Abstract: A gas turbine engine including: a high pressure turbine, a low pressure turbine, a high pressure compressor coupled to the high pressure turbine by a high pressure shaft, a propulsor and a low pressure compressor coupled to the low pressure turbine via a low pressure shaft and a reduction gearbox; wherein the high pressure compressor defines an average stage pressure ratio at cruise conditions of between 1.25 and 1.35 and consists of 10 or 11 stages; and the high pressure compressor and low pressure compressor together define a core overall pressure ratio at cruise conditions of between 40:1 and 60:1.

Abstract: A gas turbine engine comprises, in fluid flow series, a gas-generator compressor, a combustor, a gas-generator turbine, and a free power turbine. The gas-generator compressor is an axi-centrifugal compressor comprising a plurality of axial compression stages followed by a single centrifugal compression stage, wherein the International Standard Atmosphere, sea-level static (hereinafter ISA SLS) design point pressure ratio of the axi-centrifugal compressor is from 12 to 16, and a ratio of the ISA SLS pressure rise across the axial compression stages to the ISA SLS pressure rise across the centrifugal compression stage is from 0.75 to 1.

Abstract: A turboshaft gas turbine engine comprises, in fluid flow series, a gas-generator compressor, a combustor, a gas-generator turbine, and a free power turbine. The expansion ratios of the turbines are dependent upon an outlet temperature of the combustor, and an expansion relationship between the turbines is defined as the ratio of expansion ratios of the turbines over a running range of non-dimensional power outputs of the gas turbine engine. A second derivative of the expansion relationship is from 0.6 to 1.0.

Abstract: A gas turbine engine including: a high pressure turbine, a low pressure turbine, a high pressure compressor coupled to the high pressure turbine by a high pressure shaft, a propulsor and a low pressure compressor coupled to the low pressure turbine via a low pressure shaft and a reduction gearbox; wherein the high pressure compressor defines an average stage pressure ratio at cruise conditions of between 1.25 and 1.35 and consists of 10 or 11 stages; and the high pressure compressor and low pressure compressor together define a core overall pressure ratio at cruise conditions of between 40:1 and 60:1.

Abstract: A gas turbine engine comprises, in fluid flow series, a gas-generator compressor, a combustor, a gas-generator turbine, and a free power turbine. The gas-generator compressor is an axi-centrifugal compressor comprising a plurality of axial compression stages followed by a single centrifugal compression stage, wherein the International Standard Atmosphere, sea-level static (hereinafter ISA SLS) design point pressure ratio of the axi-centrifugal compressor is from 12 to 16, and a ratio of the ISA SLS pressure rise across the axial compression stages to the ISA SLS pressure rise across the centrifugal compression stage is from 0.75 to 1.

Abstract: A turbofan engine has an engine core including in flow series a compressor, a combustor and a turbine. The engine further has a fan located upstream of the engine core, has a supersonic intake for slowing down incoming air to subsonic velocities at an inlet to the fan formed by the intake, has a bypass duct surrounding the engine core, wherein the fan generates a core airflow to the engine core and a bypass airflow through the bypass duct, and has a mixer for mixing an exhaust gas flow exiting the engine core and bypass airflow exiting bypass duct. The engine further has a thrust nozzle rearwards of the mixer for discharging mixed flows, the thrust nozzle having a variable area throat. The engine further has a controller controlling the thrust produced by the engine over a range of flight operations including on-the-ground subsonic take-off and subsequent off-the-ground subsonic climb.

Abstract: A gas turbine engine comprises, in fluid flow series, a gas-generator compressor, a combustor, a gas-generator turbine, and a free power turbine. The gas-generator compressor is an axi-centrifugal compressor comprising a plurality of axial compression stages followed by a single centrifugal compression stage, wherein the International Standard Atmosphere, sea-level static (hereinafter ISA SLS) design point pressure ratio of the axi-centrifugal compressor is from 12 to 16, and a ratio of the ISA SLS pressure rise across the axial compression stages to the ISA SLS pressure rise across the centrifugal compression stage is from 0.75 to 1.

Abstract: A gas turbine engine includes an engine core having high and low pressure compressors and high and low pressure turbines. The engine further includes a fan coupled to a low pressure shaft and the low pressure turbine by a reduction gearbox. The low pressure compressor includes no more than two compressor stages, and the low and high pressure compressor together define a cruise overall core pressure ratio of between 30 and 50.

Abstract: A gas turbine engine including: a high pressure turbine, a low pressure turbine, a high pressure compressor coupled to the high pressure turbine by a high pressure shaft, a propulsor and a low pressure compressor coupled to the low pressure turbine via a low pressure shaft and a reduction gearbox; wherein the high pressure compressor defines an average stage pressure ratio at cruise conditions of between 1.25 and 1.35 and consists of 10 or 11 stages; and the high pressure compressor and low pressure compressor together define a core overall pressure ratio at cruise conditions of between 40:1 and 60:1.

Abstract: A gas turbine engine (10) and method of operation. The gas turbine engine (10) comprises a heating device configured to heat a rotor disc (32, 72) of the engine (10). A method of operation the heating device comprises detecting an engine acceleration or deceleration event, or determining that an engine acceleration or deceleration event is imminent or may be imminent. On detection of an engine acceleration or deceleration event, or in advance of the engine acceleration or deceleration event, increasing turbine rotor disc heat input to raise a temperature of the turbine rotor disc or reduce a cooling rate of the rotor disc.

