Abstract: A method of inspecting a component comprising a cellular structure bonded to a carrier substrate to form a matrix of open-ended cells is described. The method comprises: submerging the component in an inspection liquid disposed within a pressure vessel; changing the pressure within the pressure vessel; withdrawing the component from the inspection liquid; and determining whether the cells of the component are filled with the inspection liquid. A corresponding inspection apparatus is also described.

Abstract: A rigid electrical raft has electrical conductors embedded in a rigid material. The electrical conductors transmit electrical signals through the rigid electrical raft, which may form part of an electrical system of a gas turbine engine. The rigid electrical raft also has electrical heating elements embedded therein. The electrical heating elements provide heat which may be used, for example, to prevent condensation and/or ice build-up and/or to raise the temperature of electrical components to be within a desired range.

Abstract: A method of machining a nickel containing alloy gas turbine engine component (34) comprises applying a material removal gas comprising gaseous carbon monoxide at a nickel carbonyl gas forming temperature such as 50 to 60° C. to a surface of the component to form a nickel carbonyl gas, and thereby remove a surface layer from at least part of the component.

Abstract: The present invention provides a flexible hose for connecting a fuel manifold to a burner of a gas turbine engine. The flexible hose includes a metal convolute tube, an elongate member or members located in grooves formed in the inner surface of the convolute tube, and a pressure-containing sheath outside the convolute tube. The flexible tube has end connectors fluidly-tightly joined to respective ends of the hose for connection at one end of the hose to the fuel manifold of the gas turbine engine, and at the other end of the hose to the burner of the gas turbine engine.

Abstract: An oil chamber wall arrangement is described as comprising: a chamber wall surrounding a rotor; a plurality of apertures defining through-holes for a flow of oil through the chamber wall, the apertures being defined by a pair of axially separated opposing walls and a pair of circumferentially opposing walls wherein the circumferentially opposing walls are separated by a midline which is radial to the rotational axis of the rotor and at least one of the pair of circumferentially opposing walls is not parallel to the radial midline.

Abstract: This disclosure concerns an adaptive repair method, for example for aerofoil blade components. The method generates a digitised model of a physical component and uses data from the digitised model to align a part of the physical component with a nominal CAD model of the component. A combination of the digitised data and data from the nominal CAD model are then used to create an adapted model for repair of the component.

Abstract: An imaging apparatus for imaging a rotating component is shown. The imaging apparatus has a proximal end configured to be attached to the rotating component, along with a distal end. The imaging apparatus has located within it a convex mirror at the distal end, which has a reflective surface which is directed toward the proximal end and having a field of view wider than the imaging apparatus. The imaging apparatus also has located within it a camera at the proximal end, the camera being directed towards to distal end and having a field of view which includes the mirror.

Abstract: A bladed rotor assembly includes a bladed rotor and a rotor housing. The bladed rotor has a plurality of blades arranged in a circumferential array around an axis of rotation, and the rotor housing encircles the bladed rotor and has a radially inwardly facing surface. Each of the blades comprises a first, radially extending, edge, a second opposite, radially extending, edge, and a radially distal edge. The radially distal edge has a first end, a second end, and a center portion, with the first end adjoining the first, radially extending, edge, and the second end adjoining the second, radially extending, edge. A clearance between the radially distal edge and the radially inwardly facing surface is a first clearance at the first end, and decreases to a second clearance across the center portion.

Abstract: A retainer plate is provided for retaining a dovetail root of a fan blade of a gas turbine engine in a corresponding axially-extending slot in the rim of a fan disc. In use, the plate locates in a cavity formed at an end of the slot such that a first side of the plate is arranged for contact with an axial end face of the dovetail root and an opposite second side of the plate is arranged for contact with an abutment surface of the cavity to limit axial movement of the root along the slot. The retainer plate is a unitary component and has a layered structure including a first layer at the first side of the plate, a second layer at the second side of the plate, and an intermediate layer between the first and the second layers.

Abstract: A fan blade has a tip, a root, a pressure side and a suction side. The fan blade comprises a laminate body defined by a plurality of plies comprising reinforcement fibres. The angles of the fibres in the plies are arranged such that the laminate body is unbalanced.

Abstract: A gas turbine engine combustion chamber includes upstream and downstream ring structures and a plurality of circumferentially arranged combustion chamber segments. Each segment extends the full length of the combustion chamber and each segment is secured to the upstream ring structure and is mounted on the downstream ring structure. A frame structure at the downstream end of each segment has multiple spaced radially extending holes. The downstream end of each segment has an axially upstream extending groove and the downstream ring structure has an annular axially upstream extending hook which locates in the groove of each segment. A portion of the downstream ring structure abuts the frame structure of each segment. The downstream ring structure has multiple holes through the portion abutting the frame structure and segment is removably secured to the downstream ring structure by multiple fasteners locating in the holes in the segments and the downstream ring structure.

Abstract: An aircraft has an internal combustion engine and a wing including a hollow structural member. The aircraft has an electrical network including: at least one alternating current electrical generator configured to be driven by the internal combustion engine; an electrical motor configured to drive an aircraft propulsor; at least one conductor configured to electrically couple the electrical motor and the electrical generator; wherein the electrical conductor is formed of the hollow structural member of the aircraft wing.

