Abstract: A gas turbine engine includes an inlet duct to guide a core engine flow to a compressor and an engine section stator arranged in the inlet duct upstream of the compressor including vanes with leading edges defining a first annulus area in the inlet duct, a mid-span leading edge point of the engine section stator vanes being arranged at a first radius. The compressor includes a first rotor with a row of first blades with leading edges defining a second annulus area; a mid-span leading edge point of the compressor first blades being arranged at a second radius and an axial distance from the engine section stator vane mid-span leading edge point. The ratio of the second to the first annulus area is equal or greater than 0.75, and the ratio of a difference between the first and second radius to the axial distance is equal or greater than 0.23.

Abstract: An axial flux electrical machine comprises a first flux generating assembly, a second flux generating assembly, a shaft and a speed controller. The shaft has an axis of rotation. Each of the first flux generating assembly and the second flux generating assembly is rotationally located on the shaft in axial juxtaposition to one another, with the first flux generating assembly being axially separated from the second flux generating assembly by a separation distance. The speed controller is configured to modify a magnetic field generated by either of the first flux generating assembly and the second flux generating assembly so as to control a rotational speed of the electrical machine.

Abstract: A turbine blade having a body enclosing a plurality of internal channels for the passage of coolant gas received through an inlet formed in the blade root, including a gas flow separator, configured to separate coolant gas into first and second gas flows. The gas flow separator is configured to separate the first and second gas flows such that the level of contaminants in the first gas flow is lower than the level of contaminants in the second gas flow.

Abstract: A gas turbine engine for an aircraft comprises an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub; and a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft. The gas turbine engine has an engine length and a centre of gravity position measured relative to the fan, and a centre of gravity position ratio of: the centre of gravity position/the engine length is in a range from 0.43 to 0.6.

Abstract: A system for monitoring the axial position of a rotating shaft includes a phonic wheel mounted coaxially to the shaft for rotation with a circumferential row of teeth. The system includes a sensor configured to detect the passage of the teeth by generating an alternating measurement signal. First and second portions of the teeth alternate around the row and contribute respective first and second components to the alternating measurement signal. The first portion of teeth vary in height in an axial direction of the wheel such that the relative height of the first and second portions varies with axial distance across the phonic wheel, and the sensor is positioned relative to the phonic wheel such that axial displacement of the shaft causes the signal to vary the first component's amplitude relative to the second component's amplitude due to the height variation, to monitor the axial position of the shaft.

Abstract: A gas turbine engine of an aircraft includes: an engine core having a turbine including a lowest pressure rotor stage, a turbine diameter, a fan including a plurality of fan blades extending from a hub, an annular fan face at a leading edge of the fan; wherein a downstream blockage ratio is: the ? ? turbine ? ? diameter ? ? at ? ? an ? ? axial location ? ? of ? ? the ? ? lowest ? ? pressure ? ? rotor ? ? stage ground ? ? plane ? ? to ? ? wing ? ? distance and a quasi-non-dimensional mass flow rate Q defined as: Q = W ? ? T ? ? 0 P ? ? 0. ? ? A flow where: W is mass flow rate through the fan in Kg/s; T0 is average stagnation temperature of the air at the fan face in Kelvin; P0 is average stagnation pressure of the air at the fan face in Pa; and Aflow is the flow area of the fan face in m2, and wherein a Q ratio of: the downstream blockage ratio×Q is in a range from 0.005 to 0.01.

Abstract: A gas turbine engine generates noise during use, and one particularly important flight condition for noise generation is take-off. A gas turbine engine has high efficiency together with low noise, in particular the noise emanating from the front of the fan. The contribution of the fan noise emanating from the front of the engine to the Effective Perceived Noise Level (EPNL) at a take-off lateral reference point, defined as the point on a line parallel to and 450 m from the runway centre line where the EPNL is a maximum during take-off, is in the range of from 0 EPNdB and 12 EPNdB lower than the contribution of the fan noise emanating from the rear of the engine to the EPNL at the take-off lateral reference point.

