Rolls-Royce Patente

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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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A method of monitoring turbine blade creep in a gas turbine engine is provided. The method includes: receiving stereo images of a turbine blade of a row of turbine blades, the images having been obtained using a stereo borescope located in the engine adjacent the row of turbine blades; identifying same features of the blade in each of the stereo images; mapping each of the identified features by triangulation onto a 3D space to produce a 3D depth map of at least part of the blade; providing a 3D reference model of the blade; and comparing the 3D reference model with the 3D depth map to measure one or more deviations in shape of the blade to determine an amount of creep-induced distortion of the blade.

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A vane cooling system for a gas turbine engine comprises a vane (21) arranged on a stator and having a chamber (23) extending continuously from a radially inner end to a radially outer end of the vane. The vane (21) has; a radially inner inlet (24) and a radially outer inlet (25), a first cooling fluid feed (39) in communication with the radially inner inlet (24) and a second cooling fluid feed (28) in communication with the radially outer inlet (25), The first cooling fluid feed (39) has a higher pressure than the second cooling feed (28). A flow adjustment device (30) is arranged for adjusting a flow of the second cooling fluid feed into the radially outer inlet (25).

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A brush seal (400, 600, 700, 900, 900-1, 1000) includes an annular plate assembly (402, 602, 702, 902, 902-1, 1002) defining a radially inner diameter (D2, D2-1, D2-2, D2-4, D2-5, D2-6, D2-7) having an inner central axis (A2, A2-1, A2-2, A2-4, A2-5, A2-6, A2-7) and a radially outer diameter (D3, D3-1, D3-2, D3-4, D3-5, D3-6, D3-7) having an outer central axis (A3, A3-1, A3-2, A3-4, A3-5, A3-6, A3-7). The inner central axis (A2, A2-1, A2-2, A2-4, A2-5, A2-6, A2-7) and the outer central axis (A3, A3-1, A3-2, A3-4, A3-5, A3-6, A3-7) are spaced apart from each other by an offset (O1, O2, O3, O4, O5, O6, O7), such that the radially inner diameter (D2, D2-1, D2-2, D2-4, D2-5, D2-6, D2-7) and the radially outer diameter (D3, D3-1, D3-2, D3-4, D3-5, D3-6, D3-7) are eccentric with respect to each other. The brush seal (400, 600, 700, 900, 900-1, 1000) further includes a bristle pack (408, 608, 708, 908, 908-1, 1008). The bristle pack (408, 608, 708, 908, 908-1, 1008) includes a plurality of bristles (410, 610, 710, 910, 910-1, 1010) defining a bristle bore (412, 612, 712, 912, 912-2, 1012) having a bristle bore central axis (A4, A4-1, A4-2, A4-4, A4-5, A4-6). The bristle bore central axis (A4, A4-1, A4-2, A4-4, A4-5, A4-6) is aligned with the inner central axis (A2, A2-1, A2-2, A2-4, A2-5, A2-6, A2-7) and the bristle bore central axis (A4, A4-1, A4-2, A4-4, A4-5, A4-6) is offset from the outer central axis (A3, A3-1, A3-2, A3-4, A3-5, A3-6, A3-7). Moreover, the brush seal (400, 600, 700, 900, 900-1, 1000) is disposed around a rotating component (106). The bristle bore central axis (A4, A4-1, A4-2, A4-4, A4-5, A4-6) is aligned with a component axis (A1) of the rotating component (106) and the bristle pack (408, 608, 708, 908, 908-1, 1008) engages with the rotating component (106).

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