The potential applications of composites

Abstract

This paper will aim to describe the potential applications of composites to aero-engine components by reviewing related published papers as well as examining historical developments and current developments in composite use. It contains the background history of composite material development and uses, aircraft engine requirements, trends in material use, manufacturing and costs of composites, related composite projects and potential applications of composites to aero-engines. When discussing the application of composites to aero-engines, composites are categorized into four types; carbon fibre composites, ceramic matrix composites, metal matrix composites and carbon-carbon composites. In addition to discussing the particular aero-engine component applications, design considerations, processing methods, limitations and current development are discussed to some extent.

1. Introduction

1.1 Composites

Current and future aircraft demand reduced weight and increased fuel efficiency as well as reduced emissions and noise levels. Aero engines also strive to become more powerful to ensure that they continue to be competitive. Increasing use of composites is being made in the aerospace and other industries due to the beneficial properties of low weight while still providing high strength and stiffness (Anon, 1994). Due to the vast types of composite available to the designer, there are few limitations for finding a composite solution when compared to traditional materials. On the other hand, composites continue to pose a challenge in terms of reducing the cost, standardising the composite material properties and increasing the reliability in manufacturing, maintenance and repair when compared to traditional material (Armstrong, 2003). As designers have become more comfortable with composites, they are being used as standard materials in aircraft structures. As reducing the weight of aircraft becomes a priority, the reduction of weight is being extended to engine components (Ohnabe, 1999). This review will categorise composites into four main types: carbon fibre composites (CFC), ceramic matrix composites (CMC), metal matrix composites (MMC) and carbon-carbon composites (C/Cs) and look at the developments and applications of these materials with particular focus in aircraft engines.

1.2 History of composite aircraft applications

Historically composites have been used extensively in aircraft structures. Carbon fibre composites (CFC) were introduced and developed in the 1970's by the Royal Aircraft Establishment (RAE) and the British Aircraft Corporation (BAC)(Sharples, 1990). The BAC with the funding of the Ministry of Defence (MoD) and technical expertise from RAE scientists developed the design and manufacturing of CFC for military aircraft gaining expertise through their Jaguar wing, Tornado taileron and Jaguar engine bay door projects. Figure 1 shows an exploded view of the Jaguar CFC wing. The conclusions of these projects highlighted the feasibility of reduced weight and increased strength for structures made from composites. However it also concluded that more knowledge of composites was needed before they could become a commercially viable solution.

As knowledge of these materials increased, carbon fibres were and are commonly used in many parts of the aircraft structure on both military and larger commercial aircraft. In the early 1980's the rudders, ailerons, elevators, flaps and body and screw jack farings undercarriage doors of Boeing's 757 were all made out of carbon fibre Nomex, aflame resistant meta-aramidmaterial developed in the early 1960s. Airbus also made all the above out of carbon fibre Nomex with the addition of fin boxes and tail plane skins out of solid laminate carbon fibre. Military aircraft also make use of composites with the Harrier being 25% of the weight being composite. The Typhoon (previously known as the Eurofighter), the Saab Gripen, Joint Strike Fighter and the B2 bomber all have a significant amounts of their structure made of composite (Armstrong, 2003). The ALAFS (advance lightweight aircraft fuselage structure) program investigates new design methodologies that include extensive use of carbon fibre composites for applications such a unitised composite dorsal assembly to reduce part count and operating and support costs (O&S) (see figure 2).

2. Aircraft engine requirements

It is a challenge to determine aero engines design requirements as each individual part is subjected to different influences, running conditions and requirements. For each part, the designer needs to consider the structural integrity to ensure that the design can withstand the required loads. Environmental durability is also a key design requirement for aero engines with large variations in temperature, for example commercial nacelle components can be subjected to routine operating temperatures vary from -54C to 232C (Luttgeharm, 1990). The highest operating temperatures occur for a short time during takeoff while the most severe engine conditions occur for hours during cruising. As a result, creep resistance is of major importance in the hot section components. In addition aero-engines are also required to give high performance, low cycle costs, high thrust to weight ratios, light weight as well as low levels of noise and emissions.

Environmental goals to be achieved by 2020 set by the ACARE (Advisory Council for Aeronautics Research in Europe) outlines that CO2 needs to be reduced by 50% through an increase in fuel efficiency, nitrogen oxide (NOx) emissions need to be reduced by 80%, external noise needs to be reduced by 50% and to reduce the environmental impact during design, manufacturing, maintenance and disposal/recycling (Finley, 2008). VITAL European research and development program has project partners such as Airbus, Avio SpA, GKN Aerospace Services, Roll Royce, Short Brothers, composite research institute SICOMP of Sweden and the National Aerospace Laboratory (NLR) of the Netherlands and the project aims to have a 6dB reduction in engine noise and a 7% reduction in emissions (Finley, 2008). In order to strive to meet these requirements designers need to look at both engine design and engine materials. Higher pressure ratios up to 60 are needed to be achieved to obtain better thermodynamic efficiency in turning fuel energy into mechanical energy; this however provides problems in terms of increased peak compressor delivery temperature. In order to reduce the noise and optimise propulsion efficiency bypass ratios also need to increase with 15:1 involving geared fan or equivalent concepts being considered (Anon, 1994). In addition designers are exploring curved or sweeping blades in the fan and compressor making the geometry more difficult but ensuring less energy is used by creating a smoother airflow. As the bypass ratios and pressure ratios increase there will be an increase in peak combustor exit temperature. In addition the thrust-to-weight ratio will increase to increase performance and reduce weight resulting in an increase in turbine inlet temperature (TIT)(Ohnabe et al., 1999). Figure 3 shows the trends in TIT in jet engines.

Fuel consumption and noise can be reduced by improving the engines traction in the air through producing bigger fans (Finley, 2008) however increasing the size also increases the weight. In addition engine Co2 emissions is directly related to engine weight. This leads designers to use composites as they offer the same strength as traditional materials but with the benefit of reduced weight and their increased temperature capabilities allow for hotter, more efficient cycle lengths as well as providing longer component life. The Aircraft Industry are doing well to meet these targets: the GEnx engine, developed for Boeing's new 787 Dreamliner and the Boeing 747-8 aircraft makes much use of composites and is claimed to deliver 15% better fuel consumption compared to the engines it will replace as well as producing emissions approximately 94% below 2008 regulatory limits.

