A platform for a fan of a gas turbine engine includes a body portion, a forward flange at a front of the body portion that includes an engagement feature that interacts with a forward flange of a lug of a hub of a fan of a gas turbine engine, and a pin integral with a rear portion of the body portion. The pin is receivable in an aperture in a rear flange of the lug.

A rotor has formed therein a groove with which a moving blade including a dovetail portion and a platform portion connecting the dovetail portion and a blade portion engages. The groove opens in a surface intersecting the axis of rotation of the rotor, and includes a contacting portion extending obliquely with respect to the axis of rotation and coming into contact with the dovetail portion, a bottom portion of an end portion on the radially inner side of the rotor, a joining portion between the contacting portion and the bottom portion, a platform facing portion facing the platform portion, and a chamfered portion formed on an end surface in the groove-extending direction. The chamfered dimension of the joining portion is greater on the side with an acute angle between the groove and the end surface than on the side with an obtuse angle between the groove and the end surface.

The invention proposes a fan, in particular for a turbomachine of small size such as a jet engine, having a hub ratio which corresponds to the ratio of the diameter of the inner limit of the incoming air stream at the radially inner ends of the leading edges of the fan blades, divided by the diameter of the circle around which the outer ends of the fan blades pass, having a value of between 0.20 and 0.265.

A gas turbine engine for an aircraft, includes: an engine core having a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan assembly located upstream of the engine core; and a gearbox receiving an input from the core shaft and outputs drive to the fan assembly so as to drive the fan assembly at a lower rotational speed than the core shaft, wherein the fan assembly has fan blades mounted around a hub, the fan blades having blade tips defining an outer diameter of the fan assembly of from around 220 cm to around 400 cm, the hub having slots located around a rim of the hub, each slot receiving a root of a corresponding fan blade, wherein a ratio of a mass of the hub to a total mass of the fan blades is within the range of around 0.45 to around 0.7.

A vibration-damping system for turbomachine blade(s) in a turbomachine is provided. Each turbomachine blade includes a platform having a radially inner surface. The system includes a spacer axially adjacent the platform of at least one turbomachine blade in the turbomachine, e.g., adjacent a turbomachine blade stage. The spacer includes a body. A damping element coupler is on the body of the spacer, and a vibration-damping element is configured to couple to the damping element coupler. The vibration-damping element and the damping element coupler are disposed to cause the vibration-damping element to engage the radially inner surface of the platform of the turbomachine blade(s) to dampen vibration of the turbomachine blade(s), e.g., during operation of the turbomachine.

A holding device is used with a centripetal air sampling member of a rotor assembly that includes and downstream rotor discs and a centripetal air sampling member. The holding device has an annular support element with a housing configured to receive the air sampling member in a radial orientation, a first lateral extension, and a second lateral extension rigidly attached to the downstream rotor disc, extending substantially following the longitudinal axis and radially arranged externally with respect to the first extension. The holding device also includes a wedging ring axially arranged between a portion of the downstream rotor disc and the support element, the wedging ring being configured to simultaneously urge, under the action of a centrifugal force, the first extension in radial abutment against the second extension and the support element in radial abutment against a portion of the upstream rotor disc.

A compressor rotor disk for a gas turbine includes a circumferential surface; and a plurality of dovetail slots formed along the circumferential surface, each dovetail slot comprising a cutout portion formed by removing a portion of a radially outer convex surface of the dovetail slot. The compressor rotor disk can reduce stresses generated in the compressor rotor disk by centrifugal force, by reducing the centrifugal force of the compressor rotor disk through the formation of the cutout portion at the radially outer portion of the compressor rotor disk, thereby enhancing the durability and reliability of the compressor rotor disk.

A vacuum pump comprises: a pump rotor portion configured such that multiple rotor blade stages including stacking portions formed in a ring shape and turbine blades radially formed on an outer peripheral side of the stacking portions are stacked on each other; a rotor shaft to which inner peripheral surfaces of the stacking portions of the multiple rotor blade stages stacked on each other are fixed; and a communication path allowing communication between a clearance space between the rotor shaft and each stacking portion and a pump exhaust path in which the turbine blades are arranged and discharging gas in the clearance space through the pump exhaust path.

The invention relates to a turbomachine assembly (1) comprising: a first rotor module (2) comprising a first blade (20), a second rotor module (3), connected to the first rotor module (2), and comprising a second blade of smaller length than the first blade (20), and a damping device (4) comprising a first radial external surfaces (41) supported with friction against the first module (2), as well as a second radial external surface (42) supported with friction against the second module (3), so as to couple the modules (2, 3) for the purpose of damping their respective vibrational movements during operation.

A fixture assembly including a first fixture portion; a second fixture portion that interfaces with the first fixture portion; and a bladder assembly mounted to the second fixture portion to face the first fixture portion. A method of manufacturing a fan blade includes inserting a blade body and a cover into a fixture; and deploying a bladder assembly within the fixture to press the cover into the blade body.