Abstract: A turbofan engine has an engine core including in flow series a compressor, a combustor and a turbine. The engine further has a fan located upstream of the engine core, has a supersonic intake for slowing down incoming air to subsonic velocities at an inlet to the fan formed by the intake, has a bypass duct surrounding the engine core, wherein the fan generates a core airflow to the engine core and a bypass airflow through the bypass duct, and has a mixer for mixing an exhaust gas flow exiting the engine core and bypass airflow exiting bypass duct. The engine further has a thrust nozzle rearwards of the mixer for discharging mixed flows, the thrust nozzle having a variable area throat. The engine further has a controller controlling the thrust produced by the engine over a range of flight operations including on-the-ground subsonic take-off and subsequent off-the-ground subsonic climb.

Abstract: An engine for an aircraft has 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. The gearbox is arranged to receive an input from a gearbox input shaft portion of 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 is an epicyclic gearbox and has a sun gear, a plurality of planet gears, a ring gear, and a planet carrier. The planet carrier and the gearbox support each have a torsional stiffness, and a carrier to gearbox support torsional stiffness ratio is greater than or equal to 2.3.

Abstract: A thermal management system and method for regulating the dissipation of a thermal load during operation of a vehicle. The thermal management system including one or more cooling loops configured to regulate the temperature of at least one battery pack; and an energy recovery mechanism configured to recover energy dissipated upon the occurrence of a thermal runaway event. The amount of energy recovered maintains the power level at or above the level exhibited by the battery pack prior to the occurrence of the thermal runaway event. Upon the occurrence of the thermal runaway event, the energy recovery mechanism transforms the cooling loop into a Rankine cycle loop or uses a Seebeck effect to recover energy.

Abstract: A gas turbine engine for an aircraft comprises an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, wherein a compressor exit temperature is defined as an average temperature of airflow at the exit from the compressor; and a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub, each fan blade having a leading edge and a trailing edge, wherein a fan rotor entry temperature is defined as an average temperature of airflow across the leading edge of each fan blade at cruise conditions and a fan tip rotor exit temperature is 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 to fan tip temperature rise ratio is in the range from 2.845 to 3.8.

Abstract: A turbofan gas turbine engine comprises, in axial flow sequence, a heat exchanger module, a fan assembly, a compressor module, a turbine module, and an exhaust module. The fan assembly comprises 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 comprises a plurality of radially-extending hollow vanes arranged in a circumferential array with a channel extending axially between each pair of adjacent hollow vanes. The heat exchanger module has a square axial cross-sectional profile, where a side length of the square cross-section is D.

Abstract: A cooling system is provided including a two-phase pump loop and a vapor compression system. The two-phase pump loop cools a thermal load with a first coolant. The vapor compression system is configured to circulate a second coolant. The vapor compression system includes a liquid vapor separator which separates the second coolant into a liquid portion and a gaseous portion. The liquid vapor separator is a thermal energy storage for the two-phase pump loop. A condenser of the two-phase pump loop transfers heat from the first coolant to the liquid portion of the second coolant in the liquid-vapor separator.

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 turbine module, a combustor module, and an exhaust module. The machine body comprises either one or two fluid inlet apertures. The or each fluid inlet aperture is configured to allow a fluid 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 flow prior to an entry of the fluid flow into the fan module, and thence into the compressor module, the turbine module, the combustor module, and the exhaust module.

Abstract: An example hybrid aircraft propulsion system includes one or more power units configured to output electrical energy onto one or more electrical busses; a plurality of propulsors; and a plurality of electrical machines, each respective electrical machine configured to drive a respective propulsor of the plurality of propulsors using electrical energy received from at least one of the one or more electrical busses.

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: An aerofoil assembly includes a platform and a plurality of aerofoils extending radially outward from the platform. The platform has a first edge, a second edge, and a platform surface disposed between the first edge and the second edge. Each aerofoil has a leading edge proximal to the first edge and a trailing edge distal to the first edge. A pitch spacing is defined between the leading edges of adjacent aerofoils along the platform surface. A mid-pitch location is defined midway along the pitch spacing. The platform defines one or more recesses disposed between the leading edges of the plurality of aerofoils and the first edge. Each of the one or more recesses is disposed proximal to the mid-pitch location between adjacent aerofoils.

