Abstract: A system for mounting a gas turbine engine to a pylon on a wing of an aircraft. A temporary forward link, being length-adjustable, and at least one temporary rearward link, being length-adjustable, are provided. These are for temporarily attaching the gas turbine engine to the pylon. The temporary forward link and the temporary rearward link are each adapted to resist tension and compression, to maintain a positional relationship between the gas turbine engine and the pylon in the absence of adjustment of the lengths of the temporary forward link and the temporary rearward link. Adjustment of the length of the temporary links brings engine mounts into alignment with pylon mounts for service attachment of the gas turbine engine to the pylon.

Abstract: A system for mounting a gas turbine engine to a pylon on a wing of an aircraft. At least one temporary forward link, being length-adjustable, and at least one temporary rearward link, being length-adjustable, are provided. These are for temporarily attaching the gas turbine engine to the pylon. The temporary forward link and the temporary rearward link each comprise a respective winch operable to adjust pay out of a respective tension member thereby to provide length adjustment. The temporary forward link and the temporary rearward link maintain a positional relationship between the gas turbine engine and the pylon in the absence of adjustment of the lengths of the temporary forward link and the temporary rearward link. Adjustment of the length of the temporary links brings engine mounts into alignment with pylon mounts for service attachment of the gas turbine engine to the pylon.

Abstract: The blades for a rotor of a gas turbine engine are all manufactured to the same design. However, manufacturing tolerances mean that in practice each individual blade is different to the others. It is proposed to arrange the blades around the circumference of the rotor in a manner that limits excessive stress being induced in the blades due to differences in the vibration response between a given blade and its two neighbouring blades.

Abstract: An aircraft environmental control system (20) comprises a first heat exchanger (24) configured to exchange heat between air provided by a pressurised air source (13), and a working fluid of a closed cycle waste heat recovery system, and an air turbine (32) configured to receive cooled air from the first heat exchanger (24). The closed cycle waste heat recovery system comprises a waste heat recovery compressor (35), a waste heat recovery turbine (31), and a second heat exchanger (33) configured to exchange heat between working fluid of the closed cycle waste heat recovery system and air downstream of the air turbine (31). The air turbine is configured to drive a load (39), and the waste heat recovery turbine (31) is configured to drive the waste heat recovery compressor (35).

Abstract: A gas turbine engine (10) for an aircraft is disclosed. The gas turbine engine (10) comprises a mechanical power converter (42) arranged to receive an input drive (44) and produce an output drive (46). The output drive (46) has the same rotational direction as the input drive (44). The gas turbine engine further comprises a stationary supporting structure arranged to provide a stationary support for the mechanical power converter (42). The stationary supporting structure comprises one or more structural support aerofoils (24a-24h). A method (100) of supporting a mechanical power converter (42) in a gas turbine engine (10) is also disclosed.

Abstract: The blades for a rotor of a gas turbine engine are all manufactured to the same design. However, manufacturing tolerances mean that in practice each individual blade is different to the others. It is proposed to arrange the blades around the circumference of the rotor in a manner that limits excessive stress being induced in the blades due to differences in the vibration response between a given blade and its two neighbouring blades.

Abstract: A polycrystalline nickel-base superalloy is disclosed. The alloy consists essentially of a gamma matrix phase including cobalt, and a gamma prime phase including aluminium, titanium, tantalum, and niobium. The overall concentration in the alloy of cobalt is from 15 to 26 atomic percent, and the overall concentration in the alloy of aluminium, titanium, tantalum, and niobium is from 13 to 14 atomic percent. Optionally, the alloy may include one or more constituents selected from the group consisting of: boron, carbon, chromium, iron, manganese, molybdenum, tungsten, silicon, zirconium. The balance is nickel and incidental impurities. The atomic ratio of aluminium to titanium is from 4.625:1 to 6.333:1.

Abstract: The blades for a rotor of a gas turbine engine are all manufactured to the same design. However, manufacturing tolerances mean that in practice each individual blade is different to the others. It is proposed to arrange the blades around the circumference of the rotor in a manner that limits excessive stress being induced in the blades due to differences in the vibration response between a given blade and its two neighbouring blades.

Abstract: A gas turbine engine comprises a gearbox comprising a sun gear, an annulus gear, a plurality of planet gears and a carrier. Each planet gear is rotatably mounted in the carrier by a bearing. The sun gear meshes with the planet gears and the planet gears mesh with the annulus gear. The sun gear, the planet gears and the annulus gear comprise helical gear teeth. The annulus gear is secured to a surrounding structure by a radially extending member. A face of the flange nearest the middle of the annulus gear is axially spaced from the middle of the annulus gear by a first distance. The reference diameter of the annulus gear is radially spaced from a point at which the radially extending member is secured to the surrounding structure by a second distance. The ratio of first distance to second distance is between and including 0.5 and 1.2 such that misalignment between the annulus gear and the planet gears is reduced to minimise vibrations, noise, stress and wear of the gearbox.

