Rolls-Royce Patente

A method (200) for scanning a plurality of components (102). The method (200) includes providing an imaging beam source (108), an imaging beam receiver (114), and a support platform (104). The support platform (104) is configured to rotate and/or revolve relative to the imaging beam source (108) and the imaging beam receiver (114) about one or more axes. The method (200) further includes disposing the plurality of components (102) on the support platform (104). The plurality of components (102) is positioned in a gap (118) between a first geometrical figure (120) and at least one second geometrical figure (122) in a configuration that reduces the variation in material thickness penetrated at different relative angles of rotation and/or revolution. The first geometrical figure (120) and the at least one second geometrical figure (122) are concentric.

A fuel system for a hydrogen fuelled gas turbine engine (201). The fuel system comprises a main hydrogen fuel line (206) configured to provide hydrogen fuel to a combustor (203) of the gas turbine engine (201) and a closed loop heat exchange fluid line (207) comprising a heat exchange fluid. The fuel system further comprises a first heat exchanger (208) configured to exchange heat between gas turbine engine compressor bleed air and heat exchange fluid in the heat exchange fluid line (207), a second heat exchanger (209) downstream of the first heat exchanger (208) in heat exchange fluid flow configured to exchange heat between the heated heat exchange fluid and hydrogen in the main hydrogen fuel line, and a third heat exchanger (215) downstream in main hydrogen fuel flow of the second heat exchanger (209) in main hydrogen fuel flow configured to exchange heat between a further engine fluid and heated hydrogen fuel.

A fuel system for a hydrogen fuelled gas turbine engine (201). The fuel system comprises a main hydrogen fuel line (206) configured to provide hydrogen fuel to a combustor (203) of the gas turbine engine (201) and a closed loop heat exchange fluid line (207) comprising a heat exchange fluid. The fuel system further comprises a first heat exchanger (208) configured to exchange heat between gas turbine engine compressor bleed air and heat exchange fluid in the heat exchange fluid line (207), a second heat exchanger (209) downstream of the first heat exchanger (208) in heat exchange fluid flow configured to exchange heat between the heated heat exchange fluid and hydrogen in the main hydrogen fuel line, and a third heat exchanger (215) downstream in main hydrogen fuel flow of the second heat exchanger (209) in main hydrogen fuel flow configured to exchange heat between a further engine fluid and heated hydrogen fuel.

A nosecone for a gas turbine engine includes a nosecone body substantially symmetrical around a central axis. The nosecone body includes a side wall which defines a diameter of the nosecone body that increases from a forward side to an aft side of the nosecone body. The nosecone further includes an attachment assembly which attaches the nosecone body to the gas turbine engine. The nosecone includes a plurality of airstream protrusions attached to or integral with the nosecone body. A first airstream protrusion of the plurality of airstream protrusions is positioned upstream of a first fan blade to prepare an airstream flowing over the nosecone body for interaction with the first fan blade. The attachment assembly includes a plurality of attachment locations, and a first attachment location of the plurality of attachment locations is collocated with the first airstream protrusion of the plurality of airstream protrusions.

A nosecone (30) for a gas turbine engine (10) includes a nosecone body (31). The nosecone body is substantially symmetrical around a central axis (X-X). The nosecone body includes a side wall which defines a diameter of the nosecone body that increases from a forward side of the nosecone body to an aft side of the nosecone body. The nosecone includes an attachment assembly (33). The attachment assembly is configured to attach the nosecone body to the gas turbine engine. At least one of the nosecone body or the attachment assembly comprises a tear-resistant fiber.

A gas turbine engine (10) including a rear attachment (400) comprising: a plurality of angularly distributed rotational pylon connections (221, 222, 223) for coupling to a wing pylon (200) and defining a pylon connection angular extent; a plurality of angularly distributed engine connections (121, 122, 123) for coupling to a rear mount structure (117) of the gas turbine engine 10 and defining an engine connection angular extent; and a link structure extending between the pylon connections (221, 222, 223) and the engine connections (121, 122, 123); wherein the link structure, pylon connections (221, 222, 223) and engine connections (121, 122, 123) are configured to transfer a roll torque from the rear mount structure (117) to the wing pylon (200); and wherein a ratio of the engine connection angular extent to the pylon connection angular extent is at least 1.75.

There is provided a fluid flow machine (10) comprising: a casing structure (24), a turbomachine blade (311) disposed within the casing structure (24), and a sealing arrangement (330) coupled to the casing structure (24). The sealing arrangement (330) comprises (330) an abrasion portion (334) including a lattice structure (400). The abrasion portion (334) is configured to provide a seal with a tip (321) of the turbomachine blade (311). The abrasion portion (334) includes the lattice structure (400) having a lattice density that increases along a radial direction (42) of the machine (10).There is also provided a method (600) of manufacturing such a fluid flow machine (10).

