Rolls-Royce Patente

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A fuel system for a gas turbine engine (103) configured to combust hydrogen fuel. The fuel system comprises a main fuel conduit (226), a fuel pump (224) configured to operate on hydrogen within the fuel conduit (224) to provide pressurised fuel to a core combustor (236) of the gas turbine engine (103), an auxiliary combustor (228) downstream in fuel flow of the fuel pump (224), and configured to combust a portion of fuel diverted from the main fuel conduit (224) and to heat a remainder of fuel in the main fuel conduit (224), and a fuel turbine (234) downstream in fuel flow of the auxiliary combustor (228). The fuel turbine (234) is configured to be driven by the heated fuel from the auxiliary combustor (228) and configured to power the fuel pump (224). The fuel system comprises a turbine inlet valve (247) provided upstream in hydrogen fuel flow of the fuel turbine (234) and configured to control mass flow rate and I or pressure of fuel flowing into the fue turbine (234).I. A method of operation is also described.

Electrical power systems and methods of operating electrical power systems are provided. In one aspect, an electrical power system comprises first and second electrical machines, each comprising a permanent magnet rotor and a stator comprising a plurality of electrically isolated polyphase winding sets; and a control and health monitoring system configured to detect rotor demagnetization of the first and/or second electrical machine by: supplying electrical power to each polyphase winding set of the plurality of polyphase windings sets of the first electrical machine; stopping the supply of electrical power to a first polyphase winding set of the first electrical machine; measuring a first voltage across the first polyphase winding set of the first electrical machine; supplying electrical power to each polyphase winding set of the plurality of polyphase windings sets of the second electrical machine; stopping the supply of electrical power to a first polyphase winding set of the second electrical machine; measuring a second voltage across the first polyphase winding set of the second electrical machine; determining first and second magnetic flux linkage indicators based on the first and second voltages; and determining whether the permanent magnet rotor of the first and/or second electrical machine is demagnetized based on the whether a ratio of the first and second magnetic flux linkage indicators passes a threshold.

An aircraft propulsion system comprises first and second engines (100, 200) and first (134) and second electrical machines coupled to the first and second engines respectively. The first electrical machine is arranged to be driven by electrical power in order to at least partially drive the first engine. The second electrical machine is arranged to be driven by the second engine to generate electrical power. An electrical network (31) allows transmission of electrical power between the electrical machines. A controller (25) is arranged to selectively reduce fuel flow to the first engine and provide electrical power from the second electrical machine to the first electrical machine to drive the first electrical machine, based on expected properties of contrails formed in an exhaust plume of the first engine and/or the second engine. The system allows the formation of contrails by the engines to be managed in order to mitigate climatewarming effects of the contrails.

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A hydrogen-fuelled aircraft 200 comprises a fuselage 218, wings 216A, 216B and a plurality of cylindrical elongate hydrogen storage tanks 212A-I mounted on the exterior of the fuselage and disposed vertically above a plane 210 containing the wing-tips of the aircraft. An aerodynamic fairing 214 forms a continuous air-washed surface of the aircraft together with a lower portion of the fuselage 218. The hydrogen storage tanks extend generally parallel to the fuselage and are contained between the fairing and the fuselage. The aircraft 200 provides for storage of a large volume of compressed hydrogen without compromising the volume of usable space within the fuselage. The aircraft is able to carry out flight missions comparable in range and payload to those of an equivalent conventionally-fuelled aircraft.

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A method (300) for manufacturing a composite component (50) includes providing a tool (100) switchable between a deposition configuration (105) and a flange-forming configuration (107). The tool (100) includes a first portion (110) including a first deposition surface (111), a second portion (120) including a second deposition surface (121) and a pre-form support surface (122), and a flange-forming block (130) including a curved surface (131). The method (300) further includes depositing a composite material at least partially on each of the first deposition surface (111) and the second deposition surface (121) in the deposition configuration (105) to provide a composite pre-form (140). The method (300) further includes moving the second portion (120) relative to the first portion (110) to transition the tool (100) to the flange-forming configuration (107), which causes the composite pre-form (140) to at least partially engage each of the pre-form support surface (122) and the curved surface (131). The curved surface (131) defines a radius (51R) of a flange (51) of the composite component (50).

