Abstract: A gas turbine engine for an aircraft comprises, in axial flow sequence, a compressor module, a combustor module, and a turbine module, with a first electric machine being rotationally connected to the turbine module. The first electrical machine is configured to generate a maximum electrical power PEM1 (W), and the gas turbine engine is configured to generate a maximum dry thrust T (N); and a ratio S of: S=(Maximum Electrical Power Generated=PEM1)/(Maximum Dry Thrust=T) is in a range of between 2.0 and 10.0.

Abstract: An aircraft turbofan gas turbine engine includes a fan assembly, a compressor module, and a turbine module. An electric machine positioned downstream of the fan assembly is rotationally connected to the turbine module. The fan assembly is in fluid communication with the compressor module by an intermediate duct and includes a highest pressure fan stage having a plurality of fan blades defining a fan diameter. The compressor module includes a lowest pressure compressor stage having a row of rotor blades. An intermediate flow axis is defined joining a radially outer tip of a trailing edge of one of the fan blades of the highest pressure fan stage, and a radially outer tip of a leading edge of one of the rotor blades of a leading edge of a lowest-pressure compressor blade. An intermediate flow axis angle and the intermediate flow axis angle is from ?20 to ?30 degrees.

Abstract: A turbine generator comprising a turbine rotor comprising a hub and one or more blade stages. Each stage comprising a circumferential array of rotor blades in driving engagement with the hub. A turbine stator comprising a hub and one or more vane stages, each stage comprising a circumferential array of vanes. The turbine rotor and turbine stator being concentrically arranged about a common axis to define an annular flow path. The vane stages and blade stages being axially spaced along the axis and having one or more magnets arranged on the rotor. A generator stator concentrically aligned with the turbine rotor and turbine stator and one or more magnets arranged on the rotor. In use, when the turbine is driven to rotate about the axis, the or each of the magnets on the turbine rotor rotate relative to the generator stator in order to generate electric power.

Abstract: A planetary gear box for a gas turbine engine, has a planet carrier with a carrier element and a planet gear. The planet gear is mounted rotatably via a plain bearing on the carrier element. An oil feed pocket is provided in the region of an outer side of the carrier element, via which feed pocket oil can be passed into a bearing gap between the outer side of the carrier element and an inner side of the planet gear. In the circumferential region of a main load direction of the plain bearing, the carrier element is formed with at least one channel carrying transmission oil. The channel extends in the carrier element radially inside the outer side of the carrier element, extending in the axial direction of the carrier element, and includes a cross section closed with respect to the bearing gap.

Abstract: The disclosure relates to an electrical power system for providing a stabilised DC voltage to a power bus. Example embodiments include an electrical power system comprising: a DC power bus having first and second DC power bus terminals; an electrical storage unit having first and second terminals, the second terminal connected to the second DC power bus terminal; a DC:DC converter having first and second DC:AC converters and a transformer connected between the first and second DC:AC converters, the first DC:AC converter connected between the first terminal of the electrical storage unit and the first DC power bus terminal; and a controller connected to control a switching operation of one or both of the first and second DC:AC converters.

Abstract: An electrical system having a first dual-wound rotary electric machine mechanically coupled with a first gas turbine spool and having a first three-phase sub-machine and a second three-phase submachine; a second dual-wound rotary electric machine mechanically coupled with a second gas turbine spool and having a third three-phase sub-machine and a fourth three-phase submachine; and a set of N = 4 bidirectional converter circuits for conversion of alternating current (ac) to and from direct current (dc), each of which has an associated index n = (1, ... , N), and for all n, the ac side of the nth bidirectional converter circuit is connected with the nth three-phase sub-machine. The dc side of the first converter circuit is connected with the dc side of the third converter circuit, and the dc side of the second converter circuit is connected with the dc side of the fourth converter circuit.

