The present invention relates to a blade (7) comprising: —a composite material structure (17), —a blade root fastening portion (9) further comprising a shoulder (10) extending into the recess from the wall—a base (18) arranged in the recess and comprising a support member configured to abut against the shoulder (10) of the blade root fastening portion (9) and a passage (39) formed in the support member, the sections (23) of the blade root portion (22) of the composite material structure extending through the passage (39), and—a blocking part (19) arranged in the recess between the two sections (23) of the blade root portion (22) such that each section of the blade root portion (23) is pressed against the support member by the blocking part (19).

In some examples, article used as a component for a turbine engine that operates in a high temperature environment. The article may include: a ceramic or ceramic matrix composite (CMC) substrate; and a coating on the ceramic or the CMC substrate, wherein the coating defines an outer surface of the article. The coating includes a plurality of surface features defining channels on the outer surface of the article. The channels are configured to modify a flow of molten Calcia-Magnesia-Alumina Silicate (CMAS) over the outer surface of the coating in a gas flow over the outer surface of the article to reduce accumulation of the molten CMAS on the outer surface of the article.

Turbine (1) comprising a casing, an outer metal shroud (9), an inner metal shroud (5) and an annular distributor (2) having a plurality of CMC ring sectors (20), each sector comprising a mast (6), an inner platform (24), an outer platform (26) and at least one blade (28) having a hollow profile that defines an inner housing (280), the inner and outer platforms each having an opening (245, 265) communicating with said inner housing, and the mast (6) passing through said openings and the inner housing and being secured to said casing and connected to said annular sector. Each blade comprises at least one first radial shoulder (72) projecting axially towards the inside of the blade, and each mast comprises at least one second shoulder (71) projecting axially towards the outside of the mast (6) configured to radially cooperate with a first shoulder (72) and radially press the blade (28) against the mast (6).

The present invention provides an airfoil 110 with the squealer tip cooling system 50 for a turbine blade 100 at the blade tip 113, wherein the squealer tip cooling system 50 comprises a cooling passage 170 arranged within a squealer tip 117, wherein the cooling passage 170 at least partly extends toward a terminal end 74 of the squealer tip 117, and a pocket 172 at a lateral surface 75, 76 of the squealer tip 117, open externally and extending inwardly at least partly across the cooling passage 170. The pocket 172 intersects the cooling passage 170 and the pocket 172 comprises an impingement surface 70 facing the cooling passage 170, on which a cooling medium expelled through the cooling passage 170 impinges before being discharged externally through the pocket 172.

A turbomachine includes a shroud and a hub spaced apart from the shroud to channel an airflow along a direction. The turbomachine includes a plurality of airfoils coupled between the shroud and the hub. At least one airfoil of the plurality of airfoils includes a leading edge spaced apart from a trailing edge in the direction of the airflow and a pressure side opposite a suction side. The turbomachine includes at least one local vortex generator array defined on the suction side so as to extend onto the hub or the shroud. The at least one local vortex generator array is defined downstream of the leading edge.

A component with a region to be cooled having a cooling channel which is arranged and designed so as to cool the region of the component during operation by a fluid flow, wherein the cooling channel is defined by a first channel side facing the region and by a second channel side facing away from the region. The first channel side forms a larger contact surface for the cooling channel than the second channel side. An additive manufacture process can produce the component.

Disclosed is an aircraft turbine engine (10), comprising a gas generator (12) and a fan (14) arranged upstream from the gas generator (12) and configured to generate a gas inlet stream (F), part of which flows into a duct of the gas generator to form a primary stream (36), the turbine engine (10) comprising an electrical machine that is mounted coaxially downstream from the fan (14) and that comprises a rotor (62a) surrounded by a stator (62b) carried by an annular shroud (64), this shroud (64) being surrounded by a casing (40) of the gas generator that defines, with this shroud (64), a section of the flow duct for the primary stream (36), stationary vanes (42, 68) for straightening this primary stream (36) extending into this path.

A rotating machine includes a casing having a hollow shape; a rotator rotatably supported in the casing; a stator blade fixed to an inner peripheral portion of the casing; a rotor blade fixed to an outer peripheral portion of the rotator while being displaced from the stator blade in an axial direction of the rotator; a sealing device disposed between the inner peripheral portion of the casing and a tip of the rotor blade; a swirling flow generation chamber provided along a circumferential direction of the rotator on a downstream side of the sealing device in the casing in a fluid flow direction; and guiding members provided at predetermined intervals in the swirling flow generation chamber in the circumferential direction of the rotator. The guiding members each include a first guiding surface that is inclined in the circumferential direction with respect to the axial direction of the rotator.

Turbine and compressor casing abradable components for turbine engines include abradable surfaces with a zonal system of forward (zone A) and rear or aft sections (zone B) surface features. The zone A surface profile comprises an array pattern of non-directional depression dimples, or upwardly projecting dimples, or both, in the abradable surface. The dimpled forward zone A surface features reduce surface solidity in a controlled manner, to help increase abradability during blade tip rubbing incidents, yet they provide sufficient material to resist incoming hot working fluid erosion of the abradable surface. In addition, the dimples provide generic forward section aerodynamic profiling to the abradable surface, compatible with different blade airfoil-camber profiles. The aft zone B surface features comprise an array pattern of ridges and grooves.

An airfoil assembly for a gas turbine engine is disclosed and includes a platform portion defining a portion of a gas flow path and a root portion for attachment of the turbine airfoil, the root portion including a bottom surface including a bottom area and a plurality of inlets that define a total inlet area as a ratio of the inlet area to the bottom area. An airfoil extends from the platform and including a plurality of cooling air passages in communication with the plurality of inlets.