3. Trends in material use

Trends in material use have been determined by the need for the aircraft industry to meet environmental targets as well as remaining competitive in terms of performance. The main engine requirement targets include weight reduction and increased engine performance resulting in increased temperatures. This leads to trends in the penetration of advanced composite materials as outlined in fig 4.

Examining compressor materials, the increasing operating temperatures will need to be considered. Titanium alloys as they increase in temperature oxidise on the surface and also have oxygen solubility on the outside causing them to exhibit brittle behaviour. Therefore the operating environment creates a natural temperature limit on these alloys (Anon, 1994). This leads to an increased usage of resin-based composites and metal matrix composites. Figure 5 shows the trends in high temperature materials used for turbine blades. The limiting temperature of metal super alloys is approximately 1000C (Ohnabe et al., 1999). This limiting temperature has been increased by air cooling of the blades and introducing a thermal barrier coating. Ceramic matrix composites can operate at temperatures above that of super alloys and do not require air cooling contributing to performances increases. Hence they provide a viable material option to meet future targets.

These trends can already been seen in many metal parts that have been replaced with composite equivalents such as composite blocker doors resulting in weight savings of 20% and 30% for the acoustic and non-acoustic doors respectively, cascades resulting in a 30% reduction compared to the original magnesium cascade (Luttgeharm, 1990)as well as many other non-rotating engine components such as ducts which have been made from fibre composites due to the polymide matrix withstanding operating temperatures of over 260C as well as the complex fan frame of graphite-polymide to replace titanium alloy which tolerates stress levels over 100ksi and a weight reduction of 20% (Upadhya, 1994). It is an extensive task to describe all the composite components that have replaced metal parts but it suffices to say that the use of composites has made much progress on aircraft such as 757, 767, and DC-9 Super 80 and continues to make major developments in current and future aircraft such as the GE90 Engine, GEnx Engine and F136 engine.

4. Composite manufacturing and cost

Some manufacturing methods are electron-beam welding, electro-discharge machining (EDM) and electro-chemical machining (ECM), however in the current competitive aircraft industry there are many development programs set up to develop new manufacturing, assembly and design methods to lower the costs and improve the consistence of composite parts (Jeffreys & Leaney, 2000).

There have been vast improvements in the manufacturing in composites leading them to be a commercially viable option over metals. For example, the processes of resin transfer molding (RTM) has been developed to replace traditional hand lay-up and prepreg methods for fabricating fibre-reinforced resin-composites, speeding up the manufacturing process by carrying out the shaping and curing functions in one step. Dow-United technologies, a maker of composite parts in Wallingford, Connecticut, further developed RTM processes to ensure improved tolerances from +/-0.05 to +/-0.005 inch (+/-1.27mm to +/-0.127mm) as well as allowing for a higher level of complexity of part by forming complex shapes from untreated fabric plies while they are still flexible. Current RTM manufacturing processes are competitive with titanium on a cost/weight saving basis (Ashley, 1997). Other projects such as AMICC (The Automated Manufacture of Integrated Composite Components) funded by EPSRC and UK Industry, also reduced the cost of manufacturing carbon fibre composites by developing high speed tape laying (HSTL) which automates the process of layer building of varying orientation tape into laminated perform of specified thickness. When the costs were determined to the RTM aileron demonstrator HSTL provided a 45% cost saving (585) over standard hand lay biaxial fabric press (730) (Mills, 2006).

There have also been improvements in composite design software allowing for better prediction of plies using packages such as "FibreSim", from Composite Design Technologies, Wellesley, Massachusetts as well as improved resin injection simulation from packages such as "PreFlow", from Dow Chemical and Ohio State University, Columbia (Ashley, 1997).

These and other developments lead composites to offer lower total-life cycle costs compared to metals due to reduced maintenance costs, reduced manufacturing costs and improved reparability. Composites are also very reliable materials. Examples can be seen in the aileron and rudders on the Boeing 757 which were repaired using resin transfer curing resins in 1985 and did not need repairing until June 2000, after 15 years and 30000 hours of flight time (Armstrong, 2003).

5. Applications of composites to aero-engines

5.1 Carbon Fibre Composites (CFC)

Rolls Royce - 1960's

During the development of carbon fibres in the 1960's, Rolls Royce made one of the first major attempts to apply composites into aircraft engines. They designed and developed the Hyfil fan blade for the RB211 program. Although Rolls Royce had a good understanding of composites in terms of laminate layers they had a strict timescale to adhere to and did not have enough experience to apply this knowledge to rotating parts, such as the fan blade application, which have additional considerations such as foreign object damage (FOD) as well as the inherent weakness of attaching composite components (Ashley, 1997). Consequently they experienced issues regarding bird impact resistance and repeatability problems in the manufacturing process. This eventually led to the proposed carbon fibre blade being replaced with an existing titanium part. In addition, the company also went bankrupt as a result. Cost effective manufacturing methods for carbon composites were not yet developed so composites were not seen as a commercially viable option over existing familiar alloys. In addition the knowledge of composites compared to alloys was very small so designers were reluctant to change their selected materials. The trend in alloys being replaced with composite components has been driven by weight and temperature demands.

Design considerations

Carbon fibre composites pose the problem that the different materials being used have differences in thermal expansion coefficients (Adams & Wake, 1986). Designers overcome this by producing joints which compensate for physical property differences or by creating structures which are a combination of metal and carbon fibre layers (Shaw, 1993)(Allen, 1992). In a study of adhesively bonded aluminium composite structure (Halliday et al., 1999) it was possible to obtain a durable structure by altering the thickness of the carbon fibre composite layer to have a stiffness equal to that of aluminium. Distortions can de avoided by the use of room temperature curing adhesives, this also minimised residual stresses in the joint. The joint was found to be intrinsically stable when heated while immersed in water. This study showed that carbon fibre composites can perform under humid conditions.

