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Aerospace Materials Chapter: 7. Materials Developments In Aeroengine Gas Turbines
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Aerospace Materials Chapter: 7. Materials Developments In Aeroengine Gas Turbines
Champs Elysee Roldan
SOCIEDAD JULIO GARAVITO
Jun 16, 2012
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Transcript of "Aerospace Materials Chapter: 7. Materials Developments In Aeroengine Gas Turbines"
1. Chapter 7 Materials developments in aeroengine gas turbines David Clarke and Steve Bold Introduction Aeroengine gas turbines can be optimized in diﬀerent ways for diﬀerent applications, and even in diﬀerent ways for the same application. This optimization process is driven by a number of broad factors as diverse as the basic laws of physics and gas behaviour through to aircraft operational performance and the ﬁnancial needs of both the engine manufacturer and the end user of the aircraft. Through all of this runs a common factor: materials—metals, polymers and ceramics—are key to balancing all these factors for achievement of the optimum engine design and, in turn, the design needs for future engines deﬁne the need for materials development. Current engine design Key factors driving engine design In civil applications, fuel eﬃciency is a key factor but, with the three major engine manufacturers all oﬀering engines with similar fuel burn performance, secondary factors such as noise, emissions, weight and reliability become the major product diﬀerentiators. Highest engine eﬃciency is achieved through a high pressure ratio in the compressor and a large temperature rise through the combustor. Both factors are limited by the temperature capability of the materials available and the cooling technology used. Propulsive eﬃciency also depends on matching the exit velocity of the gas stream to the speed of the aircraft. For this reason, civil engines use large bypass ratios so that work is put into moving a larger mass of air more slowly than the air movingCopyright © 2001 IOP Publishing Ltd
Figure 7.1. Pressure and temperature cycles through the high thrust, high bypass Trent 800 engine. through the engine core. On take-oﬀ up to 80% of air goes through the by- pass duct. Figure 7.1 shows pressure and temperature cycles through the high thrust, high bypass Trent 800 engine. Whilst it is the obvious basis of all engine design, it is worth remember- ing that the ﬁrst criterion in the design process is having the ability to deliver the necessary thrust to ﬂy and manoeuvre the aircraft. This is particularly relevant in military applications where speciﬁc thrust (thrust/engine mass) is generally paramount, followed by fuel consumption, emissions (observa- bility), reliability, maintainability and noise. Because of the higher aircraft speed and the need for additional responsiveness, military engines have low bypass ratios and very high gas exit velocities. Pressure ratios tend to be lower giving maximum speciﬁc work but generally not maximum fuel economy. For both civil and military applications there is a common factor—cost. Product development costs and production unit costs are considered from the ﬁrst stages of engine design and create as many challenges for materials development as the engine performance parameters. Two engine designs The vast majority of civil aircraft are available ﬁtted with engines from more than one manufacturer. Figure 7.2 shows a real example of the results of twoCopyright © 2001 IOP Publishing Ltd
Figure 7.2. Comparison of turbine temperature and pressure ratio for two aeroengines designed for the same aircraft. diﬀerent manufacturers’ design solutions for one particular aircraft. Both engines were designed at the same time for the same duty and whilst both use similar basic technologies, produce the same thrust and weigh similar amounts, their design philosophies are quite diﬀerent. Physics dictates that engine thermodynamic eﬃciency increases with increasing turbine gas temperature and increasing pressure ratio (combustor entry pressure to ambient air inlet pressure). Primarily as a result of running hotter and at higher pressures, engine 2 burns around 4% less fuel than engine 1 for the same take-oﬀ thrust. The impact of raising the gas tempera- ture and pressure to achieve this higher eﬃciency, however, is that core engine components in engine 2 degrade much more quickly than those in engine 1, and engine 2 has to be removed from the aircraft twice as often for major maintenance. The ﬁnancial impact of this to the operator is that engine 1, the less technically eﬃcient engine, costs nearly 10% less to operate than engine 2. The ﬁnancial impact on the engine manufacturer is that engine Figure 7.3. Current design drivers and materials responses.Copyright © 2001 IOP Publishing Ltd
