Rolls-Royce Patente

A thermal barrier coating (110) for a substrate (102) that has a major surface (104). The thermal barrier coating (110) comprises: a bond coat layer (112) disposed on the major surface (104) of the substrate (102); a thermally grown oxide (TGO) layer (114) disposed on the bond coat layer (112); and a top coat layer (116) disposed on the TGO layer (114). The top coat layer (116) comprises a ceramic top coat material (120) that comprises: a matrix phase (122) comprising at least one compound of formula ABO<4>, AB<3>O<7> or AB<3>O<9>, where A is a rare-earth element, B is selected from the group consisting of niobium and tantalum, and O is oxygen; and an inclusion phase (124) that is dispersed in the matrix phase and comprising at least one compound of formula CO<2>, where C is selected from the group consisting of zirconium, titanium or hafnium, and O is oxygen.

A fluid flow machine comprising a first rotor (110) and a stator (130) positioned alongside the first rotor, each of the first rotor (110) and the stator (130) comprising a plurality of circumferentially distributed turbomachine blades (111, 131), wherein the stator comprises a stator shroud structure (132) comprising a stator shroud surface (133) and the first rotor comprises a first rotor shroud structure (112) comprising a first rotor shroud surface (113), the stator shroud surface and first rotor shroud surface together at least partially defining a radially inner flow surface of the fluid flow machine, wherein the first rotor shroud structure defines a first cavity (114), the first cavity being disposed radially inward of the first rotor shroud surface, and wherein the stator shroud structure comprises a first protrusion (135a), the first protrusion being disposed radially inward of the radially inner flow surface, wherein the first protrusion extends in a direction with an axial component and into the first cavity of the first rotor shroud structure.

A system (100, 200) for monitoring and optimizing a repair operation being performed on a component (102, 202) includes at least one light sensor (104, 204) disposed proximal to the component (102, 202) to collect heat signature data (110, 210) emitted from the component (102, 202). The at least one light sensor (104, 204) is disposed in an oblique orientation (D1) relative to the component (102, 202). The system (100, 200) further includes a controller (108, 208) communicably coupled with the at least one light sensor (104, 204). The controller (108, 208) is configured to receive the heat signature data (110, 210) from the at least one light sensor (104, 204), compare the heat signature data (110, 210) with a nominal range for the heat signature, and generate an output signal (112, 212) if the heat signature data (110, 210) is outside of the nominal range for the heat signature.

There is provided a busbar (300, 301, 302) comprising a plurality of tabs (312, 314, 316), a fuse link (323, 325) and a cartridge (333, 335). The plurality of tabs (312, 314, 316) are offset from one another along a separation direction (502), with each tab (312, 314, 316) being configured to electrically couple with at least one energy storage device (400). The fuse link (323, 325) extends between two adjacent tabs (312, 314, 316) of the plurality of tabs (312, 314, 316), with the fuse link (323, 325) being configured to form a gap (355) when a fault current flows along the fuse link (323, 325). The cartridge (333, 335) surrounds the fuse link (323, 325). The cartridge (333, 335) is configured to inhibit electrical arcing originating from the gap (355).

A blower compressor assembly (402) comprises: a flow modifier (412) comprising an array of nozzle guide vanes (525) and an array of diffuser vanes (515), wherein the flow modifier is movable relative to a rotor (410) between a turbine configuration and a compressor configuration; and an actuator assembly (600) for moving the flow modifier relative to the rotor between the turbine configuration and the compressor configuration, wherein the actuator assembly comprises: a first actuator (610) comprising a first chamber (613), a first valve (616) configured to selectively supply pressurised air to the first chamber (613) or vent pressurised air from the first chamber, and a piston (614) arranged to one side of the first chamber and coupled to the flow modifier such that the flow modifier is moved into the turbine configuration or the compressor configuration based on the pressure of air within the first chamber; and a second actuator, configured to controllably adjust the position of the piston when the blower compressor is in the compressor configuration.

A fuel system for an aircraft gas turbine engine (103) comprises: a main fuel conduit (206) configured to deliver hydrogen fuel to a combustor (203) of the gas turbine engine (103); a fuel metering valve (213) configured to modulate a mass flow of fuel delivered to the gas turbine engine combustor (206); and a bypass valve (209, 210) provided upstream in main fuel conduit hydrogen fuel flow of the fuel metering valve (213) and configured to vent a portion of hydrogen fuel upstream of the fuel metering valve (213). The fuel system further comprises a controller (218) configured to control the fuel metering valve (213) and the vent valve (209, 210). The controller (218) is configured to, during an engine start procedure, open the fuel metering valve (213) and open the vent valve (209, 210) to flow a first fuel portion to the combustor (206) via the fuel metering valve (213) and to vent a second fuel portion via the vent valve.

