Gas Turbine Engine Internal Air Systems Peter R N Childs 4 th August 2006 Introduction It has been … It is these flows that make the subject area interesting and an on-going challenge … Gas turbine engine technology, which is responsible as the prime mover for the vast proportion of this figure, is therefore of critical importance. The quest for fuel efficient engines, low emissions, low noise levels has put ever increasing demands on engine design. The challenge has been intensified with growing commercial pressure to meet these …

The considerable improvements in thermal efficiency and specific fuel consumption, achieved since the first successful engine test in 1937, a design based on Frank Whittle’s patent (1930), are largely due to increases in the overall pressure ratio of the thermodynamic cycle and have consequently led to higher turbine inlet temperatures. Since the early 1940s the turbine inlet temperature has risen significantly from approximately 800 oC in Whittle’s and von Ohain’s designs, to over 1600 oC in some of the current engines developed by Rolls-Royce plc (Rolls-Royce (2005)). In many high performance applications the turbine inlet temperature has now risen to such an extent that it limits the length of time that nickel alloy turbine blades can operate. The effective cooling of the turbine components, in particular blades, nozzle guide vanes and discs, is paramount. The problem is compounded by the increase in compressor outlet temperature to levels of approximately 700 oC, synonymous with the higher pressure ratios, since this supply provides the source of the cooling airflow conveyed to the turbine via the internal air system. The gas turbine engine internal air system provides cooling to various critical components, sealing for bearing chambers and flow paths and controls bearing axial loads. In order to supply the internal air system flow requirement, up to 20% of the engine core flow is extracted from the compressor. This can consume up to 5% of the fuel and it is therefore important to minimise the quantity of air required for the internal air system whilst maintaining functionality of the engine, acceptable component life, robustness and acceptable manufacture costs. For a large passenger aircraft a 1% reduction in specific fuel consumption could save 560 tonnes of fuel per annum and reduce direct operating costs by 0.5% (Smout et al. (2004)). It is therefore extremely important to be able to reduce the uncertainties associated with optimising the system. A typical internal air system will include a compressor bleed take off, transfer tubes and passages to deliver the cooling and sealing air to critical components and use of differential pressure across disc surfaces to balance bearing loads. A gas turbine engine generally comprises a series of discs for the rotating blades and a stationary casing and support structure. A common feature is the cavity formed between coaxial rotating and stationary discs, which is known as a wheelspace or a rotor- stator cavity. In addition cavities are also formed between co-rotating discs (see Long and Childs (2006)). Components requiring cooling include the combustor, which is normally cooled by the main gas path, the high pressure turbine, parts of which can reach temperatures over 1600 oC (Rolls-Royce (2005)), and shafts. Components that require sealing include turbine rotor-stator cavities, bearing chambers, transmission systems and the transfer elements delivering system air to the targeted components. Types of seal used include interstitial seals such as labyrinth, bush and rim seals, carbon seals, leaf and brush seals, hydraulic seals and static seals such as strip, cloth (Farahani and Childs (2006)) and sigma seals.

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