Abstract: A gas turbine engine for an aircraft includes an engine core including a turbine, a compressor, a core shaft, and a core exhaust nozzle, the core exhaust nozzle having a core exhaust nozzle pressure ratio calculated using total pressure at the core nozzle exit; a fan including a plurality of fan blades; and a nacelle surrounding the fan and the engine core and defining a bypass duct, the bypass duct including a bypass exhaust nozzle, the bypass exhaust nozzle having a bypass exhaust nozzle pressure ratio calculated using total pressure at the bypass nozzle exit; wherein a bypass to core ratio of: bypass ? ? exhaust ? ? nozzle ? ? pressure ? ? ratio core ? ? exhaust ? ? nozzle ? ? pressure ? ? ratio is configured to be in the range from 1.1 to 2.0 under aircraft cruise conditions.

Abstract: A gas turbine engine for an aircraft includes an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan including a plurality of fan blades, wherein a fan tip radius of the fan is measured between a centreline of the engine and an outermost tip of each fan blade at its leading edge; and a nacelle surrounding the fan and the engine core and defining a bypass exhaust nozzle, the bypass exhaust nozzle having an outer radius. An outer bypass to fan ratio of: the ? ? outer ? ? radius ? ? of ? ? the ? ? bypass ? ? exhaust ? ? nozzle the ? ? fan ? ? tip ? ? radius is in the range from 0.6 to 1.05.

Abstract: A gas turbine engine generates noise during use, and one particularly important flight condition for noise generation is take-off. A gas turbine engine is provided that has high efficiency together with low noise, in particular from the jet flow exiting the engine and the turbine. The combined contribution of the jet and the turbine to the Effective Perceived Noise Level (EPNL) at a take-off lateral reference point, defined as the point on a line parallel to and 450 m from the runway centre line where the EPNL is a maximum during take-off, is in the range of from 3 EPNdB and 15 EPNdB lower than the total engine EPNL at the take-off lateral reference point.

Abstract: A gas turbine engine for an aircraft has an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan including a plurality of fan blades extending from a hub; and a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft. The gas turbine engine has an engine length and a gearbox location relative to a forward region of the fan, and a gearbox location ratio of: gearbox location/engine length is in a range from 0.19 to 0.45.

Abstract: A gas turbine engine for an aircraft including: an engine core; a fan located upstream of engine core, fan including a plurality of fan blades; a nacelle surrounding the gas turbine engine, nacelle including an inner surface at least partly defining a bypass duct; and a bypass duct outlet guide vane extending radially across bypass duct between the engine core's outer surface and the nacelle's inner surface. An outer wall axis is defined joining a radially outer tip of a trailing edge of the bypass duct outlet guide vane and a rearmost tip of the inner surface of the nacelle, wherein the outer wall axis lies in a longitudinal plane containing the centreline of gas turbine engine, an outer bypass duct wall angle is defined as the angle between outer wall axis and centreline, and the outer bypass duct wall angle is in a range between ?15 to 1 degrees.

Abstract: A gas turbine engine generates noise during use, and one particularly important flight condition for noise generation is take-off. A gas turbine engine has high efficiency together with low noise, in particular from the jet flow exiting the engine. The contribution of the jet to the Effective Perceived Noise Level (EPNL) at a take-off lateral reference point, defined as the point on a line parallel to and 450 m from the runway centre line where the EPNL is a maximum during take-off, is in the range of from 0 EPNdB and 15 EPNdB lower than the contribution of the fan noise emanating from the rear of the engine to the EPNL at the take-off lateral reference point.

Abstract: A gas turbine engine, and an aircraft including the gas turbine engine. The gas turbine engine comprising: an engine core comprising a turbine, a compressor, and a shaft system connecting the turbine to the compressor. The shaft system is axially located by a thrust bearing located forward of the turbine, and the engine is configured such that, in the event of a shaft break which divides the shaft system into a front portion located by the thrust bearing and a rear portion unlocated by the thrust bearing, the rear portion is free to move axially rearwardly under a gas load; and the engine further comprises an axial movement sensor configured to register a shaft break when it detects the axial movement of the rear portion of the shaft system.