When applying carbon fibre composite to rotating parts it is vital to consider fatigue resistance. Carbon fibre composites demonstrate breaking strains of 2 per cent, high stiffness ranging from 220 to 400GPa and high strength of up to 5GPa (Curtis & Dorey, 1989) but the high strength and stiffness is confined to the fibre directions causing composites to be laminated. In addition carbon fibre composites have low density properties and good corrosion resistance. In addition to fibre development programs, resin manufacturers have also improved the performance of carbon composites, increasing the toughness without compromising on high-temperature properties. Also leading to increased static strength but also reduced fatigue behaviour as outlined in figure 6. Due to layers having different elastic properties, stresses may appear at inserts, holes, straight edges or free edges which can cause delamination between layers however, stress concentrators have less effect on fatigue strength. Stress concentrators produce damage zones, which if they do not damage fibres (such as cracks along the fibres within layers or inter-laminar cracking between layers) can lead to increased fatigue strength. In fact, strength may be increased by applying zero to peak load over several hundred cycles (Curtis & Dorey, 1989). However damage zones created by stress concentrations can reduce static strength by 50%. The worst fatigue loading is fully reversed axial failure (Curtis & Dorey, 1989).

Despite all the fatigue strength considerations, composite fatigue performance excels over metallic materials. Figure 7 outlines typical S-N fatigue failure of standard composite laminate against aluminium-lithium alloy both used for aerospace applications.

Past applications such as in helicopter rotor blades were driven by carbon composite fatigue resistance properties with rotor blades put into service with rated infinite lifetimes. The main difference in fatigue properties between metals are carbon composites is that they exhibit different damage mechanisms. Metals usually failure to single flaws such as cracks, or other stress concentrations whereas composites have damage zones.

Design criterion - Example: Rolls Royce intermediate casing

When considering replacing an alloy for a composite substitute there are two desired objectives: increasing the specific power without an increase in weight or decreasing the weight without a compromise on power or performance. Designers want to ensure that they achieve the optimised solution as they replace titanium alloys in the engine's "cold" section with an increased amount of fibre reinforced composites for their specific mass and adjustable stiffness (Kober et al., 2008). Old methods to determine the stiffness and strength of a structure was to carryout numerical calculations using FEM which often had to be repeated with many time consuming iterations in order to find the correct structure. Design improvements now ensure that these iterations are done automatically by an efficient algorithm. Rolls Royce used topology optimization to optimise layer thickness with specific strength constraints to determining whether to replace a titanium intermediate casing with CFRP (carbon fibre reinforced polymer). It is difficult to predict the failure of heterogeneous structures and orthotropic materials such as CFRP. The most recent failure criterion used for composites is NASA LaRC04 (Pinho et al., 2004) which can predict the type of inter-fibre fracture unlike TASAI-HILL, TSAI-WU failure criterion; alloys uses conventional Von Mises and Tresca criterion. In this case Von Mises and NASA LaRC04 were used. The intermediate casing (IMC) is located between the fan outlet guide vanes and the bypass duct and connects to pylon of the aircraft to the engine. The IMC wants for minimal weight while retaining stiffness. After optimisation, the titanium structure using Von Mises criterion produced a weight of 709g while the CFRP structure weighed 424g showing a weight reduction of about 40%. Both structures were able to carry the same load in addition to satisfying their particular damage criterion. It is however harder to quantify the weight saving potential in disturbed regions. It is also wise to avoid high curvature regions in order to decreases the stresses through the thickness of the laminate structure.

Nacelles

Composites were initially applied to simple parts such as the cowl doors on the RB211 524 as shown in figure 8 as well as now contributing for approximately 25% of the weight and 20% of the cost saving of the power plant by being applied to nacelles (Luttgeharm, 1990).

Fibre composites are now used in by-pass ducts for the General Electric F414 military engine as well as in the Rolls Royce Tay civil engines. Fan blades, outlet guide vanes, nose cone spinners, core engine fairings, annulus fillers and variable guide name rings are now all made from fibre composites.

Fan nose spinner

Roll's Rolls fan nose spinners located at the front of the engine are now made using fibre reinforced epoxy (King, 1997). Manufacturing processes of this component were fairly straightforward however entry into service took a long time. This was due to costs in quantifying the costs and weight benefits as well as manufacturing costs. This cost was provided as a result of the development in the RB211 engine which provided the costs for the certification test programme.

Fan outlet guide vanes

Rolls Royce applied T300 and M40J carbon fibres as well as fibreglass and Kevlar to the design of 74 bypass vanes on the AE3007 engine which powers the regional jet Embracer ERJ135 and 145, the Citation X business jet and unmanned vehicles such as the Global Hawk. The use of composite resulted in reduced weight and increased stiffness in the structure when compared to steel and titanium counterparts. These vanes are directly behind the inlet fan and so still make use of nickel alloy for the leading edge. Texas Composites manufactured the 6 inch by 2 inch vanes individually as shown in figure 10 using RTM technology then bonds them into 1 two-vane and 24 three-vane assemblies for one complete 74 piece set. At the start of 2003 approximately 65000 vanes had been manufactured and shipped (Berenberg, 2004).

NASA and Rolls Royce also developed commercially viable coatings through the Higher Operating Temperature Propulsion Components (HOTPC) Project in order to reduce the erosion of the vanes caused by abrasive materials and impurities in the air (Sutter et al., 2004). In addition it reduces the costs of maintenance and reduces the probability of engine failure. Using these coatings has significantly increased durability extending the component life of bypass vanes to 8 times as much of uncoated vanes. NASA's 2003 Research and development document (Sutter et al., 2004) outlined that Rolls Royce's coated fan outlet guide vanes completed almost 5000 engine cycles and over 2000 hours of engine testing with no loss of coating and minimal cracking. This was achieved despite the vanes being removed as part of a maintenance schedule after 1000 hours and being subjected to simulated extreme conditions in the Structural Dynamics Lab in Glenn to carry out further stress tests. This component shows the value in coating which will allow the component to perform for its full design life with no concerns for wear or failure.

Fan inlet case

Dow-UT has applied carbon composites to an engine fan inlet case, replacing the existing titanium assembly currently used in the F119 engine build by Pratt & Whitney to power the F-22 as shown in figure 11. As the fan inlet case supports the shape bearing to the engine case, strength was of great importance. The composite part reduced weight by 6.8kg while maintaining strength (Ashley, 1997). The fan inlet case has also had worst case load conditions statically tested. Other considerations are that the fan inlet case consists of many parts including inner bearing hub, an outer support ring, mounting lugs and many airfoil-shaped structures. By utilising RTM the composite part can be moulded to accurate final dimensions without the need for secondary machining and is moulded as one integral part. By reducing the number of parts, composites have significantly lower assembly costs when compared to titanium.

