An endwall assembly disposed at one end of a vane assembly may comprise an endwall spar that includes an cold side, an hot side, a leading edge, a trailing edge, and an axis extending from the leading edge to the trailing edge perpendicular to the leading edge. The endwall assembly may include a coversheet on the hot side of the endwall spar and a cooling channel that includes a cooling fluid inlet disposed in the endwall spar, and a cooling fluid outlet. The endwall assembly may include a structure protruding from the cold side of the endwall spar, wherein the structure is located between the cooling fluid inlet and the cooling fluid outlet along the axis; and a cooling fluid source cavity on the cold side of the wall, the cooling fluid source cavity in fluid communication with the cooling fluid channel via the cooling fluid inlet.

An air inlet into a nacelle of an aircraft turbojet engine having a de-icing device and extends along an axis X, an air stream flowing in the air inlet from upstream to downstream, the inlet comprising an inner wall and an outer wall which are connected by a leading edge, the inner wall having a plurality of air delivery lines, each air delivery line having a plurality of through-holes designed to blow elementary streams from the hot air source in order to de-ice said inner wall, the air delivery lines being parallel to one another in a cylindrical projection plane, each air delivery line having a depth P3 defined along the axis X as well as a length L3 defined along the axis Y in the cylindrical projection plane, two adjacent air delivery lines being spaced apart by a distance D3, each position along the axis Y with no more than one through-hole, the ratio of the distances L3/D3 being between 0.8 and 1.

A turbine blade including an airfoil portion having a leading edge, a trailing edge, and a pressure surface and a suction surface extending between the leading edge and the trailing edge. The airfoil portion internally forming a cooling passage, which includes first and second cooling passages, and a plurality of outflow passages each having one end which opens to a merging portion formed by connecting an end portion of the first cooling passage on a side of the trailing edge and an end portion of the second cooling passage on the side of the trailing edge, and another end which opens to the trailing edge. The first cooling passage and the second cooling passage are divided by a partition member disposed in the airfoil portion. The cooling passage includes pressure side pin fins in the first cooling passage, and suction side pin fins in the second cooling passage.

Airfoil for gas turbine engines are described. The airfoils have internal walls and internal cavities. A leading edge cavity is defined within the airfoil body along a leading edge and a leading edge feed cavity is arranged aft of the leading edge cavity. A bent leading edge rib is arranged between the leading edge cavity and the leading edge feed cavity. A main body cavity is arranged aft of the leading edge feed cavity and defined at least in part by two interior ribs that define a part of the leading edge feed cavity. The main body cavity is fluidly connected to the leading edge feed cavity by an interior fluid connection through the intersection of the two interior ribs. A shield cavity is arranged to thermally shield the leading edge feed cavity from heat pickup along the suction side of the airfoil body.

A method for coating a part having a surface that has cooling fluid openings that adjoin cooling fluid ducts inside the part. A device analyzes the surface of a part having a surface that has cooling fluid openings which adjoin cooling fluid ducts inside the part, the device being usable in the aforementioned method. The disclosed device and/or the disclosed method is used during the manufacturing and/or overhauling of parts of a turbomachine.

A blade of a high-pressure turbine of a turboshaft engine, the blade including an airfoil extending in a spanwise direction, terminating in an apex and having a suction wall and a pressure wall joined by a leading edge and joined by a trailing edge. The blade further includes an internal cooling circuit having only an upstream duct and a central chamber for cooling the blade by circulating air. The upstream duct and the central chamber are separately supplied with air. The upstream duct being dedicated to the cooling of the leading edge and the suction wall, and the central chamber being dedicated to the cooling of the pressure wall and the trailing edge and being provided with bridge elements each connecting the pressure wall and the suction wall.

A wall of a hot gas component includes a hot and a cold-gas sided surface, one film cooling hole extending from an inlet in the cold-gas sided surface to an outlet in the hot-gas sided surface and with a metering section of constant cross-section and a diffuser section extending from the metering section. The diffuser section is bordered by a diffuser bottom and two opposing diffuser side walls, has a leading region, which extends from the metering section to the outlet, lies opposite the diffuser bottom and has a constant cross-section over its entire length corresponding to an elongation of a leading region of the metering section up to the outlet. The diffuser section has two diffuser arms dividing the flow into two subflows, generating delta-vortices, a v-shaped outlet, and a v-shaped outlet opening.

An airfoil includes an airfoil section that has an airfoil wall that defines a leading end, a trailing end, and first and second sides that join the leading end and the trailing end. The first and second sides span in a longitudinal direction between first and second ends, and the airfoil wall circumscribes an internal core cavity. An arced rib extends from the first side to the second side and divides the internal core cavity into a forward cavity and an aft cavity. A cooling passage network is embedded in the airfoil wall aft of the rib and between inner and outer portions of the airfoil wall. The network includes a cooling passage leading edge and a cooling passage trailing edge. The aft core cavity has a central cavity section and a cavity lobe. The cavity lobe projects between the rib and the cooling passage leading edge.

An apparatus and method for an engine component for a turbine engine having a working airflow separated into a cooling airflow and a combustion airflow. The engine component including a wall defining an interior and having an outer surface. A tip wall spanning first and second sides of the wall to close the interior. A tip rail extending from the tip wall and having an inner tip rail surface, which in combination with the tip wall, at least partially bounds a region defining a plenum. A rim formed in at least one of the outer surface and inner tip rail surface.

A turbine blade includes: an airfoil portion extending in a blade height direction and having a pressure surface and a suction surface each of which extends between a leading and trailing edges; a shroud portion disposed on a blade tip side of the airfoil portion; a fillet portion formed by a curved surface and connected to an end portion of the shroud portion on a side of the airfoil portion; at least one first cooling hole extending along the blade height direction within the airfoil portion; at least one cooling cavity disposed at least partially within the shroud portion and communicating with the at least one first cooling hole; and a second cooling hole connected to the at least one cooling cavity and opening to a surface of the shroud portion. The airfoil portion has a reference airfoil in which a maximum blade thickness is minimum at a reference position.