A rigid electrical raft is provided to a gas turbine engine via a fusible mount arrangement. The rigid electrical raft may be a part of an electrical system of the gas turbine engine, for example a part of the electrical harness. The fusible mount is arranged to break when a predetermined load is applied. The rigid electrical raft may be attached to a fan case of the engine, and the predetermined load may be that which results from a fan blade being released from the hub. This ensures that the rigid electrical raft is protected from the load.

A containment system for an engine includes an engine case having an inner perimeter. The containment system includes a containment ring nested within the inner perimeter of the engine case and integrally formed with the engine case along a first interface and a second interface. The containment ring includes a first leg opposite a second leg, and the first interface is defined between the first leg and the engine case. The containment system includes a first plurality of perforations defined at the first interface, and the first leg of the containment ring is frangible along the first plurality of perforations to at least partially release the containment ring to protect the engine case during a containment event.

A fan casing arrangement for a gas turbine engine having a propulsive fan, the arrangement having a fan case and a fan track liner and being configured to circumscribe the fan, wherein the fan track liner is provided around the inside of fan case so as to adopt a radial position between the fan and the fan case. The fan track liner includes an elongate member which is helically-wound against the inside of the fan case in a plurality of turns. A method of installing a fan track liner in a fan casing arrangement for a gas turbine engine having a propulsive fan, the method involving: providing a fan case to circumscribe the fan, providing a flexible and elongate liner member, and helically winding the liner member against the inside of the fan case in a plurality of turns to define at least part of the fan track liner.

A turbine engine assembly includes a turbine engine case, a gearbox and a plurality of mounts connecting the gearbox to the case. Each of the mounts includes a linkage and a fuse joint. The linkage extends towards the case substantially in a first direction. The fuse joint is configured to substantially prevent movement between the gearbox and the case when the fuse joint is subject to loading less than a threshold, and to permit a constrained movement between the gearbox and the case when the loading is greater than the threshold. The fuse joint of a first of the mounts has a different configuration than the fuse joint of a second of the mounts.

A bearing arrangement rotatably supports a shaft of an aircraft engine. The bearing arrangement comprises a bearing having rolling elements disposed between inner and outer races. The inner race is affixed to the shaft. A decoupler normally structurally couples the outer race of the bearing to a stator structure of the engine. The decoupler is configured to release the bearing from the stator structure when subject to a predetermined critical load. A bumper is mounted to the stator structure and encircles the bearing. The bumper has a radially inwardly facing surface disposed in close proximity to a radially outer surface of the outer race of the bearing and defines therewith a radial gap to accommodate and constrain an orbiting motion of the rotor about the central axis of the engine after decoupling at the bearing. The bumper further has an axially forwardly facing surface which is axially spaced by a predetermined axial fore gap from a first flange projecting radially outwardly from a front end portion of the outer race of the bearing. The first flange of the outer race is axially trapped between the stator structure and the bumper. After decoupling, the bearing is free to axially and radially move within the radial gap and the axial fore gap.

The invention relates to a blade, the platform (14) of which has at least one bumper (40) which rises, on the inner face, from a connection to an intermediate area of the blade located along a longitudinal axis of the blade, between the blade root and the outer face (143), towards a free side edge (145) of the platform. The platform bumper (40) locally has a thinned portion at a distance from said free side edge of the platform.

The invention is applicable to turboshaft engine fans.

Walls of gas turbine engine casings, fan cases, and methods for forming walls, e.g., fan case walls, are provided. For example, a wall comprises a plurality of composite plies arranged in a ply layup. The wall is annular and circumferentially segmented into a plurality of regions that include at least one first region and at least one second region. The ply layup in the first and second regions is different such that the ply layup is non-axisymmetric. An exemplary fan case comprises an annular inner shell, a filler layer, an annular back sheet, and an annular outer layer. The back sheet is circumferentially segmented into a plurality of regions, including at least one first region and at least one second region, and comprises a plurality of composite plies arranged in a ply layup that is different in the first and second regions such that the ply layup is non-axisymmetric.

A method and system for establishing sets of blade frequency values for each rotating blade of a rotor assembly at two or more different points in time and determining an indication of blade health from the change in the blade frequency values is provided. Blade frequency values are determined by receiving measurements of vibratory responses from blade monitoring equipment (20) and processing via a processing device (30) vibration data as a system of rotating blades to extract a frequency of each blade. Sets of blade frequency values are compared to determine a change in the blade frequency values for each rotating blade to provide the indication of blade health.

A rotary component may comprise a first structure configured to rotate about an axis and a second structure configured to rotate about the axis. A support structure may be coupled to the first structure at a first attachment location and to the second structure at a second attachment location. The support structure may be configured to separate from the first structure and the second structure in response to a centrifugal force generated by the first structure and the second structure rotating about the axis.

A system and method for detecting the rotation breakaway of a spacer from a compressor rotor disk are disclosed herein. The method includes detecting a disk sensed feature and a spacer sensed feature on the compressor rotor disk and the spacer respectively. The method also includes comparing the timing between the two to a predetermined threshold to determine whether the relative position of the spacer to the compressor rotor disk exceeds a predetermined amount. The predetermined amount may be selected to determine whether an imbalance, rubbing, or binding can occur in the gas turbine engine or to determine whether anti-rotation features have been broken.