A design method of a center guide pin includes a step of setting a virtual center axis of a casing, a step of acquiring a center position of the center guide pin in a horizontal direction, a step of acquiring, as a first offset amount, an offset amount of the center position of the center guide pin from the virtual center axis of the casing in the horizontal direction, a step of setting a virtual center axis of a diaphragm, a step of acquiring a center position of a groove portion in the horizontal direction, a step of acquiring, as a second offset amount, an offset amount of the center position of the groove portion from the virtual center axis of the diaphragm in the horizontal direction, and a step of designing the center guide pin based on the first offset amount and the second offset amount.

A turbine engine comprising an engine core having at least a compressor section, a combustor section, and a turbine section in axial flow arrangement defining an axial direction and an engine centerline. The turbine engine further having a rotor and a stator, a carriage assembly carried by the stator, and a seal assembly biased toward the rotor.

A turbine engine having an engine core, a first seal structure, a plurality of circumferentially arranged vanes, a plurality of struts, and a second seal structure. The engine core can define a rotor and a stator having a casing. The first seal structure can include a finger seal projecting from the rotor. The plurality of circumferentially arranged vanes can include an uppers band mounted to the casing, and a lower band located radially above the finger seal. At least some of the struts of the plurality of struts can extend through the plurality of circumferentially arranged vanes. The second seal structure can form a seat with the lower band.

A blade assembly for a gas turbine engine includes a rotor, a stator, a seal plate, and a sealing member. The rotor includes a rotor blade and a rotor disc. The rotor disc defines a bucket groove which receives a cooling fluid from a first cavity upstream of the rotor. The sealing member includes a control arm. The sealing member and the rotor define a flow cavity therebetween in fluid communication with an aperture of the seal plate. The flow cavity receives the cooling fluid flowing through the bucket groove and the aperture. The control arm and the seal plate define a gap therebetween fluidly communicating the flow cavity with a second cavity between the stator and the rotor. The control arm deflects at least a portion of the cooling fluid entering the flow cavity.

A component for a gas turbine engine includes a component body formed of ceramic matrix composite lamina and has at least one hook. The at least one hook has an attachment region radially inward of the at least one hook. The attachment region is radially thinner from a hook end of the at least one hook to a remote end, and then becomes radially thicker. A slot is formed through a radial thickness of the at least one hook from the hook end in a remote direction, such that there are two sections of the attachment region. A gas turbine engine is also disclosed.

A component for a gas turbine engine includes a component body formed of ceramic matrix composite lamina and has at least one hook. The at least one hook has an attachment region radially inward of the at least one hook. The attachment region is radially thinner from a hook end of the at least one hook to a remote end, and then becomes radially thicker. A slot is formed through a radial thickness of the at least one hook from the hook end in a remote direction, such that there are two sections of the attachment region. A gas turbine engine is also disclosed.

The invention relates to an assembly for a turbomachine, comprising a stator (1) and a rotor (2) rotatable relative to the stator (1) about an axis, the rotor (2) comprising blades each ×comprising a vane (3) connected to a radially inner platform (5), a block of abradable material (6) extending radially inwardly from the radially inner platform (5) the stator (1) having a shroud comprising an annular area (9), at least one lug (12) extending radially outwardly from said annular area (9), the radially outer end of the lug (12) cooperating with the block of abradable material (6).

A bearing compartment seal for a gas turbine engine includes at least one seal ring defining an axis and having a radially inward facing sealing surface and a seal runner having a support constructed of a first material and an interface portion constructed of a second material. The interface portion includes a radially outward facing surface. A first coefficient of thermal expansion of the second material is at most approximately equal to a second coefficient of thermal expansion of the at least one seal ring.

A compressor stator vane unit includes multiple compressor stator vanes disposed at a certain interval in a circumferential direction; and an annular joint member connected with inner ends of the multiple compressor stator vanes; wherein the annular joint member constitutes an outer diameter side surface of a leakage fluid flow path provided in an inner diameter side of the joint member to communicate a high-pressure space with a low-pressure space respectively located downstream and upstream of the multiple compressor stator vanes in a fluid flow direction, and D/P is set to 0.05≤D/P≤0.2, wherein D is defined as a distance in an axial direction between an upstream end surface of the annular joint member and an upstream edge of the multiple compressor stator vanes in the fluid flow direction and P is defined as a pitch between the adjacent compressor stator vanes in the circumferential direction.

A compressor stator vane unit includes multiple compressor stator vanes disposed at a certain interval in a circumferential direction; and an annular joint member connected with inner ends of the multiple compressor stator vanes; wherein the annular joint member constitutes an outer diameter side surface of a leakage fluid flow path provided in an inner diameter side of the joint member to communicate a high-pressure space with a low-pressure space respectively located downstream and upstream of the multiple compressor stator vanes in a fluid flow direction, and D/P is set to 0.05≤D/P≤0.2, wherein D is defined as a distance in an axial direction between an upstream end surface of the annular joint member and an upstream edge of the multiple compressor stator vanes in the fluid flow direction and P is defined as a pitch between the adjacent compressor stator vanes in the circumferential direction.