A turbine blade for a turbomachine having a turbine blade wall and a fluid channel having inlet channel section on the end region leading to the cold side, outlet channel section on the end region leading to the hot side, and central channel section therebetween having a circular cross-section constant along the length. The turbine blade forms an acute angle with the surface of the turbine blade wall over which hot gas flows, and has an intermediate channel section between the inlet and central channel sections, the intermediate channel section having a larger cross-sectional area than the central channel section. The central channel section connects to the intermediate channel section forming a shoulder surface formed on a wall region of the fluid channel and, on the opposing wall region, the intermediate and central channel sections merge with one another in a linear manner with a reduced shoulder height.

A conduit through which hot combustion gases pass in a gas turbine engine. The conduit includes a wall structure having a central axis and defining an inner volume of the conduit for permitting hot combustion gases to pass through the conduit. The wall structure includes a forward end, an aft end axially spaced from the forward end, the aft end defining a combustion gas outlet for the hot combustion gases passing through the conduit, and a plurality of generally radially outwardly extending protuberances formed in the wall structure. The protuberances each include at least one cooling fluid passage formed therethrough for permitting cooling fluid to enter the inner volume. At least one of the protuberances is shaped so as to cause cooling fluid passing through it to diverge in a circumferential direction as it enters into the inner volume.

The present invention is a turbine blade (1) having a hollow blade body (2). This turbine blade (1) is provided with: cooling air holes (5) that penetrate the blade body (2) from an internal wall surface (2e) to an external wall surface (2f) thereof, and are provided with a straight tube portion (5a) that is located on the internal wall surface (2e) side of the blade body (2), and an expanded diameter portion (5b) that is located on the external wall surface (2f) side of the blade body (2); and with a guide groove (6) that is located on an internal wall of the expanded diameter portion (5b) and that guides cooling air (Y) in the expanded diameter portion (5).

An air intake of a turbojet engine nacelle includes an annular structure including an outer fairing defining an aerodynamic outer surface and an inner fairing defining an aerodynamic inner surface. The outer and inner fairings are connected upstream by an air intake lip forming a leading edge. The inner fairing includes an outer skin fixed to the air intake lip by an upstream fixation flange, and the outer skin is configured to be fixed to an outer fan casing by a downstream fixation flange. The air intake includes at least one independent acoustic panel attached to the outer skin and which includes a perforated acoustic skin and a cellular core.

Systems and methods for delivering a buffer fluid to a shaft seal of a gas turbine engine are provided. An exemplary system includes, a buffer fluid source, one or more first conduits providing fluid communication between the buffer fluid source and the shaft seal along a first route, and one or more second conduits providing fluid communication between the buffer fluid source and the shaft seal along a second route different from the first route. A heat exchanger is also disposed along the first route to facilitate heat transfer between buffer fluid in the one or more first conduits and a cooling fluid.

An assembly adapted for use in a gas turbine engine includes a carrier and a blade track segment. The carrier extends at least partway about an axis. The blade track segment is supported by the carrier radially relative to the axis to define a portion of a gas path of the assembly.

A system includes an exhaust collector tunnel (32) configured to mount inside an exhaust collector (30) of a gas turbine (12). The exhaust collector tunnel (32) has a tunnel wall (33) configured to extend around a turbine shaft (17, 19) of the gas turbine (12). The tunnel wall (33) has a variable diameter (98) along at least a portion of a length of the exhaust collector tunnel (32).

An apparatus and method for an engine component for a turbine engine comprising an outer wall having an outer surface and bounding an interior, the outer wall defining a pressure side and a suction side, extending axially between a leading edge and a trailing edge to define a chord-wise direction, and extending radially between a root and a tip to define a span-wise direction, at least one cooling supply conduit provided in the interior, and at least one cooling passage fluidly coupling the at least one cooling supply conduit to the outer surface of the outer wall, the at least one cooling passage comprising an outlet opening onto the outer surface along the leading edge, an inlet fluidly coupled to the at least one cooling supply conduit, and a curved passage defining a curvilinear centerline.

A turbine blade having a gas film cooling structure with a composite irregular groove. The turbine blade has a hollow structure, and a plurality of first grooves which are recessed grooves are provided on an outer surface thereof. A plurality of discrete holes A extending to an inner surface of the turbine blade are provided at the groove bottom of each first groove. The first groove is an irregular groove, and includes at least two portions in a depth direction. A portion having a depth H.sub.1 from the groove bottom of the first groove is a first portion, and the rest thereof is a second portion. At least one side surface of the second portion is formed by expanding in lateral direction from a corresponding side surface of the first portion.

A bearing includes a thrust gas bearing attached to a journal bearing and two or more converging-diverging orifices defined in a surface of at least one of the thrust gas bearing and the journal bearing. The converging-diverging orifices supply at least one pressurized gas to an interior of the bearing. Hydrodynamic lifting grooves are provided on the faces of the thrust gas bearing and the journal bearing and provide improved load capacity and sealing capabilities. Control over the ratios of the pressurized gases provides for additional sealing capabilities and reduced leakage. A metal mesh damper provides increased damping of the gas bearing.