume VC, the bulge and recess forming a contouring 3D. The volume of the bulge VB is greater than four times the volume VC of the recess, the connecting surface includes a flat or axisymmetric band circumferentially separating the bulge from the recess, and the bottom of the recess being formed against the intrados surface.

An aircraft engine including a core casing defining a core passage, a nacelle located radially outward from and circumferentially around the casing, and a bypass passage radially defined between nacelle and casing. A mixer mounted to the casing, including a wall having a leading edge attached to casing and defining a plurality of lobes that are circumferentially spaced apart. A damper ring surrounds the mixer at an axial location between leading edge and trailing edge of the wall, and extends circumferentially uninterrupted, and extends axially from first side to second side. The first side being spaced axially away from the leading edge, and the damper ring having an annular ring surface extending between the first and second sides and second side. The damper ring is radially mounted to the mixer wall via an annular ring surface to dampen the wall.

An aircraft engine including a core casing extending circumferentially about an axis and defining a core flow passage, a nacelle located radially outward from and around the casing, and a bypass passage defined between the nacelle and the casing. A mixer includes a peripheral wall having a leading edge and a trailing edge, the leading edge attached to the casing. A first mixer portion and a second mixer portion are circumferentially spaced apart and respectively extend away from the leading edge from a first location to a second location. The first and second mixer portions respectively have a first and second stiffness, the first stiffness greater than the second stiffness and greater than a stiffness of an upstream portion of the wall extending from the first mixer portion at the first location toward the leading edge.

An inlet guide vane assembly is provided. The inlet guide vane assembly includes a housing, a rotating shaft rotatably located in the housing, a vane portion connected to the rotating shaft, a bushing portion located between the rotating shaft and the housing, and an elastic portion located between the bushing portion and the rotating shaft.

A fan wheel for a fan is equipped with at least two fan blades with a wavy design. A fan has at least one such fan wheel. A system has at least one fan with such a fan wheel.

A stage for a compressor or a turbine in a turbine engine can include an annular row of airfoils radially extending from corresponding platforms, where each platform can include a fore edge and aft edge and each airfoil can include a leading edge and trailing edge. At least one of the platforms can have a scalloped flow surface including a bulge and a trough.

The invention relates to an axial flow fan (1) comprising a stationary hub (2); comprising an annular stationary frame (3), which surrounds the hub (2) on the outer peripheral side and which is connected to the stationary hub (2) via at least one strut (4, 5, 6); and comprising a fan wheel (8), which is in drive connection with a fan wheel drive (9), which has a rotating fan wheel hub (10), and which has a plurality of fan wheel blades (11), which each comprise a back edge (12), characterized in that the contour of the strut (4, 5, 6) is designed as sine wave shaped, and wherein the back edge (12) is preferably designed as toothed at least in a partial area (13).