There is provided a propulsion machine 10 comprising a fluid duct defined by a wall 42, 44 and a moveable member 34, 36. The propulsion machine 10 also comprises a mounting structure 92 coupled to the moveable member 34, 36, and an extendable structure 94 having a sealing surface 96 and a sensing arrangement 83. The extendable structure 94 is moveable relative to the mounting structure 92 to provide a seal between the sealing surface 96 and an opposing surface 43 of the wall 42. The sensing arrangement 83 is configured to generate one or more signals indicative of a position of the extendable structure 94 relative to the mounting structure 92.

There is provided a reheat assembly (300A, 300B, 300C) for a gas turbine engine (10). The reheat assembly (300A, 300B, 300C) comprises a jetpipe casing (310), a support duct (340), a plurality of fuel discharge ports (362), and a fuel distribution passageway (342). The jetpipe casing (310) comprises a reheat core section (320) and a reheat bypass section (330). The support duct (340) radially separates the reheat core section (320) and the reheat bypass section (330). Each fuel discharge port (362) is configured to discharge fuel received from the fuel distribution passageway (342). The fuel distribution passageway (342) is defined by an interior channel embedded within the support duct (340).

The foregoing describes a power converter apparatus (200) with at least a first power converter (101-1) and a second power converter (101-2) connected in parallel, so as to each receive an AC voltage generated by an electrical machine (104) and output a DC voltage on a DC bus (105P, 105N). Each of first and second power converters (101-1, 101-2) is coupled to respective first and second winding arrangements (102-1, 102-2) for receiving the AC voltages, and the first winding arrangement is grounded by a first impedance (201-1) and the second winding arrangement is grounded by a second impedance (201-2). The first impedance (201-1) has an effective impedance value which is different to that of the second impedance (202-1) such, in use, a zero-sequence current between the first and second power converters at least partly cancels a common-mode voltage component.

There is provided a flameholder (370) for a reheat assembly (300) of a gas turbine engine (10). The flameholder comprises a fuel atomizer (379A, 379B), an air passageway (375A, 375B) and a fuel passageway (374, 374A, 374B). The air passageway (375A, 375B) is configured to convey a flow of air (521A, 521B) to the fuel atomizer (379A, 379B). The fuel passageway (374, 374A, 374B) is configured to convey a flow of fuel (531A) to the fuel atomizer (379A, 379B). The fuel atomizer (379A, 379B) includes a mixing chamber (430) and a swirler (420).

The disclosure relates to apparatus for fuel and air injection into a combustion chamber of a gas turbine engine and to a gas turbine engine comprising such an apparatus. Example embodiments include an apparatus (200) for fuel and air injection into a combustion chamber (201) of a gas turbine engine (10), the apparatus (200) comprising: back panel (202) having front and back faces (203, 204), an aperture (205) configured to receive a fuel injector (206) and a plurality of air inlets (207) arranged around the aperture (205); a heatshield panel (208) having front and back faces (209, 210), the heatshield panel (208) having a central aperture (212) and an annular side wall (211) mounted against the front face of the back panel to form a cavity (213) between the back face of the heatshield panel and the front face of the back panel (202); and an annular swirler assembly (214) disposed in the cavity (213) between the back panel (202) and the heatshield panel (208), the annular swirler assembly (214) having a central passage (215) for flow of an air and fuel mixture into the combustion chamber (201) and a flange (216) engaged with first and second grooves (217a, 217b) in the back face (210) of the heatshield panel (208) and the front face (203) of the back panel (202) respectively.

A turbine exit diffuser for a gas turbine engine, the turbine exit diffuser comprising: an exhaust duct arranged to receive flow from a turbine; and a plurality of fingers extending axially downstream and radially away from the exhaust duct, each finger of the plurality being circumferentially spaced apart from an adjacent finger of the plurality, wherein the fingers are arranged to contact a reheater during attachment of the reheater to the turbine exit diffuser.

There is provided a fluid flow machine (10) comprising a casing structure (24), a turbomachine blade (311) disposed within the casing structure (24), and a sealing arrangement (330) coupled to the casing structure (24). The sealing arrangement (330) comprises (330) an abrasion portion (334) including a lattice structure (400). The abrasion portion (334) is configured to provide a seal with a tip (321) of the turbomachine blade (311). The abrasion portion (334) includes a filler (500) within the lattice structure (400), the filler comprising one or more selected from: a polyester, a polyimide, a cyanate ester, a siloxane, a polysiloxane or a polyepoxide.There is also provided a vehicle (200) comprising such a fluid flow machine (10).There is also provided a method (600) of manufacturing such a fluid flow machine (10).

