Search results

Based on a lecture prepared as part of the celebration of Cranfield University's 50th anniversary. After briefly reviewing the early years, including Cranfield University's entry into this technology, discusses the nature of this industry, Some of the technology drivers, including environmental concerns, are examined to provide a background against which the development and the future of the industry can be considered. This is followed by a brief survey of some of the possible new civil aero gas turbine applications over the next 50 years, both the very likely and some curiosities. Finally, the changes that are likely to occur within the industry as a result of wider economic and political trends are considered, as well as the implications for those working within the industry. The development of the civil aero gas turbine has contributed, in large measure, to today's, US$ 300 billion civil aviation industry and is rightly seen as one of mankind's major engineering achievements. A single paper cannot do justice to this industry.

WHILE the technical part of the history of the aircraft gas turbine in Great Britain presents the features of success and failure familiar in technical progress, there is another part of the history which I believe can be described as an unqualified success. I refer to the habit of collaboration which was developed between the several technical teams in my own country, between Great Britain and the United States, and, later, between Great Britain and the British Dominions.

The first gas turbine patent was granted in England to John Barber in 1791, and since then there have been numerous gas turbine inventions. These have been adequately described elsewhere, 1, 2 and I shall concern myself only with the developments which have led directly to recent British achievements.

This paper seeks to evaluate the potential of heat exchanged aeroengines for future Unmanned Aerial Vehicle (UAV), helicopter, and aircraft propulsion, with emphasis placed on reduced emissions, lower fuel burn, and less noise.

Aeroengine performance analyses were carried out covering a wide range of parameters for more complex thermodynamic cycles. This led to the identification of major component features and the establishing of preconceptual aeroengine layout concepts for various types of recuperated and ICR variants.

Novel aeroengine architectures were identified for heat exchanged turboshaft, turboprop, and turbofan variants covering a wide range of applications. While conceptual in nature, the results of the analyses and design studies generally concluded that heat exchanged engines represent a viable solution to meet demanding defence and commercial aeropropulsion needs in the 2015‐2020 timeframe, but they would require extensive development.

As highlighted in Parts I and II, early development work was focused on the use of recuperation, but this is only practical with compressor pressure ratios up to about 10. For today's aeroengines with pressure ratios up to about 50, improvement in SFC can only be realised by incorporating intercooling and recuperation. The new aeroengine concepts presented are clearly in an embryonic stage, but these should enable gas turbine and heat exchanger specialists to advance the technology by conducting more in‐depth analytical and design studies to establish higher efficiency and “greener” gas turbine aviation propulsion systems.

It is recognised that meeting future environmental and economic requirements will have a profound effect on aeroengine design and operation, and near‐term efforts will be focused on improving conventional simple‐cycle engines. This paper has addressed the longer‐term potential of heat exchanged aeroengines and has discussed novel design concepts. A deployment strategy, aimed at gaining confidence with emphasis placed on assuring engine reliability, has been suggested, with the initial development and flight worthiness test of a small recuperated turboprop engine for UAVs, followed by a larger recuperated turboshaft engine for a military helicopter, and then advancement to a larger and far more complex ICR turbofan engine.

Interest is currently being expressed in heat exchanged propulsion gas turbines for a variety of aeroengine applications, and in support of this, the aim of this paper is to evaluate the relevance of experience gained from development testing of several recuperated aeroengines in the USA in the late 1960s.

Technology status, including engine design features, performance, and specific weight of recuperated propulsion gas turbines based on radial and axial turbomachinery, that were development tested in the power range of about 300 to 4,000 hp (224 to 2,984 kW) is discussed in Part I.

A successful flight worthiness test was undertaken in the USA of a helicopter powered solely by a recuperated turboshaft engine and this demonstrated a specific fuel consumption reduction of over 25 percent compared with the simple‐cycle engine. However; in an era of low‐fuel cost, and uncertainty about the long‐term structural integrity of the high‐temperature heat exchanger, further development work was not undertaken.

The gas turbines tested were based on conventional simple‐cycle engines with essentially “bolted‐on” recuperators. Optimum approaches to minimize engine overall weight were needed in which the recuperator was integrated with the engine structure from the onset of design, and these are discussed in Part II.

Based on engine hardware testing, a formidable technology base was established, which although dated, could provide insight into the technical issues likely to be associated with the introduction of future heat exchanged aeroengines. These are projected to have the potential for reduced fuel burn, less emissions, and lower noise, and recuperated and intercooled turboshaft, turboprop, and turbofan variants are discussed in Part III.

A combustor design is a particularly important and difficult task in the development of gas turbine engines. During studies for accurate and easy combustor design, reasonable design methodologies have been established and used in engine development. The purpose of this paper is to review the design methodology for combustor in development of advanced gas turbine engines. The advanced combustor development task can be successfully achieved in less time and at lower cost by adopting new and superior design methodologies.

The review considers the main technical problems (combustion, cooling, fuel injection and ignition technology) in the development of modern combustor design and deals with combustor design methods by dividing it into preliminary design, performance evaluation, optimization and experiment. The advanced combustion and cooling technologies mainly used in combustor design are mentioned in detail. In accordance with the modern combustor design method, the design mechanisms are considered and the methods used in every stage of the design are reviewed technically.