Abstract: An engine fuel control system includes a main pump which receives fuel from a low pressure source and delivers the fuel at a first high pressure, and a variable displacement augmenter pump which receives fuel from the low pressure source and delivers the fuel at a second high pressure. The augmenter pump has a servo-controller to vary its pump flow rate. The system includes an augmentation valve which is actuatable by an electrically operated control valve to open a flow path therethrough which diverts fuel delivered by the augmenter pump or to shut the flow path. The system includes a pressure drop control valve which senses a pressure differential between the second high pressure and a reference pressure, and controls the servo-controller to vary the pump flow rate of the augmenter pump. The pressure drop control valve spills some of the fuel delivered by the augmenter pump.

Abstract: A gas turbine engine component, especially an aerofoil-sectioned nozzle guide vane (NGV), having at least one internal cooling chamber for passage of cooling air, the chamber including leading edge portion and one inlet portion via which cooling air may enter the chamber from feed source, wherein the component includes a partitioning element, e.g. curved or scoop-shaped partitioning plate or wall, provided in the chamber inlet portion and defining within the inlet portion a sub-chamber adjacent the leading edge portion, and wherein partitioning element is configured so the cooling air velocity in the sub-chamber is less than the cooling air velocity in the remainder of inlet portion. The reduced velocity of the cooling air in the sub-chamber adjacent the leading edge serves to increase pressure therein, thereby maintaining desired backflow pressure margin between the feed pressure of the cooling air delivered to the showerhead holes and the gas-path from the combustor.

Abstract: Fan containment system fitting around an array of radially extending fan blades mounted on a hub in an axial gas turbine engine. The fan containment system includes a fan case having an annular casing element for encircling the array of fan blades and a hook projecting in a radially inward direction from the annular casing element and positioned axially forward of the array of fan blades when the fan containment system fitted around fan blades. An annular fan track liner positioned substantially coaxial to the annular casing element. Clamping arrangement connects fan track liner to the hook. Clamping arrangement is configured under the condition that a fan blade impacts the fan track liner, the clamping arrangement releases connection between the hook and a portion of fan track liner so that a portion of the fan track liner can move towards the annular casing element to encourage the fan blade to impact the hook.

Abstract: A method of operating a gas turbine engine compressor. The method includes: determining an operating point of the compressor, and modulating mass flow of environmental control system input air to maintain the operating point of the gas turbine engine compressor within predetermined limits.

Abstract: A combustion staging system for fuel burners of a multi-stage combustor of a gas turbine engine is provided. The system has pilot and mains fuel manifolds respectively distributing fuel to pilot and mains stages of the burners. The system further has a plurality of check valves through which the mains manifold feeds fuel to the respective burners. The check valves are arranged to open when the fuel pressure within the mains manifold exceeds a predetermined fuel pressure relative to gas pressure in the combustor. The system further has a splitting unit which receives a metered fuel flow and controllably splits the received fuel flow into pilot and mains flows which are sent respectively to the pilot and mains fuel manifolds to perform staging control of the combustor. The splitting unit is operable to select the pilot manifold and to deselect the mains manifold for pilot-only operation in which there is a pilot supply to the combustor but no mains supply to the combustor from the burners.

Abstract: A gas turbine engine comprises a relatively high pressure compressor coupled to a relatively high pressure turbine by a relatively high pressure shaft; a relatively low pressure compressor coupled to a relatively low pressure turbine by a relatively low pressure shaft rotatable independently of the high pressure shaft; a first combustor located downstream of the high pressure compressor and upstream of the high pressure turbine; and a second combustor located downstream of the high pressure turbine, and upstream of the low pressure turbine. The engine further comprises a coupling arrangement configured to selectively transfer torque between the high pressure shaft and the low pressure shaft.

Abstract: A combustion staging system includes a splitting unit receiving a metered fuel flow and controllably splitting the received flow into pilot and mains flows for injecting at pilot and mains fuel stages performing staging combustor control. Pilot and mains fuel manifolds distribute fuel from the unit, which can select pilot-only and pilot and mains operations. A cooling flow recirculation line provides a cooling flow to the mains manifold during pilot-only operation, and a return section to collect mains manifold cooling flow. A fuel recirculating control valve open position allows the cooling flow to enter a delivery section during pilot-only operation; a shut off position prevents the cooling flow from entering the delivery section during pilot and mains operation. The unit can divert a mains flow portion into the delivery section during pilot and mains operation, the diverted portion re-joining the rest of the mains flow in the mains fuel stages.

Abstract: Apparatus for machining a component of a gas turbine engine, the apparatus including: a flexible pipe having a first end and a second end and defining a conduit for receiving a fluid; one or more first actuators arranged to enable the first end of the flexible pipe to be re-positioned relative to the second end of the flexible pipe; a turbine positioned adjacent to the first end of the flexible pipe, the turbine being arranged to receive fluid from the conduit of the flexible pipe; and a tool head coupled to the first end of the flexible pipe, the tool head including: a fastener arranged to fasten to a tool and to receive torque from the turbine.