Abstract: A blade for use in a gas turbine engine, the blade comprising a blade portion and a fir tree root portion, the blade portion and the root portion having a connected passage for allowing cooling air to flow within the blade, the fir tree root portion having an air intake on its leading edge, the air intake allowing cooling air to enter the cooling passage and wherein the fir tree root portion comprises a plurality of projections, including at least a base projection and a top projection; and wherein the air inlet located in the base projection of the fir tree root portion and wherein the air inlet comprises at least 50% of the face of the base projection of the fir tree root portion.

Abstract: An aircraft has first and second fuel sources containing fuels with different characteristics, and one or more gas turbine engines powered by the fuels and each having a staged combustion system having pilot and main fuel injectors and being operable in pilot-only and pilot-and-main ranges of operation. The gas turbine engines each have a fuel delivery regulator arranged to control fuel delivery to the pilot and main fuel injectors. The method includes: obtaining a proposed mission description; obtaining nvPM impact parameters for the gas turbine engines, the impact parameters being associated with each operating condition of the proposed mission; calculating an optimised set of one or more fuel characteristics for each flight condition of the proposed flight defined in the flight description based on the nvPM impact parameters; and determining a fuel allocation based on the optimised set of one or more fuel characteristics.

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 combustion system comprising: a combustion chamber extending in an axial direction between an inlet and an outlet, the combustion chamber configured to receive an airflow through the inlet and to discharge the airflow through the outlet; a fuel injection port configured to inject fuel into the airflow to form an air-fuel mixture; an ignition system for igniting the air-fuel mixture in the combustion chamber, the ignition system comprising an array of electrical plasma initiation points disposed downstream of the fuel injection port, and distributed radially and circumferentially around the combustion chamber, wherein each electrical plasma initiation point comprises a pair of electrodes configured to apply a voltage across an electrode gap between the pair of electrodes to produce plasma within the air-fuel mixture passing between the electrodes, thereby igniting the air-fuel mixture.

Abstract: A method of optimising the performance of combustion equipment of a gas turbine engine includes providing a fuel flow into the combustion equipment via a plurality of fuel injectors circumferentially disposed about a principal rotational axis at a plurality of injector positions; determining a plurality of temperatures of combustion gases at a plurality of circumferential positions downstream of the combustion equipment using a plurality of temperature measurement devices; ranking the plurality of circumferential positions based on the plurality of temperatures of the combustion gases determined using the plurality of temperature measurement devices; and repositioning at least some of the plurality of fuel injectors between the plurality of injector positions based at least on the ranking of the circumferential positions.

Abstract: A turbofan gas turbine engine having a bypass duct, and a bypass airflow measurement system. The bypass airflow measurement system comprises: at least one acoustic transmitter configured to transmit an acoustic waveform across the bypass duct of the gas turbine engine though which a bypass airflow passes to at least one acoustic receiver; where the at least one acoustic transmitter and the at least one acoustic receiver are located on an axial plane that is substantially perpendicular to the bypass flow. A method of measuring bypass airflow properties of a turbofan gas turbine engine is also described.

Abstract: A bladed disc system for a turbine engine having a disk portion and a plurality of blade portions which are associated with a stator section and an intercavity sealing portion, disc portion shaped such that blade portions are able to fit within firtree slot in disc portion, blade portion having aerofoil section and root section, aerofoil section having portion shaped such that they extend proximate to intercavity sealing portion, disc portion extending from portion that connects with drum to outer edge at which blade portions are connected with disc portion having width transition region in which thickness of disc increases from point at which disc connects to drum to outer edge at which it holds blade portions, and wherein width transition region has curved width transition region with radius r, and an overhanging portion which extends into the intercavity spacing between the width transition region and the intercavity sealing portion.

Abstract: A method of operating an aircraft including a gas turbine engine and a plurality of fuel tanks arranged to provide fuel to the gas turbine engine, where at least two of the fuel tanks contain fuels with different fuel characteristics. The method includes obtaining a flight profile for a flight of the aircraft; and determining a fuelling schedule for the flight based on the flight profile and the fuel characteristics. The fuelling schedule governs the variation with time of how much fuel is drawn from each tank. Fuel input to the gas turbine engine may then be controlled in operation in accordance with the fuelling schedule.

Abstract: A fuel spray nozzle includes a fuel circuit having in series a gallery, circumferentially spaced passages arranged in a row around the nozzle, and an annular spin chamber. Each passage has an inlet for receiving a respective portion of the fuel from the gallery and a metering orifice for discharging its portion of the fuel. The passages are configured such that, when the flow of liquid fuel to the inlet port is shut off, a respective differential static pressure develops across stagnant liquid fuel remaining between the inlet and the metering orifice of each passage, and the passages are further configured such that one or more selected passages develop a different differential static pressure to the remaining passages causing a flow of purging air to enter the gallery from the combustor through the selected passages and exit through the remaining passages, thereby purging the gallery and the passages of fuel.

Abstract: A gas turbine engine comprises a fan; an engine core comprising a compressor, a combustor, and a turbine coupled to the compressor through a shaft; a reduction gearbox; an Engine Section Stator (ESS) comprising a plurality of ESS vanes with an external surface washed by the core airflow; an ESS heating system adapted to heat the ESS vanes, and a temperature sensor adapted to detect the temperature of the external surface of the ESS vanes and send a signal to the ESS heating system when said temperature is below a reference temperature. Upon detection and/or inference of ice crystal conditions and receiving from the temperature sensor the signal that the temperature is below the reference temperature, the ESS heating system is activated to heat at least a portion of the external surface of the ESS vanes and promote melting and adhering of ice crystals thereto.

