Abstract: A method of manufacturing a vane arrangement for a gas turbine engine comprises providing an aerofoil having a hollow cavity with an open end and providing a support member having a stub. The method further comprises welding the aerofoil to the stub. The method yet further comprises removing material from the stub so as to define a hollow region that extends through the support member and stub to the cavity of the aerofoil.

Abstract: A method of manufacturing a component is provided. The method includes performing a machining operation by moving a rotating, abrasive grinding tool along a feed direction to remove material from the component. At least the part of the component from which the material is removed is formed of composite material. The abrasive grinding tool follows a trochoidal path along the feed direction.

Abstract: A depressurisation valve for a coolant system; comprising a main chamber having a main valve, a pilot line, and a blowdown line having a secondary valve; the main valve being located to seal a coolant line of the coolant system. The main chamber being located downstream of the cooling system, the main chamber being filled with fluid from the coolant system via a pilot line, the pressure of fluid in the main chamber acts upon a piston head of the main valve and causes the main valve to open or close dependent upon the fluid pressure in the main chamber. Fluid can escape from the main chamber via the blowdown line, which has a variable fluid pressure depending upon an operating state of the secondary valve, and wherein the secondary valve is opened automatically dependent upon the conditions within the coolant system.

Abstract: A continuum arm robot section, in which the section comprises at least three segments, featuring a base, at least one middle segment and a tip segment. Each segment corresponding to at least one interconnecting portion between pairs of neighbouring discs that form a backbone of the section of the continuum arm robot. The base segment having a greater stiffness than the at least one middle segment and tip segment and the at least one middle segment having a greater stiffness than the tip segment. Also provided is a method of adjusting the stiffness of a continuum arm. The method comprising the steps applying a first stiffening component to a segment. Applying a second stiffening component to a different segment. The first stiffening means results in the base segment having a greater stiffness than the middle segment, and the second stiffening means results in the middle segment having lower stiffness than the base segment, but higher than the tip segment.

Abstract: A fuel flow control system is configured to perform a method including: setting a fuel flow demand proportional to compressor discharge pressure; detecting an actual fuel flow; and applying an enhanced schedule for fuel flow demand while fuel flow demand is less than a predefined proportion of actual fuel flow.

Abstract: An apparatus for the detection of either or both of water and ice within a system, more specifically, but not exclusively, the detection of either or both of water or ice crystal ingestion, or ice accretion on an aerofoil, within a gas turbine engine. The apparatus and method compares the measured temperature of the first region with a second temperature value, measured at a distinct second region of the aerofoil, wherein the second radial position of the second sensor is radially offset from the first radial position of the first sensor.

Abstract: A seal segment is disclosed for a shroud ring of a rotor of a gas turbine engine. The gas turbine engine has a principal axis, there being defined an axial direction (A) parallel to the principal axis, a radial direction (R) perpendicular to the principal axis and a circumferential direction (C) perpendicular to the radial direction and the axial direction. The seal segment comprises a radially inner side and a radially outer side, wherein the seal segment comprises a circumferential side face with a slot extending axially along the seal segment and circumferentially into the seal segment. The slot is for retaining a strip seal therein. There is provided at least one relief slit formed in the seal segment from the radially outer side of the seal segment and in communication with the slot.

Abstract: A depressurisation valve for a cooling system comprising: a main chamber having a main valve, a pilot line having a secondary valve and a blowdown line; the main valve being located to seal a path of the coolant system of the nuclear reactor. The main chamber is connected to the cooling circuit via the pilot line allowing coolant to enter the main chamber, and the blowdown line allows coolant to escape from the main chamber, the pilot line having a lower fluid resistance than the blowdown line. The pressure of coolant in the main chamber maintains the main valve in a closed position, and under elevated temperature and/or pressure conditions fluid is prevented from entering the main chamber via a closure of the secondary valve on the pilot line and reduce the pressure from the valve, moving it to its open position.

Abstract: A bearing assembly for a rotatable shaft, the bearing assembly comprising: a bearing housing; a bearing located within the bearing housing having an axis of rotation and arranged to receive a rotatable shaft; a first spring bar that couples the bearing to the bearing housing, the first spring bar being configured to tune vibrations of the rotatable shaft; and a first piezoelectric actuator disposed between the bearing housing and the first spring bar, the first piezoelectric actuator being configured to extend in a first direction, wherein extension of the piezoelectric actuator in the first direction displaces the first spring bar relative to the bearing housing.

Abstract: A fuel spray nozzle including a primary atomizer to discharge a flow of swirled atomised fuel along and around a fuel spray nozzle axis. The primary atomiser includes outer air swirler disposed radially outwardly of a fuel pre-filmer channel. A secondary atomiser disposed around the primary atomiser includes secondary inner air swirler to swirl flow along an inner air channel. The secondary inner air swirler disposed radially inwardly of a secondary fuel pre-filmer channel of the secondary atomiser. A primary outer air channel defined between the primary outer swirler and the secondary inner swirler. The secondary inner air swirler include splitter wall to separate swirling flow in the secondary inner channel from the primary flow of atomised fuel. The secondary inner air swirler includes primary cap wall integral with and extending radially inwardly from the splitter wall to direct flow from the primary outer channel inwardly towards the fuel spray.

