A turbomachine blade including a body that extends mainly in a plane defined by a main axis B and a longitudinal direction, which is defined by a lower surface wall, an upper surface wall, a leading edge located at a first longitudinal end of the body and a trailing edge located at a second longitudinal end of the body, wherein the body of the blade includes a plurality of first pipes that extend mainly along the direction of the main axis B, for circulation of a gas flow, and a plurality of second pipes that extend mainly along the longitudinal direction, for circulation of a second gas flow.

High performance wedge diffusers utilized within compression systems, such as centrifugal and mixed-flow compression systems employed within gas turbine engines, are provided. In embodiments, the wedge diffuser includes a diffuser flowbody and tapered diffuser vanes, which are contained in the diffuser flowbody and which partition or separate diffuser flow passages or channels extending through the flowbody. The diffuser flow channels include, in turn, flow channel inlets formed in an inner peripheral portion of the diffuser flowbody, flow channel outlets formed in an outer peripheral portion of the diffuser flowbody, and flow channel throats fluidly coupled between the flow channel inlets and the flow channel outlets. The diffuser vanes include a first plurality of vane sidewalls, which transition from linear sidewall geometries to non-linear sidewall geometries at locations between the flow channel inlets and the flow channel outlets.

A blade for a gas turbine engine comprises an aerofoil body having a suction side, a pressure side, and a trailing edge. An internal cooling passageway is provided in the aerofoil body, and an ejection slot in fluid communication with the cooling passage and provided at the trailing edge of the aerofoil body. The ejection slot is defined between a pressure side wall and a suction side wall. Both the suction side wall and the pressure side wall include a mid-section and a trailing edge section adjacent the mid-section, and the thickness of the suction side wall and the pressure side wall reduces to define a taper with a wedge angle less than or equal to 20 degrees.

A rotor for a gas turbine having a rotor disk on which there are a plurality of rotor components distributed around the circumference. The rotor disk has a circumferential securing shoulder with a contact face. Retaining faces come to bear against the contact face, each of the retaining faces have a retaining shoulder of the respective rotor component and are designed with a form that complements the contact face. In order to optimize the bearing stresses between the retaining shoulder and the securing shoulder, the retaining face has a smaller radius than the contact face, namely the retaining radius is at least 0.99 times and at most 0.995 times the contact radius. Also provided is an axially extending aperture in the rotor component, the width of which in the circumferential direction is 25% to 75% of the rotor component width in the circumferential direction.

An impingement cooling structure is provided. The impingement cooling structure includes a flow channel formed between a first wall and a second wall facing the first wall, a plurality of impingement cooling holes disposed in the first wall such that the plurality of impingement cooling holes are spaced apart from each other along the flow channel, and a flow diverter convexly protruding from a surface of the second wall in each space between injection axes of the plurality of impingement cooling holes.

A thrust reverser for an aircraft includes a jetpipe and a thrust reverser door. The thrust reverser door has an aft end, a forward end, and an interior surface facing radially inward. The interior surface defines a continuous profile with a downstream nozzle portion directly adjacent to the aft end, a concave portion directly adjacent to the downstream nozzle portion, and a convex portion directly adjacent to the forward end and adjacent to the concave portion.

A turbine wheel (12) is provided with: a disc (22) provided so as to be capable of rotating about a center axis (C); and a plurality of blades (23) provided to a disc surface (22f) at intervals in a circumferential direction, the blades (23) causing a gas guided in from radially outward leading edges (23f) to be expelled from trailing edges (23r) disposed on one side of each of the blades along the direction of the center axis (C). In each of the blades (23), in an area on one side along the direction of the center axis (C) including the trailing edge (23r), there is provided a concave surface (27) in which a positive-pressure surface (23p) on the rear side in a rotational direction (R) is recessed forward in the rotational direction (R), causing the gas to be dispersed entirely in the radial direction of the blade (23).

A rotating machine includes a plurality of rotors. Each of the rotors includes a rotor bore protruding radially inward from a platform. A ring shaped cover plate is interfaced with each rotor bore via at least one snap. The at least one snap includes a first arm extending from the cover plate and having a convex facing contact surface, and multiple second arms extending axially from each rotor bore and having contact surfaces facing the convex surface.

A turbine component includes a hot wall, a cold wall spaced apart from the hot wall and a conduit defined between the hot wall and the cold wall. A cooling system is defined in the conduit. The cooling system includes a plurality of cooling pins, each including a first end having a first cross-sectional area and a second end having a second cross-sectional area. Each cooling pin includes a body extending between the first end and the second end, with a pin leading edge defined along the body from the first end to the second end. The pin leading edge is defined by a first diameter and a pin trailing edge is defined by a second diameter. At least one first cooling pin has the first end coupled to the hot wall and the second end coupled to the cold wall with a fillet.

A first-stage turbine vane supporting structure reduces chordal leakage between a first-stage turbine vane segment and its supporting ring. The structure includes a first-stage turbine vane segment including an outer platform, an inner platform, and a first-stage turbine vane disposed between the inner and outer platforms; an inner rail protruding from the inner platform in a longitudinal direction of the first-stage turbine vane; a supporting ring for supporting the first-stage turbine vane segment by engaging with the inner rail while facing one surface of the inner rail; a supporting member engaging with the supporting ring to define a U-shaped space bordered by three side surfaces surrounding the inner rail; and a flexible member interposed between the inner rail and one of the three side surfaces of the U-shaped space. A protrusion formed on the inner rail engages with a sealing surface of the supporting ring in an airtight manner.