A propulsion system for an aircraft. The propulsion system comprises a core, an internal fixed structure secured to the core and arranged around the core, and a nacelle surrounding the core and the internal fixed structure in which a secondary flow path is delimited between the internal fixed structure and the nacelle. The internal fixed structure has a slot which permits fluidic communication between a compartment on the inside of the internal fixed structure and the secondary flow path.

An apparatus and method for cooling a fluid within a turbine engine. A fan casing assembly for the turbine engine can include an annular fan casing with a peripheral wall having a flow path defined through the casing. A fan casing cooler includes a body to confront the peripheral wall with at least one conduit configured to carry a flow of heated fluid to convectively cool the heated fluid with a flow of air through the flow path.

An apparatus and method for cooling a fluid within a turbine engine. A fan casing assembly for the turbine engine can include an annular fan casing with a peripheral wall having a flow path defined through the casing. A fan casing cooler includes a body to confront the peripheral wall with at least one conduit configured to carry a flow of heated fluid to convectively cool the heated fluid with a flow of air through the flow path.

An assembly is provided for a turbine engine. This assembly includes a primary duct, a bleed duct, a plurality of secondary ducts, a heat exchanger and a flow regulator. The bleed duct extends from a bleed duct inlet to a bleed duct outlet. The bleed duct inlet is fluidly coupled with the primary duct. The secondary ducts are arranged in parallel between the bleed duct outlet and the primary duct. The secondary ducts include a first duct and a second duct. The heat exchanger is configured with the second duct. The flow regulator is configured to direct at least a majority of fluid flowing through the bleed duct outlet to: (A) the first duct during a first mode of operation; and (B) the second duct during a second mode of operation.

A mount assembly for attaching a heat exchanger to a casing of a gas turbine engine comprises a first attachment feature by which in use the mount assembly is attached to the casing and a second attachment feature spaced from the first attachment feature by which in use the mount assembly is attached to the heat exchanger. The first and second attachment features are joined by an elongate member. The assembly is characterised in that the elongate member is significantly larger in a length direction and a height direction than in a thickness direction, so that it is relatively stiff in the length and height directions L, H and relatively flexible in the thickness direction T; the flexibility allowing movement in use within the mount assembly to accommodate differential thermal expansion between the heat exchanger and the casing.

A mount assembly for attaching a heat exchanger to a casing of a gas turbine engine comprises a first attachment feature by which in use the mount assembly is attached to the casing and a second attachment feature spaced from the first attachment feature by which in use the mount assembly is attached to the heat exchanger. The first and second attachment features are joined by an elongate member. The assembly is characterised in that the elongate member is significantly larger in a length direction and a height direction than in a thickness direction, so that it is relatively stiff in the length and height directions L, H and relatively flexible in the thickness direction T; the flexibility allowing movement in use within the mount assembly to accommodate differential thermal expansion between the heat exchanger and the casing.

The aircraft engine can have a core gas path having a first combustor, a second gas path parallel to the core gas path, the second gas path having a second combustor, a turbine driven by the second gas path, a gearbox driven by the turbine, and a valve configured for selectively opening and closing the second gas path.

The aircraft engine can have a core gas path extending from an intake across a core compressor, and then manifolding to a plurality turbine intake paths, each turbine intake path leading to a respective turbine unit via a respective combustor unit, and gearing drivingly connecting the collective rotary power of the turbine units to at least one power output shaft. During operation, the different core-turbine units can be operated simultaneously and efficiently, or one or more of the core-turbine units can be selectively shut down while the other core-turbine units continue to operate efficiently to lower the power output.

The gas turbine engine can have a core gas path extending sequentially across a core compressor, a core combustor, and a core turbine, an auxiliary air intake path and a bypass intake path leading in parallel to the core compressor, an auxiliary compressor in the auxiliary air intake path, an auxiliary gas path downstream of the core compressor, the auxiliary gas path extending in sequence across an auxiliary combustor and an auxiliary turbine, in parallel with the core combustor and core turbine, and valves operable to control the flow through the bypass gas path and the auxiliary gas path. Accordingly, the auxiliary components can be operated to increase power output, or deactivated while allowing the core components to run efficiently while meeting a lower power output.

The aircraft engine can have a core gas path extending sequentially across a core compressor, a core combustor, and a core turbine; a boost gas path extending from an intake to the core compressor, across a boost compressor, a bypass gas path extending from the intake to the core compressor, and a bypass valve operable to selectively open and close the bypass gas path. The intake flow can be directed either across the boost gas path for increased power output, or be directed to bypass the boost gas path via the bypass gas path.