A gas turbine engine includes a compressor having a first compressor section and a second compressor section, a combustor fluidly connected to the compressor, and a turbine fluidly connected to the combustor. The turbine has a first turbine section and a second turbine section. A first shaft connects the first compressor section and the first turbine section. A second shaft connects the second compressor section and the second turbine section. A fan is connected to the first shaft via a geared architecture. A low speed accessory gearbox is interfaced with the first shaft via a mechanical interface. The low speed accessory gearbox includes a mechanical brake.

A bearing assembly for a gas turbine engine comprises a bearing; a bearing bracket, which holds the bearing and is secured by a predetermined breaking device on a connecting element, which can be connected or is connected to a support structure of the gas turbine engine; and a clutch for transmitting a torque from a first clutch element connected in a fixed manner to the rotor of the bearing to a second clutch element supported on the bearing bracket, wherein the clutch elements are spaced apart when the predetermined breaking device is intact and can be brought into contact with one another by destruction of the predetermined breaking device. A gas turbine engine and a method are furthermore provided.

Systems, methods, and devices are provided for a turbine driven pump gearbox. A turbine engine may drive a primary shaft having gears. The gears may selectably engage with gears of a secondary shaft that drives a machine (e.g., pump). By changing which gears of the primary shaft engage with which gears of the secondary shaft, a gear ratio may be changed. A power takeoff device (e.g., a generator) may be connected to the primary shaft and may be operated in reverse as a motor to rotate, slow, stop, and/or reverse rotation of the primary shaft. Brakes may be associated with one or more of the primary and secondary shafts. The power takeoff device and one or more of the brakes may be controlled to shift engagement of the shafts between different positions, changing the gear ratio and/or disengaging the shafts from each other.

An assembly is provided for an aircraft propulsion system. This assembly includes a compressor section, a combustor section, a turbine section and a flowpath extending sequentially through the compressor section, the combustor section and the turbine section. The assembly also includes a turbine rotor, a propulsor rotor and an auxiliary turbine. The turbine rotor is within the turbine section. The turbine rotor is configured to rotatably drive the propulsor rotor. The auxiliary turbine includes an auxiliary turbine rotor. The auxiliary turbine rotor is configured to rotatably drive the propulsor rotor with the turbine rotor. The auxiliary turbine is configured to receive bleed gas from the flowpath.

An aircraft assembly includes a geartrain with a first gear system and a second gear system. The first gear system includes a first sun gear, a first ring gear, a plurality of first intermediate gears and a first carrier. The first sun gear and the first ring gear are each rotatable about an axis. The first intermediate gears are between and meshed with the first sun gear and the first ring gear. Each of the first intermediate gears is rotatably mounted to the first carrier. The second gear system is interconnected with the first gear system. The second gear system includes a second sun gear, a second ring gear, a plurality of second intermediate gears and a second carrier. The second intermediate gears are between and meshed with the second sun gear and the second ring gear. Each of the second intermediate gears is rotatably mounted to the second carrier.

A first gear system includes a first sun gear, a first ring gear, first intermediate gears and a first carrier. The first intermediate gears are between and meshed with the first sun gear and the first ring gear. Each first intermediate gear is rotatably mounted to the first carrier. A second gear system includes a second sun gear, a second ring gear, second intermediate gears and a second carrier. The second ring gear is coupled to the first carrier. The second intermediate gears are between and meshed with the second sun gear and the second ring gear. Each second intermediate gear is rotatably mounted to the second carrier. A first propulsor rotor is coupled to the first carrier. A rotating structure is coupled to the first ring gear and includes a turbine rotor. The rotating structure is configured to drive rotation of the first propulsor rotor through the geartrain.

A gas turbine engine has an engine static structure and at least one component rotatable relative to the engine static structure about an engine axis of rotation. A fan is coupled to at least one component for rotation about the engine axis of rotation. An actuator is mounted to the engine static structure, wherein the actuator is activated to prevent the fan from rotation and is inactivated to allow the fan to rotate. A method for preventing rotation of a fan in a gas turbine engine is also disclosed.

A low pressure compressor section of a gas turbine engine includes low pressure compressor stages spaced apart along an axis of rotation of the low pressure compressor section. The low pressure compressor section includes at least one of a fluid conduit extending between an air inlet of the gas turbine engine and the low pressure compression stages. The fluid conduit having at least one flow diverter displaceable between a first position and a second position to modulate air through the low pressure compressor stages. And/or a compressor stage disabler is engageable with at least one of the low pressure compressor stages and configured to reduce a rotation thereof.

An aircraft turbine engine includes a rotating body comprising a compressor rotor and a turbine rotor interconnected by a rotor shaft, the turbine engine being configured to drive a member by said rotor shaft via an epicyclic reduction gear, said epicyclic reduction gear comprising a first element (50) configured to be rotationally secured to said shaft, a second element (56) rotationally secured to said member, and a third element configured to be selectively secured to a stator of the turbine engine and disengaged from the stator, wherein the turbine engine comprises driving means to rotationally drive said third element at a piloted speed when it is disengaged from said stator.

A gas turbine engine has an engine static structure. At least one component rotatable relative to the engine static structure about an engine axis of rotation. A fan is coupled to at least one component for rotation about the engine axis of rotation. An actuator is mounted to the engine static structure, wherein the actuator is activated to prevent the fan from rotation and is inactivated to allow the fan to rotate. A method for preventing rotation of a fan in a gas turbine engine is also disclosed.