Aerospace Materials Chapter: 7. Materials Developments In Aeroengine Gas Turbines

Chapter 7

Materials developments in aeroengine gas

turbines

David Clarke and Steve Bold

Introduction

Aeroengine gas turbines can be optimized in different ways for differentapplications, and even in different ways for the same application. Thisoptimization process is driven by a number of broad factors as diverse asthe basic laws of physics and gas behaviour through to aircraft operationalperformance and the financial needs of both the engine manufacturer andthe end user of the aircraft. Through all of this runs a common factor:materials—metals, polymers and ceramics—are key to balancing all thesefactors for achievement of the optimum engine design and, in turn, thedesign 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 majorengine manufacturers all offering engines with similar fuel burn performance,secondary factors such as noise, emissions, weight and reliability become themajor product differentiators. Highest engine efficiency is achieved through ahigh pressure ratio in the compressor and a large temperature rise throughthe combustor. Both factors are limited by the temperature capability ofthe materials available and the cooling technology used. Propulsive efficiencyalso depends on matching the exit velocity of the gas stream to the speed ofthe aircraft. For this reason, civil engines use large bypass ratios so that workis put into moving a larger mass of air more slowly than the air moving

Copyright © 2001 IOP Publishing Ltd

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 highthrust, 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 deliverthe necessary thrust to fly and manoeuvre the aircraft. This is particularlyrelevant 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 aircraftspeed and the need for additional responsiveness, military engines havelow bypass ratios and very high gas exit velocities. Pressure ratios tend tobe lower giving maximum specific work but generally not maximum fueleconomy.

For both civil and military applications there is a common factor—cost.Product development costs and production unit costs are considered fromthe first stages of engine design and create as many challenges for materialsdevelopment as the engine performance parameters.

Two engine designs

The vast majority of civil aircraft are available fitted with engines from morethan one manufacturer. Figure 7.2 shows a real example of the results of two

Figure 7.1. Pressure and temperature cycles through the high thrust, high bypass Trent 800

engine.


different manufacturers’ design solutions for one particular aircraft. Bothengines were designed at the same time for the same duty and whilst bothuse similar basic technologies, produce the same thrust and weigh similaramounts, their design philosophies are quite different.

Physics dictates that engine thermodynamic efficiency increases withincreasing turbine gas temperature and increasing pressure ratio (combustorentry pressure to ambient air inlet pressure). Primarily as a result of runninghotter and at higher pressures, engine 2 burns around 4% less fuel thanengine 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 coreengine components in engine 2 degrade much more quickly than those inengine 1, and engine 2 has to be removed from the aircraft twice as oftenfor major maintenance. The financial impact of this to the operator is thatengine 1, the less technically efficient engine, costs nearly 10% less to operatethan engine 2. The financial impact on the engine manufacturer is that engine

Figure 7.2. Comparison of turbine temperature and pressure ratio for two aeroengines

designed for the same aircraft.

Figure 7.3. Current design drivers and materials responses.


1 now accounts for 80% of this market sector with over 1000 engines inservice.

Engine design is thus clearly a complex optimization process andthere are many solutions to one overall requirement. The impact of thedifferent design styles and operational parameters of these two engines onthe materials needed to manufacture them is equally dramatic. Figure 7.3shows the current design drivers and material responses.

Materials

Turbine blade alloy development

Since the initial development of the gas turbine in the 1940s the temperaturecapability of the nickel alloys used for the highest temperature parts of theengine—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 inmaterial 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 castingswith control of both the longitudinal and the transverse crystal orientation.This development is summarized in figure 7.4. To achieve a further increasein temperature capability is now requiring a fifth stage of development, withthe introduction of surface coatings to reduce oxidation and corrosion, andultimately a complex system of thermal barrier coatings (TBCs) to reduce the

Figure 7.4. Temperature capability of different turbine blades since 1940.


rate of heat transfer from the blade surface to the internally cooled surfacesin its hollow core. This latter development is discussed in more detail later.

