Abstract: This invention concerns a cryo-cooled electrical conduction network. The conduction network has an electrical network divided into two or more conductive sections, each section comprising electrical equipment (24, 28). The conductive network also has a coolant network for maintaining the temperature of a coolant in each section. The electrical equipment (24, 28) and a corresponding portion of the coolant network of each section is housed in a section enclosure (10, 12, 14). The coolant network includes a coolant interface (40) located between each section, wherein the coolant interface (40) is housed in an intermediate enclosure (16, 18, 20, 22) that is isolatable from the section enclosures (10, 12, 14) in the electrical conduction network.

Abstract: A variable diffuser comprises a passage, at least two vanes disposed within the passage, and at least two pennies. The passage is defined between opposing disk faces of a hub and a tip. Each of the vanes comprises a body having a leading edge and a trailing edge. The body extends between the hub face and the disk face. Each of the pennies is coupled to a respective vane body near an edge of the penny and an actuator. Rotation of at least one penny changes the orientation of the respective vane relative to the hub face.

Abstract: An electric ducted fan for an aircraft is shown. A nacelle defines a duct that houses a propulsive fan having a fan diameter DF. An electric machine is configured to drive the fan, and has an electromagnetically active length LA and an electromagnetically active diameter DA defining an aspect ratio (LA/DA) of from 0.8 to 2. A ratio of the electromagnetically active diameter DA to the fan diameter DF (DA/DF) is from 0.45 to 0.60.

Abstract: A hybrid aircraft propulsion system. The system comprises a gas turbine engine comprising a compressor, a combustor, one or more turbines, a shaft coupled to one of the turbines, and a bypass fan mechanically driven by the shaft. The system further comprises an electrical generator mechanically coupled to the shaft, and an auxiliary propulsor mechanically coupled to an electric motor and electrically coupled to the electric generator. At maximum power, the gas turbine engine is configured to produce a turbine entry temperature at maximum power between 1800 Kelvin and 2000 Kelvin, the engine comprises a fan bypass ratio of between 4:1 and 13:1, and the generator is configured to absorb between 10% and 60% of the mechanical power generated by the turbine.

Abstract: The present disclosure concerns control a hybrid electric gas turbine system (300) for an aircraft. The system comprises an electric generator (308) and a gas turbine (309) to form a generator system, an electric motor (303) and a fan (302) to form a propulsor (301), a controller (306) and an electric storage unit (307). After receiving a command for a change in demand for thrust, the controller (306) determines an operational profile that minimises a function comprising a measure of fuel supplied to the gas turbine (309), a transfer of electric power from or to the electric storage unit (307) and a difference between measures of current and demanded thrust over a time period. The controller then operates the electric motor (303), gas turbine (309) and electric storage unit (307) according to the determined operational profile over the time period.

Abstract: An inter-blade vibration damper for a gas turbine engine has an elongate damper body. The damper body is formed as a truncated cone and has a longitudinal axis. The elongate damper body is formed as a truncated cone. A conic surface of the damper body contacts a portion of a first blade and a portion of an adjoining second blade.

Abstract: A thermochromic paint including: 1 wt % to 90 wt % binder; 2 wt % to 40 wt % filler comprising an inorganic oxide; and 0.1 wt % to 28 wt % organic compound including a chromophore. The paint provides an irreversible indication of maximum temperature.

Abstract: A variable diffuser comprises a passage, at least two vanes disposed within the passage, and at least two pennies. The passage is defined between opposing disk faces of a hub and a tip. Each of the vanes comprises a body having a leading edge and a trailing edge. The body extends between the hub face and the disk face. Each of the pennies is coupled to a respective vane body near an edge of the penny and an actuator. Rotation of at least one penny changes the orientation of the respective vane relative to the hub face.

Abstract: The present disclosure concerns an electric aircraft propulsion system. Example embodiments include an aircraft propulsion system (200) comprising: a fan (213) mounted on a central shaft (214); a first electric motor (215) mounted on the central shaft (214) and arranged to drive rotation of the fan (213) via the central shaft; and a second electric motor (216) arranged to drive rotation of the fan (213) via the central shaft (214), wherein the second electric motor (216) is coaxially mounted downstream of the first electric motor (215) and an outer diameter (D2) of the second electric motor (216) is smaller than an outer diameter (D1) of the first electric motor (215).

Abstract: Method and system for cryptographic transformation of a structured data set. The structured data set is partitioned into a first subset and a plurality of further subsets for encryption in parallel. The subsets are divided into a plurality of blocks of predetermined size. A first block for each subset is identified as well as a location of each further block in said subset relative to said first block of its subset. Cryptographic transformation of the data subsets is performed using a key according to a block chain process and an offset value for the first block of each subset from the first block of the first subset is logged. The process allows a block chain to be broken into part way in the chain. The process may allow different partitioning to be used for decryption than was used for encryption, thereby allowing parallel processing on varying numbers of computational cores.

