A manufacturing method is provided. During this method, a preform component is provided for a turbine engine. The preform component includes a substrate comprising electrically conductive material having an outer coating comprising non-electrically conductive material applied over a surface of the substrate. A preform aperture is formed in the preform component using an electrical discharge machining electrode. The preform aperture includes a meter section of a cooling aperture in the substrate. The preform aperture also includes a pilot hole in the outer coating. A diffuser section of the cooling aperture is formed in at least the outer coating using a second machining process.

A service tube assembly comprises a service tube having a threaded end portion and a ratchet hub spaced from the threaded end portion. The threaded end portion of the service tube is threadably engaged with a mating part. The assembly further comprises a locking member having a cantilever extending from a fixed end held relative to the mating part to a free end. A ratchet pawl is provided at the free end for engagement with the ratchet hub on the service tube.

A cooling structure for a turbine airfoil includes: a lattice structure body formed such that a first rib set and a second rib set arranged in a cooling passage are stacked on each other in a lattice pattern; and lattice communication portions that allow passages formed between ribs of the first rib set to communicate with passages formed between ribs of the second rib set. Each of the first and second rib sets has rib walls each including a pair of ribs that are inclined in directions opposite to each other relative to an imaginary boundary line extending in a movement direction of a cooling medium and that are in contact with each other on the imaginary boundary line. A plurality of lattice communication portions are formed between two lattice communication portions at opposite end portions of each rib that forms the rib wall.

One or more cooling systems for ventilating a turbine and rotary shaft of a gas turbine system is provided. The gas turbine system includes a gas turbine engine and a turbine exhaust collector in separate enclosures. A first cooling system includes an educator that sucks exhaust gas through a diffuser and directs it out of the turbine exhaust collector enclosure based on suction pressure created from the high velocity of exhaust gas. A second cooling system include struts that enable the exhaust gas to flow from the diffusers to a ventilation flow stack. A third cooling system includes exhaust gas sucked from an opening to a top duct based on suction pressure created from the rotation of the rotary shaft disposed about a coupling. A guideway associated with the third cooling system also directs the exhaust gas to flow to the top duct. These cooling systems are designed to increase the efficiency of ventilating the turbine and rotary shaft, prevent misalignments of the rotary shaft, which may result in thermal stresses, and allow the use of the gas turbine systems in higher ambient temperature environments.

An assembly is provided for a turbine engine. This turbine engine assembly includes a static structure, a conduit and a bracket. The static structure includes a port. The conduit extends longitudinally through the port. The bracket couples the conduit to the static structure. The bracket includes a first base mount, a second bade mount, a conduit mount, a first damper and a second damper. The first base mount is attached to the static structure. The second base mount is attached to the static structure. The conduit mount is mechanically coupled with the conduit. The first damper is between the first base mount and the conduit mount. The second damper is between the second base mount and the conduit mount.

A device for holding (101) at least one cooling tube (120) of a turbomachine casing (10) cooling system (100), the holding device including a fixing frame (104), a holding member (160) being configured to hold two cooling tubes (120), and a connection assembly (140) between the holding member (160) and a fixing frame (104), extending on either side of the frame, the connection assembly (160) comprising a connection part (150) extending through an opening (108) of the fixing frame from an outer portion (141) to the inner portion (142) of the connection assembly, the inner portion (142) being disposed between two cooling tubes (120) and secured to the holding member (160) while the outer portion (141) comprises a resilient return member (170) urged in compression towards the fixing frame by the connection part (150).

An assembly for a turbomachine through which an air flow passes, includes a stator including guide vanes extending radially in relation to a longitudinal axis, at least one inter-vane platform extending between the radially outer ends of two circumferentially consecutive guide vanes, each inter-vane platform including an outer surface that faces the axis, a heat exchanger located downstream of the stator in relation to a direction of circulation of the air flow in the turbomachine during operation, this stator including a heat exchange surface extending in the extension of an inter-vane platform. At least one inter-vane platform located in the upstream extension of the heat exchange surface is provided with at least one turbulence generator on its outer surface.

A turbine airfoil includes a body including a wall defining pressure and suction sides, and a leading edge extending between the pressure and suction sides. A cooling circuit inside the wall of the body includes at least one of: a) a suction side to pressure side cooling sub-circuit including a first cooling passage(s) extending from the suction side to the pressure side around the leading edge to a first plenum, and a plurality of first film cooling holes communicating with the first plenum and extending through the wall on the pressure side; and b) a pressure side to suction side cooling sub-circuit including second cooling passage(s) extending from the pressure side to the suction side around the leading edge to a second plenum, and a plurality of second film cooling holes communicating with the second plenum and extending through the wall on the suction side.

A flowpath component for a gas turbine engine includes a body having a leading edge and a trailing edge. A first exterior wall connects the leading edge to the trailing edge and a second exterior wall connects the leading edge to the trailing edge. At least one first internal cooling passage has a first inlet at a first end of the body. At least one second internal cooling passage has a second inlet at a second end of the body. The at least one first internal cooling passage is isolated from the at least one second internal cooling passage.

An assembly is provided for an aircraft propulsion system. This assembly includes a first drain tube, a second drain tube, a container and a gas tube. The container fluidly couples the first drain tube to the second drain tube. The container is configured to receive fluid from the first drain tube. The gas tube is fluidly coupled with the container. The gas tube is configured to direct gas into the container for propelling the fluid received within the container into the second drain tube.