Abstract: A guide vane assembly includes a guide vane and a sealing assembly slidable along a span of the guide vane relative to the guide vane and extending along a chordal axis and the span of the guide vane. The sealing assembly includes a plate configured to be disposed around at least a portion of an external surface of the guide vane. The sealing assembly further includes a manifold coupled to the plate and configured to receive pressurized fluid therein. The manifold is disposed around the plate. The sealing assembly further includes a plurality of flow tubes extending between the manifold and the plate. Each of the plurality of flow tubes are configured to receive the pressurized fluid from the manifold.

Abstract: There is provided a gas turbine engine comprising a blower system for supplying pressurised air to an airframe via an airframe port. The blower system comprises a compressor configured to receive air from a bypass duct or a core of the gas turbine engine and to discharge compressed air into a delivery line extending from the compressor to the airframe port. The blower system also comprises a heat exchanger configured to transfer heat from the compressed air to a coolant and a valve arrangement configured to switch between operation of the blower system in a baseline mode and a cooling mode, the valve arrangement being configured to: selectively divert the compressed air within the delivery line to the heat exchanger for operation in the cooling mode; and/or selectively provide the coolant to the heat exchanger for operation in the cooling mode.

Abstract: A gas turbine engine has a compression system blade ratio defined as the ratio of the height of a fan blade to the height of the most downstream compressor blade in the range of from 45 to 95. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

Abstract: There is provided a blower system for providing air to an airframe system, comprising a rotor configured to be mechanically coupled to a spool 440 of a gas turbine engine, wherein the rotor is configured to: in a blower mode, be driven to rotate by the spool to discharge air to an airframe discharge port for supply to an airframe system; and, in an engine drive mode, receive air from an external air source via an impingement port that is configured to direct the received air onto the rotor and thereby drive the rotor to rotate to drive the spool to.

Abstract: A gas turbine engine has a compression system radius ratio defined as the ratio of the radius of the tip of a fan blade to the radius of the tip of the most downstream compressor blade in the range of from 5 to 9. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

Abstract: A compressor of an aircraft air pressurization system comprises: a rotor configured to be mechanically coupled to a spool of a gas turbine engine; an oil-lubricated bearing for supporting rotation of the rotor, wherein the oil-lubricated bearing is disposed within a bearing chamber; and a seal assembly for restricting oil from the bearing chamber from reaching an inlet of the blower compressor, wherein the seal assembly comprises: a seal disposed between the bearing chamber and an air buffer chamber; an air buffer inlet configured to receive a flow of pressurized air into the air buffer chamber; and an air buffer outlet configured to allow the flow of pressurized air to be exhausted from the air buffer chamber together with oil from the bearing chamber that has passed through the seal and become entrained within the flow of pressurized air.

Abstract: There is provided a dynamic sealing assembly for a rotary machine, comprising a primary sandwich plate, a secondary sandwich plate and a bristle pack. The primary sandwich plate comprises a plurality of primary vane openings, and the secondary sandwich plate comprises a plurality of secondary vane openings. The bristle pack comprises a plurality of bristles and is disposed between the primary sandwich plate and the secondary sandwich plate. Each of the plurality of primary vane openings overlies and aligns with a respective secondary vane opening to form a vane channel for receiving a vane along a longitudinal axis of the dynamic sealing assembly. The bristle pack is configured to: provide a brush seal between each vane received within the respective vane channels and the dynamic sealing assembly; and allow relative movement between the dynamic sealing assembly and the vane received within each vane channel along the longitudinal axis.

Abstract: A gas turbine engine has a compression system blade ratio defined as the ratio of the height of a fan blade to the height of the most downstream compressor blade in the range of from 45 to 95. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

Abstract: A gas turbine engine has a compression system radius ratio defined as the ratio of the radius of the tip of a fan blade to the radius of the tip of the most downstream compressor blade in the range of from 5 to 9. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

Abstract: A gas turbine engine that has improved fuel burn provides operability and/or maintenance requirements when installed on an aircraft. The gas turbine engine is provided with a core compressor that includes twelve, thirteen or fourteen rotor stages. The gas turbine engine has a ratio of a core compressor aspect ratio divided by a core compressor pressure ratio is in the range of from 0.03 to 0.09. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

Abstract: A rotary assembly for driving spool rotation includes a rotor and a flow modifier. The rotor is mechanically coupled to a spool of a gas turbine engine. The flow modifier receives flow from and/or direct flow to the rotor. The rotary assembly permits relative movement between the rotor and the flow modifier to move between: a turbine configuration wherein the rotor receives air from an external air source to drive the spool to rotate; and a compressor configuration wherein the rotor is driven to rotate by the spool and to receive and compress air from the gas turbine engine, and discharge the compressed air for supply to the airframe system. The rotary assembly also includes controller to control relative movement between the rotor and the flow modifier through a range of turbine positions of the turbine configuration to vary a torque applied to the rotor for driving the spool.

