A rotor of a turbomachine includes a plurality of rotor blades, which each have a blade leading edge and a tip, the rotor blades forming a first group and at least one further group of rotor blades of different construction. It is provided that the rotor blades of the first group are nominal rotor blades, the rotor blades of the at least one further group each have a material cutout in the transition from the blade leading edge to the tip, the material cutout not being present in the case of the rotor blades of the first group, and the rotor blades of the first group and of the at least one further group form a non-periodic sequence in a circumferential direction.

A waste-gate valve is for opening and closing a bypass passage which bypasses an exhaust turbine of a turbocharger, and includes: a valve body disposed in the bypass passage; an open-close lever having a first insertion hole into which a valve shaft of the valve body is inserted, and being configured to open and close the bypass passage by moving the valve body; and a washer having a second insertion hole which is positioned closer to a tip of the valve shaft than the first insertion hole and into which the valve shaft is inserted, the washer being fixed to the valve shaft. The washer has a bend portion bended along an outer shape of the open-close lever.

An assembly for a turbine engine includes a combustor wall. The combustor wall includes a shell, a heat shield and an annular land. The heat shield is attached to the shell. The land extends vertically between the shell and the heat shield. The land extends laterally between a land outer surface and an inner surface, which at least partially defines a quench aperture in the combustor wall. A lateral distance between the land outer surface and the inner surface varies around the quench aperture.

An apparatus and method an engine component for a turbine engine comprising an outer wall bounding an interior and defining a pressure side and an opposing suction side, with both sides extending between a leading edge and a trailing edge to define a chord-wise direction, and extending between a root and a tip to define a span-wise direction, at least one cooling passage located within the interior, at least one cooling hole having an inlet fluidly coupled to the cooling passage and an outlet located along the outer wall.

An engine component assembly for impingement cooling. The engine component assembly includes an engine first component having a cooled surface. The engine first component having a flow path on one side of the cooled surface. A second component is a disposed adjacent to the engine first component between the flow path and the engine first component, and has a plurality of openings forming an array through the second component. The cooling flow path passes through the plurality of openings to cool the cooled surface. The second component having a surface facing the cooled surface of the engine first component. A plurality of discrete cooling features that have at least one wall that has a curved cross-section extend from the second component surface into a gap between and toward the cooled surface of the engine first component and defining an array.

Turbine distributor sector for an aircraft turbine engine, including an external annular platform sector and an internal annular platform sector, the sectors being coaxial and being connected together by blade assemblies including inner cavities cooled by gas circulation, the external platform sector including through openings of which radially internal ends open into the inner cavities, wherein the external platform sector includes inner ducts for supplying the cavities with gas, the ducts including air outlets opening into the openings and air inlets opening onto a portion of the external annular surface of the external platform sector.

A turbine blade may comprise a pressure side wall and a suction side wall opposite the pressure side wall. A tip wall may extend between the pressure side wall and the suction side wall and may comprise a first opening. A first dimension of the first opening may be greater than a second dimension of the first opening. The first dimension may be oriented in a direction extending from a leading edge of the turbine blade toward a trailing edge of the turbine blade.

A rotor for an engine is provided. The rotor comprising a rotor base part that has fastening grooves for rotor blades that are arranged in succession around a rotational axis along a circumferential direction, multiple rotor blades that are respectively supported in a form-fit manner inside a corresponding fastening groove by means of a blade root, and at least one securing element for the axial securingwith respect to a rotational axis of at least one of the rotor blades at the rotor base part. The at least one securing element has two edges that are arranged at a radial distance to one another and through which the securing element is supported in a form-fit manner at the rotor base part, on the one hand, and, on the other hand, at the at least one rotor blade.

A component for a gas turbine engine includes a gas path wall having a first surface and a second surface and a cooling hole extending through the gas path wall from the first surface to the second surface. The cooling hole includes an inlet portion having an inlet at the first surface, an outlet portion having an outlet at the second surface, and a transition defined between the inlet and the outlet. The inlet portion converges in a first direction from the inlet to the transition and diverges in a second direction from the inlet to the transition. The outlet portion diverges at least in one of the first and second directions from the transition to the outlet.

The disclosure provides a nozzle adapted for placement within a flowpath of a turbomachine. The nozzle includes an airfoil having a suction side; a pressure side opposing the suction side; a leading edge between the pressure side and the suction side; a trailing edge opposing the leading edge and between the pressure side and the suction side; a root region at a first radial end; a tip region at a second radial end opposite the first radial end; and a midspan region between the root region and the tip region. An endwall is connected with the root region or the tip region of the airfoil along the suction side, the pressure side, the trailing edge, and the leading edge. The trailing edge of the airfoil has an elliptical shape having a nominal profile substantially in accordance with Cartesian coordinate values of X, Y, and Z set forth in TABLE I.