Abstract: A direct injection fuel system is shown for injecting hydrogen fuel into a gas turbine combustor. The fuel injection system includes a plurality of fuel injector blocks. Each fuel injector block includes a fuel admission duct having an inlet for receiving hydrogen fuel from a fuel supply, an outlet for delivering hydrogen fuel into the combustor and a central axis extending from said inlet to said outlet. Each fuel injector block also includes an air admission duct located around the periphery of the fuel admission duct, having an inlet for receiving air from a diffuser and an outlet for delivering air into the combustor for mixing with the hydrogen fuel.

Abstract: A direct fuel injection system (206) is shown for injecting hydrogen fuel into a gas turbine combustor, the fuel injection system comprising a plurality of fuel injector blocks (1202). Each fuel injector block includes one or more air admission ducts (1603) for receiving air from a diffuser and an air outlet for delivering air into a mixing zone for combustion with fuel. Each fuel injector block also includes a fuel admission duct or aperture having a fuel inlet for receiving fuel (F) from a manifold, and a fuel outlet for delivering fuel into the mixing zone.

Abstract: A direct injection fuel system is shown for injecting hydrogen fuel into a gas turbine combustor. The fuel injection system includes a plurality of fuel injector blocks. Each fuel injector block includes a fuel admission duct having an inlet for receiving hydrogen fuel from a fuel supply, an outlet for delivering hydrogen fuel into the combustor and a central axis extending from said inlet to said outlet. Each fuel injector block also includes an air admission duct located around the periphery of the fuel admission duct, having an inlet for receiving air from a diffuser and an outlet for delivering air into the combustor for mixing with the hydrogen fuel.

Abstract: An annular array of turning vanes 200 is provided in a duct 100 of a gas turbine engine 10. The annular array of turning vanes 200 comprises aerodynamic vanes 220 and strut-vanes 240. The strut-vanes 240 have greater chord length and extend further axially downstream than the aerodynamic vanes 220. The leading edge of the strut-vanes 240 is upstream of the trailing edge of the aerodynamic vanes 220. The strut-vanes provide flow turning. The arrangement allows the duct 100 to be axially short.

Abstract: An annular array of turning vanes 200 is provided in a duct 100 of a gas turbine engine 10. The annular array of turning vanes 200 comprises aerodynamic vanes 220 and strut-vanes 240. The strut-vanes 240 have greater chord length and extend further axially downstream than the aerodynamic vanes 220. The leading edge of the strut-vanes 240 is upstream of the trailing edge of the aerodynamic vanes 220. The strut-vanes provide flow turning. The space to chord ratio of the aerodynamic vanes that are closest to the suction surface of a strut-vane is higher than the space to chord ratio of aerodynamic vanes that are closest to a pressure surface of the strut-vane. The arrangement allows the duct 100 to be axially short.

Abstract: A fuel injection system for a gas turbine engine comprises; a pilot fuel injector section and a main airblast fuel injector section, the main airblast fuel injector section having an aft end facing a combustion chamber. A surface of the injection system exposed to air flow through an injection system is non-axisymmetric or non-planar in a reference circumferential plane and/or is configured to generate controlled and varying acoustic impedance at or adjacent the aft end where, in use, the air flow collides with an oncoming acoustic wave.

Abstract: A fuel injector includes a central member arranged on axis of fuel injector and a first member arranged around central member. A first swirler is arranged between central member and first member. A pilot fuel injector is arranged to supply fuel onto inner surface of first member. A second member is arranged between downstream end of first member and downstream end of shroud. A second swirler is arranged between upstream end of first member and upstream end of shroud and between downstream end of first member and upstream end of second member. A third swirler is arranged between downstream end of shroud and second member. A main fuel injector is arranged to supply fuel into an annular passage between first member and second member and a fourth swirler extends through first member.

