Abstract: A manufacturing process for blades of a turbomachine, e.g. a gas turbine engine for an aircraft. In the process: a) a ceramic core piece that comprises at least two ceramic core elements and a clamping part that connects the ceramic core elements, is positioned in a wax forming device, subsequently; b) a molten wax material is applied to the outside of the ceramic core piece in the wax forming device and the wax is allowed to solidify, and subsequently; c) at least two turbomachine blades are cast using a crystallographically-oriented metal casting process and the wax and the ceramic core piece are removed.

Abstract: A shaft component, which in particular can be connected or is connected to the input or output side of a gear box in a gas turbine engine, in particular an aircraft engine, wherein the shaft component has at least two regions comprising fiber reinforced plastic, with fibers in the at least two regions differing in their composition, their geometric properties, their density, their radial position, their axial position and/or in their fiber orientation in the shaft component.

Abstract: A gas turbine engine has a compression system radius ratio defined as the ratio of the radius of the tip of a fan blade to the radius of the tip of the most downstream compressor blade in the range of from 5 to 9. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

Abstract: A method of identifying a fuel contained in a fuel tank of an aircraft and arranged to power a gas turbine engine of the aircraft is performed by processing circuitry of the aircraft and includes: obtaining at least one fuel characteristic of any fuel already present in the fuel tank prior to refuelling; determining at least one fuel characteristic of a fuel added to the fuel tank on refuelling; and calculating at least one fuel characteristic of the resultant fuel in the fuel tank after refuelling. The method may further controlling the propulsion system of the aircraft based on the calculated at least one fuel characteristic of the resultant fuel in the fuel tank after refuelling.

Abstract: A method of determining one or more fuel characteristics of an aviation fuel for powering a gas turbine engine of an aircraft includes: measuring one or more trace substance parameters of the fuel, the one or more trace substance parameters each associated with a respective trace substance in the fuel; and determining one or more fuel characteristics of the fuel based on the one or more trace substance parameters. Further, a fuel characteristic determination system, a method of operating an aircraft, and an aircraft.

Abstract: A tube gallery for a gas turbine engine includes a body. The body includes an external surface. The body also includes a plurality of channels defined in the body. Each channel includes an inlet disposed on the external surface, an outlet spaced apart from the inlet and disposed on the external surface, and a passage extending between and fluidly communicating the inlet to the outlet. The passage of each channel has a non-circular cross-sectional shape. The non-circular cross-sectional shape has a first maximum dimension along a first direction and a second maximum dimension along a second direction orthogonal to the first direction. The first maximum dimension is greater than the second maximum dimension by a factor of at least 1.2.

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: A continuum arm robot system comprising at least a first continuum arm robot and a second continuum arm robot, each continuum arm robot being controlled by its own actuator pack, and each actuator pack being coupled to a single control computer, wherein at least the second continuum arm robot comprises a releasable connection mechanism to engage in gripping the first continuum arm robot in a workspace, so as to link the at least two continuum arm robots into a single redundant robotic system with at least the second continuum arm robot providing support for the first continuum arm robot.

Abstract: There is described a fixture for supporting a plurality of gas turbine engine blades that are to be imaged. The mobile fixture comprises a plurality of interconnected mounts arranged to lie within a common plane, wherein each mount comprises at least one support surface for holding a respective blade. Support surfaces of respective mounts are oriented such that their respective normal vectors have the same angle with respect to the common plane.

Abstract: A computer-implemented method comprising: receiving data comprising two-dimensional data and three-dimensional data of a component of an engine; identifying a feature of the component using the two-dimensional data; determining coordinates of the feature in the two-dimensional data; determining coordinates of the feature in the three-dimensional data using: the determined coordinates of the feature in the two-dimensional data; and a pre-determined transformation between coordinates in two-dimensional data and coordinates in three-dimensional data; and measuring a parameter of the feature of the component using the determined coordinates of the feature in the three-dimensional data.

Abstract: A nacelle for a gas turbine engine includes a leading edge at an upstream side of the nacelle. The nacelle further includes a trailing edge at a downstream side of the nacelle. The nacelle further includes an outer surface disposed between the leading edge and the trailing edge. The nacelle further includes a concave section continuous with the outer surface and disposed proximal to the trailing edge. The concave section includes an upstream end and a downstream end spaced apart from the upstream end. The concave section is curved radially inwards relative to the outer surface between the upstream end and the downstream end.

