Abstract: A gas turbine engine for an aircraft comprises a high-pressure (HP) spool comprising an HP compressor and a first electric machine driven by an HP turbine; a low-pressure (LP) spool comprising an LP compressor and a second electric machine driven by an LP turbine; a combustion system comprising a fuel metering unit; and an engine controller configured to, in response to a change of a power lever angle setting indicative of a deceleration event, reduce fuel flow to the combustion system by the fuel metering unit, and to operate the first electric machine in a generator mode to reduce the HP spool rotational speed and engine core mass flow.

Abstract: A gas turbine engine for an aircraft comprises an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, wherein a compressor exit temperature is defined as an average temperature of airflow at the exit from the compressor; and a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub, each fan blade having a leading edge and a trailing edge, wherein a fan rotor entry temperature is defined as an average temperature of airflow across the leading edge of each fan blade at cruise conditions and a fan tip rotor exit temperature is defined as an average temperature of airflow across a radially outer portion of each fan blade at the trailing edge at cruise conditions. A core to fan tip temperature rise ratio is in the range from 2.845 to 3.8.

Abstract: A gas turbine engine for mounting to an airframe of an aircraft comprises an engine core; a fan located upstream of the engine core; a bifurcation spanning a bypass duct defined between the engine core and a nacelle surrounding the gas turbine engine, the bifurcation comprising aerodynamically shaped fairings defining an interior space therebetween; and a cabin blower system arranged in the interior space of the upper bifurcation.

Abstract: A gas turbine engine for an aircraft includes an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor; and a fan located upstream of the engine core. The engine core further includes three bearings arranged to support the core shaft including a forward bearing and two rearward bearings, with a minor span defined as the distance between the two rearward bearings. A minor span to turbine length ratio is equal to or less than 1.05.

Abstract: An apparatus for ultrasonically inspecting a component of a gas turbine engine at a centre frequency required for ultrasonic inspection of the component. The apparatus comprises an ultrasonic transducer and a waveguide. The ultrasonic transducer and the waveguide are mounted in a holder, the holder is movably mounted in a carrier, and the carrier is movably mounted in a frame. The waveguide has a first end, a second end and a surface. The waveguide is circular in cross-section, the surface of the waveguide between the first end and the second end is concave, and the waveguide reduces in diameter from the first end to the second end. The length of the waveguide is equal to or greater than 0.5 times and equal to or less than 3 times the wavelength of an ultrasonic signal at the centre frequency required for the ultrasonic inspection.

Abstract: There is disclosed an exhaust nozzle for a gas turbine engine, the exhaust nozzle comprising a frame extending along a longitudinal axis. The exhaust nozzle comprises a convergent petal pivotably attached at a convergent pivot point to the frame and extending axially downstream and radially inward from the frame, a follower roller fixed to the convergent petal on a radially outer side of the convergent petal, and a cam defining a working surface configured to engage the follower roller to react a force from the convergent petal. The cam is movable along a travel in an axial direction to actuate radial movement of the follower roller to pivot the convergent petal. The cam defines a concave working surface such that a contact angle between the follower roller and the cam varies along the travel to thereby vary a radial component of the force reacted by the cam.

Abstract: Fault-tolerant electric drive systems including a machine having a rotor and a stator having coils arranged in pairs. Each coil in each pair separated by 180 degrees, a first phase (?A) having one of the coil pairs and a phase drive circuit connected therewith, a second phase (?B) having a second one of the coil pairs and a second phase drive circuit connected therewith, a third phase (?C) having a third one of the coil pairs and a third phase drive circuit connected therewith, and a fourth phase having a fourth one of the coil pairs and a fourth phase drive circuit connected therewith. Further included is a controller connected with the first, second, third and fourth phase drive circuits to control operation thereof.

Abstract: An axial flow turbomachine (102) for producing thrust to propel an aircraft is shown. The turbomachine has an inner duct (202) and an outer duct (204), both of which are annular and concentric with one another. An inner fan (206) is located in the inner duct, and is configured to produce a primary pressurised flow (P). An outer fan (207) is located in an outer duct, and is configured to produce a secondary pressurised flow (S). The outer fan has a hollow hub (208) through which the inner duct passes. The outer fan has a diameter of from 2.5 to 3.5 times a diameter of the inner fan.

Abstract: An exhaust nozzle assembly for a gas turbine engine. The assembly includes concentrically arranged inner mixer and outer exhaust nozzles, the exhaust nozzle extending axially downstream of said mixer nozzle. A centre-body is axially mounted within and extends axially downstream from the mixer nozzle. A core flow duct is defined by the mixer nozzle and the centre-body, the core flow duct having a core exit area. An exhaust duct is defined at least in part by the exhaust nozzle downstream of the mixer nozzle, the exhaust duct having an exhaust exit area. The mixer nozzle includes a mixer cowl which is axially-translatable along the centre axis and the exhaust nozzle includes an exhaust cowl which is either axially-translatable along or angularly-adjustable relative to the centre axis. The assembly further includes an actuation mechanism and the mixer cowl and exhaust cowl are movable by the actuation mechanism.

Abstract: A star-configuration epicyclic gearbox is shown. It comprises a sun gear (112) for connection with a first shaft (112), a plurality of planet gears (113) intermeshed with the sun gear and located in a static carrier (114), and a ring gear (115) for connection with a second shaft (116) and which is intermeshed with the plurality of planet gears. It further comprises one or more electric machines (117) drivingly connected with a respective one of the plurality of planet gears.