Current developments

General electric developed the first commercial use of composites to engine parts with the use of composite fan blades on their GE90 engine which powered the Boeing 777 in 1995. In addition to this composite component, as part of their 9.7 million dollar contract with GE, Dow-UT manufactured the redesigned composite fan platform components which provide a flow-path surface between the fan blade airfoils (Ashley, 1997). The use of 22 components used in each engine contributed to the engine weight reduction.

GE has applied the use of carbon fibre and epoxy resin composite materials to its GEnx jet engine in the front fan case and fan blades (Anon., 2005). These composites have been manufactured using the same resins, fibres and manufacturing processes as on the GE90 mentioned above. Using these materials have been a significant part in increasing the engine performance with it achieving a stable operation in the range of normal operating limits of 2000 to 80500lbs of power at the GE Peebles, Ohio testing centre. The application of composites to the fan case has decreased the engine weight by 160kg and the overall weight saving potential in using the composites is a 360kg weight reduction for a two engine aircraft, such as the Boeing's 787 Dreamliner, Boeing 747 - 8 and Airbus' initial A350 aircraft, that the GEnx will power (Anon., 2005).

Much testing was carried out to ensure the engine could perform under the range of environmental conditions. Temperature testing carried out in February 2007 where the engine was exposed to -4F concluded that the GEnx can perform during severe icing cloud conditions. The structural integrity of the composite fan case was tested on July 3 2007 by releasing a fan blade at maximum speed. The fan case contained the released fan blade while retaining structural integrity. Impact tests consisting of hailstone tests carried out in May 2007 - where ice balls of one and two inch diameters were shot into the engine; rain storm tests - where severe flying conditions are simulated by firing a generated rain storm cloud into the engine; and bird impact tests occurring during take-off conditions were carried out. The GEnx engine successfully passed all these tests in addition to others carried out such as cross wind testing leading the engine to be certified in March 2008 after GE submitted 180 reports which outlined the analysis and testing carried out of every component (GE, 2009).

Other carbon composite developments can be seen in the Volvo aero group who are a partner in GE GEnx engine and carryout testing of load-carrying structures in aircraft engines (Anon., 2005). In addition they also invested SEK50 million in the research and development of composites over an 18 month period. Their main objectives is to reduce the weight of the engine to increase efficiency through developing new manufacturing methods for existing stationary components such as compressor shells, turbine housings and frames as well as developing and manufacturing a fan module as part of its role in VITAL (Finley, 2008). Some potential new manufacturing methods include welding a frame together as oppose to casting using a single material as this allows for a mix of different types of material such as mixing carbon fibre with plastics. This would possibly result in a structure that is stronger (due to carbon fibre) and 10-30% lighter (due to carbon fibre mixed with plastics) than aluminium.

5.2 Ceramic Matrix Composites (CMC)

Material properties

The first type of ceramics known as monolithic ceramics, exhibit brittle behaviour which leads to failure at the limit of the elastic region. Small defects from manufacturing processes such as grinding can act as stress concentrations leading to the initiation of cracks and sudden failure due to their brittle nature. Therefore surface finish is of great importance with companies such as RRACC producing surface finished down to 0.02 microns RA (Angus, 1993). Due to their brittle nature monolithic ceramics are unsuitable materials for aircraft engine components. By the addition of fibres to the ceramic to produce ceramic matrix composites (CMC), the structure becomes more damage tolerant making CMC a suitable material for aircraft engine material selection. The addition of the fibres increases the ductile properties of the material by introducing crack deflection and load transfer mechanisms which allow for improved energy absorption at the onset of failure (Mason et al., 1993) and (Ohnabe et al., 1999). The properties of the CMC are determined by the fibres and also the interface between the matrix and the fibres. CMC display a linear elastic behaviour similar to monolithic ceramics up to the point when microcracks start to develop, however once the behaviour becomes non-linear and irreversible, CMC demonstrate a much more controlled failure process. This controlled failure means that matrix microcracking provides a useful design stress for predicting failure behaviour before the material actually fails. This allows the designer to determine stress limits for the onset of cracking as well as determining more realistic predictions of the life of the material component. Figure 12 shows the stress-strain curves for monolithic toughened ceramics and ceramic matrix composites.

Companies such as Rolls Royce focus on silicon carbide reinforced glass ceramic composites (SiC/SiC) which are one of many ceramic materials such as silicon carbide and silicon nitride fibres which exhibit excellent oxidation resistance compared to carbon fibres resulting in increased structural stability at higher temperatures (Mason et al., 1993). Temperature is also an important consideration. At temperatures above 700C, oxidation of the carbon interface caused by exposed fibre ends, rapid diffusion through the matrix microcracks and oxygen diffusion through the porous matrix results in brittle behaviour and a consequent loss in strength. In addition the fibres start to deteriorate at temperatures above 1200C. To resolve these issues a protective coating is applied. Much research has been and is being carried out to find improved interface layer materials and improved protective coatings. The presence of matrix microcracks limits these materials to lower stress components (Mason et al., 1993). The types of loading on CMC is complex and consist of mechanical loading, stresses due to gas pressure at contact points, lads imposed by other components as well as thermal stresses arising from differences in material expansion coefficients and temperature gradients across sections. Despite these limitations CMCs still provide better specific stiffness and strength at higher operating temperatures, increased cycle temperature and reduced cooling compared to conventional alloys as shown in figure 13. Using CMC, there are possibilities to use high pressure air that would be used for cooling to be reused in the cycle which would increase thrust and efficiency as well as reducing the overall cost (Angus, 1993).

Design, processing and manufacturing considerations

CMCs have anisotropic properties which leads them to be difficult to design with as established design methodologies for metal components cannot be used. Separate design methodologies have had to be developed for CMCs (Mason et al., 1993). In addition, the current cost of producing CMCs is not commercially viable and will need to be reduced before parts even such as nozzle petals are considered for commercial use. Despite much development from companies such as RRACC who developed and patented a ceramic composite design, ceramic composites still have long process times and slow convergence to the new design concept (Angus, 1993). The latter has been further improved by new design optimisation which uses an efficient algorithm to automatically carry out design iterations.

