94•GT-475 Or Printed in Its Publications

THE AMERICAN SOCIETY OF MECHANICAL ENGINEERS 345 E. 47th St., New York, N.Y. 10017 The Society shall not be responsible for statements or opinions advanced In papers or discussion at meetings of the Society or of its Divisions or Sections, 94•GT-475 or printed in its publications. Discussion is printed only if the paper is pub- lished in an ASME Journal, Papers are available from ASME for 15 months after the meeting. Printed in U.S.A. Copyright © 1994 by ASME Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1994/78873/V005T12A012/2405346/v005t12a012-94-gt-475.pdf by guest on 26 September 2021 CURRENT AND FUTURE MATERIALS IN ADVANCED GAS TURBINE ENGINES

G. A. Kool Materials Department National Aerospace Laboratory NLR Amsterdam, The Netherlands 11111111111111111111111 )

ABSTRACT as during development of new, untried aero engine materials. Gas turbine engines are constructed of components with From the operator's point of view, safety and cost are two major excellent strength and stiffness, a minimum density, a high concerns in engine operation. These two points will be discussed temperature capability for long times, and at affordable cost. in more detail. Material behaviour under service conditions and Metallic materials are the centrepiece in fulfilling these material costs must be considered extensively where conventional requirements. Future gas turbine engines will have to have higher materials are replaced and new materials are implemented. thrust-to-weight ratios, better fuel efficiencies and still lower costs. This will require new and advanced lightweight materials Safety with higher temperature capabilities. Engine problems account for only a small share of air transport This paper discusses some of the presently applied materials in accidents [I], namely 12 % between 1959 and 1989 (Fig. 1). the fan, compressor and turbine sections of gas turbines, and Considering the period 1976-1983 for transport aircraft, only a reviews the material developments that are occurring and will be little more than 25 percent of the failures involved discs (Fig. 2). necessary for the near and long term futures. Because of the storage of kinetic energy in disc fragments, disc failures produce the most serious consequences. Of the 52 cases recorded, 12 were classified category three: significant damage to INTRODUCTION . In civil and military aircraft the propulsion is based on the gas turbine engine. In principle the gas turbine ingests air from the atmosphere and compresses it several times in the compressor. engine-related aircraft accidents Fuel is added and the mixture is burned giving a high pressure, 1959-1989 high velocity gas stream. Part of the energy in the gas stream is used to rotate a turbine section which in turn drives the engine lam and inappopriate compressor. However, the largest part or the energy can be used crew response to drive a propeller, a fan, or give thrust by itself. 25% The goals for the gas turbine engines of the 21st century are significantly higher thrust to weight ratios, better fuel efficiencies uncontained and lower life cycle cost. These general requirements translate engine failures into the need for materials with increased strength and stiffness, 48% reduced density, and higher temperature capability for longer maintenance-related times. Major concerns once these materials are available will be human talon 8% the design, development, manufacture, testing and inspection, and the repair of required components at affordable cost. cone= failure modes This paper identifies some presently applied engine materials in N. (inducing simulaneous the fan, compressor and turbine sections, and reviews material losses at thrust on more than developments for the near and far term. one engine) thrust reverser 8% 4%

ENGINE MATERIAL REQUIREMENTS Safety and costs have to be considered continuously during the Fig I Engine problems account for 12 percent of air transport selection and application of current, well-known materials as well accidents during 1959- 1989 Met 13

Presented at the International Gas Turbine and Aeroengine Congress and Exposition The Hague, Netherlands — June 13-16, 1994

uncontained failures commercial transports 1976-1983

122 blades 52 discs 28 spacers

cat. 3 cat. 4

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cal. 2 46 cat. 1 69

20 15 category 1: nacelle damage category 2: minor aircraft damage category 3: significant aircraft damage category 4: severe aircraft damage

Fig. 2 Uncontained engine failures [Ref 1]