Abstract: There is provided a dynamic sealing assembly for a rotary machine, comprising a primary sandwich plate, a secondary sandwich plate and a bristle pack. The primary sandwich plate comprises a plurality of primary vane openings, and the secondary sandwich plate comprises a plurality of secondary vane openings. The bristle pack comprises a plurality of bristles and is disposed between the primary sandwich plate and the secondary sandwich plate. Each of the plurality of primary vane openings overlies and aligns with a respective secondary vane opening to form a vane channel for receiving a vane along a longitudinal axis of the dynamic sealing assembly. The bristle pack is configured to: provide a brush seal between each vane received within the respective vane channels and the dynamic sealing assembly; and allow relative movement between the dynamic sealing assembly and the vane received within each vane channel along the longitudinal axis.

Abstract: A compressor variable angle measurement system for guiding the positioning variable vanes supported on a penny of a compressor of a gas turbine engine. The system comprising a gauge assembly that is connectable to a computing device. the gauge assembly comprises a base plate and a clamp arm. The gauge assembly is configured to removably grip a variable vane between three vane contact portions of the baseplate and the vane contact portion of the clamp arm and on the leading edge vane engaging portion and the trailing edge vane engaging portion of the base plate, the stagger angle of the variable vane with respect to the radial setting pin being determined by the computing device from measurements made by an inertial measurement unit.

Abstract: A gas turbine engine has a compression system blade ratio defined as the ratio of the height of a fan blade to the height of the most downstream compressor blade in the range of from 45 to 95. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

Abstract: There is provided a gas turbine engine comprising a blower system for supplying pressurised air to an airframe via an airframe port. The blower system comprises a compressor configured to receive air from a bypass duct or a core of the gas turbine engine and to discharge compressed air into a delivery line extending from the compressor to the airframe port. The blower system also comprises a heat exchanger configured to transfer heat from the compressed air to a coolant and a valve arrangement configured to switch between operation of the blower system in a baseline mode and a cooling mode, the valve arrangement being configured to: selectively divert the compressed air within the delivery line to the heat exchanger for operation in the cooling mode; and/or selectively provide the coolant to the heat exchanger for operation in the cooling mode.

Abstract: A gas turbine engine has a compression system blade ratio defined as the ratio of the height of a fan blade to the height of the most downstream compressor blade in the range of from 45 to 95. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

Abstract: There is provided a blower system for providing air to an airframe system, comprising a rotor configured to be mechanically coupled to a spool 440 of a gas turbine engine, wherein the rotor is configured to: in a blower mode, be driven to rotate by the spool to discharge air to an airframe discharge port for supply to an airframe system; and, in an engine drive mode, receive air from an external air source via an impingement port that is configured to direct the received air onto the rotor and thereby drive the rotor to rotate to drive the spool to.

Abstract: A gas turbine engine has a compression system radius ratio defined as the ratio of the radius of the tip of a fan blade to the radius of the tip of the most downstream compressor blade in the range of from 5 to 9. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

Abstract: A gas turbine engine has a compression system blade ratio defined as the ratio of the height of a fan blade to the height of the most downstream compressor blade in the range of from 45 to 95. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

Abstract: A compressor variable angle measurement system for guiding the positioning variable vanes supported on a penny of a compressor of a gas turbine engine. The system comprising a gauge assembly that is connectable to a computing device. the gauge assembly comprises a base plate and a clamp arm. The gauge assembly is configured to removably grip a variable vane between three vane contact portions of the baseplate and the vane contact portion of the clamp arm and on the leading edge vane engaging portion and the trailing edge vane engaging portion of the base plate, the stagger angle of the variable vane with respect to the radial setting pin being determined by the computing device from measurements made by an inertial measurement unit.