A blade (1) for a turbomachine, including an impact chamber (2), a single impulse body (3) being situated in the impact chamber, a clearance (S1+S2) between the impulse body and the impact chamber being at least 10 μm and/or at most 1.5 mm in at least one direction.

A blade (1) for a turbomachine, including an impact chamber (2), a single impulse body (3) being situated in the impact chamber, a clearance (S1+S2) between the impulse body and the impact chamber being at least 10 μm and/or at most 1.5 mm in at least one direction.

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.

Identifying effects of geometric variations of physical parts may include measuring one or more surfaces of a physical part in three-dimensions, analyzing measurements of the physical part to determine a geometric variation from a baseline model, modifying an existing computational fluid dynamics (CFD) mesh for the baseline model based on the geometric variation using a mesh metamorphosis algorithm to create a target mesh for the physical part, and analyzing the target mesh using a CFD analysis to determine an effect (e.g., an effect on unsteadiness) in a system caused by the geometric variation.

Identifying effects of geometric variations of physical parts may include measuring one or more surfaces of a physical part in three-dimensions, analyzing measurements of the physical part to determine a geometric variation from a baseline model, modifying an existing computational fluid dynamics (CFD) mesh for the baseline model based on the geometric variation using a mesh metamorphosis algorithm to create a target mesh for the physical part, and analyzing the target mesh using a CFD analysis to determine an effect (e.g., an effect on unsteadiness) in a system caused by the geometric variation.

A turbomachine assembly includes a casing, first and second rotors, and a damper. The first rotor includes a disk and blades and is movable in rotation relative to the casing. The second rotor is movable relative to the casing around a longitudinal axis. The damper damps a movement of the first rotor relative to the second rotor. The damper includes first to third parts. The first part bears on the first rotor in a first area extending over a first angular sector around the longitudinal axis and applies a first centrifugal force on the first rotor. The second part bears on the first rotor in a second area that is smaller than the first angular sector and extends over a second angular sector around the longitudinal axis. The third part bears on the second rotor and applies a second centrifugal force on the second rotor.

A turbomachine assembly includes a casing, first and second rotors, and a damper. The first rotor includes a disk and blades and is movable in rotation relative to the casing. The second rotor is movable relative to the casing around a longitudinal axis. The damper damps a movement of the first rotor relative to the second rotor. The damper includes first to third parts. The first part bears on the first rotor in a first area extending over a first angular sector around the longitudinal axis and applies a first centrifugal force on the first rotor. The second part bears on the first rotor in a second area that is smaller than the first angular sector and extends over a second angular sector around the longitudinal axis. The third part bears on the second rotor and applies a second centrifugal force on the second rotor.

During operation, a bladed rotor disk typically experiences out-of-plane vibration which can result in deterioration and/or cracking at the interface between adjacent shrouds of the turbine blades. In an embodiment, slots are formed at the end of a labyrinth seal segment of each shroud. Preloaded spring strips are inserted through the slots to couple adjacent shrouds while preventing the natural frequency of the turbine blades from drifting to the operating speed range and/or providing vibration damping to the untuned blade mode.

An annular parts assembly for mounting onto a shaft of an aircraft engine is provided. The assembly comprises a first annular body having a surface defining a plurality of pulling features extending from a remainder of the surface, the pulling features circumferentially spaced apart on the surface. The assembly comprises a second annular body defining a balancing ring, the balancing ring concentric with the first annular body, the balancing ring having a plurality of protrusions and circumferential spaces between adjacent ones of the plurality of protrusions, the circumferential spaces accommodating the pulling features such that the pulling features of the first annular body and the protrusions intercalate.

An annular parts assembly for mounting onto a shaft of an aircraft engine is provided. The assembly comprises a first annular body having a surface defining a plurality of pulling features extending from a remainder of the surface, the pulling features circumferentially spaced apart on the surface. The assembly comprises a second annular body defining a balancing ring, the balancing ring concentric with the first annular body, the balancing ring having a plurality of protrusions and circumferential spaces between adjacent ones of the plurality of protrusions, the circumferential spaces accommodating the pulling features such that the pulling features of the first annular body and the protrusions intercalate.