Similar to CFC, ceramic fibres can be processed into complex shapes using textile methods of weaving and braiding. The chemical vapour infiltration (CVI) process, which is used in the manufacture of C/C composites, has been developed for the manufacturing of ceramic composites such as SiC/SiC. This process has been developed from the chemical vapour deposition (CVD) process used in electronics and coatings industries. The CVI process as outlined by Mason (Mason et al., 1993), involves insertion of a fibrous preform inside a furnace where the precursor gases, commonly methyltrichlorosilane (MTCS), diffuse into it and reacts to form the matrix. By diluting MTCS in hydrogen, the matrix i.e. silicon carbide increasingly fills the spaces between the fibres as the reaction happens, densifying the composite. The process is carried out at temperatures between 1000 and 1200C and HCI is produced as a waste product. In other processing methods such as the Pulse CVI process, the precursor gas is pumped into the reactor, allowed to react for a short time and then removed. This allows for the processing of complex shapes while increasing the efficiency of infiltration. Other manufacturing methods such as thermal gradient methods exist but these have their own problems. The problems surrounding CVI is in the difficulty in infiltrating into the centre of the preform as well as the control of gas flows within the CVI reactor. In addition the waste gases are corrosive and require neutralising in a scrubbing unit before being exhausted. These issues result in alterations which increase the process time and therefore the cost. Rolls Royce is currently trying to optimise the CVI process before considering other methods such as thermal gradient methods or pulse CVI (Mason et al., 1993).

Due to demanding environment, increases in engine operating temperatures as well as higher emissions targets work still needs to be carried out to improve fibre stability at high temperatures, improved oxidation resistant materials for matrix interface as well as fabrication methods that give consistent product. Other considerations as outlined by Angus (Angus, 1993) such as quality assurance, non-destructive testing, joining between different material interfaces such as ceramic/ceramic or ceramic/metal, ease of repair and emissivity will also continue to be important.

Design considerations - Example: Engine Igniters

Ceramic composites materials are used at the igniters tips of high energy igniters used extensively in aero-engines. Engine ignition conditions favour high temperatures and high energy release during sparks. Ceramic composites, in addition to operating well under these conditions, have a longer lifespan than diffusion igniters. Igniters operating life must range from 40 to 4000 hours depending on the application therefore determining the lifespan is of great importance (Blanchart & Jankowiak, 2006). When determining the life of igniters to consider the degradation mechanisms of CMC materials as well as mechanical problems in the ceramic assembly arising from thermal stresses and bonding requirements of different materials in the component need to be considered. Blanchart P (Blanchart & Jankowiak, 2006) examines these issues with particular focus on particle-reinforced ceramic matrix composites with a volume fraction of 50 to 70% micro-sized SiC grains as the conductive phase and SiAlON as the matrix phase.

In addition to thermal expansion coefficients and temperature gradients of different materials, thermal stresses and surface erosion can be caused by thermal stresses as a result of engine operation and the extreme transient temperatures exhibited during the start up and shut down of the engine which rapidly heat or cool the outside surface of igniters. This causes a large temperature gradient to form across the outside and the centre, therefore the igniters needs to be designed and developed to allow for these thermal stresses. High oxidation resistance is needed for igniters. Oxidation of composites causes a breakdown voltage from a local insulating layer on SiC grains. However the rate of breakdown is fairly low allowing SiC-SiAlON to be selected as suitable materials for extreme environment applications. In the particular study outlined by Blanchart, although large stresses are induced by the air atmosphere at 850C, the breakdown voltage increased by only 30% after 8 hours as shown in figure 14. The global oxidation reactions in air indicate an oxidation process controlled by diffusion mechanisms as it follows a parabolic curve.

Gas pressure in combustion chambers reaches ranges of 10 to 40 bars. Higher pressures result in an increase relative erosion at the same energy level as a result of reduced electron mean free path in the plasma and a reduction in the extend of the secondary ionisation phenomenon. Therefore pressure gradients and high pressures are important considerations when designing engine igniters. Spark energy and high temperature are also of importance as the energy delivered during sparks varies with the peak voltage and requirements from the engine manufacturer. In addition the voltage also has to increase as the material characteristics change as a result of long working period at high temperatures; temperatures of 300-1000C are reached. CMCs preformed well under increases in pressure and increases in energy as outlined in figure 15, where a 30% increase in relative erosion was exhibited as the energy increased from 1 to 5J at a higher pressure of 12 bars.

Ceramic composites provide the high energy discharge properties required for aero-engine igniters. The application of ceramic composites to igniters is well developed as the spark formation mechanisms involved which are related to ceramic compositions; microstructure characteristics and surface morphology are well known. Further development lies in extended the operating life and operating capabilities of ceramics as igniters by carrying out further investigations to investigate problems associated with degradation mechanisms and increasing operating temperatures of future engines.

Convergent nozzle Flap

CMCs were applied to nozzle flaps by replacing the current super alloy with 3D woven SiC polymer derived fibers combine with SiC matrix by the combined process of CVI and PIP. Environmental degradation was evaluated by determining the elastic modulus after engine testing. The elastic modulus was determined by the use of non-destructive inspections techniques that involve impact sound. The nozzle flaps were tested in a demonstrator engine under sea level conditions and performed for over 30 hours showing satisfactory results. In addition to successful performance the use of CMCs lead to a weight reduction of 50% over the current super alloys (Ohnabe et al., 1999).

Acoustic Liners

Ohnabe outlines the development of CMCs to exhaust nozzles (Ohnabe et al., 1999). The application of CMCs to exhaust nozzle has been employed to ejector suppressers in SST/HST. The main target in the development of mixer ejector nozzle is reducing noise levels (in line with Annex 16 Chapter 3 ICAO regulations) but without reducing the thrust. The ceramic composite applied to the acoustic liner was porous Al2O3 combined with 10% SiC whiskers.

