Rolls-Royce Patente

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A gas turbine engine (10) for an aircraft comprises an engine core (11) comprising a turbine (19), a combustor (16), a compressor (14), and a core shaft (26) connecting the turbine to the compressor. The engine comprises a fan (23) located upstream of the engine core and arranged to be driven by the core shaft. The engine comprises a nacelle (21) surrounding the fan and the engine core and defining a bypass duct (22) located radially outside of the engine core, where the bypass ratio, defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core at cruise conditions, is at least 4. The engine comprises an engine heat management system (1700) and an actuator (1704) configured to actuate at least one valve (1704a) within that system. The engine comprises a fuel supply system (152) arranged to supply fuel to fueldraulically drive the actuator. The actuator is configured to actuate said valve so as to enable non-binary position adjustment between an open valve position and a closed valve position.

In one example, a method for forming a coating system including an environmental barrier coating (EBC) and abradable coating on a substrate. The method may include depositing the EBC on a ceramic or ceramic matric composite (CMC) substrate to form an as-deposited EBC; heat treating the as-deposited EBC following the deposition of the as-deposited EBC on the substrate to form a heat treated EBC; and subsequently depositing an abradable coating on the heat treated EBC to form an as-deposited abradable coating. The as-deposited abradable coating may be heat treated to form a heat treated abradable coating.

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An afterburner arrangement (200) comprising: an internal casing (214) and an external casing (216) defining a bypass pathway (218) between them; a mounting strut (228) forming a structural connection between the internal casing and the external casing; and A plurality of fuel nozzles (222/224/226) associated with the mounting strut, wherein the mounting strut at least partly houses a corresponding plurality of fuel pathways (250/254/258) to provide fuel to the respective fuel nozzles.

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There is provided a blower system 400 for providing air to an airframe system 450, comprising a rotor 410 configured to be mechanically coupled to a spool 440 of a gas turbine engine 10, wherein the rotor 410 is configured to: in a blower mode, be driven to rotate by the spool 440 to discharge air to an airframe discharge port 426 for supply to an airframe system 450; and, in an engine drive mode, receive air from an external air source 470 via an impingement port 424 that is configured to direct the received air onto the rotor and thereby drive the rotor to rotate to drive the spool 440 to.

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A system (100) for measuring loading on a test specimen (102). The system (100) includes the test specimen (102) arranged between a first loading bar (104) and a second loading bar (106). The system (100) further includes a first loading unit (112) and a second loading unit (116) configured to apply a first load (114) and a second load (118) to the first and second loading bars (104, 106), respectively. The system (100) further includes a first clamp (120) and a second clamp (130) configured to hold the first and second loading bars (104, 106) against the first and second loads (114, 118), respectively. The system (100) further includes a clamp actuating unit (140) configured to selectively release at least the first clamp (120). The clamp actuating unit (140) further includes a controller (150) configured to electrically actuate at least one first electromechanical transducer (142) from a retained state (P1) to a released state (P2) to release the first clamp (120), such that the first loading bar (104) applies a first loading wave (122) to the test specimen (102).

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Apparatus for supporting at least a part of an engine, the engine comprising: an engine stand configured to support at least a part of an engine; an accelerometer coupled to the engine stand and configured to: measure acceleration of the engine stand, and generate accelerometer data; and a first processor configured to receive the accelerometer data from the accelerometer.

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Reconfigurable permanent magnet electrical machines and aircraft power and propulsion systems comprising such electrical machines are described. In one aspect, an aircraft power and propulsion system comprises: a gas turbine engine; a permanent magnet electrical machine comprising a rotor drivingly coupled to a spool of the gas turbine engine, and a stator comprising windings controllably switchable between a star configuration and a delta configuration; a DC electrical network; an AC:DC power electronics converter, an AC side of which is connected to terminals of the stator windings of the electrical machine and a DC side of which is connected to the DC electrical network; and a control system. The control system is configured to monitor at least one operating condition of the power and propulsion system and to control the switching of the stator windings between the star configuration and the delta configuration based on the at least one operating condition.

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