Abstract: A journal bearing comprising a first component and a second component, the first and second components being arranged to rotate relative to one another in normal use. The first component comprises a first body, a first layer forming a functional surface facing the second component, and a first subsurface layer between the body and the layer. The first subsurface layer is formed of a material having a first coefficient of thermal expansion in the radial direction, and the first body is formed of a material having a second coefficient of thermal expansion in the radial direction. The first coefficient of thermal expansion is lower than the second coefficient of thermal expansion.

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 edge, 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 a ratio of a leading edge radius of each of the plurality of first rotor blades at 0% span height to a minimum leading edge radius is comprised between 1 and 1.50.

Abstract: An aircraft propulsion system comprises a propulsor an electric motor coupled to the propulsor. The electric motor comprises a surface mounted permanent magnet electric machine comprising a rotor mounted radially inward of a stator. A Motor Diameter Ratio (MDR) is defined as an inner diameter (Dstator,in) of the motor stator in metres divided by an outer diameter (Dstator,out) of the motor stator in metres. The MDR of the electric motor stator is within 10% of the value given by the equation: MDR = 0.83 × ( Torq ? u ? e L a ? c ? t ? i ? v ? e × 1 U fan , tip ) 0 . 0 ? 8 where: T is the maximum torque rating of the electric motor in kilonewton metres; Lactive is the active length of the motor stator in metres; and Utip is the tip speed of the propulsor at the maximum rated torque of the electric motor in metres per second. In the described embodiments, the MDR is less than 1.

Abstract: An electric machine (101) for use in an aircraft is shown. The electric machine comprises a casing (104) containing electromechanical components, a shaft (106) which extends outside of the casing, a seal (107) to seal the casing around the shaft, and a depressurisation system (102) configured to depressurise the casing below an external pressure to prevent electrical breakdown within gas in the casing.

Abstract: Gas turbine engine including a nacelle and an engine core within the nacelle. The engine core defines a principal rotational axis along its length. The engine core and nacelle define a bypass passage therebetween. The gas turbine engine further includes a cooling system including a cooling duct, which duct defines an inlet for receiving bypass air from the bypass passage at an upstream location and an outlet for discharging the bypass air at a downstream location. The cooling duct extends, relative to the principal axis, axially and circumferentially around a section of the engine core. The cooling duct comprises: first portion that extends at least axially relative to the principal rotational axis; second portion downstream of the first portion that extends circumferentially around the engine core relative to the principal rotational axis; and third portion downstream of second portion that extends at least axially relative to the principal rotational axis.

Abstract: A system for supplying lubricant to a component of a gas turbine engine having a fan shaft is provided. The system includes a pump drivably couplable to the fan shaft for pumping lubricant to the component. The pump includes an inlet for receiving lubricant from a lubricant source, an outlet for outputting lubricant to the component and a swashplate movable between at least a first position and a second position. The system also includes a swashplate actuator for actuating the swashplate between the first position and the second position according to whether the fan shaft is rotating in a forward direction or a reverse direction opposite to the forward direction.

Abstract: An isostatic pressing canister for use in manufacturing a component by isostatic pressing is described. The canister comprises: a canister wall enclosing an internal cavity, the canister wall having an internal surface facing the internal cavity; a filling point for filling the internal cavity with powdered material, the filling point comprising a hole in the canister; and a structure supported within the internal cavity and located between the hole and the portion of the internal surface opposite the hole. A portion of the internal surface of the canister wall opposite the hole is shaped so that, during the filling of the internal cavity, powdered material falling from the hole and towards the portion of the internal surface is deflected away from the portion of the internal surface. A method of manufacturing a component using the canister is also described.

Abstract: In one example, an article including a metallic substrate; and an abradable coating on the metallic substrate, the abradable coating including a plurality of abradable layers in an alternating arrangement with a plurality of CMAS resistant layers, wherein the plurality of CMAS resistant layers are configured to react with molten CMAS to form a stable phase.

Abstract: There is disclosed an arrangement for separating an air-oil mixture in an oil distribution system of an engine. The arrangement comprises a distribution channel of the system, which defines a fluid flowpath suitable for receiving the air-oil mixture, and an ultrasonic transducer configured to insonate the fluid flowpath.