Abstract: Apparatus for controlling a power generation system, the apparatus comprising a controller configured to: identify a trigger indicative of a future change in electrical power output by the power generation system to a first power level; control the power generation system to change electrical power output to a second power level in response to the trigger, the second power level being equal to, or different to the first power level; and control supply of at least a portion of the electrical power output from the power generation system at the second power level to an electrical energy storage system to charge the electrical energy storage system

Abstract: Structural rod supports are provided in vanes in a gas turbine assembly in order to provide support to the vanes and an annular seal. Structural rods couple to the turbine case at one end and the annular seal at an opposite end.

Abstract: A combustion staging system is provided for fuel injectors of a multi-stage combustor of a gas turbine engine. The system has a splitting unit which receives a metered total fuel flow and controllably splits the metered total fuel flow into out-going pilot and mains fuel flows to perform pilot-only and pilot-and-mains staging control of the combustor. The system further has pilot and mains fuel manifolds which respectively receive the pilot and mains fuel flows, the mains fuel manifold being split into a primary line and a servo line such that each line receives a respective portion of the mains fuel flow. The system further has a plurality of mains flow control valves which distribute the mains fuel flow from the mains fuel manifold to mains discharge orifices of respective injectors of the combustor, both the primary line and the servo line extending to the mains flow control valves before reuniting.

Abstract: A gas turbine engine includes: an engine core with a turbine, compressor, and core shaft; a fan upstream the engine core; a gearbox that can receive an input from the core shaft and output drive to a fan shaft to drive the fan at a lower rotational speed than the core shaft. A fan-gearbox axial distance is greater than or equal to 0.35 m. The fan shaft has: (i) a radial bending stiffness ratio of a radial bending stiffness of the fan shaft at the fan input to a radial bending stiffness of the fan shaft at the gearbox output that is greater than or equal to 6.0×10?3; and/or (ii) a tilt stiffness ratio of a tilt stiffness of the fan shaft at the fan input to a tilt stiffness of the fan shaft at the gearbox output that is greater than or equal to 2.5×10?2.

Abstract: A coupling arrangement for a gas turbine engine. The arrangement comprises first, second and third members. The first member has a first threaded mating surface extending in a first direction (X) and a flange extending in a direction generally normal to the first direction (X). The second member has a second threaded mating surface extending in the first direction (X) and a flange extending in a direction generally normal to the first direction (X), the flanges of the first and second members engaging against one another. The third member has a third threaded mating surface configured to engage against the first threaded mating surface, and a fourth threaded mating surface configured to engage against the second threaded mating surface.

Abstract: A system for providing active flow control in an aircraft having a gas turbine engine. The system includes an environmental control system that includes a cabin blower system having a compressor operable to compress a fluid delivered by a fan section of the gas turbine engine to generate a pressurised fluid for use by the environmental control system. The environmental control system is fluidicly connected to an active flow control system via a fluid supply line, for allowing the pressurised fluid generated by the compressor to be supplied to the active flow control system so that it can be ejected from the aircraft across an exterior surface of a movable control element of the aircraft.

Abstract: A gas turbine engine for an aircraft includes: an engine core with a compressor system including a first, lower pressure, compressor, and a second, higher pressure, compressor; an inner core casing provided radially inwardly of the compressor blades of the compressor system; and an outer core casing surrounding the compressor system, the inner core casing and the outer core casing defining a core working gas flow path therebetween. The outer core casing includes: a first flange connection arranged to allow separation of the outer core casing at an axial position of the first flange connection, the first flange connection having a first flange radius, A gas path radius is defined as the outer radius of the core gas flow path at the axial position of the first flange connection, and a gas path ratio of: first ? ? flange ? ? radius gas ? ? path ? ? radius is equal to or greater than 1.10.

Abstract: A gas turbine engine for an aircraft and an engine core including: a compressor system, a first lower pressure compressor, a second, higher pressure compressor; an outer core casing surrounding the compressor system. The outer core casing includes a first flange connection arranged to allow separation of the outer core casing at an axial position of the first flange connection, the first flange connection having a first flange radius, wherein the first flange connection is the connection that is downstream of an axial position defined by the axial midpoint between the mid-span axial location on the trailing edge of the most downstream aerofoil of the first compressor and the mid-span axial location on the leading edge of the most upstream aerofoil of the second compressor, and a front mount arranged to be connected to a pylon.

Abstract: An annular combustor for a gas turbine engine including inner and outer combustor walls, wherein each wall defines an annulus and the inner wall is radially inward of the outer wall. The combustor includes a primary zone where the inner and outer combustor walls converge in a downstream direction, and a secondary zone downstream of the primary zone. In the secondary zone, the inner and outer walls are arranged to converge at a different rate to the primary zone, are non-convergent or are divergent in a downstream direction, a rate of change of radial width of the combustor is different in the zones. A transition is provided from the primary zone to the secondary zone. A plurality of combustor cooling tiles lines the inner and outer walls. One or more of the tiles are arranged to extend from the primary to secondary zone and across the transition from the zones.