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A cleaning system (100) for cleaning gas paths (210) in an engine core (211) of a gas turbine engine (200) is provided. The cleaning system (100) includes a source (102) of an engine cleaning liquid (104); an engine cleaning mist forming unit (106) that vapourises the engine cleaning liquid (104) to form an engine cleaning mist (114) and delivers the engine cleaning mist (114) into the engine core (211) of the gas turbine engine (200); at least one delivery device (108) configured to deliver the engine cleaning liquid (104) to the engine cleaning mist forming unit (106); a pump (110) configured to draw the engine cleaning mist (114) through the engine core (211) to clean the gas paths (210) within the engine core (211); and a mist collecting arrangement (112) including a condensing chamber (116). The mist collecting arrangement (112) is configured to collect the engine cleaning mist (114) that has passed through the engine core (211) and condense the collected engine cleaning mist (114) in the condensing chamber (116). A method (600) of cleaning an engine core (210) of a gas turbine engine (200), and a gas turbine engine (200), are also described.

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There is provided a fuel management system (400A) for a gas turbine engine. The fuel management system comprises: a fuel tank (41) configured to store fuel for the gas turbine engine; a fuel supply line (410) configured to supply fuel from the fuel tank to a combustor (16) of the gas turbine engine; a reheat fuel supply line (412) configured to supply fuel from the fuel tank to a reheat of the gas turbine engine, the reheat fuel supply line extending from a reheat branching point (411) on the fuel supply line to the reheat; a fuel supply pump (451) disposed along the fuel supply line upstream of the reheat branching point; a reheat pump (460) disposed along the reheat fuel supply line, the reheat pump configured to pressurise fuel to a reheat delivery pressure for delivery to the reheat; and a reheat recirculation line (480) configured to recirculate fuel from the reheat fuel supply line to a location upstream of the fuel supply pump, the reheat recirculation line extending from a reheat recirculation branching point (471) on the reheat fuel supply line downstream of the reheat pump.

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A gas turbine engine for an aircraft comprises, in axial flow sequence, a heat exchanger module, and a core engine. The core engine comprises, in axial flow sequence, an intermediate-pressure compressor, a high-pressure compressor, a high pressure turbine, and a low-pressure turbine. The high-pressure compressor is rotationally connected to the high-pressure turbine by a first shaft, and the intermediate-pressure compressor is rotationally connected to the low-pressure turbine by a second shaft. The heat exchanger module is in fluid communication with the core engine by an inlet duct. The heat exchanger module comprises a central hub and a plurality of heat transfer elements extending radially outwardly from the central hub and spaced in a circumferential array, for transfer of heat energy from a first fluid contained within the heat transfer elements to an inlet airflow passing over a surface of the heat transfer elements prior to entry of the airflow into an inlet to the core engine. The gas turbine engine further comprises a first electric machine. The first electric machine is rotationally connected to the first shaft, the first electric machine is positioned downstream of the heat exchanger module, the first electric machine has a maximum electrical power output of between 180kW and 500kW, the first electric machine has an axial length in the range of between 0.20m and 0.45m, and the first electric machine has a maximum diameter in the range of between 0.15m and 0.39m.

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A method of electron beam welding comprising splitting the output of an electron beam welder into two components, a pre/post heat ring and a fusion spot by rastering between the pre/post heat ring and a fusion spot, applying the two outputs to a workpiece that is to be welded, traversing the two outputs along the desired weld path, and wherein the fusion spot lies within the pre/post heat ring and travels in tandem with and inside the pre/post heat ring. The pre/post heat ring may be annular. The fusion spot may be located at the centre of the pre/post heat ring. The output of the electron beam welder may comprise 1 to 100,000 discrete points. The beams may be deflected using the EB welder deflector coils. The time the electron beam spends on each discrete point may be identical.

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