Abstract: A fuel staging system for a gas turbine engine has a plurality of fuel injectors each having a mains burner. The system has a mains manifold connected to a mains delivery line and configured to distribute fuel from the mains delivery line to the mains burner of each of the plurality of fuel injectors, and a check valve disposed in the mains delivery line upstream of the mains manifold. The check valve is configured to permit flow of fuel from the mains delivery line to the mains manifold when the pressure of fuel in the mains delivery line exceeds a threshold pressure.

Abstract: A combustion staging system for fuel injectors of a multi-stage combustor of a gas turbine engine. The system includes plural fuel injectors, each having respective pilot and mains injection stages. It further includes a splitting unit which, to perform staging control of the combustor, receives a metered fuel flow and, for pilot and mains operation, controllably splits the received fuel flow into a pilot flow for injecting at the pilot stages of the injectors and a mains flow for injecting at the mains stages of the injectors, and for pilot-only operation, controllably splits the received fuel flow into a first part of the pilot flow for injecting at the pilot stages of a first portion of the injectors and a second part of the pilot flow for injecting at the pilot stages of a second portion of the injectors.

Abstract: A fuel injector comprises a fuel feed arm and a fuel injector head and the fuel injector head has a fuel passage. The fuel feed arm has a fuel supply passage in fluid communication with the fuel passage in the fuel injector head. A first light guide extends through the fuel feed arm and has a distal end arranged to direct light into the fuel passage in the fuel injector head. A light source is arranged to supply light into a proximal end of the first light guide. A second light guide extends through the fuel feed arm and has a distal end arranged to receive transmitted light in the fuel passage in the fuel injector head and a light receiver is arranged to collect light at a proximal end of the second light guide.

Abstract: A method of controlling turbine tip clearance includes measuring turbine speed; measuring turbine temperature; measuring parameters indicative of current operating conditions; determining limits for the turbine speed and turbine temperature; calculating target tip clearance from the turbine speed, turbine temperature and parameters, to optimise turbine efficiency within the turbine speed and turbine temperature limits; and controlling turbine tip clearance apparatus to the calculated target tip clearance.

Abstract: An electrical machine has a variable reluctance rotor, and a stator formed as an annular array of stator segments. The reluctance of the rotor-to-stator magnetic flux path varies with rotor position whereby the stator segments are magnetically energizable to rotate the rotor. The stator segments are arranged in the array such that, when energized to rotate the rotor, they produce an unbalanced force on the rotor. The machine further has a compensator including one or more balancing segments which are configured to be magnetically energizable to produce a balancing force on the rotor which balances the unbalanced force. The reluctance of the rotor-to-compensator magnetic flux path is substantially invariant with rotor position.

Abstract: A method of controlling two or more sources of electrical power to supply electrical power to a shared bus. Each source of electrical power has a rated power Pr, a maximum overload power Pm, and is operable along a multi-voltage droop slope, the multi-voltage droop slope including a share region and a catch region, the share region being defined by an upper voltage limit, Vhigh and a base voltage Vbase, and the catch region being defined by the base voltage Vbase and a lower voltage limit Vlow.

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 of: the ? ? turbine ? ? diameter ? ? at ? ? an ? ? axial location ? ? of ? ? the ? ? lowest ? ? pressure ? ? rotor ? ? stage a ? ? distance ? ? from ? ? a ? ? ground ? ? plane ? ? to ? ? the ? ? wing is in the range from 0.2 to 0.3.

Abstract: A hydraulic seal arrangement for a rotating machine comprises an annular housing disposed around a shaft having an axis of rotation, the annular housing being relatively rotatable with respect to the shaft and having a sealing fluid zone within a radially outer portion of an interior of the annular housing. A fin extends radially from the shaft into the annular housing, with at least a portion of the fin extending radially into the sealing fluid zone. In use, the sealing fluid is centrifugally accelerated before being directly introduced into the sealing fluid zone by a sealing fluid inlet so as to generate a swirl component upon its introduction.

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 is in a range between 0.8 rpm/mm to 5 rpm/mm.

Abstract: A fuel injector comprising a first air swirler passage and a second air swirler passage extending axially through the fuel injector and arranged to direct air through the fuel injector, a splitter arranged between the first air swirler passage and the second air swirler passage and comprising a first splitter surface having a first divergent portion which is divergent in the downstream direction, a second splitter surface located radially inward of the first splitter surface and having a second divergent portion which is divergent in the downstream direction, a third splitter surface located radially inward of the first and second splitter surface, and a first connecting surface extending between the second and third splitter surfaces, wherein a first cavity is formed between the first and second splitter surfaces, and the second divergent portion comprises at least one opening in fluid communication with the first cavity.

Abstract: A fuel injector including: a plurality of air swirler passages; at least one fuel supply passage arranged to supply fuel into at least one of the air swirler passages; and at least one cavity separating an exterior of the fuel supply passage from a body of the fuel injector; wherein the cavity is at least partially filled with a thermally insulating material.

Abstract: A mount assembly for attaching a heat exchanger to a casing of a gas turbine engine comprises a support bracket and a pin body, the pin body laterally moveable in use relative to the support bracket. The assembly further comprises a spring that in use can apply a restoring force to the pin body opposed to its lateral movement relative to the support bracket.

