A tool for performing inspection and/or maintenance operations on an engine defines a longitudinal direction and a tangential direction. The tool includes a base extending along the longitudinal direction and including a body, a first extension member extending from the body in the tangential direction at a first location, and a second extension member extending from the body in the tangential direction at a second location. The second location is spaced from the first location along the longitudinal direction. The tool also includes a pivot member rotatably coupled to the base and moveable between an insertion position in which the pivot member is oriented generally along the longitudinal direction and a deployed position in which the pivot member is oriented away from the longitudinal direction.

A tip shroud, comprising a plurality of tip shoes encircling a rotor assembly, in a turbine may deform due to thermal gradients experienced during operation of the turbine. This can make it difficult to remove the tip shroud during disassembly of the turbine. In an embodiment, to facilitate consistent and reliable removal of the tip shroud during each disassembly of the turbine, one or more, including potentially all, of the tip shoes of a tip shroud may be provided with one or more radially protruding puller hooks. Each puller hook enables an axial force to be transferred by one or more tools to an axially inner surface of the puller hook, to thereby produce axial movement of the tip shoe out of the tip shroud.

A hydrostatic advanced low leakage seal configured to be disposed between relatively rotatable components. The seal includes a base. The seal also includes a shoe extending circumferentially. The seal further includes a radially outer beam operatively coupling the shoe to the base. The seal yet further includes a radially inner beam operatively coupling the shoe to the base, wherein one of the radially inner beam and the radially outer beam is oriented to be angled relative to the other of the radially inner and outer beam.

A blade includes an airfoil and a root having diverging walls. The diverging walls are made of a ceramic matrix composite material. A reinforcement element is provided between the diverging walls.

A method of assembling a gas turbine engine comprising the steps of providing a casing having an insertion aperture in its outer surface. A guide vane is inserted through the insertion aperture. The guide vane is secured to the outer surface of the casing such that the guide vane can be serviced from an outer part of the casing.

A turbine blade includes: an airfoil body; a cooling passage extending along a blade height direction inside the airfoil body; and a plurality of turbulators disposed on an inner wall surface of the cooling passage and arranged along the cooling passage. The airfoil body has a first end portion and a second end portion which are opposite end portions in the blade height direction. A passage width of the cooling passage in a suction-pressure direction of the airfoil body at the second end portion is greater than a passage width of the cooling passage at the first end portion. A height of the plurality of turbulators increases from a first end portion side to a second end portion side in the blade height direction.

A turbine vane of a turbine engine is described. The turbine vane includes a blade and a root. The root includes a stilt having lateral flanks with a curvilinear profile. The stilt includes a frangible zone suitable for undergoing a breakage of the stilt if radial forces higher than a threshold are exerted on the vane, in particular centrifugal forces during an overspeed state of the turbine. The frangible zone includes at least one oblong frangibility recess formed on at least one of the lateral flanks of the stilt, the oblong recess extending in an axial direction of the stilt along a longitudinal axis parallel to or included in a minimum cross-sectional plane which contains a minimum cross-section of the stilt.

A nozzle segment for a gas turbine engine comprises an outer band having a cooling air inlet, an inner band having a first cooling air outlet, and a nozzle airfoil comprising a cooling flow passage arranged to receive the cooling air as a cooling air stream. A first channel and a second channel are arranged within the cooling flow passage. A deflector is arranged to divide the cooling air stream into a first cooling air stream in the first channel and a second cooling air stream in the second channel, respectively. The deflector deflects the first cooling air stream obliquely to a suction sidewall in the first channel, wherein the first channel is configured to transport the first cooling air stream along the first channel in a swirly flow.

A turbomachine airfoil element includes an airfoil having pressure and suction sides spaced apart from one another in a thickness direction and joined to one another at leading and trailing edges. The airfoil extends in a radial direction a span that is in a range of 0.83-0.93 inch (21.1-23.6 mm). A chord length extends in a chordwise direction from the leading edge to the trailing edge at 50% span and is in a range of 0.80-0.90 inch (20.4-22.9 mm). The airfoil element includes at least two of a first mode with a frequency is 3135±10% Hz, a second mode with a frequency is 6634±10% Hz, a third mode with a frequency is 13260±10% Hz, a fourth mode with a frequency is 16838±10% Hz and a fifth mode with a frequency is 17980±10% Hz.

A gas turbine engine is provided including a turbine section including a turbine having a plurality of first speed turbine rotor blades; a compressor section including a compressor having a plurality of first speed compressor rotor blades and a plurality of second speed compressor rotor blades; a gearbox; and a first spool rotatable by the plurality of first speed turbine rotor blades, the first spool coupled to the plurality of first speed compressor rotor blades for driving the plurality of first speed compressor rotor blades in a first direction and to the plurality of second speed compressor rotor blades across the gearbox for driving the plurality of second speed compressor rotor blades in a second direction, opposite the first direction.