A fuel injector for a combustor (15, 300) of a gas turbine engine (10) is described. The fuel injector comprises a fuel injection device (360) having a fuel injection port (362) configured to inject fuel into the combustor, wherein the fuel injection device is configured to inhibit cavitation of the fuel. Also disclosed is a combustor (15) for a gas turbine engine (10), a reheat assembly (300) for a gas turbine engine (10), a gas turbine engine (10), and an aircraft (200).

An aerofoil component for a high-pressure or intermediate-pressure turbine (16, 17) of a multi-spool gas turbine engine (10). The component having an aerofoil member (31, 41) which, in use, spans a working gas annulus of the gas turbine engine. The aerofoil member has pressure side and suction side aerofoil surfaces which each extend from a leading edge to a trailing edge of the aerofoil member such that transverse cross sections through the aerofoil member provide respective aerofoil sections. The spacing between the leading and trailing edges on the midspan aerofoil section (MS) defines a midspan axial chord length (C<x>) of the aerofoil member. The aerofoil member are arranged, in use, in a circumferential row around the annulus with plural, equally spaced, identical aerofoil members whereby the circumferential spacing of the aerofoil members at the trailing edges of their midspan aerofoil sections (MS) defines a midspan pitch (S) of the aerofoil member. At midspan: the value of S/C<x> is in the range from 1.4 to 1.6; and the uncovered turning angle is 16° or more.

An apparatus (200) for in-situ application of an engineering coating (52) to one or more components (50). The apparatus has a head section (300) adapted to selectively apply the engineering coating (52) to the components (50). The head section (300) has an applicator end (302) and a coupling end (304) opposite to the applicator end (302). The apparatus has a body section (400) coupled to the head section (300) at the coupling end (304) and adapted to actuate the head section (300). The body section (400) has a first stage actuator (402) adapted to actuate the head section (300) to swivel about a first axis (X-X') and a second stage actuator (404) adapted to actuate the head section (300) and the first stage actuator (402) together to rotate about a second axis (Y-Y'). The second axis (Y-Y') is orthogonal to the first axis (X-X'). The apparatus has one or more gas channels (500) provided along the body section (400) to the head section (300).

The foregoing describes a reheat assembly for a gas turbine engine. The reheat assembly comprises a support duct section, a vaporiser and a heat shield. The support duct section comprises a core side and a bypass side. During operation of the engine, the core side faces a core flow and the bypass side that faces a bypass flow. The vaporiser is attached to the bypass side of the support duct and comprises a plenum configured to cool core flow introduced into the plenum through an inlet to the vaporiser. The heat shield is configured to face the core side of the support duct section to shield the plenum from heat emitted by the core flow. The heat shield is spaced from core side of the support duct section to form a channel between the support duct section and the heat shield that extends upstream of the inlet.

An apparatus (200) for in-situ application of an engineering coating (52) to components (50) has a head section (300) for applying the engineering coating (52) and a multi-part body section (400) for controlling the application process. The body section (400) has a first body member (402), a second body member (404), and a third body member (406) interconnected with each other. The first body member (402) includes a predetermined pathway (410) that guides movement and is connected to the head section (300). The second body member (404) is engaged with the first body member (402) and includes at least one hollow channel (412). The third body member (406) houses an actuator (420), an engagement pin (414) for connecting with the pathway, and at least one coupling member (418) for interacting with the at least one hollow channel (412). The actuator (420) ensures precise movement of the head section (300). The apparatus (200) further includes one or more gas channels (500) to provide necessary gases to the head section (300).

The foregoing describes a reheat assembly (300) for a gas turbine engine (100). The reheat assembly (300) comprises: a support duct section (321') comprising a core side (322') for facing a core reheat region (232), and a bypass side for facing a bypass reheat region (231); a support duct (21') comprising a plurality of circumferentially spaced inlets (380), each inlet (380) being configured to communicate the bypass reheat region (231) with the core reheat region (232); a fuel supply system (360) comprising a plurality of bypass fuel injection ports (363), each bypass fuel injection port (363) being associated with an inlet (380) of the plurality of inlets (380); wherein each bypass fuel injection port (363) is configured to discharge fuel into the bypass reheat region (231) for transit through the associated inlet into the core reheat region (232).

