A fan case assembly is adapted to extend around blades of a fan rotor included in a gas turbine engine. The fan case assembly includes an annular case that extends around an axis, a fan track liner coupled to the annular case and extending circumferentially at least partway about the axis, and a bolting arrangement that couples the fan track liner to the annular case.

A rotor assembly (30) comprising a rotor (32) having an annular array of rotor blades (34), the rotor mounted to a shaft (38). A phonic wheel (40) coupled to the shaft. A speed sensor (44) axially aligned with the phonic wheel and configured to measure voltage (V), amplitude of the voltage being proportional to clearance (46) between the sensor and phonic wheel. A processor (48) configured to: receive the voltage measurement; derive shaft speed (ω) from the voltage measurement; identify modulation of the voltage amplitude at a frequency which is an integer multiple of the shaft speed; compare voltage amplitude to a threshold; and output a rotor damage signal based on the comparison.

There is described a shaft shear event detection method. The method comprises obtaining a demodulated waveform of a shaft oscillation wave superimposed on a shaft rotational speed signal, comparing the amplitude to an amplitude threshold, detecting oscillation when the amplitude threshold is exceeded for a plurality of samples, and detecting a shaft shear when oscillation continues for a predetermined time limit.

A roller bearing arrangement for a gas turbine engine. The roller bearing arrangement includes a fan shaft, and a stub shaft connected to the fan shaft. The roller bearing arrangement further includes a plurality of roller bearing elements positioned between a first axial bearing surface created on a radially outer surface of the stub shaft and a second axial bearing surface of a static structure, the roller bearing arrangement further including a first snubber positioned between the radially outer surface of the fan shaft and a radially inner surface of the stub shaft, the first snubber being spaced apart from the radially inner surface of the stub shaft or the radially outer surface of the fan shaft so as to limit a radial movement range of the stub shaft.

A gas turbine fan casing made of composite material with a fibrous reinforcement includes a plurality of superimposed turns of a strip-shaped fibrous texture having a three-dimensional weaving between a plurality of layers of warp yarns and a plurality of layers of weft yarns, the fibrous reinforcement being densified by a matrix. The fibrous texture includes at least one lateral section of variable thickness in which the weft yarns have a size or a count different from the size or the count of the weft yarns of the plurality of layers of weft yarns present in the remainder of the fibrous texture.

A gas turbine engine for an aircraft that includes a nacelle, a fan, an engine core, a bypass duct extending between the engine core and the nacelle and guiding a bypass airflow through the bypass duct, and at least one non-structural strut extending in the radial direction within the bypass duct, wherein the non-structural strut includes an outside wall acting as a heat exchanger, and wherein the outside wall includes first transport means configured to transport in the outside wall at least one fluid to be cooled. It is provided that the non-structural strut further includes second transport means configured to transport a fluid to be heated, wherein the first transport means and the second transport means are configured such that the fluid to be heated is heated by the at least one fluid to be cooled and the at least one fluid to be cooled is cooled both by the bypass airflow and the fluid to be heated.

A gas turbine engine for an aircraft that includes a nacelle, a fan, an engine core, a bypass duct extending between the engine core and the nacelle and guiding a bypass airflow through the bypass duct, and at least one non-structural strut extending in the radial direction within the bypass duct, wherein the non-structural strut includes an outside wall acting as a heat exchanger, and wherein the outside wall includes first transport means configured to transport in the outside wall at least one fluid to be cooled. It is provided that the non-structural strut further includes second transport means configured to transport a fluid to be heated, wherein the first transport means and the second transport means are configured such that the fluid to be heated is heated by the at least one fluid to be cooled and the at least one fluid to be cooled is cooled both by the bypass airflow and the fluid to be heated.

Method for monitoring a turbomachine comprising a stator and a rotor, device, system, aircraft and computer program product. The method comprises a step of acquiring an input signal which represents a deformation of the stator or rotor of the turbomachine. The input signal has been acquired by a deformation gauge which is attached to the stator or rotor. The input signal comprises a first component which represents deformations of the stator or rotor which are caused by rotation of the rotor vanes in relation to the stator, and a second component which represents deformations which are caused by different elements of the rotor vanes. The method comprises a step of re-sampling the input signal in order to obtain a re-sampled input signal comprising a predefined whole number of samples per revolution of the turbomachine rotor. The method also comprises a step of processing the re-sampled input signal, comprising: filtering the re-sampled input signal, making it possible to acquire a filtered input signal in which the second component is attenuated, and separating the filtered input signal into a plurality of third components, each third component representing a contribution to the deformation caused by a respective associated vane. The method also comprises a step of detecting, in a third component, contact between the vane associated with the third component and the stator.

A barrier device is provided for an inlet cowl for an aircraft engine including an inner barrel circumferentially surrounding an opening in the inlet cowl formed along an axis of rotation of the aircraft engine, through which air passes to the aircraft engine, the inner barrel including a face sheet disposed on a radially inward side of the inner barrel relative to the axis. The barrier device includes a containment doubler of the inner barrel, disposed on a radially outward side of the inner barrel, and a blade fragment barrier including one or more strips disposed between the containment doubler and the face sheet, so as to extend circumferentially at least partially around the opening and to occupy a radial distance between the containment doubler and the face sheet.

A containment system for a gas turbine engine includes an outer containment ring having a first projection that extends radially inward and an inner containment ring having at least a second projection that extends radially outward toward the outer containment ring. The second projection is offset and spaced apart from the first projection to define a chamber in a first state. The containment system includes a lattice defined within the chamber that spans the chamber, and the lattice is integrally formed with each of the outer containment ring and the inner containment ring. A density of the lattice varies in at least one of an axial direction and a radial direction. The second projection is configured to interlock with the first projection in a second state.