A fuel system for an aircraft gas turbine engine (103) comprises:a main fuel conduit (206) configured to deliver hydrogen fuel to a combustor (203) of the gas turbine engine (103);a fuel metering valve (213) configured to modulate a mass flow of fuel delivered to the gas turbine engine combustor (206); anda variable pressure control valve (207) upstream of the fuel metering valve (213), the pressure control valve (207) being configured to control pressure of fuel flow delivered to the fuel metering valve (213).

An electrical machine having a monitoring apparatus, the monitoring apparatus comprising at least one first fibre optic cable (207) having a periodic structure within at least a section of the first fibre optic cable, the first fibre optic cable being coupled to an outcoupling optical set up (208) comprising at least one first lens (208), and an interrogator (210) connected to a second fibre optic cable and at least one second lens (209), the light transmitted through the first fibre optic cable (207) is manipulated by the first lens and transmitted to the second lens (209) which is coupled to the second fibre optical cable which transfers a light signal to the interrogator (210), and wherein a portion of the first fibre optical cable is connected to a component (202) within the electrical machine.

A data recorder comprising: at least one processor; at least one memory comprising computer readable instructions; the at least one processor being configured to read the computer readable instructions to cause performance of: receiving one or more of: sensor configuration data; or sensor calibration data from a memory of a connector connected between the data recorder and a sensor; reading the connector to receive a digital signal; processing the received digital signal; and controlling storage of data in the processed digital signal in the at least one memory of the data recorder.

A connector for connecting a sensor to a data recorder, the connector comprising: circuitry configured to receive a signal from the sensor and to transmit a digital signal to the data recorder; and a memory storing one or more of: sensor configuration data or sensor calibration data.

A system (100) for repairing a damaged blade (102) of a gas turbine engine (10) includes a sacrificial build platform (106, 106-1, 106-2) configured to be positioned and fused with the damaged blade (102) proximal to at least one stub portion (108, 108-1, 108-2) of the damaged blade (102). The at least one stub portion (108, 108-1, 108-2) protrudes from a root (110) of the damaged blade (102) and defines a deposition surface (112). The sacrificial build platform (106, 106-1, 106-2) includes a base (114) defining a pair of cavities (116a, 116b). The at least one stub portion (108-2) includes a pair of projections (108a, 108b), each cavity (116a, 116b) is configured to receive a corresponding projection (108a, 108b) from the pair of projections (108a, 108b). The system (100) further includes an additive manufacturing system (150) that is configured to deposit material on the base (114) of the sacrificial build platform (106, 106-1, 106-2) and the deposition surface (112) of the at least one stub portion (108, 108-1, 108-2) to form a base structure (118) of a repaired blade (122).

A thermal barrier coating (110) for a substrate (102) that has a major surface (104). The thermal barrier coating (110) comprises: a bond coat layer (112) disposed on the major surface (104) of the substrate (102); a thermally grown oxide (TGO) layer (114) disposed on the bond coat layer (112); and a top coat layer (116) disposed on the TGO layer (114). The top coat layer (116) comprises a ceramic top coat material (120) that comprises: a matrix phase (122) comprising at least one compound of formula ABO<4>, AB<3>O<7> or AB<3>O<9>, where A is a rare-earth element, B is selected from the group consisting of niobium and tantalum, and O is oxygen; and an inclusion phase (124) that is dispersed in the matrix phase and comprising one or more of alumina and a compound of formula A<3>AlO<12>, AAlO<3> or A<4>Al<2>O<9>, where A is a rare-earth element, Al is aluminium, and O is oxygen.

An aircraft (10) comprises a wing (14a, 14b) and a fuel tank (26a, 26b) mounted above the wing (14a, 14b), spaced from an upper surface (18) of the wing (14a, 14b). A first portion of a (38) lower surface (36) of the tank (26a, 26b) comprises a concave profile extending in a longitudinal direction above a convex upper surface portion (18) of the wing (14a, 14b).

A gas turbine engine 108A comprises an engine core 201 having first 222 and second 224 engine core exhaust paths arranged to pass first and second portions respectively of the mass flow of the engine's core exhaust mass. A heat-exchange system comprises a recuperator system 230A-D disposed within the first engine core exhaust path and arranged to transfer heat from said first portion to a buffer fluid, and a heat exchanger arranged to transfer heat from the buffer fluid to fuel within a fuel path arranged to convey fuel to the engine's combustor 208. The engine provides for heat to be recovered from the engine's core exhaust flow to the engine's fuel supply, thus improving thermal efficiency, but without significantly impeding the engine core exhaust flow or presenting the significant fire or explosion risk associated with a recuperator arranged to heat fuel directly.

