Abstract: A method of controlling electrical power supplied to a component of a vehicle, the method comprising: receiving a signal comprising information associated with an operating condition of a gas turbine engine; determining whether a parameter exceeds a predetermined threshold value using the information in the received signal; and controlling a reduction in electrical power supplied to a component of a vehicle from a generator of the gas turbine engine if the parameter exceeds the predetermined threshold value.

Abstract: A turbomachine for a gas turbine engine, comprising: a rotor disc arranged to rotate about a fixed central axis. A plurality of rotor blades connected to the rotor disc, wherein the rotor blades extend within a gas flow path. A rotating annulus is arranged to rotate in a fixed rotational alignment relative to the rotor disc. An outer surface of the rotating annulus is arranged to define part of the boundary of the gas flow path. A gas turbine engine for an aircraft and a method of defining the boundary of a gas flow path within a turbomachine is also disclosed.

Abstract: A gas turbine engine comprising a fan, an intermediate pressure compressor, a high pressure compressor, a high pressure turbine, a low pressure turbine and a gearbox. The low pressure turbine is arranged to drive the intermediate pressure compressor via a shaft and an intermediate pressure compressor drive shaft and the fan via the shaft, a gearbox input shaft, the gearbox and a gearbox output shaft. The intermediate pressure compressor drive shaft has a V-shaped cross-section comprising a base, a first limb and a second limb extending from the base. An end of the first limb of the intermediate pressure compressor drive shaft is removably connected to the intermediate pressure compressor. An end of the second limb of the intermediate pressure compressor drive shaft is removably connected to the gearbox input shaft and the base of the intermediate pressure compressor drive shaft is removably connected to the shaft.

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 manufacturing an austenitic iron alloy comprising placing austenitic iron alloy powder in a can, evacuating air and other gases from the can, supplying nitrogen gas into the can, sealing the can and then hot isostatically pressing the austenitic iron alloy powder in the can to diffuse the nitrogen into the austenitic iron alloy powder and to produce a nitrogen enriched austenitic iron alloy bar and removing the can from the nitrogen enriched austenitic iron alloy bar.

Abstract: A turbofan engine has: an engine core producing a core exhaust flow; an upstream fan; a bypass duct carrying a bypass air flow produced by the fan; an exhaust assembly. The turbofan engine afterburner system has: a core plug; second core flow duct, and entrance(s) from the first core flow duct for diversion into the second of a portion of the core exhaust flow, and exit for re-joining the diverted portion of the core exhaust flow to the undiverted portion of the core exhaust flow; door(s) open during normal operation of the engine but closable, during a reheat operation, to block the entrances to the second core flow duct; and fuel injector(s) and igniter(s) operable, during the reheat operation, to inject and ignite fuel within the second core flow duct, providing a reheat flow into the core exhaust flow from the exit of the second core flow duct.

Abstract: A bracket for mounting a first component to a second component, the bracket comprising at least one mounting boss defining a through-hole having an opening axis in axial direction for receiving a fixing used to attach the bracket to one of the first and second components; wherein the mounting boss comprises a fixing support face around the through-hole on a surface of the bracket, the fixing support face configured to bear a head of the fixing; the mounting boss is configured to resist compressive forces exerted on the bracket by the head of the fixing; and the mounting boss comprises a cavity in a space projected from the fixing support face in the axial direction of the through-hole.

Abstract: The present disclosure relates to a geared gas turbine engine for an aircraft. Example embodiments include a gas turbine engine for an aircraft including: 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; and a gearbox that receives an input from the core shaft to drive the fan at a lower rotational speed than the core shaft, the gearbox having a gear ratio of around 3.4 or higher, wherein the gas turbine engine is configured such that a jet velocity ratio between a first jet velocity exiting from a bypass duct of the engine and a second jet velocity exiting from an exhaust nozzle of the engine core is within a range from around 0.75 to around 0.82.

Abstract: A gas turbine engine includes a fan mounted to rotate about a main longitudinal axis; an engine core, including a compressor, a combustor, and turbine coupled to the compressor through a shaft; and reduction gearbox; wherein the compressor includes a plurality of stages, each stage including a respective rotor and stator, a first stage of the plurality of stages being arranged at an inlet and including a first rotor with a plurality of blades; each blade extending chordwise from a leading edge to a trailing edge, and from root to tip for a span height, wherein 0% of the span height corresponds to the root and 100% of span height corresponds to tip; wherein the ratio of a maximum leading edge radius of the first rotor blades to a minimum leading edge radius of the first rotor blades is included between 2.2 and 3.5.

