A centrifugal fan impeller structure includes a hub and multiple blades extending from an outer circumference of the hub. Each blade has a fixed end, a free end, an upper end face, a lower end face, a front end face and a rear end face. The front and rear end faces define therebetween a thickness. The thickness ranges from 0.05 mm to 0.15 mm. The are two inclinations between the upper and lower end faces of the blade, whereby the thickness of the blade is minified to increase the number of the blades. By means of the inclinations, the blades are not subject to breakage in the demolding process.

A fiber preform for a turbine engine blade and also a single-piece blade suitable for being formed using such a preform, a rotor wheel, and a turbine engine including such a blade, the fiber preform being obtained by three-dimensional weaving and comprising a first longitudinal segment suitable for forming a blade root (21), a second longitudinal segment extending the first longitudinal segment upwards and suitable for forming an airfoil portion (22), a first transverse segment extending transversely from the junction between the first and second longitudinal segments and suitable for forming a first platform (23), and a first stiffener strip extending downwards from the distal edge of the first transverse portion and suitable for forming a first platform stiffener (25).

A seal segment assembly for a turbomachine, in particular a gas turbine, including a first seal carrier and a second seal carrier, that are adjacently disposed in the circumferential direction, the first seal carrier having a first carrier base and at least one first sealing member that is joined to the first carrier base, and the second seal carrier having a second carrier base and at least one second sealing member that is joined to the second carrier base, the first sealing member and the second sealing member being formed by a plurality of cavities, that are adjacently disposed in the circumferential direction and in the axial direction, in particular evenly spaced, the cavities extending in the radial direction from the particular carrier base. The first carrier base and the second carrier base are intercouplable or are intercoupled in the circumferential direction by a mating connection assembly.

The invention relates to a turbocompressor (1) comprising a compressor housing (2) and a compressor wheel (4) with blades (5). The compressor wheel (4) is rotatably mounted relative to the compressor housing (2) and is arranged such that the exposed upper edges of the blades (5) are spaced from a compressor housing (2) wall (3) facing the blade upper edges across a head gap (7), wherein both the upper edges of the blades (5) as well as the housing wall (3) have at least one recess (11, 13) and at least one elevation (10, 14) over the respective Meridian contour, said recess and elevation interacting locally such that the head gap (7) defines a Z-shaped course in the region of the recesses (11, 13) and the elevations (10, 14) when viewed on a Meridian plane.

A seal assembly between a transition seal structure associated with a downstream end of a transition duct and a vane seal structure associated with an upstream end of a vane structure in a first row vane assembly of a gas turbine engine includes a seal structure. The seal structure includes inner and outer seal members, each having a radially extending first leg and an axially extending second leg that provide each seal member with an L or V-shape. The seal members are arranged in a nested relationship with one of the seal members being positioned between the first and second legs of the other seal member. The first and/or second legs of the inner and outer seal members is/are received in a corresponding slot defined at least in part by one of the transition seal structure and the vane seal structure.

The present disclosure is directed to a system and method for repairing an abradable material coated on a casing of a gas turbine engine. The system includes an articulating guide configured to fit into an access port of the gas turbine engine. Further, the articulating guide has a proximal end and a distal end. The system also includes a repair tool configured at a distal end of the articulating guide. The repair tool includes a body having a proximal end and a shaped distal end, with the shaped distal end extending away from the body. Thus, the shaped distal body is configured to trench out an area of the abradable material comprising a defect. The system also includes a filler material for filling the trenched out area.

A platform for use between adjacent propulsion rotor airfoils joined to a rotor disk to provide an inner flowpath boundary includes: an axially extending I-beam supporting a radially outer skin having a flowpath surface; the I-beam including an inner I-flange disposed at an inner edge of an axially extending I-web, and an outer I-flange disposed at an outer edge of the I-web; the I-beam including a laterally-extending forward end flange at a forward end of the I-web, and a laterally-extending aft end flange at an aft end of the I-web; and the radially outer skin disposed on top of and joined to the radially outer I-flange such that the forward end flange and the aft end flange abut the outer skin.

A turbine blade that includes an airfoil defined by a concave shaped pressure side outer wall and a convex shaped suction side outer wall that connect along leading and trailing edges and, therebetween, form a radially extending chamber for receiving the flow of a coolant. The turbine blade further may include a rib configuration that partitions the chamber of the airfoil into radially extending flow passages. A first flow passage may include a first side on which turbulators are positioned, wherein each of the turbulators comprises a canted configuration.

A turbine wheel in a turbine engine is provided. The wheel includes a disk having a plurality of blades with roots that are axially engaged and radially retained in longitudinal slots in the outer periphery of the disk, the blades having platforms that extend circumferentially end to end and that are radially facing longitudinal teeth of the disk that define the slots. Protection members for protecting the flanks of the slots and made out of sheet metal are mounted on the teeth of the disk and held thereon by co-operating shapes, the protection members bear radially against the platforms of the blades in order to oppose circumferential tilting of the blades, and at least partially close the radial gaps between the teeth and the platforms of the blades.

A thermal shielding arrangement for a turbine probe comprises a heat shield having first and second mating portions axially engaged in overlapping relationship around a probe extending through an air cavity between an exhaust case and a turbine housing. The first mating portion is provided on a radially outer surface of the turbine housing. The second mating portion projects radially inwardly from a radially inner surface of the exhaust case.