Abstract: An electric ducted fan for an aircraft is shown. A nacelle defines a duct that houses a propulsive fan having a fan diameter DF. An electric machine is configured to drive the fan, and has an electromagnetically active length LA and an electromagnetically active diameter DA defining an aspect ratio (LA/DA) of from 0.8 to 2. A speed reduction device is located between the electric machine and the fan, and has a reduction ratio of at least 3:1. A ratio of the electromagnetically active diameter DA to the fan diameter DF (DA/DF) is from 0.3 to 0.5.

Abstract: A fluid ejector having: a diffuser duct for mixing of a first gas stream with a second gas stream having faster velocity than the first gas stream, wherein the diffuser duct extends along an axial direction; and wherein, in cross-section perpendicular to the axial direction, the diffuser duct is defined between a first wall extending around the axial direction, and a second wall extending around the axial direction, radially within the first wall.

Abstract: A method for manufacturing a component having a defined geometry includes: a) defining a pre-component geometry including interim shape elements and additional, sacrificial elements for supporting interim elements; b) on a base plate, depositing multiple layers of a powder including a material from which the pre-component will be manufactured; c) sintering the powder to form the pre-component to the defined geometry; d) removing at least some of the sacrificial elements from the pre-component; e) subjecting the remaining pre-component from step d) to a HIP step; and f) removing remaining sacrificial elements from the pre-component product of step e) to provide a component to the defined component geometry. In the definition of the pre-component geometry, the interim elements differ from the corresponding final shape elements in the defined component geometry such that during the HIP step, the interim shape elements adjust to form final shape elements in the defined component geometry.

Abstract: A gas turbine engine fan blade apparatus is disclosed comprising a reverse pitch rotatable mounting and an aerofoil part. The aerofoil part comprises radially inner and radially outer portions that are separated by an interface extending from a leading edge of the aerofoil part to a trailing edge of the aerofoil part. The reverse pitch rotatable mounting locates the radially inner and radially outer portions with respect to one another. The reverse pitch rotatable mounting and separation at the interface permit relative pitch change rotation between the radially inner and radially outer portions such that one of the radially inner and radially outer portions is rotatable to a reverse pitch configuration by comparison with the other portion.

Abstract: A fan blade is provided with an aerofoil portion for which, at radii between 20% and 40% of the blade span, the location of the position of maximum thickness along the camber line is at less than a defined percentage of the total length of the camber line. For all cross-sections through the aerofoil portion at radii greater than 70% of the blade span, the location of the position of maximum thickness along the camber line is at more than a defined percentage of the total length of the camber line. The geometry of the fan blade may result in a lower susceptibility to flutter.

Abstract: Apparatus for insertion into a cavity of an object, the apparatus comprising: a first tube comprising a first end, a second end and a first cavity extending between the first end and the second; a member positioned within the first cavity of the first tube and configured to enable an action to be performed; a heat pipe including a first end, a second end, and a second cavity extending between the first end and the second end, the first tube being positioned within the second cavity of the heat pipe; and an actuator configured to move the first tube relative to the heat pipe.

Abstract: A louvre offtake arrangement for a gas turbine engine includes a first duct, where a primary flow flows; a second duct, defining an offtake, connected to the first duct at an inlet; and a plurality of louvres arranged at the inlet. At least one louvre protrudes from the second duct into the first duct to divert part of the primary flow towards the second duct.

Abstract: A method of detecting a shaft break in a turbine engine is provided. The shaft connects a torque-exerting turbine of the engine to a torque-receiving unit of the engine. The method includes: receiving repeated measurements of the rotational frequency of the torque-receiving unit and the rotational frequency measurements of the turbine; determining from the measurements an amount of shaft twist between the turbine and the torque-receiving unit; comparing the amount of twist with a threshold twist value; and signalling a shaft break when the amount of twist exceeds the threshold twist value and a pre-condition is met. The pre-condition includes determining from the measurements that the turbine rotational speed is greater than the simultaneous torque-receiving unit rotational speed by a threshold speed value.

Abstract: There is disclosed a method for spatiotemporal registration between a first image stream and a second image stream. A first registration frame set derived from the first image stream and a second registration frame set derived from the second image stream are defined. A spatial correlation analysis is conducted based on the first registration frame set and the second registration frame set to determine a spatial registration offset. The first registration frame set and the second registration frame set are spatially registered based on the spatial registration offset. A temporal correlation analysis is conducted based on the first registration frame set and the second registration frame set after spatial registration to determine a temporal registration offset. The first image stream and the second image stream are spatially and temporally registered based on the spatial registration offset and the temporal registration offset.

Abstract: A method locates and maintains a core in fixed space relationship within the interior of a shell mould. The method provides at least one pin extending into the core with at least one axial end of the pin protruding from the core. A wax pattern having an outer surface is formed by encasing the core and at least one protruding axial end of the pin in wax such that at least one protruding axial end of the pin terminates at the outer surface of the wax pattern. A shell mould is formed around the wax pattern such that, upon removal of the wax pattern, and in a subsequent casting process for production of hollow metal components, at least one protruding axial end of the pin abuts the shell mould, thus fixing the pin and maintaining a position of the core relative to the mould. The protruding axial end of the pin has an enlarged head.