Rolls-Royce Patente

A probe (100) for non-destructive testing of a component (50). The probe (100) includes a head (102) configured to transmit a signal or a wave to a surface (52) of the component (50) for non-destructive testing of the component (50). The head (102) includes a first head end (104) facing away from the component (50) and an opposing second head end (106) facing the component (50). The probe (100) further includes a tube (108) spaced apart from the head (102) and including a first tube end (110) distal to the head (102) and an opposing second tube end (112) proximal to the head (102). The tube (108) defines an internal passage (114) therein extending between the first tube end (110) and the second tube end (112). The probe (100) further includes a flexible actuator (116) disposed between the head (102) and the tube (108). The flexible actuator (116) is configured to be actuated between a non-actuated state (S1) and an actuated state (S2).

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A system for inspecting a complex component, the system comprising: a flexible inspection device having a driver for controlling the motion of the flexible inspection device by manipulating the relative positions of a plurality of joints within the device, the flexible device having a sensor on its distal end; a feed mechanism for controlling the insertion or retraction of the flexible device into or from the complex component; a component driver for driving the movement of an aspect of the complex component from at least a first position to a second position; and a computer running a computer program, the computer interfacing with the flexible inspection device, the feed mechanism, and the component driver, so as to link the operation of the flexible inspection device, the feed mechanism and the component driver through a single program.

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An air-breathing turbojet engine (101) for a hypersonic vehicle is shown. The engine comprises a pump for pumping a cryogenic fuel, an inlet (102) configured to compress inlet air by one or more shocks, a cooler (103) to cool the compressed inlet air using the cryogenic fuel, and a turbo-compressor (104) to compress the air further. A combustor (105) receives compressed cooled air from the turbo-compressor and a first portion of the cryogenic fuel for combustion. A first turbine (106) expands and is driven by combustion products, and a second turbine (107) expands and is driven by a second portion of the cryogenic fuel. The first turbine and the second turbine drive the turbo-compressor via a shaft. An afterburner (109) receives combustion products from the first turbine and the second portion of the cryogenic fuel from the second turbine for combustion therein.

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A thermal management system (100) for a gas turbine engine (10) includes a plurality of inlet and outlet vanes (104, 106) extending from a bypass inner wall (110) towards an outer cowl (120), a plurality of first removable shims (144), a plurality of second removable shims (146), and a third removable shim (148). Each first removable shim (144) extends between and engages the outer cowl (120) and a corresponding inlet vane (104). Each second removable shim (146) extends between and engages the outer cowl (120) and a corresponding outlet vane (106). The third removable shim (148) extends between and engages a heat exchanger (108) and the outer cowl (120). Each of the plurality of first removable shims (144), each of the plurality of second removable shims (146), and the third removable shim (148) is removable for positionally adjusting the outer cowl (120) relative to the bypass inner wall (110).

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A gas turbine engine (10) for an aircraft with an engine core (11) comprising a turbine (17, 19), a combustor (16) and a compressor (14, 15), the turbine comprising a first turbine (19) and a second turbine (17) and the compressor comprising a first compressor (14) and a second compressor (15), a first core shaft (26) connecting the first turbine to the first compressor, a second core shaft (27) connecting the second turbine to the second compressor, the second turbine, second compressor, and second core shaft being arranged to rotate at a higher rotational speed than the first core shaft, the gas turbine engine further comprising a fan (23) comprising a plurality of fan blades and a gearbox (30) that receives an input from the first core shaft (26) and outputs drive to the fan so as to drive the fan at a lower rotational speed than the first core shaft, wherein the second turbine comprises at least one ceramic matrix composite component and the mass of ceramic matrix composite in the second turbine is in the range of from 2% to 15% of the total mass of the second turbine.

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An engine system comprises a first fuel store of a first fuel, a second fuel store of a second fuel, an engine arranged to produce mechanical power by combustion or oxidation of fuel in the engine, a fuel distribution system arranged to deliver fuel from the first and second fuel stores to the engine, the first fuel delivered at a first flow rate, the second fuel delivered at a second flow rate, the first and second flow rates contributing to a total flow rate of fuel to the engine; and a control system arranged to control the relative fractions of the total flow rate of fuel to the engine represented by the first and second flow rates according to the required power output of the engine such that the relative fraction of the total flow rate of fuel to the engine represented by the second flow rate increases with increasing required power output of the engine for at least part of a period of operation of the engine, wherein the first and second fuels are selected such that operation of the engine using only the second fuel results in a lower engine temperature than operation of the engine using only the first fuel, when producing the same amount of mechanical power and/or wherein the second fuel has a lower specific energy than the first fuel and/or wherein the second fuel produces more water during combustion than the first fuel per unit of fuel energy released during combustion.

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