Abstract: A method of operating a gas turbine engine. The method includes determining a first compressor non-dimensional rotational speed set point, determining a current first compressor rotational speed and first compressor inlet temperature, and transferring power between the first and second shafts such that the first shaft non-dimensional rotational speed matches the set point.

Abstract: A gas turbine engine. The engine includes a first compressor coupled to a first turbine by a first shaft, the first turbine having first and second turbine stages. A first combustor is provided downstream of the first compressor and upstream of the first stage of the first turbine. A second combustor is provided downstream of the first stage of the first turbine, and upstream of the second stage of the first turbine. A further turbine is provided downstream of the first turbine, and is coupled to a further compressor by a further shaft.

Abstract: An aircraft gas turbine engine comprises a high pressure compressor driven by a high pressure turbine via a high pressure shaft, a first combustor provided downstream of the high pressure compressor and upstream of the high pressure turbine, a low pressure compressor driven by a low pressure turbine via a low pressure shaft, the low pressure compressor being configured to provide air to the high pressure compressor and to a bypass flow. The low pressure turbine comprises at least first and second turbine stages. The engine further comprises a second combustor provided downstream of the first stage of the low pressure turbine and upstream of the second stage of the low pressure turbine. The engine comprises a shaft coupling arrangement configured to transfer power between the high and low pressure shafts.

Abstract: A gas turbine engine comprises a relatively high pressure compressor coupled to a relatively high pressure turbine by a relatively high pressure shaft; a relatively low pressure compressor coupled to a relatively low pressure turbine by a relatively low pressure shaft rotatable independently of the high pressure shaft; a first combustor located downstream of the high pressure compressor and upstream of the high pressure turbine; and a second combustor located downstream of the high pressure turbine, and upstream of the low pressure turbine. The engine further comprises a coupling arrangement configured to selectively transfer torque between the high pressure shaft and the low pressure shaft.

Abstract: A gas turbine auxiliary power unit system for an aircraft comprises an aircraft cabin provided with a pressurised air supply, a gas turbine engine comprising a compressor assembly, and a controller. The pressurised air supply is provided to the aircraft cabin through an aircraft cabin air inlet, and exhausted from the aircraft cabin through an aircraft cabin air exhaust, while the gas turbine engine is provided with an inlet air flow from the aircraft cabin air exhaust. The controller is operable to control a flow rate of a fuel supplied to the gas turbine engine responsive to a power demand on the auxiliary power unit, and also to restrict an inlet air flow into the gas turbine engine, so as to maintain an air pressure inside the aircraft cabin within a pre-determined range.

Abstract: A combustor for a gas turbine engine which includes a flame tube made of a ceramic material and a metal casing which surrounds the flame tube. The combustor further includes a thermal insulation layer between the flame tube and the casing which protects the casing from the high temperatures produced by combustion in the combustor and a loading structure which holds the flame tube in a state of axial and radial compression at all engine operating conditions.

Abstract: A gas turbine engine. The engine includes a first compressor coupled to a first turbine by a first shaft, the first turbine having first and second turbine stages. A first combustor is provided downstream of the first compressor and upstream of the first stage of the first turbine. A second combustor is provided downstream of the first stage of the first turbine, and upstream of the second stage of the first turbine. A further turbine is provided downstream of the first turbine, and is coupled to a further compressor by a further shaft.