A blower system (400) for supplying pressurized air to an airframe of an aircraft comprises: an air inlet (422) for receiving air from an air source; an air outlet (426) for supplying pressurized air to an airframe; a blower compressor (420) in fluid communication with the air inlet and the air outlet, the blower compressor comprising a first rotor (420a) and a second rotor (420b), wherein the first and second rotors are independently rotatable in order to compress air received from the air inlet and deliver the compressed air to the air outlet; a variable transmission (430) comprising a first output (436) for driving the first rotor, a second output (438) for driving the second rotor, and first, second and third inputs (432, 434, 435); and first and second electrical machines (464, 465), each comprising an electric rotor and an electric stator, wherein: the first input of the variable transmission is configured to be mechanically coupled to a spool (440) of a gas turbine engine (10); the second input of the transmission is mechanically coupled to the electric rotor of the first electrical machine; the third input of the transmission is mechanically coupled to the electric rotor of the second electrical machine; and the electric stators of the first and second electrical machines are electrically connected to an electrical power management system (466) configured to control the supply of electrical power to the first and second electrical machines in order to control the speed of the first and second outputs of the variable transmission.

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A feed mechanism for inserting and/or retracting a flexible robotic arm from an area, the feed mechanism comprising a housing, a passageway that extends about a central longitudinal axis through the length of the housing, a drive portion and a rotational portion, the drive portion not being fixedly connected within the housing and having at least a pair of drive wheels coupled to a drive motor, the drive wheels are connected to a mounting frame with the mounting frame being connected to pivotable linkages with at least one of the linkages being connected to a spring, such that the expansion or contraction of the spring allows the drive wheels to move relative to the longitudinal axis, and the rotational portion being coupled to a motor, the rotational motor having a gear system that connects to the drive portion and causes a rotation of the drive portion about the central longitudinal axis within the housing.

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A combustor assembly (100) for a gas turbine engine includes a combustor shell (110) extending along a shell axis (111). The combustor shell (110) includes a shell wall (112) extending circumferentially around the shell axis (111) and a meter panel (116) connected to an upstream end (113) of the shell wall (112). The shell wall (112) and the meter panel (116) together define a combustion space (165) therebetween. The combustor assembly (100) further includes: a tile (120) disposed within the combustor shell (110) and including a tile flange portion (125) extending radially with respect to the shell axis (111); a heatshield (140) connected to the meter panel (116) and extending at least radially towards the tile (120); a fastener (150) connecting the tile flange portion (125) to the meter panel (116); and a compartment (160) thermally shielded from the combustion space (165). Either the heatshield (140) and the tile (120) together form the compartment (160) or the tile (120) alone forms the compartment (160). The fastener (150) is spaced apart from the combustion space (165) and at least partially disposed within the compartment (160).

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An aircraft comprises a hydrogen-fuelled propulsion system, a plurality of like generally cylindrical hydrogen storage tanks and a conveying system arranged to convey hydrogen from the hydrogen storage tanks to the hydrogen-fuelled propulsion system. The aircraft further comprises a fuselage having a cargo bay (502) including one or more (510A-G) of the plurality of hydrogen storage tanks, the longitudinal axes (511A-G) of the one or more hydrogen storage tanks within the cargo bay extending parallel to the longitudinal axis (501) of the fuselage and lying in one or more planes (595, 597) extending across the width dimension of the cargo bay. The hydrogen storage tanks within the cargo bay have a common aspect ratio R in the range 42 ≤ R ≤ 25.7, allowing the volume of space with the cargo bay occupied by stored hydrogen to be maximised or approximately maximised.

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There is provided a dynamic sealing assembly for a rotary machine 400, comprising a primary sandwich plate 710, a secondary sandwich plate 720 and a bristle pack 730. The primary sandwich plate 710 comprises a plurality of primary vane openings 712, and the secondary sandwich plate 720 comprises a plurality of secondary vane openings 722. The bristle pack 730 comprises a plurality of bristles and is disposed between the primary sandwich plate and the secondary sandwich plate. Each of the plurality of primary vane openings 712 overlies and aligns with a respective secondary vane opening 722 to form a vane channel 715A, 715C, 7150 for receiving a vane 515A-515E along a longitudinal axis 703 of the dynamic sealing assembly 700. The bristle pack 730 is configured to: provide a brush seal between each vane 515A-515E received within the respective vane channels 715A, 715C, 7150 and the dynamic sealing assembly 700; and allow relative movement between the dynamic sealing assembly 700 and the vane 515A-515E received within each vane channel 715A, 715C, 7150 along the longitudinal axis 703.

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Disclosed is a fuel system for a gas turbine engine having a core exhaust and a combustion section. The system comprises a fuel pump for fluid communication with a fuel reservoir; a driving turbine for driving the fuel pump; and an air feed to drive the driving turbine, the air feed being fluidly connectable to the core exhaust of the gas turbine engine. Also disclosed is a fuel system comprising: a fuel pump for fluid communication with a fuel reservoir; a fuel line from the fuel pump for fluid communication with the combustion section of the gas turbine engine; and an air feed to a heat exchanger, the air feed being fluidly connectable to the core exhaust of the gas turbine engine, wherein the fuel line is in thermal communication with the air feed within a heat exchanger.

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