Bladed disk

When designing rotating bladed discs it is important to consider the toughness of the component in being able to redistribute stresses as well as the peak strain and mean hoop stress. In addition, the differences in strength/density ratios between the rotating blade and static components are also critical. CMCs offer the advantage of being lightweight compared to conventional alloys. When applied to bladed discs, this reduction in weight results in lower loads such as shaft load and bearing compartment load which not only results in a weight reduction of the overall system but also improvements to the systems performance. Bladed disks were made of continuous Si-Ti-C-O Tyrannoe fibre reinforced SiC matrix composites (made of three dimensionally woven fabrics), developed by Ube Industries, Japan. The fibres were densified using a chemical vapour infiltration (CVI) process followed by polymer pregnation and pyrolysis (PIP) process (Ohnabe et al., 1999). This CMC was also applied to a 3D bladed disc developed in the AMG project (Araki et al, 1998).

The fibre reinforced SiC matrix composite exhibited a 500Mpa maximum tensile strength at room temperature. The stress-strain properties are shown in figure 17. Considering the stress-strain behaviour, the mechanical behaviour of the composite rotating disk was predicted using finite element analysis (FEA) and compared to testing results.

The results from placing two different thickness discs in a spin test rig showed that a 22mm thick SiC/SiC disk at room temperature exhibits the same strength as a 5mm thin SiC/SiC disc in a spin test rig with a burst speed of 32800 rpm (Suzumura et al., 1996). The measured strain correlated well with the predicted FEA values. In addition, the CMC material used exhibited good burst behaviour, with the stresses redistributed to a value of the mean equivalent to the ultimate tensile strength and good correlation between the burst behaviour of the two disks (Ohnabe et al., 1994).

Combustor Liner

CMCs have been applied to a combustor liner in the AMG project (Nishio et al., 1966) and the HYPR project (Kinoshita et al., 1995). The prototype CMC liner developed in the AMG project was made of SI-TI-C-O fibre combined with SiC and glass matrix and was processed by filament winding and PIP as shown in figure 18. The small scale CMC ram combustor liner pieces developed and fabricated on a trial for the HYPR project were made of C/SiC and SiC/SiC.

Turbine nozzle vane

Research has been carried out on the application of CMCs to turbine nozzle vanes. Methods of slip casting to form the complicated shapes required has being carried out using SiC whiskers and silicon nitride powder (Si3N4) (Sasaki et al., 1996), as shown in figure 19.

Some other potential applications of CMC inter-turbine transition ducts, fasteners (Hiromatsu, 1996), frame folders (Kashiwgi et al., 1993) and heat seals. Before CMC become a commercially viable option the cost of manufacturing and processing needs to reduce and improvements to the thermal stability and interface layers need to be made to cope with increasing temperatures.

Current Developments

GE and Rolls Royce have applied the use of CMC to components of the third-stage, low pressure turbine of their F136 which will power the F-35 Lightning ii Joint Strike Fighter (JFK). The vane tests being carried out on the F136 are the first introduction of CMC to the hot combustor and turbine sections as well as CMCs being more commercially viable. The components are made of silicon carbide ceramic fibres and were processed using CVI methods to produce a 100% CMC density without excess silicon (Norris, 2009). A ceramic resin heat resistant coating was applied after manufacture which allows the CMC structures to operate at temperatures of up to 1200C which is higher than conventional alloys such as nickel alloys and allows for reduced cooling air. In addition, CMC materials provide one third the weight of nickel and half the weight of titanium. By having a reduction in weight and a reduction in required cooling, the jet engine can run more efficiently or at a higher thrust. Early tests on a pre-production standard F136 LPT vane showed very little deformation or failure however some performance advantage is sacrificed by the joint required between the CMC vane and the metallic rotor to transfer load and compensate for the differential in coefficient expansion between the two materials.

Future rig tests of a CMC made shroud on a T700 turboshaft and land-based GE turbines will increase material and design knowledge. The CMC based combustor liner is predicted to reduce cooling requirements by 50%, improving efficiency and reduce nitrous oxide emissions by approximately 20%. In addition, CFM International will evaluate several CMC components during the Leap X demonstration program in 2010 which is aimed at developing technology for next-generation engines for narrow body aircraft (Angus, 1993).

5.3 Metal matrix composites (MMC)

Material properties

Metal matrix composites (MMC) can be divided into two distinctive types: particulate composites which a have structure where the foreign particles within their structure are uniformly distributed and composites where there are long fibres that run through the metal. The main metals used are titanium and aluminium or steels. Reinforcements used consist of carbide or oxide particulates such as silicon carbide or aluminium oxide, or for long fibres, reinforcements of alumina or silicon carbide are used (Butler, 2005). Both types of MMC allow for mechanical properties to have a range of tuning unlike conventional monolithic alloys (Butler, 2005). Both types of MMC are high in strength and stiffness as well as being lighter than conventional alloys. Figure 20 illustrates some mechanical property ranges of aluminium and titanium based matrix composites at room temperature.

MMC can also operate at higher temperature to alloys and exhibit good wear resistance (Charles, 1992b). This is becoming an increasing important consideration as engine thrust to weight ratios increase. Titanium based metal matrix composites have high strength and high modulus fibres leading to 50% weight reductions when compared to monolithic superalloys, making them an ideal material for material selection for propulsion system components (Pank & Jackson, 1993). Figure 21 shows the strength of titanium and aluminium based matrix composites against temperature.

The mechanical properties of metal matrix composites at high temperatures has continued to improve as aluminium and titanium based matrix continue to be developed and understood. Figure 22 illustrates the ultimate tensile strength, yield strength and percentage elongation as functions of temperature of titanium matrix reinforced composites tested in 2003 (Valente, 2003).

Design, processing and manufacturing considerations

Due to the high strength, high stiffness, improved elevated temperature properties and good wear resistance, MMC provide designers with a material that has improved structural efficiency, improved structural reliability and reduced maintenance costs when compared to carbon fibre reinforced composites (CFRP) (Charles, 1992b). MMC can be used for their excellent specific strengths allowing for lighter designs without compromising strength and stiffness or can be applied to areas deeper into the engine where combustion takes place due to their high temperature characteristics. In addition these characteristics can be tuned for a particular application, so for example the MMC expands at the same rate as the adjacent or connecting materials (Butler, 2005).MMC perform most effectively in the direction of the long fibres, similar to CFC however they are a strong metal alloy rather than highly brittle epoxy used in CFC. This leads to the significance of off-axis loading to be reduced although designs still need to consider the orientation of long fibres and the directions in which stresses are acting. MMC also do not have good drape characteristics compared with CFRC which makes them unsuitable for complex moulds (Butler, 2005).