Abstract: A method and apparatus for applying a cladding layer to a surface of a component uses a cladding tool having a maximum reach less than the size of the surface. Geometry of the surface is segmented into a plurality of tessellated segments, each of which has a peripheral extent determined by a maximum reach of the cladding tool. A nominal tool subpath for each tessellated segment is generated, and then combined to generate a nominal tool path for depositing the cladding layer on the surface. The surface is clad using the nominal toolpath, including a process of adjusting the nominal tool path to an adjusted tool path that accounts for dimensions of the bead to be deposited by the tool to match an edge of the bead to be deposited with an edge of a previously deposited bead.

Abstract: A turboelectric generator system includes a gas turbine engine which includes, in fluid flow series, a gas-generator compressor, a combustor, a gas-generator turbine, and a variable-speed free power turbine. The system further comprises a variable-frequency electric machine rotatably connected with the free power turbine and a power converter configured to convert a variable frequency electrical output from the electric machine to a fixed frequency output.

Abstract: A lubrication system for a gas turbine engine, the system comprising: a gearbox, the gearbox comprising a sump; an oil tank; a primary gearbox lubrication system configured to pump oil from the oil tank to lubricate the gearbox with a gearbox primary feed; a secondary gearbox lubrication system, configured to lubricate the gearbox with oil from the sump when the oil level in the sump reaches a predetermined level; wherein the system is configured to increase the oil level in the sump to at least the predetermined level in response to a failure of the primary gearbox lubrication system.

Abstract: A nacelle for a gas turbine engine having a longitudinal centre line includes an intake lip disposed at an upstream end of the nacelle. The intake lip includes a crown and a keel. The crown includes a crown leading edge and the keel includes a keel leading edge. The crown leading edge and the keel leading edge define a scarf line therebetween. The scarf line forms a scarf angle (?scarf) relative to a reference line perpendicular to the longitudinal centre line. A fan casing is disposed downstream of the intake lip and includes a casing leading edge. The casing leading edge defines a droop line normal to the casing leading edge. The droop line forms a droop angle (?droop) relative to the longitudinal centre line. A relationship between the droop angle (?droop) and the scarf angle (?scarf) is given by: ?droop=?scarf/1.5±1 degree.

Abstract: A monitoring system for monitoring one or more properties associated with a rotating shaft is provided. The system includes a first phonic wheel which is mounted coaxially to the shaft for rotation therewith, the first phonic wheel comprising a circumferential row of teeth. The system further includes a first sensor configured to detect the passage of the row of teeth of the first phonic wheel by generating a first alternating measurement signal. The system further includes a processor unit configured to determine the durations of successive first speed samples. Each first speed sample is a block of n successive cycles of the first alternating measurement signal, where n is an integer, and in which the beginning of each cycle is a zero-crossing point from the previous cycle and the end of each cycle is the corresponding zero-crossing point to the next cycle.

Abstract: A supersonic gas turbine engine for an aircraft that comprises a nacelle, a fan, an engine core including a primary duct configured to guide a core airflow through the engine core, a bypass duct extending between the engine core and an engine casing and configured to guide a bypass airflow through the bypass duct, an intake located upstream of the fan, and a tertiary airflow duct extending between the engine casing and the nacelle and configured to guide a tertiary airflow. The intake is configured to extract air from the intake and guide it to the tertiary airflow duct in which the extracted air flows as tertiary airflow. It is provided that at least one heat exchanger is mounted in the tertiary airflow duct.

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: There is disclosed an exhaust nozzle for a gas turbine engine. The exhaust nozzle comprises a frame extending along a longitudinal axis, and a convergent petal pivotably attached at a convergent pivot point to the frame and extending axially downstream and radially inward from the frame. The exhaust nozzle comprises a follower roller fixed to the convergent petal on a radially outer side of the convergent petal, and a cam defining a working surface configured to engage the follower roller to react a force from the convergent petal. The cam is movable along a travel in an axial direction to actuate radial movement of the follower roller to pivot the convergent petal. The cam defines a convex working surface such that a contact angle between the follower roller and the cam varies along the travel to thereby vary an axial component of the force reacted by the cam.

Abstract: A method of reassembling a gas turbine engine including steps to support a core engine on a core stand and support a fan case on a fan stand that includes a base frame coupled a fan case frame by a hinge. The fan case frame is arranged to (i) rotate about an axis of the hinge between abutting the base frame and being perpendicular to the base frame, and (ii) tilt about a yaw axis perpendicular to the hinge, which is located in a plane of the fan case frame. The method includes rotating and tilting the fan case frame to align the fan case and core engine, translating at least one of the core engine and the fan case into axial abutment, coupling the fan case and the core engine to reassemble the gas turbine engine, and decoupling the fan case from the fan stand.

Abstract: An axial flow turbomachine (102) for producing thrust to propel an aircraft is shown. The turbomachine has an inner duct (202) and an outer duct (204), both of which are annular and concentric with one another. An inner fan (206) is located in the inner duct, and is configured to produce a primary pressurised flow (P). An outer fan (207) is located in an outer duct, and is configured to produce a secondary pressurised flow (S). The outer fan has a hollow hub (208) through which the inner duct passes. A swept area of the outer fan is from 2 to 20 times greater than a swept area of the inner fan.