Abstract: There is provided a continuum robot comprising a first end, a second end and an elongate body, a sensor arranged at the first end and a cooling jacket adjacent the sensor. The cooling jacket comprises a cavity containing a cooling medium for absorbing heat from the sensor.

Abstract: A component for a gas turbine engine, comprising: first and second walls; a coolant channel defined by the space between the first and second walls; and a first rib extending between the first and second walls to the end of the coolant channel in a coolant flow direction, such that the coolant channel is bifurcated in the coolant flow direction.

Abstract: Turbine vane assemblies incorporating both metallic and ceramic matrix composite materials are provided in the present disclosure. The turbine vane assemblies further include interface components that allow for differing rates of thermal expansion in the ceramic matrix composite components and the metallic components.

Abstract: A synchronous electrical machine includes a stator having a circumferential row of teeth carrying stator windings. The electrical machine further includes a coaxial rotor having a circumferential row of poles carrying field windings. Each pole has first and second side flanks and a tip surface which extends continuously there between to form an air gap to the teeth of the stator. On a transverse cross-section through the machine, one of the side flanks forms a leading edge of the pole, and the other side flank forms an opposite, trailing edge of the pole. Each pole carries a row of permanent magnets which extends across the tip surface from one of the side flanks to the other side flank, the permanent magnets compensating for armature magnetic reaction. The radial thickness of the permanent magnets increases smoothly with progressive distance across the tip surface from one of the side flanks to the other side flank.

Abstract: A turbine vane assembly adapted for use in a gas turbine engine includes a support strut and a turbine vane arranged around the support strut. The support strut is made of metallic materials. The turbine vane is made of ceramic matrix composite materials to insulate the metallic materials of the support strut.

Abstract: A planetary gearing includes a sun gear rotating about a rotation axis and driven by a sun shaft; planet gears driven by the sun gear; a ring gear engaging the planet gears; and a plurality of planet pins that respectively include an outer-side abutment surface having an axially forward end and an axially rearward end. Respectively, one planet pin is arranged inside a planet gear, and the planet pin and the planet gear form a lubricated journal bearing. At an axially forward face side and/or axially rearward face side, each planet gear forms a recess that extends inside the planet gear starting from the face side. The planet pins respectively form a crowning at their abutment surface such that their outer diameter decreases from a maximum outer diameter towards at least one axial end of the abutment surface, and has a minimum at the axial end.

Abstract: A gas turbine engine may include a combustor having an inner wall and an outer wall defining a combustion chamber there between. The inner wall and the outer wall may each have at least one opening into the combustion chamber. The gas turbine engine may also include at least one mobile conduit through which a cooling fluid may flow. The mobile conduit may pass through the combustion chamber from the at least one opening in the outer wall to the at least one opening in the inner wall. The gas turbine engine may further include a first joint and a second joint fluidly connecting the mobile conduit to the at least one opening in the inner wall and the at least one opening in the outer wall, respectively. The first joint and the second joint may enable multiple degrees of freedom of the mobile conduit within the combustion chamber.

Abstract: A turbine section adapted for use in a gas turbine engine includes a turbine case, a turbine wheel, and a turbine vane assembly. The turbine case is made from metallic materials. The turbine wheel is housed in the case. The turbine vane assembly is fixed to the case and configured to smooth and redirect air moving along a primary gas path of the turbine section.

Abstract: The disclosure relates to a ceramic matrix composite (CMC) aerofoil. Example embodiments include an aerofoil comprising first and second tubular CMC cores (302, 303) extending along a longitudinal axis of the aerofoil; and an outer CMC layer surrounding the first and second tubular CMC cores (302, 303) and defining an outer shape of the aerofoil having leading and trailing edges, wherein fibres within a wall of the second tubular CMC core extend to the trailing edge of the aerofoil.

Abstract: A power system includes a synchronous electrical generator having a rotor driven by a shaft; a permanent magnet signaling generator, coupled to the shaft; and an angle computation unit configured to calculate a rotor angle or load angle based on a voltage from the permanent magnet signaling generator and a voltage from the synchronous electrical generator.

