A gas turbine engine (10) includes a compressor section, a turbine section, and an accessory gearbox (100). A turning unit (200) for the gas turbine engine includes an output assembly (204) configured to be mechanically coupled to the gas turbine engine, and an electric motor (202). The electric motor is operable to rotate, through the output assembly, one or more components of the compressor section or the turbine section at a rotational speed less than about fifty revolutions per minute during a shut-down condition of the gas turbine engine.

An output assembly for an accessory system of a gas turbine engine includes a first rotating member extending along an axis and including a mechanical connector for coupling the output assembly to an accessory system. The output assembly also including a second rotating member extending along the axis, the second rotating member coupled to the first rotating member at a first axial position and at a second axial position, the second rotating member coupled to the first rotating member at the second axial position through a one-way clutch. Additionally, the first rotating member of the output assembly further includes an extension member extending at least partially between the first axial position and the second axial position, the extension member including a designed fail point.

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.

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 vaned ring for a turbomachine stator includes an outer shell, a vane having a vane head, and two platforms arranged circumferentially on either side of the vane head, and fixed on a radially inner face of the outer shell by detachable fasteners. The vane head has first conical seatings bearing on respective second conical seatings of the two platforms so as to prevent a displacement of the vane radially inwards and along the circumferential direction. The vaned ring includes a frangible pin including a base mounted on the outer shell and a head projecting radially inwards with respect to the outer shell and having a head surface forming, vis--vis a top surface of the vane head, a stop opposing a displacement of the vane radially outwards.

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).

The invention relates to a fan, in particular for an aircraft cooling unit, comprising a wheel (128) and a shaft line (136) for driving the wheel about an axis (A), said wheel including a hub (138) supporting an annular row of blades (140), and means (142) for connection to said shaft line that are housed inside said hub. The fan is characterized in that the connection means include at least one meltable safety element (158) designed to be broken and to rotationally disengage at least a portion of the connection means from the shaft line when a torque for driving the wheel that is transmitted by the shaft line exceeds a specific threshold.

A portable jig and fixture having at least one base plate, a column assembly removably coupled to the at least one base plate, a slide assembly removably coupled to the column assembly, a power feed head removably coupled to the slide assembly, and a spindle unit removably coupled with the slide assembly having a motor removably engaged it. At least one cutter assembly is removably engaged with the spindle unit and at least one drill jig is also removably engaged with the spindle unit. Also, a turbomachine having complementary structure on the compressor discharge casing can engage with the at least one base plate alignment mounting structure on the portable jig and fixture.

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.

An airfoil defining a span extending between a root and a tip and a chord at each point along the span extending between a leading edge and a trailing edge. The airfoil includes a blade extending between the root and tip and extending between the leading edge and trailing edge. The blade includes a pressure side and a suction side defining a thickness therebetween at each point along the span. The blade defines a first chord reduction on at least one of the leading edge or trailing edge along a first portion of the span. Further, the blade defines a frangible line extending from the first chord reduction at least partially along the chord at a point along the span within the first portion of the span.