1 now accounts for 80% of this market sector with over 1000 engines in service. Engine design is thus clearly a complex optimization process and there are many solutions to one overall requirement. The impact of the diﬀerent design styles and operational parameters of these two engines on the materials needed to manufacture them is equally dramatic. Figure 7.3 shows the current design drivers and material responses. Materials Turbine blade alloy development Since the initial development of the gas turbine in the 1940s the temperature capability of the nickel alloys used for the highest temperature parts of the engine—high pressure turbine blades, discs and the combustion chamber— has increased by around 4008C to nearly 12008C. This 4008C increase has involved four major stages of development in material and manufacturing technologies, from forged alloys to cast systems, then directionally solidiﬁed (DS) castings to eliminate creep problems asso- ciated with transverse grain boundaries, and now to single crystal castings with control of both the longitudinal and the transverse crystal orientation. This development is summarized in ﬁgure 7.4. To achieve a further increase in temperature capability is now requiring a ﬁfth stage of development, with the introduction of surface coatings to reduce oxidation and corrosion, and ultimately a complex system of thermal barrier coatings (TBCs) to reduce the Figure 7.4. Temperature capability of diﬀerent turbine blades since 1940.Copyright © 2001 IOP Publishing Ltd
rate of heat transfer from the blade surface to the internally cooled surfaces in its hollow core. This latter development is discussed in more detail later. An increase of 408C in turbine blade temperature as highlighted between the two engines in the example in ﬁgure 7.2 denotes a signiﬁcant increase in materials technology level. Typically this represents the diﬀerence between two technology levels, e.g. a directionally solidiﬁed to a single crystal alloy, or a single crystal to a system of single crystal plus oxidation-resistant coat- ings. Whilst this level of operating temperature increase gives a signiﬁcant fuel burn beneﬁt (4% is around the maximum diﬀerence seen between diﬀer- ent engines designed for a single airframe application) it is only of real value if this can be matched by comparable reliability between the two units. The need for a particular level of materials technology or a signiﬁcant advance is thus created as the fundamental engine architecture and thermodynamic cycle are set. It is obviously possible to operate engines at low temperatures but only with a deﬁned performance penalty and in an increasingly compe- titive market this is unacceptable. Continuous performance enhancements, only deliverable by ongoing materials developments, are essential. Evolutionary versus revolutionary development Historically the majority of materials developments for gas turbines have been evolutionary, as in the gradual evolution of each nickel alloy technology described above. Both nickel and titanium alloys, however, are now at a level of development where further small advances are of limited beneﬁt and are increasingly expensive to achieve. Engine design now demands revolutionary developments in the ﬁelds of increased temperature capability and reduced component mass. The primary materials development programmes to achieve this are almost all based around either composite materials or, increasingly often, a composite structure, integrating the beneﬁts and proper- ties of various materials systems into a single component. The following sections discuss some speciﬁc examples of this approach. Nickel based metal/ceramic structures The majority of turbine components are cooled by air from the compressor. This increases component life but reduces the engine eﬃciency, since the engine pressure ratio is eﬀectively being progressively decreased as air is bled oﬀ to these various sections. The eﬀect of this is to increase the engine fuel burn compared with the theoretical minimum. In practice this results in increased cost for the operator. Whilst the reduction in eﬃciency as a result of these cooling bleeds is very small, the overall ﬁnancial eﬀect to a large ﬂeet operator may be very signiﬁcant. A 1% reduction in fuel burn to a typical long haul operator of ten Boeing 747-400s would reduce annual fuel bills by well over $1 million.Copyright © 2001 IOP Publishing Ltd