Abstract: A gas turbine engine for an aircraft including: engine core including a turbine; and fan including a plurality of fan blades extending radially from a hub, each fan blade having a leading and trailing edge. Turbine includes a lowest pressure turbine stage having a row of rotor blades each extending radially and having a leading and trailing edge. A fan-turbine radius difference is measured as radial distance between: a point on a circle swept by a radially outer tip of the trailing edge of each of the rotor blades of the lowest pressure stage of the turbine; and a point on a circle swept by a radially outer tip of the leading edge of each of fan blades; and a fan speed to fan-turbine radius ratio defined as: the ? ? maximum ? ? take ? - ? off ? ? rotational ? ? speed ? ? of ? ? the ? ? fan fan ? - ? turbine ? ? radius ? ? difference ? ? ( 120 ) is in a range between 0.8 rpm/mm to 5 rpm/mm.

Abstract: A gas turbine engine for an aircraft includes an engine core having a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan having a plurality of fan blades, wherein a fan tip radius of the fan is measured between a centreline of the engine and an outermost tip of each fan blade at its leading edge; and a nacelle surrounding the fan and the engine core and defining a bypass exhaust nozzle, the bypass exhaust nozzle having an inner radius. An inner bypass to fan ratio of: the ? ? inner ? ? radius ? ? of ? ? the ? ? bypass ? ? exhaust ? ? nozzle the ? ? fan ? ? tip ? ? radius is in the range from 0.4 to 0.

Abstract: A gas turbine engine generates noise during use, and one particularly important flight condition for noise generation is take-off. A gas turbine engine that has high efficiency provides low noise, in particular from the fan and the turbine that drives the fan. Values are defined for a noise parameter NP that results in a gas turbine engine having reduced combined fan and turbine noise.

Abstract: A gas turbine engine generates noise during use, and one particularly important flight condition for noise generation is take-off. A gas turbine engine has high efficiency together with low noise, in particular from the bypass flow exiting the engine. The average velocity of the flow at the exit to the bypass duct is in the range of from 200 m/s to 275 m/s at a take-off lateral reference point, defined as the point on a line parallel to and 450 m from the runway centre line where the Effective Perceived Noise Level (EPNL) is a maximum during take-off.

Abstract: A gas turbine engine for an aircraft, including engine length, is arranged to be mounted beneath a wing of the aircraft, and has an engine core having a turbine, a compressor, and a core shaft, the turbine having a lowest pressure rotor stage; a fan located upstream of the engine core, the fan having a fan tip radius; and a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft. A downstream blockage ratio is defined as: the ? ? turbine ? ? diameter ? ? at ? ? an ? ? axial location ? ? of ? ? the ? ? lowest ? ? pressure ? ? rotor ground ? ? plane ? ? to ? ? wing ? ? distance and an engine blockage ratio of: ( 2 × the ? ? fan ? ? tip ? ? radius ? / ? the ? ? engine ? ? length ) the ? ? downstream ? ? blockage ? ? ratio is in the range from 2.5 to 4.

Abstract: An epicyclic geartrain includes a debris collection arrangement. The epicyclic geartrain comprises a sun gear, a plurality of planet gears, with the plurality of planet gears being supported by a planet torque ring, and a ring gear. The planet gears meshingly surround the sun gear, and the ring gear meshingly surrounds the planet gears. At least one of the sun gear, the plurality of planet gears, and the ring gear, is provided with a plurality of permanent magnet portions, and a debris collection element. The plurality of magnet portions is arranged as a circumferential array across a side of the corresponding gear. The debris collection element extends along the side face of the gear. The debris collection element is slidably positioned against the side face, such that rotation of the gear causes the debris collection element to remove any ferromagnetic debris particles that are magnetically attached to the side face.