Abstract: Disclosed is an exhaust nozzle for a gas turbine engine, the exhaust nozzle comprising an outer frame extending along a longitudinal direction, a convergent petal pivotably attached to the frame and extending axially downstream and radially inward from the pivot, radially within the frame, and a sealing hinge arrangement between an upstream member and a downstream member of the exhaust nozzle. One of the upstream member or the downstream member defines a cylindrical socket having an opening along a cylinder axis which receives a corresponding cylindrical hinge element the other of the downstream member or upstream member, where the upstream member is defined by the frame and the downstream member is the convergent petal; or the exhaust nozzle further comprises a divergent petal downstream of the convergent petal and pivotably attached to the convergent petal, the upstream member being the convergent petal and the downstream member being the divergent petal.

Abstract: There is disclosed an exhaust nozzle for a gas turbine engine, the exhaust nozzle comprising a frame extending along a longitudinal axis. The exhaust nozzle comprises a convergent petal pivotably attached at a convergent pivot point to the frame and extending axially downstream and radially inward from the frame, a follower roller fixed to the convergent petal on a radially outer side of the convergent petal, and a cam defining a working surface configured to engage the follower roller to react a force from the convergent petal. The cam is movable along a travel in an axial direction to actuate radial movement of the follower roller to pivot the convergent petal. The cam defines a concave working surface such that a contact angle between the follower roller and the cam varies along the travel to thereby vary a radial component of the force reacted by the cam.

Abstract: A joint assembly joins first and second components about a common axis. The first component has a first end portion having a radially outwardly facing surface shaped to fit radially inside a second surface of a hollow second end portion of the second component to form an interface between the opposing first and second surfaces. The first and second surfaces have a concavity extending laterally with respect to the axis such that when the first and second surfaces are opposingly arranged the opposing concavities define a cavity at the interface. A retaining member is insertable into the cavity at the interface to prevent axial separation of the first and second components. One of the first and second end portions has a free end protrusion axially spaced from the concavity and the other of the first and second end portions has an axially extending recess arranged to receive the free end.

Abstract: Maraging steel alloys are disclosed. The alloys are produced by microalloying of the maraging steel alloy to form carbides at prior austenite grain boundaries to increase Zener drag. A particular example alloy consists essentially of, by weight, 7.4 to 8.4 percent nickel, 7.6 to 8.6 percent chromium, 8.4 to 9.4 percent cobalt, 1.8 to 2.2 percent molybdenum, 2 to 2.6 percent tungsten, 1.6 to 2 percent aluminium, 0.05 to 0.08 percent carbon, a carbide former selected from the group consisting of: niobium at a concentration of 0.25 to 0.28 percent; titanium, at a concentration of 0.2 to 0.28 percent; and vanadium, at a concentration of 0.21 to 0.4 percent; the balance being iron and incidental impurities.

Abstract: A gas turbine engine has a core engine including: a compressor, combustion equipment which receives compressed air from the compressor, a circumferential row of nozzle guide vanes, and a turbine. The nozzle guide vanes defines a throat receiving hot working gases from the combustion equipment into the turbine. The gas turbine engine further has an air system which is switchably operable between an on-position which opens a diversion pathway along which a portion of the compressed air exiting the compressor bypasses the combustion equipment to join the hot working gases at re-entry holes located between the nozzle guide vanes and a rotor at the front of the turbine, thereby increasing the semi-dimensional mass flow ?(T30)0.5/(P30) of the core engine at the exit of the compressor, and an off-position which closes the diversion pathway, thereby decreasing the semi-dimensional mass flow of the core engine at the exit of the compressor.

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 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: A collision avoidance method for a water-based vessel, including: obtaining real time data relating to the path of two or more vessels; identifying a collision risk between the two or more vessels; determining if the collision risk is above a predetermined threshold, wherein when the collision risk is above the predetermined threshold, determining one or more collision avoidance manoeuvres on the basis of historical navigational data which corresponds to the real-time data; providing the one or more collision avoidance manoeuvres to an operator of the one or more vessels.

Abstract: A turboelectric generator system includes a gas turbine engine which includes, in fluid flow series, a gas-generator compressor, a combustor, a gas-generator turbine, and a variable-speed free power turbine. The system further comprises a variable-frequency electric machine rotatably connected with the free power turbine and a power converter configured to convert a variable frequency electrical output from the electric machine to a fixed frequency output.

Abstract: A gauge comprising an elongate rod and a head. The rod has an elongate axis. The head has a transverse dimension perpendicular to the elongate axis of the rod. The gauge has an insertion configuration and an inspection configuration, wherein the transverse dimension of the head is greater in the inspection configuration than in the insertion configuration. The head comprises two head portions each with a longitudinally-extending outer surface and a longitudinally-extending inner surface and, in the insertion configuration, the inner surfaces of the head portions are at least partly in abutment with each other. The gauge can be used in a method of measuring an internal dimension of a bore within a component.