Abstract: An apparatus for controlling flow of coolant into an inter-stage cavity of a turbomachine is described. The cavity is bounded by a first turbine stage, a second turbine stage axially displaced along a common axis of rotation with the first turbine stage, and an annular platform bridging a space between the axially displaced first and second turbine stages. An annular plenum chamber is arranged inboard of the annular platform, the annular plenum chamber having one or more inlets for receiving coolant and one or more outlets exiting into the cavity, whereby, in use, coolant is delivered into the cavity at an increased pressure compared to coolant entering the plenum chamber at the inlet. The apparatus is beneficially arranged immediately upstream (with respect to the flow of a working fluid through the turbomachine) of an inter-stage seal assembly.

Abstract: The present invention provides a projectile for use in a simulated fan-blade-off ballistic test. The projectile has an ellipsoid body having a blind axial bore extending from a first axial end. The blind axial bore is for housing a weight adjustment body which can be used to modify the weight and/or centre of gravity of the projectile. A sealing plug may seal the weight adjustment body within the axial bore and an insert may be provided to fix the position of the weight adjustment body and/or to control the sliding of the weight adjustment body within the axial bore.

Abstract: A gas shield arrangement for a gas turbine engine includes a first gas shield that is arranged to cover an inlet of a gas flow path of the gas turbine engine and a second gas shield that is arranged to cover a first outlet of the gas flow path of the gas turbine engine. The gas shield arrangement further includes a reservoir that holds fluid and a pump that pumps the fluid from the reservoir to one of a first valve and a second valve and that returns at least some of the pumped fluid to the reservoir from the other of the first valve and the second valve.

Abstract: A method of controlling a reciprocating internal combustion engine comprising: a cylinder defining a cavity having a first end and a second end; and a piston moveable within the cavity of the cylinder between the first end and the second end, the method comprising: receiving at least a first signal; determining a quantity of liquid air to be injected using at least the received first signal; controlling injection of the determined quantity of liquid air into the first end of the cavity at a first time when the piston is closer to the first end than the second end.

Abstract: The present invention provides a multi-stage compressor for a gas turbine engine. The compressor has: a first outer casing, a second outer casing radially outward of the first outer casing, and a first bleed plenum one or more second bleed plenums located between the first and second outer casings and arranged to receive, in use, bleed flows of compressed air from respective stages of the compressor and to send the bleed flows to respective ports in the second outer casing. The first bleed plenum overlaps the, or each, second bleed plenum such that the, or each, second bleed plenum fluidly communicates with its port via a respective duct which, on extending between an off-take from the second bleed plenum to the port, passes through the first bleed plenum. The, or each, duct is configured to accommodate relative movement between the first and second outer casings.

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 comprising a gearbox that receives an input from a core shaft and outputs drive to the fan to drive the fan at a lower rotational speed than the core shaft. The gearbox is an epicyclic gearbox comprising planet gears, each of which is rotatable about its own axis on a respective pin with a journal bearing formed between each planet gear and its respective pin. A first oil supply system is arranged to provide an oil supply to the journal bearing via a first journal bearing supply channel formed through the pin. A second oil supply system is arranged to provide an oil supply to the journal bearing via a second journal bearing supply channel formed through the pin, the second oil system being different to the first oil system. The second journal bearing supply channel is parallel to the first journal bearing supply channel.

Abstract: A gas turbine engine for an aircraft comprises a gearbox that receives an input from a core shaft and outputs 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 comprising planet gears, each of which is rotatable about its own axis on a respective pin, with a journal bearing formed between each planet gear and its respective pin. A first oil supply system is arranged to provide an oil supply to the journal bearing via a first journal bearing supply channel formed through the pin. A second oil supply system is arranged to provide an oil supply to the journal bearing via a second journal bearing supply channel formed through the pin, the second oil system being different to the first oil system.

Abstract: An assembly for receiving in a radially extending bucket groove (31) of a turbine disc (22), the assembly including a turbine blade having an elongate body (26) of aerofoil cross-section, a root portion (24) at one end of the elongate body (26) and a tip at the other. The root portion (24) configured to be retained in the bucket groove (31). A heat shield (27) configured to be received between an end of the root portion (24) and a radially inner surface of the bucket groove (31) and to interlock with the root portion (24) in a manner which deters separation in a radial direction.

Abstract: An apparatus for the simultaneous casting of multiple components using a directional solidification process includes; a pouring cup arranged on a centerline, an array of moulds encircling the pouring cup and centre line, an array of feed channels extending from the pouring cup to a top end of each mould, and a heat deflector. The heat deflector comprises a wall arranged between the array of moulds and the centerline extending along the length of the moulds and in thermal contact with the moulds.