In 1992, D Charles outlined that although two British companies, Alcan and BP, were developing continuous fibre reinforcements and were at the forefront of expansion, the scope of using MMC commercially still remained limited due to high material costs as well as limited development appropriate component fabrication technology (Charles, 1992b). Since then, due to a strive for lighter, stronger aircraft as well as relatively high budgets, there have been major developments in manufacturing processes as well as in the metal matrix and the particulate and long fibre materials used. Particulate MMC exhibit behaviour similar to monolithic alloys which allowing them to be processed and machined in the same way. However the machining cost of particulate MMC can be high if ceramic particulates are used as this may require diamond-tipped tooling (Butler, 2005). Although the fabrication manufacturing costs of MMC are higher than alloys due to the fabrication complexity, the process allows the material to be produced to a very high finish reducing the wastage and need for additional machining resulting in overall reduced costs when compared to alloys. One method of processing MMC is to hot press, plasma sprayed performs (Valente, 2003).

Potential application to aircraft engines

Due to the temperature limitations of alloys and other current metallic materials used, there is much scope for the applications of MMC to aircraft engines due to their benefits in terms of weight reduction as well as their operation at high temperatures. Rolls Royce aim to apply MMC to a significant proportion of future aircraft engines. Companies such as Textron Specialty Materials, US have produced fibre-reinforced titanium with strength three times larger than nickel based superalloys at temperatures up to 815C.

Textron have used titanium MMC with silicon carbide reinforcement in the form of a loop to manufacture compressor discs. The use of MMC could reduce the weight of a compressor disc from 16 to 1kg. There has also been much research and investigation into fabricating MMC blades. Many engine technology programmes such as IHPTET (Integrated high-performance turbine engine technology) (Charles, 1992b) use MMC. The possible applications of MMC to military aircraft has been shown in figure 23, which shows that MMC could also be applied to vanes and rotors as well as LP and HP casing and bypass ducts.

Current two dimensional convergent-divergent (2DCD) nozzles have independently variable divergent and convergent sections which optimize thrust vectoring capabilities and engine performance (Pank & Jackson, 1993). The gross thrust and pressure ratios on the engine are dependent on the varying convergent and divergent cross sectional areas of the nozzle. Therefore it is critical for the mechanical members of the nozzle to retain stiffness over time with varying temperatures in order for effective control of the cross sectional areas. MMC would be a suitable material for this application as the structure stiffness and strength of the nozzle could be optimized with a specified weight.

Current developments

It has been discussed that MMC can be made using a variety of metals such as titanium, aluminium or steels with carbide or oxide particulates or long fibres which may be alumina or silicon carbide. These composites can be stronger, stiffer, lighter or more heat resistant that conventional alloys. The lower weight allows cost savings throughout the components lifetime in addition to increasing engine performance and reducing fuel consumption of the overall aircraft to which it is applied. Companies such as Qinetiq are currently focussing on developing processes for titanium-based long fibre MMC (Butler, 2005). They have developed a process of applying a uniform coating of silicon carbide to a fine tungsten wire and depositing a uniform coating of titanium matrix alloy to the outside. Fibres can be grouped and arranged before fusing together resulting in the fibres being packed in a very regular array and making for a very light and strong material. This material can then be joined with titanium seamlessly which provides reinforcing seams throughout the component (Butler, 2005). MMC materials have proved to be suitable materials for aircraft engine design selection and are currently in the approval and certification process for use in aviation. The first examples are expected to fly within the next five years (Butler, 2005). Some companies at the forefront of MMC application are Dunlop Aerospace who produce commercially viable MMCs which have properties that make them suitable to replace heavier titanium alloys as well as Rolls Royce who are investigating the application of titanium based MMC to replace titanium alloys which are now reaching their operational temperature limits.

5.4 Carbon-carbon composites (C/Cs)

The main area for the application of carbon-carbon composites (C/Cs) is in spacecraft engines due to the extreme temperatures experienced in the various engine parts. C/Cs are being considered for similar reasons to using MMC, CMC and CFRC in that they can offer a significant weight reduction which results in an increase in engine performance and fuel savings. Weight can be seen as being even more critical in spacecraft than for aircraft, so investigating the viability of C/Cs is essential. The Institute of Space and Astronautical Science, Japan have studied a two-stages-to-orbit (TSTO) type space plane for a future space transportation (Goto et al., 2003) and determined that the most suitable candidate for the first stage propulsion system of a perfectly recycled TSTO space plane was an air turbo ramjet engine with expanded cylinder (ATREX). They considered the application of C/Cs for the turbine disk of the ATREX and focused particularly on fly out behavior. Figure 24 outlines the high temperature structures of the ATREX and the liquid flow.

C/Cs have low densities of less than 2000 kg/m3 and retain high toughness and high strength up to temperatures above 2273K. However, existing research highlights that C/Cs can fail by oxidation causing a rapid degradation of strength. For this reason many industries are trying to use SiC matrix composites (CMC as discussed in section 5.2) as oxidation coatings for these materials are fairly developed. C/Cs offer an advantage over CMC in terms of higher strengths and maximum operating temperatures and for this particular application, oxidation does not pose a problem as sever temperatures are only experienced for less than 5 minutes a flight (Goto et al., 2003).

Understanding the fracture behavior is of great importance when applying C/Cs to rotating parts. Kogo et al has recently studied the fracture behavoiur of a 2D C/C laminated disk (Kogo et al., 1998) and showed that even for small radius ratios of ri/ro=0.2, where ri is the inner radius and ro is the outer radius, the fracture observed correlated with the average hoop stress criterion. For 3D disks, micro-fractures are usually observed before total fracture in rotational tests, this can also cause flown-out fragments causing extensive growth in vibration. The disk used in this study was heat treated at 2573K to attain a bulk density of 1800kg/m3. In this study, the fracture speed of the 3D C/Cs was 516m/s which above the suitable limit for use in ATREX.

