Combustor apparatus
An integral guide vane combustor assembly for a gas turbine engine comprises an annular array of radially extending guide vane combustors. The gas turbine engine comprises, in axial flow sequence, a compressor assembly, the integral guide vane combustor assembly, a turbine assembly, and an exhaust assembly. Each of the guide vane combustors comprises, in axial flow sequence, a guide vane portion, and a combustor body portion. The guide vane portion is formed integrally with the combustor body portion, and the guide vane portion is configured to direct a gas flow exiting the compressor assembly into the combustor body portion.
1 . An integral guide vane combustor assembly for a gas turbine engine, the gas turbine engine comprising, in axial flow sequence, a compressor assembly, the integral guide vane combustor assembly, a turbine assembly, and an exhaust assembly, the integral guide vane combustor assembly comprising an annular array of radially extending guide vane combustors;
wherein each guide vane combustor comprises, in axial flow sequence, a guide vane portion, and a combustor body portion where a combustion occurs, the guide vane portion being formed integrally with the combustor body portion, and the guide vane portion being configured to direct a gas flow exiting the compressor assembly into the combustor body portion;
wherein the guide vane portion is configured to direct the gas flow entering the guide vane portion to proceed at a first offset angle relative to a longitudinal axis of the gas turbine engine, and is further configured to direct the gas flow exiting the guide vane portion to proceed at a second offset angle relative to the longitudinal axis of the gas turbine engine,
the second offset angle is less than the first offset angle, and an absolute value of the second offset angle is not zero, and
the first offset angle is a positive value; and
wherein,
in the annular array of radially extending guide vane combustors, each radially extending guide vane combustor is aligned from a frontend of the guide vane portion to a backend of the combustor body portion along one central longitudinal axis, the front end of the guide vane portion receiving pressurized gas exiting the compressor assembly, and after combustion, combusted gas exiting the combustor body portion via the backend,
the one central longitudinal axis is tilted with respect to the longitudinal axis of the gas turbine engine so that the backend of the combustor body portion is disposed radially more outward than the frontend of the guide vane portion, and
the annular array of radially extending guide vane combustors comprises between 6 and 16 guide vane combustors.
2 . The integral guide vane combustor assembly as claimed in claim 1 , wherein the annular array of radially extending guide vane combustors are circumferentially equally spaced.
3 . The integral guide vane combustor assembly as claimed in claim 1 , wherein each guide vane combustor is formed by an additive layer manufacturing process.
4 . The integral guide vane combustor assembly as claimed in claim 1 , wherein the absolute value of the second offset angle is not zero and is not greater than five.
5 . A gas turbine engine comprising:
in axial flow sequence, a fan assembly, a compressor assembly, an integral guide vane combustor assembly, a turbine assembly, and an exhaust assembly,
wherein the integral guide vane combustor assembly comprises an annular array of radially extending guide vane combustors, wherein each guide vane combustor comprises, in axial flow sequence, a guide vane portion, and a combustor body portion where a combustion occurs, the guide vane portion being formed integrally with the combustor body portion, and the guide vane portion being configured to direct a gas flow exiting the compressor assembly into the combustor body portion,
wherein the guide vane portion is configured to direct the gas flow entering the guide vane portion to proceed at a first offset angle relative to a longitudinal axis of the gas turbine engine, and is further configured to direct the gas flow exiting the guide vane portion to proceed at a second offset angle relative to the longitudinal axis of the gas turbine engine,
the second offset angle is less than the first offset angle, and an absolute value of the second offset angle is not zero, and the first offset angle is a positive value, and
the integral guide vane combustor assembly further comprises a plurality of fuel swirl nozzles for the guide vane combustor; and
wherein,
in the annular array of radially extending guide vane combustors, each radially extending guide vane combustor is aligned from a frontend of the guide vane portion to a backend of combustor body portion along one central longitudinal axis, the front end of the guide vane portion receiving pressurized gas exiting the compressor assembly, and after combustion, the combusted gas exiting the combustor body portion via the backend,
the one central longitudinal axis is tilted with respect to the longitudinal axis of the gas turbine engine so that the backend of the combustor body portion is disposed radially more outward than the frontend of the guide vane portion, and
the annular array of radially extending guide vane combustors comprises between 6 and 16 guide vane combustors.
6 . The gas turbine engine as claimed in claim 5 , wherein the fan assembly comprises a plurality of fan blades extending radially from a hub, the plurality of fan blades defining a fan diameter (D FAN ), and wherein the fan diameter D FAN is within a range of 0.3 m to 2.0 m.
7 . The gas turbine engine as claimed in claim 6 , wherein the fan diameter D FAN is within the range of 0.4 m to 1.5 m.
8 . The gas turbine engine as claimed in claim 6 , wherein the fan diameter D FAN is within the range of 0.7 m to 1.0 m.
9 . The gas turbine engine as claimed in claim 5 , further comprising:
an outer casing, the outer casing enclosing the sequential arrangement of the fan assembly, the compressor assembly, and the turbine assembly,
an annular bypass duct being defined between the outer casing and the sequential arrangement of the compressor assembly, and the turbine assembly, a bypass ratio being defined as a ratio of a mass air flow rate through the bypass duct to a mass air flow rate through the sequential arrangement of modules, and wherein the bypass ratio is less than 4.0.
10 . The gas turbine engine as claimed in claim 5 , wherein the fan assembly has two or more fan stages, at least one of the fan stages comprising a plurality of fan blades defining a fan diameter D FAN .