FIG. 1 shows a gas turbine engine 10. Gas turbine engines 10 are mounted on an aircraft 100 in pairs, as shown in FIG. 2. The engine 10 comprises, in axial flow series, an air intake duct 11, an intake fan 12, a bypass duct 13, an intermediate pressure compressor 14, a high pressure compressor 16, a combustor 18, a high pressure turbine 20, an intermediate pressure turbine 22, a low pressure turbine 24 and an exhaust nozzle 25. The fan 12, compressors 14, 16 and turbines 20, 22, 24 all rotate about the major axis of the gas turbine engine 10 and so define the axial direction of gas turbine engine.
Air is drawn through the air intake duct 11 by the intake fan 12 where it is accelerated. A significant portion of the airflow is discharged through the bypass duct 13 generating a corresponding portion of the engine 10 thrust. The remainder is drawn through the intermediate pressure compressor 14 into what is termed the core of the engine 10 where the air is compressed. A further stage of compression takes place in the high pressure compressor 16 before the air is mixed with fuel and burned in the combustor 18. The resulting hot working fluid is discharged through the high pressure turbine 20, the intermediate pressure turbine 22 and the low pressure turbine 24 in series, where work is extracted from the working fluid. The work extracted drives the intake fan 12, the intermediate pressure compressor 14 and the high pressure compressor 16 via shafts 26, 28, 30. The working fluid, which has reduced in pressure and temperature, is then expelled through the exhaust nozzle 25 and generates the remaining portion of the engine 10 thrust.
FIG. 2 shows an aircraft 100 by powered by a pair of gas turbine engines 10. The aircraft 100 comprises an aircraft pneumatic system comprising an environmental control system (ECS) powered by high pressure air provided by a bleed air system (BAS). Bleed air systems generally comprise cabin bleed offtakes 32, 34 which duct air from the compressor 14, 16 for use in the aircraft pneumatic system, such as the (ECS) and wing de-icing. The ECS provides cabin air to the cabin interior at a required temperature, pressure and flow rate. The BAS system comprises a fan air heat exchanger 36 which exchanges heat between relatively cool fan air provided by a fan air duct 38, and the relatively hot cabin bleed air provided from the ports 32, 34. Once cooled by the heat exchanger 36, the cabin bleed air is then passed to an air cycle machine (not shown), where the cabin bleed air is processed to obtain the desired heat, pressure and flow rate, before being passed to the cabin of the aircraft 100.
In the prior example shown in FIG. 3, low pressure and high pressure cabin bleed offtakes 32, 34 are provided. It is generally desirable to extract bleed air from the low pressure bleed offtake 32 (i.e. one near the front of the engine), since air taken from the low pressure bleed offtake 32 has been compressed to a lesser extent compared to air taken from the high pressure bleed offtake 34. Consequently, a given mass of air bled from the low pressure bleed offtake 32 represents a smaller energy loss to the thermodynamic cycle of the engine 10 compared to the same mass of air taken from the high pressure bleed offtake 34, and so the specific fuel consumption (SFC) of the engine 10 will be greater (i.e. more fuel will be burned for a given thrust) where air is bled from the high pressure bleed offtake 34.
The pneumatic system of the aircraft 100 further comprises one or more engine core compressor handling bleed offtakes 40. The handling bleed offtake 40 is operated by an engine controller (FADEC 42), which opens and closes the handling bleed offtakes 40 on a predetermined schedule to ensure that the respective core compressor 14, 16 does not stall or surge during operation. An outlet of the offtake 40 communicates with a low pressure area, such as the bypass duct 13, such that flow from the bleed offtake 40 can be dumped overboard. In general, a core compressor handling bleed offtake is provided at a high pressure compression stage of each core compressor 14, 16.
However, current systems are non-optimal, in the sense that relatively high pressure cabin bleed air must be used to provide sufficient pressure and flow to the ECS at some engine conditions, such as when the engine is at low power. Consequently, at least one of the cabin bleed offtakes 32, 34 must be located at a relatively high pressure stage of the engine compressor 14, 16, and must be utilised extensively during engine operations, particularly when the engine is operated at low thrust, and therefore low engine overall pressure. Such arrangements are relatively complex, and may result in excessive thrust specific fuel consumption (SFC), since the cabin bleed air must be provided at relatively high pressure in such prior arrangements. In some cases, the minimum thrust that can be achieved by the engine in flight (known as “flight idle thrust”) is limited by the overall pressure ratio required to operate the ECS system. Consequently, in such circumstances, the engines 10 are operated at a higher thrust than would be required for maintaining the desired flight profile and engine operability alone, thereby again resulting in increased SFC. A further disadvantage of current systems, is that the high pressure handling bleed air is simply vented to the bypass duct, and therefore largely lost to the thermodynamic cycle of the engine, save for a small amount of thrust generated by the bleed air.
In an unrelated problem, aircraft gas turbine engines are a major source of aircraft noise. There is a continued effort to reduce aircraft noise, particularly during the take-off, approach and landing phases of flight. A major source of engine noise during low power operation of a gas turbine engine (i.e. during approach and landing) is the efflux from the handling bleed valve. The high levels of noise are thought to be caused by the high pressure, high temperature air turbulently mixing in the engine bypass flow.
The present invention describes an aircraft pneumatic system and a method of operating an aircraft pneumatic system which seeks to overcome some or all of the above problems.
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