Abstract: A composite storage tank comprises a composite wall enclosing a gas storage volume and defining a cylindrical portion of the tank. The composite wall incorporates first and second sets of metallic fibres each of which is susceptible to embrittiement by hydrogen and has ends extending through the exterior surface of the composite wall. By measuring the electrical resistances of the metallic fibres, a measure of the amount of hydrogen that has leaked through the composite wall over a period of time, and the present physical condition of the tank, may be determined. The approximate axial and azimuthal coordinates of a particular leakage point may also be determined.

Abstract: In general, techniques are described regarding a rotary servo for actuators. A servo assembly includes a cylindrical outer sleeve including ports, a cylindrical outer spool annularly disposed within the cylindrical outer sleeve, a stepper motor mechanically coupled to the cylindrical outer spool, and an actuator mechanically coupled to compressor variable geometry that controls compression provided by a compressor. The cylindrical outer spool includes channels configured to provide fluidic interconnection between the ports and a cylindrical inner spool, where the cylindrical inner spool is annularly disposed within the cylindrical outer spool, and the cylindrical inner spool includes grooves configured to provide fluidic interconnection through the channels of the cylindrical outer sleeve. The stepper motor is configured to rotate the cylindrical outer spool within the cylindrical outer sleeve to deliver a fluid to and thereby actuate the actuator to control the compressor variable geometry.

Abstract: A method includes predicting a composition of calcium-magnesium-aluminum-silicate (CMAS) to be encountered by a high temperature mechanical system during use of the high temperature mechanical system. The method further includes selecting a composition of a CMAS-resistant barrier coating layer based at least in part on the predicted composition of CMAS. The CMAS-resistant barrier coating layer includes a base composition and at least one secondary oxide selected based on the predicted composition of CMAS. The at least one secondary oxide includes at least one of an oxide of a divalent element, an oxide of a trivalent element, or an oxide of a tetravalent element. The CMAS-resistant barrier coating layer comprises greater than 0 mol. % and less than about 7 mol. % of the at least one secondary oxide.

Abstract: An electric machine comprises a stator winding located within a stator housing. The stator housing comprises an inlet chamber configured to communicate with a supply of coolant, and an outlet. The stator housing contains a baffle configured to separate the stator winding and outlet from the inlet chamber. The baffle includes at least one aperture therethrough configured to supply an impingement flow of coolant to the stator winding.

Abstract: A turbofan gas turbine engine includes, in axial flow sequence, a heat exchanger module, a fan assembly, a compressor module, a turbine module, and an exhaust module. The fan assembly includes a plurality of fan blades defining a fan diameter. The heat exchanger module is in fluid communication with the fan assembly by an inlet duct, and the heat exchanger module including a plurality of heat transfer elements for transfer of heat from a first fluid contained within the heat transfer elements to an airflow passing over a surface of the heat transfer elements prior to entry of the airflow into an inlet to the fan assembly. Each heat transfer element may be individually and independently fluidly isolated from the remaining heat transfer elements.

Abstract: A turbofan gas turbine engine includes, in axial flow sequence, a heat exchanger module, a fan assembly, a compressor module, a turbine module, and an exhaust module. The fan assembly includes fan blades defining a fan diameter. The heat exchanger module is in communication with the fan assembly by an inlet duct, and the heat exchanger module further includes radially-extending hollow vanes arranged in a circumferential array, with a channel extending axially between hollow vanes. Each hollow vane accommodates at least one heat transfer element to transfer heat from a first fluid contained within the or each heat transfer element to a corresponding vane airflow passing through the hollow vane and over a surface of the or each heat transfer element. Each hollow vane further includes a flow modulator configured to regulate airflow in proportion to total airflow entering the heat exchanger module in response to a user requirement.