Abstract: An axial flow turbomachine (102) for producing thrust to propel an aircraft is shown. The turbomachine has an inner duct (202) and an outer duct (204), both of which are annular and concentric with one another. An inner fan (206) is located in the inner duct, and is configured to produce a primary pressurised flow (P). An outer fan (207) is located in an outer duct, and is configured to produce a secondary pressurised flow (S). The outer fan has a hollow hub (208) through which the inner duct passes. The inner fan is configured to have, in operation, a rate of rotation of from 3 to 8 times that of the outer fan.

Abstract: An axial flow turbomachine (102) for producing thrust to propel an aircraft is shown. The turbomachine has an inner duct (202) and an outer duct (204), both of which are annular and concentric with one another. An inner fan (206) is located in the inner duct, and is configured to produce a primary pressurised flow (P). An outer fan (207) is located in an outer duct, and is configured to produce a secondary pressurised flow (S). The outer fan has a hollow hub (208) through which the inner duct passes. The inner fan is configured to have, in operation, a tip speed of from 1 to 3 times that of the outer fan.

Abstract: A bi-directional direct current (DC) solid state power controller (SSPC) architecture and control method. The SSPC protects a DC distribution system by isolating both the positive and negative buses independently in case of short circuit or ground fault. The SSPC architecture includes two self-heal interleaved capacitors and includes a fast, soft-charging control technique that provides line-isolated charging of the DC bulk capacitor to avoid inrush current when powering up the DC distribution system. The soft-charging function alternately charges one of the two interleaved capacitors, while the other capacitor discharges to the DC bulk capacitor. Repetitive switching results in a charging and discharging process that increases the voltage of the DC bulk capacitor prior to powering up the DC distribution system, while keeping the DC power source isolated from the load.

Abstract: A monitoring system includes first and second, axially adjacent, phonic wheels formed from respective axially adjacent portions of a unitary annular body and mounted coaxially to the shaft for rotation therewith. The first and second phonic wheels have respective first and second circumferential rows of teeth. The system further includes a sensor configured to detect the passage of the first row of teeth by generating a first alternating measurement signal component, and to detect the passage of the second row of teeth by generating a second alternating measurement signal component. The sensor generates a signal having both the first and the second alternating measurement signal components when axially positioned midway between the first and second phonic wheels. The teeth of the first and second rows are configured such that the first alternating measurement signal component has an identical frequency to the second alternating measurement signal component.

Abstract: The present invention provides a rotor, an axial flux electrical machine and a hybrid-electrical or electrical air craft. The rotor for the axial flux electrical machine comprises first sections of a first material and second sections of a second material. The first sections and second sections are arranged in alternating order and ring-shape. The arrangement is characterized by third sections which form interface areas between the first sections and second sections, whereas the third sections are comprising the first material and the second material in such a manner that the first section and the second section are connected force-fitted.

Abstract: Gearboxes for aircraft gas turbine engines, in particular arrangements for journal bearings such gearboxes, and related methods of operating such gearboxes and gas turbine engines. A gearbox for an aircraft gas turbine engine includes: 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: A coolant for a subtractive machine process, the coolant comprising a phase change material which changes from a solid state to a liquid state as a result of frictional heat generated during the subtractive machine process of the machine tool on a component to be machined, and returns to the solid state as the component cools.

Abstract: A rotor assembly adapted for a gas turbine engine includes a shaft, a wheel, and a retaining nut. The shaft extends along an axis and includes a first tapered surface. The wheel is arranged circumferentially around the shaft and includes a second tapered surface. The retaining nut is fastened to the shaft and applies an axial force to the wheel to couple the wheel with the shaft.

Abstract: A boundary layer ingestion fan system for location aft of the fuselage of an aircraft is shown. It comprises a nacelle defining a duct, and a fan located within the duct. The fan comprises a hub arranged to rotate around a rotational axis (A-A) and a plurality of blades attached to the hub, each of which has a span from a root at the hub defining a 0 percent span position (rhub) to a tip defining a 100 percent span position (rtip) and a plurality of span positions therebetween (r?[rhub, rtip]). A hub-tip ratio of the fan, defined as the ratio of the diameter of the hub to the diameter of the fan measured at the leading edge of the blades, is from 0.45 to 0.55.

Abstract: There is provided a method of controlling a robot within an environment comprising: i) receiving, from a 3D scanner, data relating to at least a portion of the environment for constructing a 3D point cloud representing at least a portion of the environment; ii) comparing the 3D point cloud to a virtual 3D model of the environment and, based upon the comparison, determining a position of the robot; then iii) determining a movement trajectory for the robot based upon the determined position of the robot. Also provided is a control apparatus and a robot control system.