Abstract: A gas turbine engine for an aircraft includes an engine core including a turbine, compressor, and core shaft, wherein a compressor exit temperature is defined as an average temperature of airflow at the exit from the compressor at cruise conditions and a core entry temperature is defined as an average temperature of airflow entering the engine core at cruise conditions. A fan rotor entry temperature is defined as an average temperature of airflow across the leading edge of each fan blade at cruise conditions. A core compressor temperature rise is defined as the compressor exit temperature divided by the core entry temperature. A fan root temperature rise is defined as the core entry temperature divided by the fan rotor entry temperature. A core compressor to fan root temperature rise ratio is in a specified range.

Abstract: An engine core including turbine, compressor, and core shaft connecting the turbine to the compressor, wherein a compressor exit temperature has an average airflow; and a fan upstream including a plurality of fan blades extending from a hub, each fan blade having a leading and trailing edge, wherein a fan rotor entry temperature has an average airflow across the leading edge of each blade at cruise conditions and fan tip rotor exit temperature has an average temperature of airflow across a radially outer portion of each blade at the trailing edge cruise conditions. A fan tip temperature rise as: the ? ? fan ? ? tip ? ? rotor ? ? exit ? ? temperature the ? ? fan ? ? rotor ? ? entry ? ? temperature .

Abstract: A gas turbine engine includes an engine core including a turbine, compressor, and core shaft connecting the turbine to compressor, wherein a compressor exit temperature defined as an average temperature of airflow at exit from compressor at cruise conditions and a core entry temperature defined as an average temperature of airflow entering engine core at cruise conditions, and a fan located upstream of the engine core, wherein a fan rotor entry temperature defined as an average temperature of airflow across leading edge each fan blade at cruise conditions and fan tip rotor exit temperature defined as an average temperature of airflow across a radially outer portion of each fan blade at the trailing edge at cruise conditions. A core compressor temperature rise defined as: the ? ? compressor ? ? exit ? ? temperature the ? ? core ? ? entry ? ? temperature .

Abstract: A gas turbine engine for an aircraft includes an engine core with a turbine, compressor, and core shaft connecting the two; and a fan upstream of the core with a plurality of blades extending from a hub each with a leading and trailing edge, wherein fan tip radius is between the engine centreline and each blade's leading edge outermost tip and hub radius is between the engine centreline and the hub's outer surface at each blade's leading edge radial position, the ratio of hub to tip radius between 0.2 and 0.285. A fan rotor entry temperature is the average temperature of airflow across the leading edge of each blade at cruise conditions and a fan rotor exit temperature is an average temperature of airflow across a radially outer portion of each blade at the trailing edge at cruise conditions, the ratio of entry to exit temperature between 1.11 and 1.05.

Abstract: An electrode guide assembly for an EDM process includes a guide tube having a first end and a second end, a fluid feed portion, a fluid return portion, and an electrode. The guide tube has a set of at least two supporting protrusions, with the protrusions projecting radially inwardly from an inner diametral surface of the guide tube. The electrode is slidably accommodated within the guide tube, with an outer diametral surface of the electrode abutting against the set of supporting protrusions. The first end of the guide tube is in fluid communication with the second end of the guide tube to thereby provide a fluid feed channel, and the second end of the guide tube is in fluid communication with the first end of the guide tube to thereby provide a fluid return channel.

Abstract: A gas turbine engine for an aircraft includes an engine core including a first, lower pressure, turbine, a first compressor, and a first core shaft connecting the first turbine to the first compressor; and a second, higher pressure, turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor, and a fan located upstream of the engine core and comprising a plurality of fan blades extending from a hub. First and second turbine entrance and exit temperatures are defined as average temperature of airflow at the entrance or exit to the respective turbine at cruise conditions. A low pressure turbine temperature change is defined as: the ? ? first ? ? turbine ? ? entrance ? ? temperature the ? ? first ? ? turbine ? ? exit ? ? temperature .

Abstract: A gas turbine engine for an aircraft includes an engine core including a first, lower pressure, turbine, a first compressor, and a first core shaft connecting the first turbine to the first compressor; and a second, higher pressure, turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor, and a fan located upstream of the engine core and including a plurality of fan blades extending from a hub. A low pressure turbine temperature change is defined as: the ? ? first ? ? turbine ? ? exit ? ? temperature the ? ? first ? ? turbine ? ? entrance ? ? temperature . A fan tip temperature rise is defined as: the ? ? fan ? ? tip ? ? rotor ? ? exit ? ? temperature the ? ? fan ? ? rotor ? ? entry ? ? temperature .

Abstract: A mount assembly for attaching a heat exchanger to a casing of a gas turbine engine comprises a first attachment feature by which in use the mount assembly is attached to the casing and a second attachment feature spaced from the first attachment feature by which in use the mount assembly is attached to the heat exchanger. The first and second attachment features are joined by an elongate member. The assembly is characterised in that the elongate member is significantly larger in a length direction and a height direction than in a thickness direction, so that it is relatively stiff in the length and height directions L, H and relatively flexible in the thickness direction T; the flexibility allowing movement in use within the mount assembly to accommodate differential thermal expansion between the heat exchanger and the casing.