Abstract: An aircraft comprises a machine body. The machine body encloses a turbofan gas turbine engine and a plurality of ancillary systems. The turbofan gas turbine engine comprises, in axial flow sequence, a heat exchanger module, a fan assembly, a compressor module, a combustor module, a turbine module, and an exhaust module. The machine body comprises a single fluid inlet aperture, with the fluid inlet aperture being configured to allow a fluid cooling flow to enter the machine body and to pass through the heat exchanger module. The heat exchanger module is configured to transfer a waste heat load from the gas turbine engine and the ancillary systems to the fluid cooling flow prior to an entry of the entire fluid cooling flow into the fan module.

Abstract: A method of inspecting a surface of a component, e.g. a turbine or compressor blade of a gas turbine engine. The method comprises (a) providing a probe for inspecting the component surface; (b) defining a reference surface that is offset from the component surface; (c) moving the probe so as to contact a plurality of discrete spaced apart inspection points on the component surface, each contact of the probe with an inspection point comprising a first movement of the probe from the reference surface to the inspection point; (d) retracting the probe from the component surface after each contact with an inspection point; and (e) inspecting the component surface each time the probe contacts an inspection point.

Abstract: A turbofan gas turbine engine includes, in axial flow sequence, a heat exchanger module, a fan assembly, a compressor module, a turbine module, and an exhaust module. The fan assembly includes a plurality of fan blades defining a fan diameter (D). The heat exchanger module is in fluid communication with the fan assembly by an inlet duct, and the heat exchanger module includes a plurality of radially-extending hollow vanes arranged in a circumferential array with a channel extending axially between each pair of adjacent hollow vanes. An airflow entering the heat exchanger module is divided between a set of vane airflows through each of the hollow vanes and a set of channel airflows through each of the channels.

Abstract: A method of manufacturing a composite laminate. The method comprises providing a base layer, providing a discontinuous reinforcing patch on the base layer, and providing a top layer over the base layer and discontinuous reinforcing patch. Also, a composite laminate having a discontinuous reinforcing patch interposed between a base layer and a top layer. The discontinuous reinforcing patch comprises a patterned nanomaterial layer with nanomaterial-filled zones and vacant zones.

Abstract: A computer-implemented method of enabling optimisation of derate for a propulsion system of a vehicle, the method comprising: determining a derate for the propulsion system of the vehicle using: an algorithm; a vehicle model defining path constraints for the vehicle through space; a propulsion system model defining parameters of the propulsion system; an objective function defining one or more objectives; and controlling output of the determined derate.

Abstract: A propulsion system for an aircraft includes a gas turbine engine and a fuel tank, wherein the fuel includes at least a proportion of a sustainable aviation fuel—SAF—having a density between 90% and 98% of the density, ?K, of kerosene and a calorific value between 101% and 105% the calorific value CVK, of kerosene. The engine includes a combustor; and a fuel pump arranged to supply a fuel thereto at an energy flow rate, C, the pump being arranged to output fuel at a volumetric flow rate, Q, the percentage of fuel passing through the pump not provided to the combustor being referred to as a spill percentage. The fuel include X % SAF, where X % is in the range from 5% to 100%, and has a density, ?F, and a calorific value CVF.

Abstract: Electrical power systems for distributing electrical power in aircraft are described. One such electrical power system comprises: an electrical power source configured to output a number R?2 of dc power channels, each dc power channel having a respective index r=(1, . . . , R); and a group of N dc load channels connected to the R dc power channels by a switching arrangement, wherein each dc load channel has a respective index n=(1, . . . , N) and R<N?2R. The switching arrangement is operable to connect the r-th power channel to the n-th load channel according to a relationship r(n).

Abstract: A gas turbine engine configured with an engine core. A fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox arranged to receive an input from the core shaft and to output drive to the fan so as to drive the fan at a lower rotational speed than the core shaft. The gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier on which the planet gears are mounted, the planet carrier having an effective linear torsional stiffness and the gearbox having a gear mesh stiffness between the planet gears and the ring gear. Additionally, the product of the effective linear torsional stiffness of the planet carrier and the gear mesh stiffness between the planet gears and the ring gear is greater than or equal to 5.0×1018N2m?2.