There is provided a fluid flow machine (10) comprising a casing structure (24) extending around an axial direction (41) of the fluid flow machine (10) and a turbomachine blade (311-315) disposed within the casing structure (24) The casing structure (24) includes an axially extending slot (331-335) having an axial extent (32) at least partially overlapping with an axial extent (31) of a tip (321-325) of the turbomachine blade (311-315), wherein the slot (331-335) has an angular extent (29) of no more than 36 degrees. The casing structure (24) includes a circumferentially extending groove (341, 342) having an angular extent (28) of at least 72 degrees. The groove (341, 342) is axially offset from the slot (331-335).

A method (200) for scanning a component (102) includes providing a filter (104) and the component, the filter at least partially surrounding the component. The filter comprises a multi-phase material (140) that has an attenuation coefficient (C1) that is lower than an attenuation coefficient (C2) of a material of the component (102) to be scanned. The method further includes disposing the component and the filter on a support platform (120), and providing an imaging beam source (108) and an imaging beam receiver (114). The support platform (120) is configured to rotate and/or revolve relative to the imaging beam source (108) and the imaging beam receiver (114) about one or more axes. The method includes generating, via the imaging beam source (108), an imaging beam (110) that passes through the component and the filter. The method includes attenuating, via the, a scatter beam (111) that is produced upon irradiation of the component (102) with the imaging beam (110).

A method (100) of performing spectral analysis on a sample (220) identified within a gas turbine engine (10). The method (100) involves: obtaining (102) an optical spectrum (240) of the sample (220); normalising (104) the optical spectrum (240) to obtain a normalised optical spectrum (301); selecting (106) a characteristic parameter that characterises the normalised optical spectrum (301); determining (108) a characteristic parameter value associated with the normalised optical spectrum (301) based on the characteristic parameter; comparing (110) the characteristic parameter value with one or more predetermined reference values corresponding to the characteristic parameter; and determining (112) a presence of at least one compound in the sample (220) based on the comparison between the characteristic parameter value and the one or more predetermined reference values. The method can determine whether there is a need to disassemble the gas turbine engine for maintenance.

An apparatus (100) for measuring a clearance (C) between a first component (102) and one or more areas of interest (104) of a second component (106). The apparatus (100) includes an imaging beam source (108) configured to generate an imaging beam (110) that passes between the first component (102) and the second component (106). The apparatus (100) includes a first tube (112) configured to guide the imaging beam (110) from the imaging beam source (108) to the one or more areas of interest (104) of the second component (106). The apparatus (100) includes an imaging beam receiver (114) configured to receive the imaging beam (110). The imaging beam receiver (114) is configured to generate an image (116) in response to receiving the imaging beam (110). The image (116) depicts the clearance (C) between the first component (102) and the one or more areas of interest (104) of the second component (106).

A method of inspecting a component of a gas turbine engine for an aircraft, in which a housing circumferentially surrounds the component so as to define an annular space between the housing and the component. The method comprises guiding a guide line into the annular space through an opening to the annular space and around the component in a circumferential direction, and withdrawing the guide line through the opening, such withdrawal moving a probe that is connected to the guide line in a circumferential direction around the component for performing an inspection of the compressor.

A device (100) for wear testing includes a frame (106), a rotary actuator (108), and a linear actuator (110) spaced apart from the rotary actuator (108) along a longitudinal axis (102) of the frame (106). The device (100) has first and second specimen holding units (112, 140). The first specimen holding unit (112) includes a seat (124) connected to the linear actuator (110), a plate (114) engaging with the seat (124), and a first thrust bearing (136) connected to the plate (114) and at least one first diaphragm plate (134) connected to the frame (106). The second specimen holding unit (140) includes a shaft (142) connected to the rotary actuator (108) and a second thrust bearing (150) connected the frame (106). The rotary actuator (108) is configured to rotate a second specimen (60) about the longitudinal axis (102). The linear actuator (110) is configured to move a first specimen (50) along the longitudinal axis (102), such that the first specimen (50) engages with the second specimen (60).

A power event manager controller for a gas turbine engine power system, the gas turbine engine power system comprising at least one gas turbine engine, a thermal management system and at least one of a generator, an energy storage system, the power event manager controller receiving an input power demand, and receiving input relating to the mission plan, and inputs from each of the systems within the gas turbine engine power system, based upon the inputs from the from the power demand and the mission plan the power event controller defines a series of constraints, and wherein the power event manager controller utilises an optimiser function to obtain control reference trajectories which minimise an objective cost function of modelled states, whilst being subject to the series of constraints.

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 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.

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 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 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 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).