A turbine shroud assembly adapted for use in a gas turbine engine includes a blade track segment and a carrier. The blade track segment includes a runner that extends circumferentially partway around a central axis and an attachment post that extends radially from the runner. The carrier is coupled with the attachment post to support the blade track segment relative to the central axis.

A system includes a foreign object damage (FOD) screen configured to be disposed upstream of an air intake of a gas turbine engine and to keep debris from entering the air intake. The FOD screen is configured to extend across a fluid flow path extending through the air intake into the gas turbine engine. The FOD screen includes a flexible, woven fabric made of a non-metal material and configured to absorb and dissipate energy from the debris, and the flexible, woven fabric includes a tensile strength ranging between 2700 megapascals (mPa) and 3700 mPa.

A steam valve includes a valve body having a valve seat in a steam channel, a valve stem configured to close the steam channel by coming in contact with the valve seat and to open the steam channel by moving from the valve seat and a guide portion configured to slidably guide the valve stem in a moving direction thereof. The guide portion has a guide body that encloses the valve stem from a radial outer side of the valve stem and a guide bush that is detachably fixed inside the guide body, has an inner surface which abuts and slides on an outer circumferential surface of the valve stem, and is formed of a material having corrosion resistance.

A turbine intermediate casing for a gas turbine comprising outer and inner wall elements arranged in series in circumferential direction and delimiting a flow passage for exhaust gas; strut fairing elements arranged in radial direction between respective outer and inner wall elements; an outer casing encompassing the outer and inner wall elements; and an annular mounting structure encompassing a bearing region of a shaft of the gas turbine. The strut fairing elements comprise a space for accommodating carrier elements which extend from the annular mounting structure through the flow passage to the outer casing exhaust gas flowing through the flow passage being directed by the strut fairing elements around the carrier elements. The outer and inner wall and strut fairing elements comprise ceramic fiber composite materials, the outer and inner wall elements being connected to the outer casing and the annular mounting structure respectively.

A combined fan for cooling a heat source includes a hub and multiple blades. The hub has a first side and a second side opposite to each other. The blades are distributed along an outer circumferential surface of the hub, and each of the blades includes a first part extending from the hub and a second part connected to the first part. Each of the second parts protrudes beyond the first side or the second side, and each of the second parts is formed by multiple strip structures. Therefore, each of the strip structures can be closer to a surface of the heat source to dissipate the heat more effectively.

A gas turbine engine including a drive shaft that drives a propulsor. A frame which supports the drive shaft is a K-frame bearing support. A gear system is connected to the drive shaft. The gear system includes a gear mesh that defines a gear mesh lateral stiffness and a gear mesh transverse. A flexible support supports the gear system that defines a flexible support transverse stiffness and a flexible support lateral stiffness. The flexible support lateral stiffness is less than 8% of the gear mesh lateral stiffness.

A gas turbine engine for an aircraft including: an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan including a plurality of fan blades; a gearbox that can receive an input from the core shaft, and can output drive to a fan shaft via an output of the gearbox so as to drive the fan at a lower rotational speed than the core shaft; and a fan shaft mounting structure arranged to mount the fan shaft within the engine, the fan shaft mounting structure including at least two supporting bearings connected to the fan shaft. A fan-gearbox axial distance is defined as the axial distance between the output of the gearbox and the fan axial centreline, the fan-gearbox axial distance being greater than or equal to 0.35 m.

A thrust reverser system for a turbine engine includes a support structure, a transcowl, a door, a lock, and a first elastic element. The transcowl is mounted on the support structure and is translatable between a stowed position, a deployed position, and an over-stow position. The door is pivotally coupled to the support structure and is rotatable between at least a first position, a second position, and a third position. The lock is movable between a locked position, to prevent transcowl translation toward the deployed position, and an unlocked position, to allow transcowl translation toward the deployed position. The lock is only able to move to the unlocked position when the transcowl is in the over-stow position. The first elastic element is disposed within the stowed position aperture and, when engaging both the support structure and the transcowl, supplies a force to the transcowl.

The present invention describes a ring fastener (10) comprising a discontinuous band (11) having a first end (14) and a second end (16). The first end (14) comprises a pair of spaced-apart elongate finger members (15a,15b) (having a first finger member (15a) and a second finger member (15b)), and the second end (16) comprises an elongate tongue member (17). The finger members (15a,15b) and the tongue member (17) are interlockable, such that when interlocked, the band (11) comprises a continuous circumference. The ring fastener (10) further comprises a closed configuration wherein the finger members (15a,15b) and the tongue member (17) are substantially aligned with a circumferential surface (13) of the band (11); and an open configuration wherein the first finger member (15a) is radially and/or axially deflected from its substantially aligned position, so that it is angled with respect to the circumferential surface (13) of the band (11). The invention also describes a method of assembling a bore (40) and a component (42), and the use of such an assembly.

A damping device for being situated between a housing wall of a housing of a thermal gas turbine and a casing ring is provided. The casing ring includes an area radially internal with regard to a rotation axis of a rotor of the thermal gas turbine and facing rotating moving blades of the gas turbine. The damping device includes at least sectionally a porous damping structure. A method for manufacturing this type of damping device as well as to a thermal gas turbine, in particular an aircraft engine, in which this type of damping device is situated in a housing of the gas turbine between a housing wall and a casing ring are also provided.