Silicon impregnation processes of CVI (chemical vapor infiltration) and liquid Si infiltration were carried out to improve the adhesion between fibres and try and prevent against fly-out behavior. However, in this study, fly out behavior occurred and was determined to be as a result of unstable debonding between fiber bundles caused by sharply terminated sections as a result of machining. In addition, simultaneous to the fly-out behavior, there is a slight increase in shaft vibration as a result of rotation imbalance increases as the speed. Figure 25 illustrated the effect of sharply machined point on fly-out behavior.

To try and reduce the effects of fly out behaviour, the strength between the fibre bundles can be improved to try and reduce debonding effects by impregnation of more carbides or carbon matrices. A finer fiber texture could also be used in order to decrease bundle thickness. In this study, liquid Si impregnation was found to be an effective method for preventing fly-out behaviour.

The main improvement lies in the machining methods of C/Cs. 3D C/Cs have a coarse bundle structure which means that any edge formed using conventional milling techniques may not contain a reinforcing fiber and will be prone to fracture. If conventional machining methods are to be used, their parameters such as tip radius, operating temperatures and cutting speed need to be optimised to ensure that a sharp turbine blade is achieved.

6. Related Composite Projects

The scope of composites goes beyond just aviation applications with composites being applied to many other applications. MMC have been applied to spacecraft in the production of thermally stable space structures. DWA Composites and Materials Concepts Inc. Have both produced graphite-reinforced aluminium structures for NASA. The space shuttle, HOTOL and NASP was the first major use of MMC in re-useable aircraft giving a 44% weight saving over aluminium and reducing required thermal insulation due to composites lower thermal conductivity (Charles, 1992b). MMC can also be applied to guided weapons, in parts such as fins, missiles and microwave packaging for microwave devices (Charles, 1992b) as well as for satellites and communication systems (Upadhya, 1994).

In addition to jet engines, composites are being used in wind turbines as well as in oil and gas applications. Recently, on the 3rd July 2009, the Technical University of Munich (TUM) and GE Global research signed a memorandum of understanding to set up a world class manufacturing centre of excellence for carbon composites. The site, based at TUM campus in Garching, near Munich in Germany will focus on the automated manufacturing of complex composite structures. This will enable composite structures to be manufactured at a reduced cost as well as increasing the quality, reliability and rate of production. By improving the manufacturing process, new applications can be investigated such as the development of longer advanced wind blades for wind turbines to increases wind capture and developing stronger risers with no reduction of weight to enable high pressure deep-sea oil exploration and production (Anon., 2009). Figure 26 outlines the history of the major composite projects in Japan which focused on the development of composites for applications of land and marine gas turbines and gas generators in addition to aero-engines (Ohnabe et al., 1999).

In order for the application of composites to be successful, a bigger market is needed. Aviation research of composites for the application to aero-engine parts should also be applied to other areas to ensure the viability of these new materials. A good example of this is the production of ceramic turbochargers in the Japanese motor industry (Anon, 1994). Anon (Anon, 1994) and Butler (Butler, 2005) both comment on the benefits of aviation collaborating with the motor industry resulting in increases in the use of these materials and also increasing the market of using these materials. There are many core suppliers and end users of composites in the UK; Qinetiq have estimated that the aerospace market for composites, specifically MMC could be as high as $200 million (123 million); collaboration with motor industry would increase this.

7. Conclusion

Composites are commonly used in aircraft structures, but are now being introduced into jet engines in aero-components in both the "cold" and "hot" sections as weight reduction becomes a priority. All the components in current and future engines needs to operate under individual influences, running conditions and requirements. The structural integrity, environmental durability, operating temperatures and creep resistance are some important material property considerations {{12 Anon 1994}}. In addition aero-engines are also required to give high performance, low cycle costs, high thrust to weight ratios (Charles, 1992b) as well as demanding reduced weight and reduced emissions and noise levels as well as increases in fuel efficiency.

Composites are suitable materials for aero-engine applications due to the beneficial properties of low weight while still providing high strength and stiffness. Composite fatigue performance excels over metallic materials. CFC demonstrating breaking strains of 2 per cent, high stiffness ranging from 220 to 400GPa and high strength of up to 5 GPa (Curtis & Dorey, 1989), CMC have excellent energy absorption at the onset of failure demonstrating a much more controlled failure process and MMC are high in strength and stiffness as well as being lighter than conventional alloys (Butler, 2005). Composites also have the added benefit of having tuneable mechanical properties unlike conventional monolithic alloys making composites suitable for designing for a particular component or to the material properties of existing structures (Butler, 2005). With research and development into composites, such as development of the composite matrix, resin improvements as well as improvements to manufacturing processes, composites exhibit excellent corrosion and wear resistance. CFC have increasing toughness without compromising on high-temperature properties, CMC have excellent oxidation resistance compared to CFC resulting in increased structural stability at higher temperatures and MMC in addition to CMC can also operate at higher temperature compared to alloys (Charles, 1992b). The application of composites to aero-engine parts allows the engine performance to continue to develop, for example, by replacing conventional alloys with CMC in the hot section, there is the possibility to use high pressure air that would be used for cooling to be reused in the cycle increasing thrust and efficiency as well as reducing the overall cost. Also composites can be used in areas deeper into the engine where combustion takes place due to their weight reduction without compromising on strength or stiffness. Although C/C composites have not been applied in many areas of aero-engines, due to the weight reduction benefit and temperature resistance, C/C are being considered for application to spacecraft engines due to the extreme temperatures experienced in the various engine parts (Goto et al., 2003).

Although composites have got a long way in being applied to aero-engines, due to demanding environment with increases in engine operating temperatures as well as higher emissions targets, fibre stability, improved oxidation resistant materials for matrix interface and fabrication methods that give consistent product need to continue to be developed. Other considerations as outlined by Angus (Angus, 1993) such as quality assurance, non-destructive testing, ease of repair and emissivity will also continue to be important. Designers need to continue to become comfortable with composites and for them to be considered for material selection, in addition to material properties, commercial viability is of great importance. As the cost continues to reduce, the composite material properties are standardised and there are increases in the reliability of manufacturing, maintenance and repair when compared to traditional material, composites will continue to be applied to more and more aero-engine parts as well as other applications.

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Imperial College London - Mechanical Engineering

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