Abstract: A control system for use in controlling a cabin blower system. The cabin blower system includes a gas turbine engine spool, a cabin blower compressor powered by the spool and arranged in use to compress fluid used in a cabin of an aircraft, and one or more control mechanisms via which the control system controls the power extracted by the cabin blower compressor from the spool. The control system is arranged in use to control the power extracted from the spool by the cabin blower compressor in accordance with one or more primary control parameters. The control system is arranged in use to alter the spool power extracted by the cabin blower compressor by comparison with the power that would have been extracted in accordance with the primary control parameters alone, in response to modifications in a secondary control parameter indicative of the commencement or occurrence of an engine transient.

Abstract: A bearing arrangement (30) comprises an inner support annulus (33), an outer support annulus (31), and a plurality of rollers (34) therebetween. At least one of the inner and the outer support annulus (33, 31) comprises one or more radially extending rigid supports (37) and one or more damping sectors (38a, 38b, 38c). The one or more rigid supports (37) are circumferentially interposed between one or more damping sectors (38a, 38b, 38c). Each damping sector (38a, 38b, 38c) comprises at least one damping member (39) comprising an elongate flexible member held rigidly at one or both ends thereof, such that vibrational movement of the inner or outer support annulus (33, 31) causes bending of the elongate flexible members (39).

Abstract: A method of reducing rotor bow in a high pressure rotor of a gas turbine engine that has in axial flow a low pressure rotor and a high pressure rotor. The method involves storing bleed air from the gas turbine engine when the engine is running to provide stored pneumatic energy; and using that stored pneumatic energy after the engine has been shut-down to rotate the high pressure rotor at a speed and for a duration that reduces rotor bow. A gas turbine engine wherein rotor bow in the high pressure rotor after engine shut-down has been reduced by carrying out the aforesaid method is also disclosed.

Abstract: A power system, including: a synchronous electrical generator having a rotor; and an angle computation unit configured to: determine a rotor angle in a steady state period of the synchronous electrical generator, determine a change in rotor angle in a transient period of the synchronous electrical generator, and estimate the rotor angle in the transient period based on the steady state rotor angle and the change in rotor angle.

Abstract: A cooled gas turbine engine component comprises a wall which has a plurality of effusion cooling apertures extending there-through from a first surface to a second surface. The apertures are arranged at an angle to the second surface and each aperture has an inlet in the first surface and an outlet in the second surface. Each aperture has a metering portion and a diffusing portion arranged in flow series and each metering portion is elongate and the width is greater than the length of the metering portion. Each diffusing portion increases in dimension in the length from the metering portion to the outlet. Each outlet has a rectangular shape in the second surface of the wall. Each inlet has an elongate shape in the first surface of the wall and the inlet in the wall is arranged substantially diagonally with respect to the outlet in the wall.

Abstract: A vehicle simulator including: a base; a plurality of moveable components; and one or more physical and/or power connectors for coupling the components to the base. The components are positionable on the base with the connectors coupling the components to the base in a first arrangement to simulate a first vehicle type. The components are moveable to and positionable on the base with the connectors coupling the components to the base in a second arrangement different to the first arrangement to simulate a second vehicle type.

Abstract: A method of analysing a component formed from a metal alloy to identify a possible defect, wherein the metal alloy comprises a first crystal grain region and the possible defect comprises a second crystal grain region aligned to a different axis to the first crystal grain region, the method comprising the steps of: obtaining a first image of the component illuminated using a first polarisation state of light, the first image comprising first polarisation data; obtaining a second image of the component illuminated using a second polarisation state of light different to the first polarisation state, the second image comprising second polarisation data; determining a difference in polarisation data for plural pixels of the first image between each pixel of the first image and a corresponding pixel of the second image; and identifying pixels corresponding to the second crystal grain region based on the difference in polarisation data.

Abstract: A gas turbine engine, in particular an aircraft engine, includes: a turbine connected via an input shaft device to a gearbox device having a sun gear, a planet carrier having a plurality of planet gears attached thereto, and a ring gear, the sun gear is connected to the input shaft device, the planet carrier or the ring gear is connected to a propulsive fan via an output shaft device of the gearbox device, with a rear carrier bearing device radially between the planet carrier and a static structure on the input side of the gearbox device, an inter-shaft bearing system being positioned radially between the input shaft device and the planet carrier of the gearbox device. The input shaft device having a high rigidity or the input shaft device having a diaphragm section.

Abstract: An aerofoil has a leading edge, a trailing edge, a suction surface and a pressure surface. The leading edge includes apertures extending through the aerofoil from the suction surface to the pressure surface. The apertures define a first row spaced a distance (L1) of between 2 and 6 cm from the leading edge in a chordal direction (C).

Abstract: An engine for an aircraft includes an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox. The gearbox receives an input from a gearbox input shaft portion of the 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 including a sun gear, a plurality of planet gears, a ring gear, and a planet carrier on which the planet gears are mounted. The carrier and the gearbox input shaft each have a torsional stiffness, and a carrier to gearbox input shaft torsional stiffness ratio is greater than or equal to 70.

Abstract: A gas turbine engine, in particular an aircraft engine, includes: a turbine connected via an input shaft device to a gearbox device having a sun gear, a planet carrier having a plurality of planet gears attached thereto, and a ring gear, the sun gear is connected to the input shaft device, the planet carrier or the ring gear is connected to a propulsive fan via an output shaft device of the gearbox device, with a rear carrier bearing device radially between the planet carrier and a static structure on the input side of the gearbox device or a front carrier bearing device radially between the planet carrier and a static structure on the output side of the gearbox device.