Abstract: A turboshaft gas turbine engine comprises, in fluid flow series, a gas-generator compressor, a combustor, a gas-generator turbine, and a free power turbine. The expansion ratios of the turbines are dependent upon an outlet temperature of the combustor, and an expansion relationship between the turbines is defined as the ratio of expansion ratios of the turbines over a running range of non-dimensional power outputs of the gas turbine engine. A second derivative of the expansion relationship is from 0.6 to 1.0.

Abstract: A turboshaft gas turbine engine comprises, in fluid flow series, a gas-generator compressor, a combustor, a gas-generator turbine, and a free power turbine. The gas-generator turbine is a single stage turbine having a specific work of from 200 to 350 kilojoules per kilogram, said specific work being the power output of the turbine per unit mass flow therethrough.

Abstract: A gas turbine engine comprises, in fluid flow series, a gas-generator compressor, a combustor, a gas-generator turbine, and a free power turbine. The gas-generator compressor is an axi-centrifugal compressor comprising a plurality of axial compression stages followed by a single centrifugal compression stage, wherein the International Standard Atmosphere, sea-level static (hereinafter ISA SLS) design point pressure ratio of the axi-centrifugal compressor is from 12 to 16, and a ratio of the ISA SLS pressure rise across the axial compression stages to the ISA SLS pressure rise across the centrifugal compression stage is from 0.75 to 1.

Abstract: Electrical machine apparatus comprising: a rotor having a longitudinal axis and being arranged to rotate about the longitudinal axis in a first circumferential direction, the rotor defining one or more conduits for receiving fluid therein, the one or more conduits including an inlet and an outlet; and a fluid guide defining a first aperture arranged to direct fluid in a second direction towards the rotor, the second direction having a positive circumferential component in the first circumferential direction.

Abstract: Electrical machine apparatus comprising: a rotor having an axis of rotation and defining a cavity therein; and a conduit positioned within the cavity of the rotor, the conduit comprising an inlet arranged to receive a fluid and an outlet arranged to exhaust the fluid, the inlet having a first radial distance from the axis of rotation and the outlet having a second radial distance from the axis of rotation, the first radial distance being greater than the second radial distance.

Abstract: A turboshaft gas turbine engine comprises, in fluid flow series, a gas-generator compressor, a combustor, a gas-generator turbine, and a free power turbine. The rotational speeds of the turbines are dependent upon an outlet temperature of the combustor, and a speed relationship between the turbines is defined as the ratio of rotational speeds of the turbines over a running range of non-dimensional power outputs of the gas turbine engine. A first derivative of the speed relationship is from 0.9 to 1.

Abstract: An auxiliary power unit (10) for an aircraft (1) comprises a gas turbine engine comprising an engine core (12). The engine core comprises a core compressor (14) and core turbine (18) coupled by a core shaft (24). The auxiliary power unit further comprises a load spool comprising a load compressor (30), a load turbine (20) and a permanent magnet electric machine (28), each being coupled by a load shaft (26). The load shaft (26) and core shaft (24) are configured to rotate independently of one another.

Abstract: An electrical machine has a stator carrying electrical windings which protrude at opposite ends of the stator to form respective, ring-shaped end windings. The electrical machine further has a rotor having a plurality of magnetic field-producing elements for producing a rotor magnetic field which interacts with a stator magnetic field produced by the windings. The electrical machine further has a coolant bath for holding liquid coolant. The electrical machine further has a heat sink in thermal contact with at least one of the end windings, the bath and the heat sink being configured such that the heat sink is immersed in the coolant held in the bath. The heat sink defines one or more fluid pathways configured such that vapour bubbles, formed when coolant in contact with the heat sink boils, escape by rising through the heat sink.

Abstract: A control system for use in controlling a cabin blower system. The cabin blower system includes a gas turbine engine spool, a cabin blower compressor powered by the spool and arranged in use to compress fluid used in a cabin of an aircraft, and one or more control mechanisms via which the control system controls the power extracted by the cabin blower compressor from the spool. The control system is arranged in use to control the power extracted from the spool by the cabin blower compressor in accordance with one or more primary control parameters. The control system is arranged in use to alter the spool power extracted by the cabin blower compressor by comparison with the power that would have been extracted in accordance with the primary control parameters alone, in response to modifications in a secondary control parameter indicative of the commencement or occurrence of an engine transient.

Abstract: A turbofan engine has an engine core including in flow series a compressor, a combustor and a turbine. The engine further has a fan located upstream of the engine core, has a supersonic intake for slowing down incoming air to subsonic velocities at an inlet to the fan formed by the intake, has a bypass duct surrounding the engine core, wherein the fan generates a core airflow to the engine core and a bypass airflow through the bypass duct, and has a mixer for mixing an exhaust gas flow exiting the engine core and bypass airflow exiting bypass duct. The engine further has a thrust nozzle rearwards of the mixer for discharging mixed flows, the thrust nozzle having a variable area throat. The engine further has a controller controlling the thrust produced by the engine over a range of flight operations including on-the-ground subsonic take-off and subsequent off-the-ground subsonic climb.