To reduce the need for cooling air, many turbine components now incorporate ceramic materials. The latest developments of high pressure (HP) turbine blades use nickel alloys coated with a graded series of layered materials to give a component which relies for its operation on the synergy of properties from the metal substrate and ceramic coating. The single crystal nickel base alloy is ﬁrst coated with a plasma sprayed MCrAlY overlay coating (M represents Ni or Co). The chromium and aluminium provide oxidation resistance while the presence of yttrium improves scale adhesion. In addition the layer acts as a bond coat to prevent spalling of the outer coating layer—the thermal barrier coating. The thermal barrier coating is a low thermal conductivity ceramic which restricts the ﬂow of heat from the gas stream to the metal blade. This maximizes the beneﬁt obtained from blade cooling and oﬀers a potential increase in operating temperature of over 1008C. Thermal barrier coatings have been used in the combustion chambers of the RB211 since 1975 but it is only with advances in bond coat technology and ceramic deposition techniques that they can be reliably used on critical rotating parts where coating failure could lead to premature component removal. Coatings are applied by electron beam physical vapour deposition to develop the columnar grain microstructure necessary to resist thermal and mechanical strains, particularly around blade leading and trailing edges. Figure 7.5 shows a coated blade from the Trent 800. Full exploitation of coating technology requires the coating to be seen as an integral part of the component from the start of the design process. The coating system must be compatible with the requirements of aerodynamics, Trent 800 HP blade + thermal barrier coating TBC usage is a balance between : Material properties Strain tolerance Oxidation/corrosion Life prediction and Manufacturing issues Coating thickness Hole closure Coating structure and integrity Process control Figure 7.5. Trent 800 high pressure (HP) turbine blade, showing cooling holes and thermal barrier coating (TBC).Copyright © 2001 IOP Publishing Ltd
mechanical integrity and blade cooling and must be created in a cost-eﬀective manner suitable for mass production. All of these considerations have led to the unique, cost-eﬀective thermal barrier coating system designed for use in the Rolls-Royce Trent engines. Thermal barrier coatings are discussed in more detail in chapter 22. Nickel/ceramic hybrid combinations are also being developed on a macro scale for mechanically integrated large static structures. Here a simple metallic unit forms the major load carrying element and a semi-structural ceramic shell, or more commonly a ﬁbre reinforced ceramic composite, forms the high temperature surfaces of the component. This approach has been used in turbine blade tip seals on the Trent engine, and exhaust structures for military engines such as the EJ200. Titanium aluminides and titanium metal matrix composites Unlike turbine components, compressor components are not normally cooled. They operate at gas stream temperatures which reach around 650ÿ7508C at the compressor exit. The primary design driver here is to increase pressure ratio and this largely drives further increases in this exit temperature. Compressor discs and blades are largely formed from titanium alloys, with much of the historical alloy development emphasis having been on developing defect tolerance and increased temperature capability. The latter is exempliﬁed in alloys such as IMI834 which are used in applications up to 6308C—almost twice the capability of the routine Ti-6Al-4V at 3508C. Moving to these higher temperature alloys has allowed weight savings in compressor modules of over 15%. Alternatives to titanium are primarily steel and nickel alloys with the consequent mass increases. Major performance improvements are now being made by either using intermetallic (titanium aluminide) materials or silicon carbide ﬁbre reinforced titanium composites. Gamma titanium aluminides are the most advanced of the intermetallics with half the density of current titanium alloys. Primary beneﬁts are very high speciﬁc stiﬀness and inherent non-burning chemistry. Temperature capability is currently around 7508C, making the system suitable for back end compressor and turbine applications. Their non-burning nature makes them particularly suitable for stator vanes, where titanium alloy use is limited by the risk of titanium ﬁres. Application of aluminides could lead to whole engine weight savings of up to 4%. These materials will also give reduced engine life cycle costs and improved engine/airframe functionality since reducing mass on one element of an engine structure has a consequent knock-on eﬀect on surrounding components—lower-mass blades lead to lower-mass discs and shafts, reduced stiﬀness casings etc. Titanium aluminides are discussed further in chapter 17.Copyright © 2001 IOP Publishing Ltd
