Abstract: An aircraft including first and second like gas turbine engines having like combustion apparatus each of which includes a respective plurality of fuel injectors arranged in an annular array and having a first and second sets of fuel-emitting apertures arranged to emit fuel normally to the plane of the array and in a direction having a component in the plane of the array directed towards an adjacent fuel injector. The aircraft further includes a fuel system which for any given engine increases the proportion of the total fuel flow rate to that engine which is provided to the second set of apertures during lighting or re-lighting of that engine's combustion apparatus, or upon detection of aircraft manoeuvring likely to increase the risk of flame-out. The aircraft provides faster and more reliable lighting and re-lighting, and additional resistance to flame-out, especially for use of hydrogen fuel.

Abstract: A fuel system (200) for a hydrogen-fuelled gas turbine engine (201) comprises a main fuel conduit (217) configured to conduct hydrogen fuel from a hydrogen storage unit (204) to a core combustor (206) of the gas turbine engine (201), a pre-heater (218) comprising an auxiliary combustor (222) configured to combust hydrogen fuel diverted from the main fuel conduit (217) to heat hydrogen fuel in the main fuel conduit (217). A first exhaust passage (228) is configured to conduct combustion products from the auxiliary combustor (222) to the core combustor (206) of the gas turbine engine (201).

Abstract: An aircraft includes an engine system which includes a gas turbine engine and a fuel system. The engine has combustion apparatus having an annular array of alternating primary and secondary fuel-injectors. The fuel system includes a fuel store and a controller. The controller is arranged to receive one or more signals indicative of one or more of (i) starting of the gas turbine engine; (ii) flame-out of the combustion apparatus; and (ii) a manoeuvring of the aircraft associated with a risk of flame-out of the combustion apparatus or preparation for such manoeuvring; and in response thereto to control the fuel system to commence supply of fuel from the fuel store to the secondary fuel-injectors of the combustion apparatus. Compared to known aircraft, engine lighting and relighting are achieved for more rapidly and reliably and the aircraft is less susceptible to flame-out, especially where hydrogen fuel is used.

Abstract: A fuel system (300) for a gas turbine engine (203). The fuel system comprises a fuel offtake (302) configured to divert a portion of hydrogen fuel from a main fuel conduit (217), a burner (218) configured to burn the portion of hydrogen fuel diverted from the main fuel conduit (217) and at least first and second heat exchangers (306, 310). The first heat exchanger (306) is configured to transfer heat from exhaust gasses produced by the burner (218) to hydrogen fuel in the main fuel conduit (217) and the second heat exchanger (310) is provided upstream in hydrogen flow of the first heat exchanger (306) and is configured to transfer heat from a further heat exchange fluid to hydrogen fuel in the main fuel conduit (217). In an embodiment, the further heat exchange fluid is compressor bleed air bled from a core compressor of the gas turbine engine.

Abstract: A combustion chamber comprises an upstream end wall structure and inner and outer annular wall structures. The upstream end wall structure comprises an upstream wall and a plurality of circumferentially arranged heat shields secured to the upstream wall. The upstream wall has a plurality of circumferentially spaced fuel injector apertures. Each heat shield has radially outer and radially inner ends and a fuel injector aperture aligned with a corresponding fuel injector aperture in the upstream wall. The radially outer and inner ends of each heat shield have outer and inner rails spacing the heat shield from the upstream wall. The radially outer and inner ends of each heat shield have first and second pluralities of circumferentially spaced apertures extending there-through and through the associated outer and inner rails to direct coolant over the surface of the outer and inner annular wall structures to form respective films of coolant.

Abstract: A gas turbine engine combustion chamber arrangement includes an annular outer wall and annular inner wall including an upstream and a downstream row of tiles. The downstream end of each tile in the upstream row of tiles has a rail extending towards and sealing with the outer wall and a lip extending in a downstream direction towards the tiles in the downstream row of tiles. The outer wall has a row of apertures to direct coolant onto the lips of the tiles in the upstream row of tiles. The downstream end of each tile in the upstream row of tiles has a plurality of fences extending in a downstream direction from the rail and each fence extends from the outer surface of the lip towards the outer wall. The row of apertures in the annular outer wall is arranged to direct coolant onto the lip between two circumferentially adjacent fences.

Abstract: A gas turbine engine combustion chamber arrangement includes an annular outer wall and annular inner wall including an upstream and a downstream row of tiles. The downstream end of each tile in the upstream row of tiles has a rail extending towards and sealing with the outer wall and a lip extending in a downstream direction towards the tiles in the downstream row of tiles. The outer wall has a row of apertures to direct coolant onto the lips of the tiles in the upstream row of tiles. The downstream end of each tile in the upstream row of tiles has a plurality of fences extending in a downstream direction from the rail and each fence extends from the outer surface of the lip towards the outer wall. The row of apertures in the annular outer wall is arranged to direct coolant onto the lip between two circumferentially adjacent fences.