Abstract: The disclosure relates to a valve for a fuel system having a body with at least one inlet and one outlet, the inlet fluidly connected to a pressurised fuel source in use. A shuttle is mounted within the body, the shuttle having a cavity of fixed volume and movable between a first position where fluid is permitted to flow through the inlet and is prevented from flowing through the outlet and a second position where fluid is prevented from flowing through the inlet and is permitted to flow through the outlet. A piston is configured to engage the fluid within the shuttle cavity to move the shuttle between the first and second position. A biasing mechanism biases the shuttle towards the first position and where the shuttle moves towards the second position when the fluid within the shuttle reaches a critical pressure.

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 method of forming a protective sheath for an aerofoil component includes: providing a first sheath portion and a second sheath portion, the first sheath portion and the second sheath portion each comprising an inner surface, an outer surface and an end surface between the inner and outer surfaces and having a sacrificial flange at its distal end; positioning the first sheath portion and second sheath portion so that the inner surface of the first sheath portion abuts against the inner surface of the second sheath portion with the end surfaces of the first and second sheath portions aligned to form a mating edge; and joining the first sheath portion to the second sheath portion by welding along the mating edge, wherein the sacrificial flanges are completely consumed and a curved outer profile is formed.

Abstract: Described is a turbine arrangement of a gas turbine engine, the turbine comprising: a rotor comprising a disc which incorporates a plurality of circumferentially spaced turbine blades, each blade having an aerofoil extending for a platform and a root portion securing the blade to the disc, a lock plate located at an axial end of the root portion of the blade, the lock plate comprising: a plate body having an inward facing surface which is located adjacent the blade root and an outward facing surface, radially inner and outer edge portions, and first and second circumferential edges; a head located at the radially outer edge portion of the plate body; a foot portion located at the radially inner portion of the plate body; a projection extending from the outwardly facing surface away from the turbine blade root in use, the projection being located adjacent to the foot portion, wherein the foot portion and head portion are located within slots provided at least partially by the disc and turbine blade respectivel

Abstract: An investment casting pattern including: a consumable body formed of: a first body portion; and a second body portion; and a fastener arranged to form an interface between the first body portion and the second body portion, and to locate the first body portion relative to the second body portion, wherein the fastener is formed of a different material than the first body portion and the second body portion.

Abstract: Apparatus for providing thermal energy to an article, the apparatus comprising: a first flexible heater configured to emit infrared radiation; and a first flexible member comprising a material configured to absorb the infrared radiation emitted by the first flexible heater and to generate thermal energy from the absorbed infrared radiation, the first flexible member being configured to transfer the generated thermal energy to the article through thermal conduction.

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 ? ? ( 120 ) is in a range between 0.8 rpm/mm to 5 rpm/mm.

Abstract: A slot is provided in an endwall of a flow passage, for example between two stator vanes or rotor blades of a gas turbine engine. The length direction of the flow passage is aligned substantially with the main flow through the flow passage. The alignment of the slot means that the “over-turned” boundary layer flow can be extracted through the slot but with minimal impact on the mainstream flow.

Abstract: A rotor assembly with a rotor hub, a plurality of permanent magnets positioned circumferentially around at least a portion of the rotor hub, and a coating on at least a portion of the plurality of permanent magnets. The coating forming a retaining band that circumferentially extends around the rotor hub and the plurality of permanent magnets. The coating includes a nano-crystalline layer including a metal or metal alloy and defines an average grain size of less than about 50 nanometers (nm).

Abstract: A gas turbine engine includes a fan mounted to rotate about a main longitudinal axis; an engine core, including a compressor, a combustor, and turbine coupled to the compressor through a shaft; and reduction gearbox; wherein the compressor includes a plurality of stages, each stage including a respective rotor and stator, a first stage of the plurality of stages being arranged at an inlet and including a first rotor with a plurality of blades; each blade extending chordwise from a leading edge to a trailing edges, and from root to tip for a span height, wherein 0% of the span height corresponds to the root and 100% of span height corresponds to tip; wherein the ratio of a maximum leading edge radius of curvature of the first rotor blades to a minimum leading edge radius of curvature of the first rotor blades is included between 2.2 and 3.5.