An increase of 40 8C in turbine blade temperature as highlighted betweenthe two engines in the example in figure 7.2 denotes a significant increase inmaterials technology level. Typically this represents the difference betweentwo 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 significantfuel burn benefit (4% is around the maximum difference seen between differ-ent engines designed for a single airframe application) it is only of real valueif this can be matched by comparable reliability between the two units. Theneed for a particular level of materials technology or a significant advance isthus created as the fundamental engine architecture and thermodynamiccycle are set. It is obviously possible to operate engines at low temperaturesbut 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 havebeen evolutionary, as in the gradual evolution of each nickel alloy technologydescribed above. Both nickel and titanium alloys, however, are now at a levelof development where further small advances are of limited benefit and areincreasingly expensive to achieve. Engine design now demands revolutionarydevelopments in the fields of increased temperature capability and reducedcomponent mass. The primary materials development programmes toachieve 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 followingsections 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 theengine pressure ratio is effectively being progressively decreased as air isbled off to these various sections. The effect of this is to increase the enginefuel burn compared with the theoretical minimum. In practice this resultsin increased cost for the operator. Whilst the reduction in efficiency as aresult of these cooling bleeds is very small, the overall financial effect to alarge fleet operator may be very significant. A 1% reduction in fuel burnto a typical long haul operator of ten Boeing 747-400s would reduceannual fuel bills by well over $1 million.


To reduce the need for cooling air, many turbine components nowincorporate ceramic materials. The latest developments of high pressure(HP) turbine blades use nickel alloys coated with a graded series of layeredmaterials to give a component which relies for its operation on the synergyof properties from the metal substrate and ceramic coating. The single crystalnickel base alloy is first coated with a plasma sprayed MCrAlY overlaycoating (M represents Ni or Co). The chromium and aluminium provideoxidation resistance while the presence of yttrium improves scale adhesion.In addition the layer acts as a bond coat to prevent spalling of the outercoating layer—the thermal barrier coating. The thermal barrier coating isa low thermal conductivity ceramic which restricts the flow of heat fromthe gas stream to the metal blade. This maximizes the benefit obtainedfrom blade cooling and offers a potential increase in operating temperatureof over 1008C.

Thermal barrier coatings have been used in the combustion chambers ofthe RB211 since 1975 but it is only with advances in bond coat technologyand ceramic deposition techniques that they can be reliably used on criticalrotating parts where coating failure could lead to premature componentremoval. Coatings are applied by electron beam physical vapour depositionto develop the columnar grain microstructure necessary to resist thermal andmechanical 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 asan integral part of the component from the start of the design process. Thecoating system must be compatible with the requirements of aerodynamics,

TBC usage is abalance between :

Material propertiesStrain toleranceOxidation/corrosionLife prediction

and ManufacturingissuesCoating thicknessHole closureCoating structureand integrityProcess control

Trent 800 HP blade + thermal barriercoating

Figure 7.5. Trent 800 high pressure (HP) turbine blade, showing cooling holes and thermal

barrier coating (TBC).


mechanical integrity and blade cooling and must be created in a cost-effectivemanner suitable for mass production. All of these considerations have led tothe unique, cost-effective thermal barrier coating system designed for use inthe Rolls-Royce Trent engines. Thermal barrier coatings are discussed inmore detail in chapter 22.

Nickel/ceramic hybrid combinations are also being developed on a macroscale for mechanically integrated large static structures. Here a simple metallicunit forms the major load carrying element and a semi-structural ceramic shell,or more commonly a fibre reinforced ceramic composite, forms the hightemperature surfaces of the component. This approach has been used inturbine blade tip seals on the Trent engine, and exhaust structures for militaryengines such as the EJ200.

Titanium aluminides and titanium metal matrix composites

Unlike turbine components, compressor components are not normallycooled. They operate at gas stream temperatures which reach around650�750 8 C at the compressor exit. The primary design driver here is toincrease pressure ratio and this largely drives further increases in this exittemperature.

Compressor discs and blades are largely formed from titanium alloys,with much of the historical alloy development emphasis having been ondeveloping defect tolerance and increased temperature capability. Thelatter is exemplified in alloys such as IMI834 which are used in applicationsup to 630 8 C—almost twice the capability of the routine Ti-6Al-4V at350 8C. Moving to these higher temperature alloys has allowed weightsavings in compressor modules of over 15%. Alternatives to titanium areprimarily steel and nickel alloys with the consequent mass increases. Majorperformance improvements are now being made by either using intermetallic(titanium aluminide) materials or silicon carbide fibre reinforced titaniumcomposites.

