A nested segment assembly and a method of additively manufacturing the same are provided. In one example aspect, the nested segment assembly includes a first component segment and a segment component segment positioned adjacent the first component segment. The first component segment has an end face and a tongue projecting outwardly from the end face. The second component segment defines a groove. The groove is defined at an end face of the second component segment that is adjacent the end face of the first component segment. The first component segment and the second component segment are additively printed such that at least a portion of the tongue of the first component segment is nested within the groove defined by the second component segment. Segmented components having a tesla valve formed therebetween are also provided.

A multi-flow turbojet engine generally includes an upstream ducted fan driven by a gas generator, the gas generator having a first and a second compressor, which are coaxial, an inlet case forming a support for the rotors of the upstream ducted fan and of the first compressor, an inter-compressor case downstream of the inlet casing and forming a support for the rotors of the second compressor, attachment means for thrust take-up rods, which attachment means are arranged on the inter-compressor case. The turbojet engine includes at least one axial stress transmission rod for connecting the inlet case to the inter-compressor case.

Isolation seals for gas turbine engines are described. The isolation seals include a first interface member configured to be fixedly attached to a first case structure, the first interface member being a full-hoop structure, a housing configured to be mounted to a second case structure, a connector pin arranged within the housing and moveable relative to the housing, and a second interface member located on an end of the connector pin, the second interface member being a full-hoop structure. The first interface member and the second interface member are engageable to form a seal therebetween.

A gas turbine engine component includes an outer diameter endwall, an inner diameter endwall spaced radially inward of the outer diameter endwall, and at least one body supported between the outer and inner endwalls for rotation about an axis. The body includes an outer diameter surface spaced from the outer diameter endwall by a first gap and an inner diameter surface spaced from the inner diameter endwall by a second gap. The outer and inner diameter surfaces and the outer and inner diameter endwalls are configured such that the first and second gaps remain generally constant in size as the body rotates about the axis.

The invention relates to a turbomachine comprising a low-pressure shaft (27) configured to drive a fan shaft (32) in rotation by means of a coupling assembly comprising a first shaft (1) on which are formed a plurality of first axial grooves (7), a second shaft (2) on which are formed a plurality of second axial grooves (12), and a coupling device (4) including a plurality of rolling elements (13) and an annular cage (14) positioned between the first shaft (1) and the second shaft (2), the rolling elements (13) being positioned between one of the first axial grooves (7) and one of the second axial grooves (12) to couple in rotation the first and the second shaft (1, 2), each first and second groove (7, 12) having a first (16) and a second (17) substantially planar surface inclined with respect to one another and extending along the axis.

Damper stacks, rotor blades, and turbomachines are provided. A rotor blade includes a main body including a shank and an airfoil extending radially outwardly from the shank. The rotor blade further includes a damping passage defined in the main body, the damping passage extending radially through the main body. The rotor blade further includes a damper stack disposed within the damping passage, the damper stack including a plurality of damper pins, each of the plurality of damper pins in contact with a neighboring damper pin.

Damper stacks, rotor blades, and turbomachines are provided. A rotor blade includes a main body including a shank and an airfoil extending radially outwardly from the shank. The rotor blade further includes a platform surrounding the main body, the platform comprising a slash face. The rotor blade further includes a damper stack disposed at the slash face and extending generally along an axial direction. The damper stack includes a plurality of damper pins, each of the plurality of damper pins being in contact with a neighboring damper pin.

An aspect of the present disclosure is directed to a seal assembly. The seal assembly includes a fluid bearing wall including a bearing face defining a fluid opening through the fluid bearing wall. The seal assembly further includes a tilt member is extended from and attached to the fluid bearing wall. The tilt member defines a fluid passage therethrough in fluid communication with the fluid opening at the bearing face of the fluid bearing wall. The seal assembly still further includes a seal body surrounding the tilt member. A cavity is defined between the tilt member and the seal body. The seal assembly further includes a spring member coupled to the seal body and the fluid bearing wall. A passage is defined between the fluid bearing wall and the seal body.

A fan blade is provided and comprises a leading edge, an attachment root extending aft of the leading edge, and a trench formed in a surface of the attachment root. An attachment root is also provided. The attachment root comprises a leading edge, a dovetail extending aft of the leading edge, and a trench formed in a surface of the dovetail. A gas turbine engine is also provided. The gas turbine engine comprises a compressor section configured to rotate about an axis, a combustor aft of the compressor section, and a turbine section aft of the compressor section and configured to rotate about the axis. A fan may be disposed forward of the turbine section and include a blade. The blade may have a trench formed in an attachment root.

A guide vane segment for a turbomachine stage that has an inner ring segment (10) and a plurality of guide vanes (20) that are configured on the inner ring segment; a detuning region (V) of the inner ring segment extending in each case from a trailing edge (21) of at least one guide vane (20), in particular of at least 50% of the guide vanes (50), axially toward a leading edge (22) of the guide vane (20) over at most 50% of a width (B) of the inner ring segment (10) and/or at most 5 mm and/or circumferentially on both sides, in each case over at most 25% of a spacing (A) between adjacent guide vanes (20), at and/or in which at least one cavity (112) is configured which contains at least one impulse element (100) with clearance of motion for providing impact contacts.