Abstract: An annular array of turning vanes 200 is provided in a duct 100 of a gas turbine engine 10. The annular array of turning vanes 200 comprises aerodynamic vanes 220 and strut-vanes 240. The strut-vanes 240 have greater chord length and extend further axially downstream than the aerodynamic vanes 220. The leading edge of the strut-vanes 240 is upstream of the trailing edge of the aerodynamic vanes 220. The strut-vanes provide flow turning. The space to chord ratio of the aerodynamic vanes that are closest to the suction surface of a strut-vane is higher than the space to chord ratio of aerodynamic vanes that are closest to a pressure surface of the strut-vane. The arrangement allows the duct 100 to be axially short.

Abstract: An annular array of turning vanes 200 is provided in a duct 100 of a gas turbine engine 10. The annular array of turning vanes 200 comprises aerodynamic vanes 220 and strut-vanes 240. The strut-vanes 240 have greater chord length and extend further axially downstream than the aerodynamic vanes 220. The leading edge of the strut-vanes 240 is upstream of the trailing edge of the aerodynamic vanes 220. The strut-vanes provide flow turning. The arrangement allows the duct 100 to be axially short.

Abstract: A fuel injector includes a central member arranged on axis of fuel injector and a first member arranged around central member. A first swirler is arranged between central member and first member. A pilot fuel injector is arranged to supply fuel onto inner surface of first member. A second member is arranged between downstream end of first member and downstream end of shroud. A second swirler is arranged between upstream end of first member and upstream end of shroud and between downstream end of first member and upstream end of second member. A third swirler is arranged between downstream end of shroud and second member. A main fuel injector is arranged to supply fuel into an annular passage between first member and second member and a fourth swirler extends through first member.

Abstract: A fuel injection system for a gas turbine engine (10) comprises; a pilot fuel injector section (22, 23) and a main airblast fuel injector section (25, 26, 27, 28, 29), the main airblast fuel injector section having an aft end (29) facing a combustion chamber (30). A surface of the injection system exposed to air flow through an injection system is non-axisymmetric or non-planar in a reference circumferential plane and/or is configured to generate controlled and varying acoustic impedance at or adjacent the aft end where, in use, the air flow collides with an oncoming acoustic wave.

Abstract: A combustion chamber head of a gas turbine has a substantially annular combustion chamber outer wall 18 as well as a substantially annular combustion chamber inner wall 42 and several burners 6 distributed around the circumference. The combustion chamber head 5 has an inflow-side wall 13 which together with a wall 14 facing the combustion chamber 7 forms a combustion chamber head volume 15. The inflow-side wall 13 is provided with at least one inflow opening 32, the wall 14 facing the combustion chamber 7 is provided with at least one outflow opening 17 for connecting the combustion chamber head volume 15 to the combustion chamber 7, and at least one cooling air duct 29 is provided in the wall 14 facing the combustion chamber 7. A method for cooling and damping of the combustion chamber head is also disclosed.

Abstract: A combustion chamber for a gas turbine engine comprising at least one Helmholtz resonator having a resonator cavity and a resonator neck in flow communication with the interior of the combustion chamber. The cavity extends around at least part of the neck and is spaced apart therefrom to define a cooling chamber therebetween. The cooling chamber allows the neck and the cavity to be cooled with a flow of cooling air. The sunken neck allows the resonator to be located within enclosures with minimal volume without compromising on the damping properties afforded.

Abstract: A combustion chamber suitable for a gas turbine engine is provided with at least one Helmholtz resonator having a resonator cavity and a resonator neck in flow communication with the chamber interior. The resonator neck is provided with at least cooling holes extending through its wall for improved damping and cooling, at least one of the holes is directed towards the resonator cavity.

Abstract: A gas turbine engine combustor, eg of annular combustor configuration, has air inlets, typically in the dilution zone, which incorporate vortex generators operative to trip the inlet jet flow causing stabilizing vortex flow at the lateral edges of the jet. The vortex generators can comprise an inwardly directed abrupt change of section between forward and rearward parts of the air inlets. The rearward section of the air inlets can have a configuration operative to reinforce the jet stabilizing effect of the vortex flow. The combustor exhibits improved uniformity of jet and crossflow mixing.