Abstract: Electrical power systems having variable-voltage DC busses and methods of controlling voltage settings of such busses. One electrical power system comprises: a variable-voltage DC bus; a number N?2 of electrical machines, each electrical machine connected to the variable-voltage DC bus via one of a corresponding number N of converters, each electrical machine and corresponding converter having an index n=(1, . . . , N); and a controller configured to select a voltage setting Vdc_bus for the variable-voltage DC bus and to provide control signals to the converters to control the voltage setting of the variable-voltage DC bus according to the selected voltage setting Vdc_bus. The controller configured to select a voltage setting Vdc_bus greater than or equal to a minimum voltage requirement Vdc_min for the bus. The controller is configured to determine the minimum voltage requirement Vdc_min using present operating speeds of each of the N electrical machines.

Abstract: A stator for a radial-flux rotary electric machine includes a yoke with a plurality of teeth, each tooth extending from a root at the yoke to a tip. The stator further includes a plurality of coils, each one of which being located on a respective one of the plurality of teeth, the coils being sized such that a slot fill factor is between zero and unity to define axial voids in the slots. The stator further includes a plurality of magnetic sealing members, each one of which being engaged with the tips of adjacent teeth to circumferentially and axially seal the axial voids in the slots for retaining an axial flow of coolant therein, the magnetic sealing members being composed of a magnetic material to provide a flux path between adjacent tooth tips to increase an amount of leakage flux.

Abstract: A computer-implemented method controls input of at least a portion of a first training data set into a first machine learning algorithm. The first training data set includes data quantifying damage to a first compressor and data quantifying a first operating parameter of the first compressor. The first machine learning algorithm is executed, and data quantifying the first operating parameter is received as an output of the first machine learning algorithm. The first machine learning algorithm is trained using the received data output from the first machine learning algorithm and data quantifying the first operating parameter of the first compressor. The trained first machine learning algorithm is configured to enable determination of operability of a second compressor of a gas turbine engine.

Abstract: A system for capture of atmospheric carbon dioxide. The system includes a carbon dioxide capture device having a capture medium operable to capture atmospheric carbon dioxide under a first set of thermodynamic conditions, and to release carbon dioxide under a second set of thermodynamic conditions different to the first set of thermodynamic conditions. The first set of thermodynamic conditions includes a first temperature range and the second set of thermodynamic conditions includes a second temperature range. A heat engine is configured receive a working fluid at an inlet temperature and arranged to discharge at least a portion of the working fluid at a discharge temperature in the second temperature range or higher, and a heat transfer arrangement is configured to convey heat from discharged working fluid to the capture medium to release carbon dioxide therefrom.

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, and a gearbox. The turbine is the lowest pressure turbine of the engine, and has a turbine length that is the distance between the root of the most upstream blade of the turbine at its leading edge and the root of the most downstream blade of the lowest pressure turbine at its trailing edge. The engine core further includes three bearings arranged to support the core shaft. The three bearings include a forward bearing and two rearward bearings, with a minor span being defined as the axial distance between the two rearward bearings. A ratio of the minor span to a turbine length is equal to or less than 1.05.

Abstract: Gearboxes for aircraft gas turbine engines, in particular to arrangements for journal bearings such gearboxes, and to related methods of operating such gearboxes and gas turbine engines. Example embodiments include a gearbox for an aircraft gas turbine engine, the gearbox including: a sun gear; a plurality of planet gears surrounding and engaged with the sun gear; and a ring gear surrounding and engaged with the plurality of planet gears, each of the plurality of planet gears being rotatably mounted around a pin of a planet gear carrier with a journal bearing having an internal sliding surface on the planet gear and an external sliding surface on the pin.

Abstract: An aerospace battery may include a housing, a retaining seat disposed in the housing, and a battery pack core supported by the retaining seat in the housing. The battery pack core may include at least one cell district comprising a plurality of pouch cell batteries and at least one flexible cold plate disposed between at least two adjacent pouch cell batteries of the plurality of pouch cell batteries. The retaining seat may define a fluid delivery channel configured to couple to an inlet of a cooling channel of the flexible cold plate and a fluid return channel configured to couple to an outlet of the cooling channel of the flexible cold plate.

Abstract: A gas turbine engine for an aircraft includes a staged combustion system having pilot fuel injectors and main fuel injectors. The gas turbine engine further includes a fuel delivery regulator arranged to control delivery of fuel to the pilot and main fuel injectors, and a fuel characteristic determination module configured to determine one or more fuel characteristics of the fuel being supplied to the staged combustion system. A controller is configured to determine a staging point defining the point at which the staged combustion system is switched between pilot-only operation and pilot-and-main operation, the staging point being determined based on the determined one or more fuel characteristics, the controller being configured to control the staged combustion system according to the determined staging point.

Abstract: In a gas turbine engine of the type having a high-pressure (HP) spool and a low-pres sure (LP) spool, methods of increasing surge margin and compression efficiency at a given thrust are provided. One method increases compression efficiency and includes transferring mechanical power from the HP spool to the LP spool to reduce a corrected speed of a HP compressor therein and raise a working line of a LP compressor therein. Another method increases surge margin and includes transferring mechanical power from the LP spool to the HP spool to increase a corrected speed of a HP compressor therein and lower a working line of a LP compressor therein.

Abstract: A method of operating a gas turbine engine including an engine core including a turbine, compressor, combustor to combust a fuel, and core shaft connecting the turbine and compressor; a fan upstream of the engine core; a fan shaft; a gearbox that receives an input from the core shaft and outputs drive to the fan via the fan shaft; a primary oil loop system to supply oil to the gearbox; and a heat exchange system. The method includes controlling the heat exchange system to adjust fuel viscosity to be lower than or equal to 0.58 mm2/s on entry to the combustor at cruise conditions.