A turbomachine turbine blade, includes a platform, a vane, a cooling cavity supplying a plurality of cooling outlets provided along the trailing edge, two radially adjacent cooling outlets defining therebetween a rib. At least one cooling hole is formed in the thickness of at least one rib and/or in the thickness of a portion of the trailing edge fillet located in the axial extension of at least one rib, so as to ensure fluid communication for a cooling flow between the inside and the outside of the blade for cooling the at least one rib.

The present application discloses a blade, comprising a blade tip, a blade root, a leading edge, and a trailing edge, wherein the leading edge and the trailing edge each extend from the blade tip to the blade root; the blade may rotate around a rotation axis, and the rotation axis and a normal plane of the rotation axis perpendicularly intersect at the foot of the perpendicular; a projection of the leading edge on the normal plane along the rotation axis is a first curve, and the first curve has an even number of inflection points. The blade of the present application can reduce noise and improve aerodynamic performance when the blade rotates.

A profile for a trailing edge of an airfoil, the trailing edge including flanges, and the profile is U or V shaped, is made of composite material and includes: sections joined at a first profile end and defining an inner space, and recesses arranged along the two sections towards the inner space and complementary to each other between the two sections; wherein the profile is configured to follow a theoretical aerodynamic contour of the airfoil so that the two sections partially cover the trailing edge and the flanges are housed inside the inner space, and wherein each of the plurality of recesses defines at least two joining points through which the two flange of the trailing edge and the two sections are configured to be joined each other.

A centrifugal blower includes a centrifugal fan and a separation cylinder. The centrifugal fan has a separation plate. The separation cylinder is disposed inward of the blades in the radial direction of the centrifugal fan. The separation plate has an inner end surface extending from the one side to the other side in the axial direction at a position of an inner end in the radial direction. The separation cylinder has a separation cylinder end surface extending from the one side to the other side in the axial direction at a position of an end on the other side in the axial direction. A height of one of the separation cylinder end surface and the inner end surface in the axial direction is larger than a height of the other of the separation cylinder end surface and the inner end surface in the axial direction.

Disclosed in the present application is a blade and an axial flow impeller using same. The blade comprises a blade tip, a blade root, a leading edge, a trailing edge, an upper surface and a lower surface. The upper surface and lower surface are disposed opposite each other; the blade tip, the blade root, the leading edge and the trailing edge surround the upper surface and the lower surface, and connect the upper surface and the lower surface. The blade is rotatable about a rotation axis, the rotation axis being perpendicular to a normal plane. The blade tip comprises a blade tip base part and a blade tip trailing part, the blade tip base part being close to the leading edge, and the blade tip trailing part being close to the trailing edge and being bent upwards relative to the blade tip base part. An angle of attack of a chord of the blade tip trailing part is greater than an angle of attack of a chord of the blade tip base part, wherein the angle of attack is an acute included angle between the chord and the normal plane. The blade of the present application can provide a large air volume, and has higher static pressure and higher efficiency.

The present technique presents a turbomachine component having an airfoil e.g. a vane of a gas turbine. The airfoil wall defines an internal space which includes a first and a second cooling channels having a first and a second impingement inserts, that define a first main and a first peripheral flow channels in the first cooling channel and a second main and a second peripheral flow channels in the second cooling channel, respectively. Impingement jets ejected from the main flow channels via impingement holes of the corresponding impingement inserts are received in the corresponding peripheral flow channels. A channel connecting conduit conducts a flow of the cooling air from the first cooling channel to the second cooling channel. The channel connecting conduit includes an inlet connected to an outlet of the first cooling channel, and an outlet connected to an inlet of the second cooling channel.

A hybrid three-layer system is presented. The hybrid three-layer system includes a two-layer composite system and an additively manufactured third layer comprising a lattice structure. The composite layer system includes a metallic substrate, a structured surface, and a thermal protection system. The structured surface may be additively manufactured onto the metallic substrate and includes structured surface features formed to project above the metallic substrate. Each of the structured surface features are separated from adjacent structured surface features by grooves. The thermal protection coating may be thermally sprayed onto the structured surface and is bonded to each of the structured surface features. The lattice structure is in contact with a surface of the metallic substrate of the composite layer system.

A turbine engine with at least a compressor section, combustor section, turbine section and a set of airfoils. The airfoils include geometric characteristics to create a high contraction ratio (CR), a low blade turning (BT) at a radially inward location the airfoil, a low solidity, or a low aspect ratio (AR).

An airfoil for a turbine engine having a working airflow separated into a cooling airflow and a combustion airflow, the airfoil comprising a wall defining an interior and having an outer surface over which flows the combustion airflow, the outer surface defining a first side and a second side extending between a leading edge and a trailing edge to define a chord-wise direction; at least one cooling conduit located within the interior and fluidly coupled to the cooling airflow. A primary cooling passage having at least one inlet fluidly coupled to the at least one cooling conduit, a primary outlet on the outer surface. A passage connecting the at least one inlet to the primary outlet, the passage separated into a first portion and a second portion. The primary outlet spaced from the trailing edge a predetermined distance.

Airfoils and core assemblies for such airfoils are described. The airfoils include a leading edge cavity defined, in part, by a leading edge interior rib and a trailing edge cavity defined, in part, by a trailing edge interior rib. A plurality of pressure side cavities are defined by pressure side skin cavity walls with at least one pressure side skin cavity wall not extending to the suction side wall. A plurality of suction side cavities are defined by suction side skin cavity walls with at least one suction side skin cavity wall not extending to the pressure side wall. A main body cavity extends between the leading edge interior rib and the trailing edge interior rib and the plurality of side cavities are arranged in a staggered pattern to define the bounds of the main body cavity.