A turbine shroud assembly adapted for use with a gas turbine engine includes a carrier, a seal segment, and a mount assembly. The carrier is configured to be coupled to a turbine case. The seal segment is shaped to define a gas path boundary of the shroud assembly. The mounting assembly is configured to couple the seal segment to the carrier.

A turbofan having a dilution rate of at least 10 and including a fan having a disc provided with blades at the periphery thereof, a distance between the head of the blades and the housing of the fan being less than or equal to ten millimeters; a primary flow space and secondary flow space that are concentric; a turbine, housed in the primary flow space and in fluid communication with the fan; and a reduction mechanism coupling the turbine and the fan.

A method of manufacturing a motoring system for a gas turbine engine including the steps of: forming a mechanical shaft fuse, the mechanical shaft fuse including a plurality of through holes; forming an outer housing; installing a reduction gear train into the outer housing, the reduction gear train having an input and an output; operably connecting an electric motor to the input; operably connecting a clutch to the output using the mechanical shaft fuse, the clutch in operation engages and disengages the reduction gear train; operably connecting a starter to the clutch, the starter having an output shaft; and operably connecting an accessory gearbox to the output shaft of the starter. The mechanical shaft fuse in operation shears when torque on the mechanical is greater than or equal to a selected value.

A frangible strut for a gas turbine engine includes a first end to couple to a first structure of the gas turbine engine, and a second end opposite the first end. The second end is to couple to a second structure of the gas turbine engine. The second end includes an outer portion coaxial with and spaced radially apart from an inner portion. The inner portion is to receive a fastening device to couple the second end to the second structure. The outer portion is coupled to a body of the frangible strut. The inner portion is interconnected to the outer portion by a plurality of frangible members that are spaced apart about a perimeter of the inner portion, and each of the frangible members are configured to release the inner portion from the outer portion. The frangible strut includes the body interconnecting the first end and the second end.

Lubricant pump systems and associated methods for aircraft engines are provided. The method includes receiving an input torque, dividing the input torque between a first load path receiving a first portion of the input torque, and a second load path receiving a second portion of the input torque. A first lubricant pump of the aircraft engine is driven via the first load path using the first portion of the input torque. A second lubricant pump of the aircraft engine is driven via the second load path using the second portion of the input torque. When a malfunction of the second lubricant pump occurs, the method includes ceasing to drive the first lubricant pump and the second lubricant pump using the input torque.

A turbine shroud assembly adapted for use with a gas turbine engine includes a carrier, a seal segment, and a mount assembly. The carrier is configured to be coupled to a turbine case. The seal segment is shaped to define a gas path boundary of the shroud assembly. The mounting assembly is configured to couple the seal segment to the carrier.

A turbomachine includes a fan shaft driven by a turbine shaft via a device for reducing a speed of rotation. The turbomachine includes an uncoupling device interposed between the reduction device and the turbine shaft. The uncoupling device is configured to uncouple the reduction device and the turbine shaft in response to the exceeding of a determined resistant torque exerted by the reduction device on the turbine shaft.

A turbine engine has a fan and a low pressure compressor that rotate at a common speed and in a common direction. A fan drive turbine drives a gear reduction to, in turn, drive the low pressure compressor and the fan at a speed which is slower than a speed of the fan drive turbine. A combustor intermediate the low pressure compressor and the fan drive turbine and a thrust bearing mount the fan drive turbine, the thrust bearing being aft of a location of the combustor. A shear section in a drive connection connecting the fan drive turbine to the gear reduction is weaker than other portions of the drive connection. The shear section is aft of the thrust bearing.

An aircraft gas turbine engine (10) comprises a main engine shaft (22, 23), a main engine shaft bearing arrangement (36, 44, 49, 50) configured to rotatably support the main engine shaft (22, 23) and an electric machine (30) comprising a rotor (34) and a stator (32). The rotor (34) is mounted to the main engine shaft (22, 23) and is rotatably supported by the main engine shaft bearing arrangement (36, 44, 49, 50), and the stator (32) is mounted to static structure (46) of the gas turbine engine (10).

A gas turbine engine has a bearing housing mounted to structure linked to the engine mounting pads. The bearing housing has bearing housing flanges with bearing housing flange openings aligned with attachment flange openings of the structure. Some of the bearing housing flange openings are slots. A first group of fasteners extends through one of the attachment flange openings and through one of the bearing housing flange openings. A second group of fasteners extends through one of the attachment flange openings and through one of the slots. The first group of fasteners are sacrificial fasteners configured to fracture in response to a load on the bearing housing exceeding a fracture load. The bearing housing is displaceable relative to the structure after fracture of the sacrificial fasteners via the second group of the fasteners moving within and relative to respective ones of the slots.