Gamma titanium aluminides are the most advanced of the intermetallicswith half the density of current titanium alloys. Primary benefits are veryhigh specific stiffness and inherent non-burning chemistry. Temperaturecapability is currently around 750 8 C, making the system suitable for backend compressor and turbine applications. Their non-burning nature makesthem particularly suitable for stator vanes, where titanium alloy use is limitedby the risk of titanium fires. Application of aluminides could lead to wholeengine weight savings of up to 4%. These materials will also give reducedengine life cycle costs and improved engine/airframe functionality sincereducing mass on one element of an engine structure has a consequentknock-on effect on surrounding components—lower-mass blades lead tolower-mass discs and shafts, reduced stiffness casings etc. Titaniumaluminides are discussed further in chapter 17.


Knock-on effects in reducing mass of associated components areparticularly significant in rotating bling (bladed ring) structures made intitanium metal matrix composite (Ti MMC). The bling is an integral struc-ture of titanium blades on a titanium metal matrix composite. The reinforcedring increases hoop stiffness by 100% and strength by 50%. This removes theneed 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 figure7.6. Rolls-Royce Allison has successfully run the first demonstration oftitanium metal matrix composite blings in an engine. Titanium metalmatrix composites are discussed in more detail in chapter 18.

The weight savings in the bling itself are greatly increased by the optionspresented for major changes to the surrounding structures, e.g. the possibilityof using a larger diameter (and hence stiffer) shaft, since the shaft diameterconstraint imposed by the disc bores is effectively removed. This geometricchange opens up the possibility of replacing the steel shaft with a titaniumunit, with dramatic weight benefits. Steel shafts are currently used toensure adequate torsional and bending stiffness, which must be achievedwithin 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 enhancedmaterials capabilities in more diverse fields. Engine design has evolved for thepast 50 years around a largely unchanged group of mechanical technologies,and these have needed a standard set of materials with optimized mechanical

Figure 7.6. Weight savings with blisc and bling designs using titanium metal matrix

composites (MMCs).


and structural properties. Competitive new designs now require revolution-ary changes in the engine mechanical design, and this is driving materialsdevelopment in new fields (for the engine manufacturers) such as electricalrather than mechanical properties. Figure 7.7 shows the future expecteddesign drivers and the material responses.

The more-electric engine (MEE)

Engines are heavily dependent on mechanical/hydraulic actuation and drivemechanisms to operate pumps, variable nozzles and vanes, etc. These allrequire mechanical drive systems from the engine core. To increase reliabilityand reduce cost and weight, these will be replaced with small electric motordrives. This will ultimately require an electrical generator mounted within theengine core in an environment of over 3008C. Permanent magnets necessaryfor use in this machine will currently only operate up to 2208C, so that theunit will have to be insulated and cooled with air, similar to current turbinestructures, until new magnet capabilities are available. Again, this use ofcooling air will detract from the self-same improvements in efficiency beingmade 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 currentlymounted on electronics in a single heavily-insulated and fireproof case onthe outside of the engine. Distribution of the systems around the enginewill increase reliability and reduce cost and weight, but for the maximumbenefit will require high temperature electronics based on materials such as

Figure 7.7. Future design drivers and material responses.


silicon carbide (silicon based semiconductors cannot be reliably operatedabove around 1008C). Silicon carbide based electronics systems have beendemonstrated at over 4008C for military applications but are currently notcost effective for commercial use.

Materials development for gas turbines is ongoing in these fields and themore conventional areas and is presenting a more diverse set of requirementsthan ever before. Many of these materials, however, will only ever fill limitedniche applications in gas turbines and similar high speed, high temperaturemachines. As a result perhaps the biggest and most common materialschallenge for the future will be the one currently faced by high temperaturepolymer composites and structural ceramic composites—how to develop andbring to production a highly specialized, low volume material quickly andcost effectively. The answer must lie in collaboration across the industrysupply chain and the research and development institutes to simplifyprocesses, to automate aspects of materials assessment, and above all toinnovate and develop new, creative solutions in manufacturing, and tothen apply these quickly with controlled levels of risk.


Business

Aerospace Materials Chapter: 7. Materials Developments In Aeroengine Gas Turbines