Abstract: A method for detecting a fault in a gas supply system for a gas turbine engine. The method comprises: measuring a pressure difference between two gas supplying ducts of the gas supply system; comparing the measured pressure difference to a threshold level; and determining a fault based on the measured pressure difference and the comparison.

Abstract: A gas turbine engine includes a spool which rotates in use, a static structure and an air bearing, the air bearing being provided at an interface between the spool and the static structure.

Abstract: A method of encapsulating an electrical winding in an encapsulant composition. The method includes: (i) providing an electrically insulated wire; (ii) applying the encapsulation composition to the electrically insulated wire to encapsulate the wire; and (iii) winding a further layer of wire around the encapsulated wire. The electrical encapsulant composition includes a silica containing inorganic-organic nano-hybrid matrix, and particles of refractory ceramic suspended in the inorganic-organic nano-hybrid matrix.

Abstract: A method of repairing a metallic component (formed from a first material) by powder feeding laser deposition, comprises the step of depositing a plurality of first repair layers onto a repair surface of the component to form a first repair zone, the first of the plurality of first repair layers comprising a mixture of A/B by weight of the first material and a second material, each nth successive one of the plurality of first repair layers comprising a change in the proportion of the second material in the mixture, the last of the plurality of first repair layers comprising a mixture of C/D by weight of the first material and the second material.

Abstract: A die for moulding a core by a PIM process, the core having at least one internal feature, the die including; a first die part defining a first portion of an outer surface of the core; a second die part defining a second portion of the outer surface of the core; and an internal feature forming element for defining the surface of an internal feature of the core; wherein the internal feature forming element incorporates a temperature control circuit.

Abstract: There is provided a burner for a gas turbine engine, the burner comprising a radially inner pilot fuel flow passage surrounded by a radially outer main fuel flow passage. The main fuel flow passage is interposed between concentrically arranged radially inner and radially outer air flow passages. The inner and outer air flow passages are in fluid communication with one another via at least one diverting passage at an upstream end of the burner. The burner further comprises at least one control duct connectable to a reduced pressure/vacuum source for selectively reducing the air pressure in the vicinity of the diverting passage such that air flow is selectively diverted from the inner air flow passage to the outer flow passage via the diverting passage.

Abstract: A method of evaluating a composite sample comprises providing a flatbed scanner comprising a scanning head movable along a scanning direction; providing a composite sample comprising a plurality of plies arranged on planes parallel to a longitudinal direction and comprising a scanning plane containing ply edges, wherein the ply edges are oriented parallel to the longitudinal direction; arranging the composite sample on the flatbed scanner such that the scanning plane is parallel to the scanning direction and the ply edges of the scanning plane are perpendicular to the scanning direction; moving the scanning head along the scanning direction and perpendicularly to the ply edges to scan the scanning plane of the composite sample and obtain an image of the scanning plane; and evaluating the image to identify possible defects in the composite sample.

Abstract: A bladed rotor arrangement comprising a rotor, a plurality of rotor blades, at least one seal plate and a plurality of lock plates. Each rotor blade locates in a corresponding axially extending slot in the rotor. A plurality of lock plates are arranged at a first axial end of the rotor, the radially outer ends of the lock plates engage grooves on the platforms of the rotor blades, and the radially inner ends of the lock plates engage a circumferentially extending groove. The seal plate is arranged at the first axial end of the rotor, the radially inner ends of the lock plates engage the circumferentially extending groove defined by the seal plate. The seal plate has at least one projection extending axially from the seal plate and arranged to locate in a slot in a first axial end of the root of a rotor blade.

Abstract: A method of producing a curved plate by additive layer manufacture is provided. The method includes building up the plate layer-by-layer such that a first layer forms a bottom edge of the plate and the plate is additively extended upwards therefrom by subsequent layers such that a last layer forms an opposing top edge of the plate. In this way, spaced, first and second lateral edges of the plate each grow layer-by-layer to extend between respective ends of the bottom and top edges. The plate thus built up arches over from the bottom edge to adopt a curved shape on cross-sections through the plate normal to the bottom edge.

Abstract: An aerofoil structure with a hollow cavity is manufactured by diffusion bonding and superplastic forming. Outer panels are formed of a first material; a membrane is formed of a second material. Stop-off material is applied to preselected areas on at least one side of the membrane or of one of the panels so as to prevent diffusion bonding between the panels and the membrane at the preselected areas. The panels and the membrane are arranged in a stack and a diffusion bonding process is performed to bond together the first and second panels and the membrane to form an assembly. A superplastic forming process is performed at a forming temperature to expand the assembly to form the aerofoil structure. The forming temperature is selected so that the second material undergoes superplastic deformation at the forming temperature and the first material does not undergo superplastic deformation at the forming temperature.

Abstract: A method of manufacturing a magnetically graded material, including depositing a steel filler material to a substrate, and applying a directed energy source to first and second regions of the filler material to thereby join the filler material to form a joined material. The energy source is directed to the first region while the first region is provided with an inert shield gas such that the material of the first regions includes a magnetic phase, and the energy source is directed to the second region while the second region is provided with a nitrogen containing shield gas to thereby impart an non-magnetic phase to the joined material.