Rolls Royce Plc

The foregoing describes a power converter apparatus with at least a first power converter and a second power converter connected in parallel, so as to each receive an AC voltage generated by an electrical machine and output a DC voltage on a DC bus. Each of first and second power converters is coupled to respective first and second winding arrangements for receiving the AC voltages, and the first winding arrangement is grounded by a first impedance and the second winding arrangement is grounded by a second impedance. The first impedance has an effective impedance value which is different to that of the second impedance such, in use, a zero-sequence current between the first and second power converters at least partly cancels a common-mode voltage component.

A fuel injector for a combustor ( 15, 300 ) of a gas turbine engine ( 10 ) is described. The fuel injector comprises a fuel injection device ( 360 ) having a fuel injection port ( 362 ) configured to inject fuel into the combustor, wherein the fuel injection device is configured to promote cavitation of the fuel. Also disclosed is a combustor ( 15 ) for a gas turbine engine ( 10 ), a reheat assembly ( 300 ) for a gas turbine engine ( 10 ), a gas turbine engine ( 10 ), and an aircraft ( 200 ).

There is provided a flameholder for a reheat assembly of a gas turbine engine. The flameholder comprises a fuel atomizer, an air passageway and a fuel passageway. The air passageway is configured to convey a flow of air to the fuel atomizer. The fuel passageway is configured to convey a flow of fuel to the fuel atomizer.

There is provided a propulsion machine comprising a fluid duct defined by a wall and a moveable member. The propulsion machine also comprises a mounting structure coupled to the moveable member, and an extendable structure having a sealing surface and a sensing arrangement. The extendable structure is moveable relative to the mounting structure to provide a seal between the sealing surface and an opposing surface of the wall. The sensing arrangement is configured to generate one or more signals indicative of a position of the extendable structure relative to the mounting structure.

A fuel injector for a combustor of a gas turbine engine is described. The fuel injector includes a fuel injection device having a fuel injection port configured to inject fuel into the combustor, wherein the fuel injection port includes an elongate aperture. A combustor for a gas turbine engine, a reheat assembly for a gas turbine engine, a gas turbine engine, and an aircraft.

A turbine exit diffuser for a gas turbine engine, the turbine exit diffuser comprising: an exhaust duct arranged to receive flow from a turbine; and a plurality of fingers extending axially downstream and radially away from the exhaust duct, each finger of the plurality being circumferentially spaced apart from an adjacent finger of the plurality, wherein the fingers are arranged to contact a reheater during attachment of the reheater to the turbine exit diffuser.

There is provided a reheat assembly for a gas turbine engine. The reheat assembly comprises a jetpipe casing, a support duct, a plurality of fuel discharge ports, and a fuel distribution passageway. The jetpipe casing comprises a reheat core section and a reheat bypass section. The support duct radially separates the reheat core section and the reheat bypass section. Each fuel discharge port is configured to discharge fuel received from the fuel distribution passageway. The fuel distribution passageway is defined by an interior channel embedded within the support duct.

A gas turbine engine including a rear attachment comprising: a plurality of angularly distributed rotational pylon connections for coupling to a wing pylon and defining a pylon connection angular extent; a plurality of angularly distributed engine connections for coupling to a rear mount structure of the gas turbine engine 10 and defining an engine connection angular extent; and a link structure extending between the pylon connections and the engine connections; wherein the link structure, pylon connections and engine connections are configured to transfer a roll torque from the rear mount structure to the wing pylon; and wherein a ratio of the engine connection angular extent to the pylon connection angular extent is at least 1.75.

A geared gas turbine engine comprises a heat management system configured to provide lubrication and cooling to a power gearbox and turbomachinery bearings, and comprising a pipe assembly adapted to provide a lubricant flow to the power gearbox and turbomachinery bearings to remove the heat generated by the power gearbox and turbomachinery bearings, an air-lubricant heat exchanger to dissipate a first amount of heat, and a fuel-lubricant heat exchanger to dissipate a second amount of heat wherein the heat management system is configured to provide the first amount of heat and the second amount of heat such that at cruise conditions a proportion of heat generated by the gearbox and the turbomachinery and dissipated to air is in the range of from 0.35 to 0.80.

An aerofoil component for a high-pressure or intermediate-pressure turbine of a multi-spool gas turbine engine having an aerofoil member which spans a working gas annulus of the gas turbine engine. The aerofoil member has pressure side and suction side aerofoil surfaces which each extend from a leading to a trailing edge such that transverse cross sections through the aerofoil member provide respective aerofoil sections. The spacing between the leading and trailing edges on the midspan aerofoil section defines a midspan axial chord length. The aerofoil member are arranged in a circumferential row around the annulus with plural, equally spaced, identical aerofoil members whereby the circumferential spacing of the aerofoil members at the trailing edges of their midspan aerofoil sections defines a midspan pitch of the aerofoil member. At midspan: the value of S/C x is in the range from 1.4 to 1.6; and the uncovered turning angle is 16° or more.