Abstract: A gas turbine engine includes an engine core including: a compressor system including first, lower pressure compressor, and second, higher pressure compressor; and an outer core casing. The engine includes a front mount arranged for connection to a pylon; and a fan located upstream of the engine core. The outer core casing includes a first flange connection that: is arranged to allow separation of the outer core casing at an axial position thereof, and is the first flange connection downstream of an axial position defined by the axial midpoint between the mid-span axial location on trailing edge of the most downstream aerofoil of first compressor and mid-span axial location on leading edge of the most upstream aerofoil of the second compressor. A front mount position ratio of: axial ? ? distance ? ? between ? ? the ? ? first ? ? flange ? ? connection and ? ? the ? ? front ? ? mount first ? ? flange ? ? radius is equal to or less than 1.18.

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 comprises a fan mounted to rotate about a main longitudinal axis; an engine core, comprising in axial flow series a compressor, a combustor, and a turbine coupled to the compressor through a shaft; a reduction gearbox that receives an input from the shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the shaft; wherein the compressor comprises a first stage at an inlet and a second stage, downstream of the first stage, comprising respectively a first rotor with a row of first blades and a second rotor with a row of second blades, the first and second blades comprising respective leading edges, trailing edges and tips, and wherein the ratio of a maximum leading edge radius of the first blades to a maximum leading edge radius of the second blades is greater than 2.8.

Abstract: A gas turbine engine includes an engine core including: a compressor system including first, lower pressure compressor, and second, higher pressure compressor; and inner and outer core casings that define a core working gas flow path (A) therebetween, which has an outer radius that defines a gas path radius. The outer core casing includes a first flange connection that: has a first flange radius, is arranged to allow separation of the outer core casing at an axial position thereof, and is the first flange connection that is downstream of an axial position defined by the axial midpoint between the mid-span axial location on trailing edge of the most downstream aerofoil of first compressor and mid-span axial location on leading edge of the most upstream aerofoil of the second compressor. A gas path ratio of: first ? ? flange ? ? radius gas ? ? path ? ? radius is equal to or greater than 1.10.

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 method of obtaining vibration data relating to a rotating shaft, the method comprising the steps of: receiving an output of a speed probe adjacent a phonic wheel coaxially coupled to a rotating shaft, the speed probe being configured to produce an output with a magnitude dependent upon a distance between the speed probe and the phonic wheel; determining an amplitude modulation of the output; and deriving vibration data relating to the rotating shaft from the determined amplitude modulation.

Abstract: A gas turbine engine for an aircraft includes: an engine core with: a compressor system including a first, lower pressure, compressor, and a second, higher pressure, compressor; and an outer core casing surrounding the compressor system. The gas turbine engine further includes a fan located upstream of the engine core with a plurality of fan blades and having a fan diameter. The outer core casing includes a first flange connection arranged to allow separation of the outer core casing at an axial position of the first flange connection which has a first flange radius, wherein the first flange connection is downstream of an axial position defined by the axial midpoint between the mid-span axial location on the trailing edge of the most downstream aerofoil of the first compressor and the mid-span axial location on the leading edge of the most upstream aerofoil of the second compressor.

Abstract: A gas turbine engine includes an engine core including: a compressor system including first, lower pressure, compressor, and a second, higher pressure, compressor; and an outer core casing surrounding the compressor system and including a first flange connection arranged to allow separation of the outer core casing at an axial position of the first flange connection, wherein the first flange connection is the first flange connection that is downstream of an axial position defined by the axial midpoint between the mid-span axial location on the trailing edge of the most downstream aerofoil of the first compressor and the mid-span axial location on the leading edge of the most upstream aerofoil of the second compressor; a nacelle surrounding the engine core and defining a bypass duct between the engine core and the nacelle; wherein an axial midpoint of the radially outer edge is defined as the fan OGV tip centrepoint.

Abstract: A gas turbine engine for an aircraft includes an engine core including a turbine, compressor, and core shaft connecting turbine to compressor; a fan located upstream of the engine core and including a plurality of fan blades each having a leading and trailing edge. The turbine includes a lowest pressure turbine stage having a row of rotor blades, each rotor blades extending radially and having a leading and trailing edge. The engine has a fan tip axis that joins a radially outer tip of the leading edge of a fan blade and the radially outer tip of the trailing edge of a rotor blade of the lowest pressure stage. The fan tip axis lies in a longitudinal plane which contains a centreline of engine. A fan axis angle is defined as the angle between fan tip axis and centreline, and is in a range between 10 and 20 degrees.

Abstract: An 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; a fan shaft; a gearbox that receives an input from the core shaft and outputs drive to the fan via the fan shaft so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox having a sun gear, a plurality of planet gears, a ring gear, and a planet carrier on which the planet gears are mounted; and a gearbox support arranged to mount the gearbox within the engine. The fan shaft, core shaft, gearbox and the gearbox support together may form a transmission.

Abstract: A blade has a root portion and elongate portion extending from the root portion to a tip. The elongate portion has an aerofoil-shaped cross section having leading and trailing edges and suction and pressure sides. The tip may include a gutter defining squealer. The squealer has a wall extending from the trailing edge and along a substantial portion of the tip perimeter. A main trailing edge cooling channel extends within the elongate portion in a direction from root to tip adjacent the trailing edge and exiting into the gutter. A gallery channel is arranged just behind the gutter and extends from an open end intersecting the main trailing edge cooling channel to a closed end located just behind a trailing edge apogee. Cooling channels extend from the gallery channel and through the squealer wall. The gallery channel diameter is greater at the open end than the closed end.