A connector system for a gas turbine engine having an engine core and a fan located upstream of the engine core and configured to provide a bypass airflow around the engine core, wherein the connector system comprises: an engine core mount block, configured to be mounted to the surface of the engine core and comprising a plurality of fluid conduits passing through the engine core mount block; and a rigid conduit configured to be supported by connection of a first end of the rigid conduit to the engine core mount block and/or by connection of a second end to a connection point in the gas turbine engine arranged radially outward of the bypass airflow; wherein the rigid conduit comprises a plurality of fluid conduits extending from the first end to the second end of the rigid conduit; and when the first end of the rigid conduit is connected to the engine core mount block, each of the fluid conduits in the engine core mount block is in fluid communication a respective fluid conduit in the rigid conduit.

The invention concerns a shaft component, which can be connected or is connected to the input or output side of a gear box in a gas turbine engine, in particular an aircraft engine, wherein the shaft component has partially a region comprising fiber reinforced plastic, the fibers in this region being arranged only in an angular range of +/40 to 50, in particular of +/42 to 48, most particularly +/45, in relation to the main axis of rotation of the shaft component. The invention also concerns a method for producing a shaft component and a gas turbine engine.

A drive shaft component, which can be connected or is connected to the input side of a gear box in a gas turbine engine, in particular an aircraft engine, characterized in that the drive shaft component has partially a region including fiber reinforced plastic, the fibers in this region being arranged only in an angular range of +/40 to 50, in particular of +/42 to 48, most particularly +/45, in relation to the main axis of rotation of the drive shaft component. The invention also concerns a method for producing a drive shaft component and a gas turbine engine.

A control system for a gas turbine engine includes an engine core, the engine core including combustion equipment, a turbine, a compressor, and a core shaft connecting the turbine to the compressor. The control system includes at least one variable stator vane for controlling the angle at which gas enters the engine core, and there is a bypass passage within the engine core for directing gas flow to bypass the combustion equipment.

An anti-rotation assembly may include a first shaft, a second shaft, and a locking sprocket. The first shaft may have a first crenelated end rim, the second shaft may have a second crenelated end rim, and the locking sprocket may be configured to engage the first crenelated end rim and the second crenelated end rim to anti-rotatably couple the first shaft to the second shaft. That is, the locking sprocket may be configured to prevent relative rotation between the two shafts. The anti-rotation assembly may further include an axial retaining ring configured to axially retain the locking sprocket.

An anti-rotation assembly may include a first shaft, a second shaft, and a locking sprocket. The first shaft may have a first crenelated end rim, the second shaft may have a second crenelated end rim, and the locking sprocket may be configured to engage the first crenelated end rim and the second crenelated end rim to anti-rotatably couple the first shaft to the second shaft. That is, the locking sprocket may be configured to prevent relative rotation between the two shafts. The anti-rotation assembly may further include an axial retaining ring configured to axially retain the locking sprocket.

An aircraft system is provided that includes a gas turbine engine, a gearbox and at least one compressor. The gas turbine engine includes a compressor section, a turbine section, a combustor section and a rotating structure. The combustor section is fluidly coupled with and between the compressor section and the turbine section. The rotating structure includes a compressor section rotor within the compressor section and a turbine section rotor within the turbine section. The at least one compressor includes a compressor rotor rotatably driven by the rotating structure through the gearbox. The gas turbine engine may be dedicated to powering the at least one compressor.

An aircraft system is provided that includes a gas turbine engine, a gearbox and at least one compressor. The gas turbine engine includes a compressor section, a turbine section, a combustor section and a rotating structure. The combustor section is fluidly coupled with and between the compressor section and the turbine section. The rotating structure includes a compressor section rotor within the compressor section and a turbine section rotor within the turbine section. The at least one compressor includes a compressor rotor rotatably driven by the rotating structure through the gearbox. The gas turbine engine may be dedicated to powering the at least one compressor.

A highly efficient gas turbine engine includes the fan of the gas turbine engine driven from a turbine via a gearbox, such that the fan has a lower rotational speed than the driving turbine, thereby providing efficiency gains. The efficient fan system is mated to a core that has low cooling flow requirements and/or high temperature capability, and which may have particularly low mass for a given power.

A highly efficient gas turbine engine includes the fan of the gas turbine engine driven from a turbine via a gearbox, such that the fan has a lower rotational speed than the driving turbine, thereby providing efficiency gains. The efficient fan system is mated to a core that has low cooling flow requirements and/or high temperature capability, and which may have particularly low mass for a given power.