The improved performances and strict emission limits of gas turbine engines require the application of advanced technologies when designing combustors. The optimized design mechanism and reasonable performance evaluation methods are very important in reducing experiments and increasing the effectiveness of the design.

This paper provides a comprehensive review of the design methodology for the advanced gas turbine engine combustor.

This paper is an overview of aero gas turbine engine starting systems and discusses the system design considerations and integration with other aircraft systems.

Review of a range of recent publications on the subject, aiming to provide an introduction to modern aero gas turbine engine starting systems.

Provides basic information on starter types and their limitations, and why some starter types are more favoured in modern installations. The effects of altitude and temperature are discussed which may not be initially considered as variables affecting aero start systems.

This paper provides further information on the starting systems of modern aero gas turbines and the considerations associated with integration and efficiency.

To advance the design of heat exchanged gas turbine propulsion aeroengines utilising experience gained from early development testing, and based on technologies prevailing in the 1970‐2000 time frame.

With emphasis on recuperated helicopter turboshaft engines, particularly in the 1,000 hp (746 kW) class, detailed performance analyses, parametric trade‐off studies, and overall power plant layouts, based on state‐of‐the‐art turbomachinery component efficiencies and high‐temperature heat exchanger technologies, were undertaken for several engine configuration concepts.

Using optimised cycle parameters, and the selection of a light weight tubular heat exchanger concept, an attractive engine architecture was established in which the recuperator was fully integrated with the engine structure. This resulted in a reduced overall engine weight and lower specific fuel consumption, and represented a significant advancement in technology from the modified simple‐cycle engines tested in the late 1960s.

While heat exchanged engine technology advancements were projected, there were essentially two major factors that essentially negated the continued study and development of recuperated aeroengines, namely again as mentioned in Part I, the reduced fuel consumption was not regarded as an important economic factor in an era of low‐fuel cost, and more importantly in this time frame very significant simple‐cycle engine performance advancements were made with the use of significantly higher pressure ratios and increased turbine inlet temperatures. Simply stated, recuperated variants could not compete with such a rapidly moving target.

Establishing an engine design concept in which the recuperator was an integral part of the engine structure to minimise the overall power plant weight was regarded as a technical achievement. Such an approach, together with the emergence of lighter weight recuperators of assured structural integrity, would find acceptance around the year 2000 when there was renewed interest in the use of more efficient heat exchanged variants towards the future goal of establishing “greener” aeroengines, and this is discussed in Part III of this paper.

IT was generally believed during early servicing experience of gas turbines that nearly every defect could result in power losses, but the last eight years have provided operational experience which has helped in the design, and improvements have made the gas turbine a trouble‐free and comparatively reliable unit. There are still defects which may arise, but it can safely be stated that they have no bearing on the power output. The customary teething troubles of early days have been overcome and power units are now available having an overhaul life of 500 hours. This figure is being increased rapidly and in the near future, the engine life between overhauls may exceed 1,000 hours. The overhaul of an engine provides an engine with a complete life cycle and the engine goes out into service embodied with improved design details, thus making it more reliable than before. The overhaul procedure consists of stripping, washing, crack testing, inspection, building and testing the engine, and a brief comment on each has been made in these notes, supported by a typical overhaul shop layout. The gas turbine power plant operating costs consist of primary and overhead charges, but in our discussion the overhead charges pertaining to overhaul will not be considered. These notes are a generalization of a variety of gas turbines and do not relate to any particular design.

In a modern gas turbine‐powered aircraft, engine material cost is the major single contributor to direct maintenance and overhaul cost. Due to the large expenditures involved, even a small percentage improvement in this field will result in the achievement of large savings. By concentrating on the reduction of both scheduled and unscheduled removals, this cost can be minimized. An analysis of the experience gained with the application (o gas turbine engines, of traditional overhaul life development programmes shows that these are of no benefit in determining the limitations of the engine. Experience and control of the total time or take‐off cycles of components, rather than time between overhaul, is the most important factor in the operational development of gas turbine engines. Overhaul life development can be divided into two phases: the initial programme applied to new engines with little operational background, and the developed programme applied to engines with substantial operating experience. In the initial programme, the sampling of two engines every 100 hours from a declared plateau can provide the required information and produces the required stagger in total time of components. In the developed programme large steps in overhaul life can be declared without further sampling. Life development of critical engine components cannot be done in service and requires manufacturer's cyclic testing ahead of airline experience until fail‐safe designs are achieved. Early failure detection of gas turbine engines is a Held of great potential return which to date has been only barely explored. Both engine manufacturers and operators need to develop tools and techniques for effective early failure detection quickly, to gain the large benefits which this field offers. Examples of the application of some early failure detection designs and techniques in Trans‐Canada Air Lines (T.C.A.) are given. These are in the field of visual examination with increased access to the engine, X‐ray examination of flame tubes with the engine installed in the aircraft, more effective use of the oil system to give warning of failure, engine vibration monitoring and automatic flight performance recording.