Figure 7.6. Weight savings with blisc and bling designs using titanium metal matrix composites (MMCs). Knock-on eﬀects in reducing mass of associated components are particularly signiﬁcant in rotating bling (bladed ring) structures made in titanium metal matrix composite (Ti MMC). The bling is an integral struc- ture of titanium blades on a titanium metal matrix composite. The reinforced ring increases hoop stiﬀness by 100% and strength by 50%. This removes the need for the heavy bore of the disc, giving weight savings of the order of 40% over a conventional titanium blisk (bladed disc) design, as shown in ﬁgure 7.6. Rolls-Royce Allison has successfully run the ﬁrst demonstration of titanium metal matrix composite blings in an engine. Titanium metal matrix composites are discussed in more detail in chapter 18. The weight savings in the bling itself are greatly increased by the options presented for major changes to the surrounding structures, e.g. the possibility of using a larger diameter (and hence stiﬀer) shaft, since the shaft diameter constraint imposed by the disc bores is eﬀectively removed. This geometric change opens up the possibility of replacing the steel shaft with a titanium unit, with dramatic weight beneﬁts. Steel shafts are currently used to ensure adequate torsional and bending stiﬀness, which must be achieved within the very low diameter allowed for a conventional shaft geometry. Future design drivers Future engine designs demand still further increases in the basic design para- meters of temperature and pressure, but are increasingly requiring enhanced materials capabilities in more diverse ﬁelds. Engine design has evolved for the past 50 years around a largely unchanged group of mechanical technologies, and these have needed a standard set of materials with optimized mechanicalCopyright © 2001 IOP Publishing Ltd
Figure 7.7. Future design drivers and material responses. and structural properties. Competitive new designs now require revolution- ary changes in the engine mechanical design, and this is driving materials development in new ﬁelds (for the engine manufacturers) such as electrical rather than mechanical properties. Figure 7.7 shows the future expected design drivers and the material responses. The more-electric engine (MEE) Engines are heavily dependent on mechanical/hydraulic actuation and drive mechanisms to operate pumps, variable nozzles and vanes, etc. These all require mechanical drive systems from the engine core. To increase reliability and reduce cost and weight, these will be replaced with small electric motor drives. This will ultimately require an electrical generator mounted within the engine core in an environment of over 3008C. Permanent magnets necessary for use in this machine will currently only operate up to 2208C, so that the unit will have to be insulated and cooled with air, similar to current turbine structures, until new magnet capabilities are available. Again, this use of cooling air will detract from the self-same improvements in eﬃciency being made by the new technology. Distributed control software With the move to the MEE, control systems will also become more distrib- uted around the engine into smaller units. All control software is currently mounted on electronics in a single heavily-insulated and ﬁreproof case on the outside of the engine. Distribution of the systems around the engine will increase reliability and reduce cost and weight, but for the maximum beneﬁt will require high temperature electronics based on materials such asCopyright © 2001 IOP Publishing Ltd
silicon carbide (silicon based semiconductors cannot be reliably operated above around 1008C). Silicon carbide based electronics systems have been demonstrated at over 4008C for military applications but are currently not cost eﬀective for commercial use. Materials development for gas turbines is ongoing in these ﬁelds and the more conventional areas and is presenting a more diverse set of requirements than ever before. Many of these materials, however, will only ever ﬁll limited niche applications in gas turbines and similar high speed, high temperature machines. As a result perhaps the biggest and most common materials challenge for the future will be the one currently faced by high temperature polymer composites and structural ceramic composites—how to develop and bring to production a highly specialized, low volume material quickly and cost eﬀectively. The answer must lie in collaboration across the industry supply chain and the research and development institutes to simplify processes, to automate aspects of materials assessment, and above all to innovate and develop new, creative solutions in manufacturing, and to then apply these quickly with controlled levels of risk.Copyright © 2001 IOP Publishing Ltd