Abstract: An oil tank filling system for filling the oil tank of a gas turbine engine that includes an aft core cowl. The system includes: an oil tank that is located within a core of the engine; an oil access port located on the aft core cowl; an oil tank filling pipe that connects the oil access port to a tank filling port on the oil tank; and a hydraulic lock that prevents oil entering the oil tank once the oil tank contains a predetermined volume of oil. A method of filling the oil tank of a gas turbine engine and a gas turbine engine that includes the oil tank filling system are also disclosed.

Abstract: A lift rotor arrangement (100) for a VTOL aircraft (200). The lift rotor arrangement (100) comprises: a fairing (6) mounted on a wing segment (10); and first and second rotor blades (17, 18) mounted on a first shaft (4) extending vertically from the fairing (6). The first shaft (4) is movable between an extended position in which the first and second rotor blades (17, 18) are vertically spaced above the wing segment (10) and are rotatable to provide vertical lift, and a retracted position in which the first and second rotor blades (17, 18) are rotationally-fixed with the first rotor blade (17) stowed within the wing segment (10). The blades (17, 18) may be rotatable around an axis substantially perpendicular to the axis of the respective first shaft (4) so as to act as ailerons/elevons in the retracted position.

Abstract: A method of inserting a pin in a bore of an article may include first providing an article with a bore, the bore having an inside diameter. The method may then include providing a pin having a hollow body, the hollow body having a diameter with a profile with a pattern of voids. The method may then include applying an axial force on an end of the hollow body to cause the diameter of the pin to reduce in size to a point where the diameter of the pin is less than the inside diameter of the article. The method may further include aligning the pin inside of the bore of the article, and then removing the axial force on the end of the hollow body.

Abstract: A shaft assembly for a gas turbine engine is provided. The shaft assembly comprises: a first shaft having an outer surface; a first coupling ring disposed around the outer surface of the first shaft, an inner surface of the first coupling ring being coupled to the outer surface of the first shaft; a second shaft having an inner surface; and a second coupling ring disposed around the inner surface of the second shaft, an outer surface of the second coupling ring being coupled to the inner surface of the second shaft, wherein an outer surface of the first coupling ring is configured to mate with an inner surface of the second coupling ring, such that concentricity of the first and second shafts is maintained at the shaft assembly by virtue of the mating of the first and second coupling rings. Methods of assembling and re-assembling a shaft assembly are also provided.

Abstract: There is disclosed a tool (102) for forming a composite component having a curved body and an integral flange from a pre-form (200), the tool comprising: a curved body portion 104 having a lay-up surface (110); and a forming assembly (105) having a lay-up configuration and a forming configuration. The forming assembly comprises: a plurality of forming elements (106) each having a lay-up surface (120) and a primary flange-forming surface (122), the forming elements (106) being radially outwardly moveable between the lay-up configuration and the forming configuration, in which the forming elements are circumferentially spaced from one another; and a plurality of filler elements (107) each having a secondary flange-forming surface (156), the filler elements (107) being arranged to move into the circumferential gaps between the forming elements in the forming configuration so as to form a substantially continuous flange-forming surface.

Abstract: An assembly of a servo pump and a hydraulic motor. The assembly has a housing which contains the pump and the motor. The motor has a rotating body which rotates under the motive power of a pressurised motor liquid flow. The motor has a high pressure region which receives the pressurised motor liquid flow, and a low pressure region through which the motor liquid flow leaves the motor. The pump also has a rotating body. The pump has a low pressure region which receives servo liquid flow to be pumped by its rotating body, and a high pressure region through which pressurised servo liquid flow leaves the pump. Each rotating body is mounted on a respective journal.

Abstract: A method of defining at least one scan parameter for an x-ray scan of a single crystal structure, the method comprising: determining a target orientation of the structure for the scan; and defining different non-zero levels of x-ray exposure for different parts of a scan area based on either or both of the target orientation and characteristics of the structure; and, defining the scan area so that substantially all x-rays of the scan are directed to the structure in the target orientation.

Abstract: A tip clearance control (TCC) system 100 is provided to control the gap 176 between the tips of turbine blades 172 of a gas turbine engine 10 and the casing 174 within which they rotate. The inlet 120 to the TCC system is provided in a bifurcation panel 110 that extends across a bypass duct 22 of the gas turbine engine 10. The inlet is provided on a first major surface 112 of the bifurcation panel 110 that is defined such that the direction (R) that points through the bifurcation panel from the first major surface 112 to a second major surface 114 corresponds to the fan rotation direction. This provides a particularly effective and/or efficient TCC system 100.

Abstract: A fuel injector for a gas turbine engine, the injector comprising a main fuel flow circuit and a pilot fuel flow circuit. The pilot fuel flow circuit comprises at least one convoluted pilot fuel flow passage extending continuously from a pilot fuel feed at an upstream end of the injector to a pilot fuel spray nozzle inlet. The at least one pilot fuel flow passage comprises at least two axially-extending portions extending between the upstream end of the injector towards a downstream end of the injector and joined by an elbow portion proximal the downstream end of the injector. In use, pilot fuel can flow in opposing axial directions within the at least two axially-extending portions, the axially-extending portions being in thermal communication with the main fuel flow circuit.