Engine components Turbine lost in flight section — Combustion Fan section section ( Compressor section

forward

Accessory drive section

Fig. 3 CF6-6 engine cutaway [Ref. 2]

the aircraft resulting in damage to the primary structure, rapid POLYMER-MATRIX COMPOSITES (PMCs) depressurization, fire, slight injuries to the passengers. Three Extensive experience with composites has been gained over the were classified category four: severe damage to the aircraft past decade. Kevlar aramid fibre containment casings have been ending with crash, loss of the aircraft, serious or fatal injury to standard on CF6-engines since 1980. The outlet guide vanes for the passengers. the CF6-88C2, 88 in number, have been made with graphite Disc separations are usually fatigue-related, whether induced by epoxy since 1985. At present GE is researching a composite engine operating cycles (e.g. start/stops) or high-cycle dynamic forward fan for the 6E90, the GE engine for the next generation modes of the rotor or rotor/stator interactions, sometimes of of jetliners. At Pratt and Whitney, composites are standard on aerodynamic origin. The origin of fatigue cracking and fracture commercial engines for guide vanes and for the ducting where often lies in the design of the part or in material deficiencies. One acoustic liners are required, as in the area of the fan tips. Fibrous Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1994/78873/V005T12A012/2405346/v005t12a012-94-gt-475.pdf by guest on 26 September 2021 of the rare major accidents which was caused by a material defect materials have good damping characteristics for noise. Rolls- occurred in the summer of 1989 in Sioux City, Iowa, USA, [2]. Royce are looking at PMR-15 for core fairings now made with After burst of a CF6-6 fan disc (Fig. 3) from the tail engine of a titanium on R821I-engines, as well as graphite-epoxy bypass DC-I0 the engine component fragments hit the aircraft's hydraulic ducts on Rolls' Tay-engine, used in the Fokker 100. control systems. The disc failure from the tail section engine was Like the GE90 fan, most of today's production composites are followed by an emergency landing of a virtually uncontrollable made with epoxy resins. The epoxy resin serves as the binder, or aircraft, which crashed at the airport 45 minutes after the disc matrix material that holds high-strength graphite (or other) fibres failure. A semi-circular fatigue crack with a radius of 14.5 mm in place. Epoxy-matrix/carbon fibre composites can be used at resulted in the disc burst. Fatigue cracking originated from a relatively low service temperatures to a maximum of 120-150 °C. cavity (1.4 mm long and 0.4 mm deep) in the Ti-6A1-4V disc This is far to low for the more critical, hot segments of a jet bore. This cavity was located in an area with a hard phase engine. The NASA-developed polyimide PMR-15 is capable of inclusion containing microcracks and microporosity. withstanding thousands of hours of use at temperatures between The above illustrates that high standards or even higher ones 290 and 345 °C. Almost all high-temperature parts now in are required for new materials in design, development, production are found on military engines and made of graphite- manufacturing processes and inspections. Safety is not negotiable polyimide composites [4]. GE has used PMR-15 resin for the and cannot be traded in order to improve other factors. main duct of the F404 engine, which powers the F-18 fighter, since 1988. More than 250 engines with PMR-15 ducts have been Costs built. Also in production at GE are PMR-15 air splitters and inner The primary goal of an engine manufacturer and operator is. ducts for the F110 for the F-I6 fighter. Pratt and Whitney is in safe and reliable operation. Ultimately this can be translated into production with a graphite-polyimide aerodynamic exhaust feather satisfied customers and financial benefits. The second objective is on its F100-PW-229 for the F-15 and F-16 fighters. To realize to minimize the operational cost. The main cost factors for long the full advantage of PMCs in aircraft engines, however, new range civil engine operation are [3]: composite materials must be developed with: • faio[fl depreciation, interest and insurance. The - improvements in the stability of polymer matrices coupled capital cost is significantly influenced by the number of spare with improvements in polymer/fibre interfaces engines and modules required to support operation. Factors to _ reduced proressing costs decrease cost are: - oxidation resistant coatings that will enable PMC use at - low removal rates temperatures up to 425 °C - short shop turnaround times Polymer research at NASA Lewis [5] has produced major - pooling of spare engines advances in high-temperature polymer-matrix composites. The new polymer V-CAP, which is given a recently developed • Maintenance costs (II %'r material and labour; both on-wing nitrogen-postcure treatment, has a useful lifetime double that of and in the shop. In general, maintenance costs can be related PMR-I1-50 composite and five times that of PMR-15 (Fig. 4). to the number of components in an engine. In case of KSSU This lifetime is based on 10 % weight loss in air at 370 °C. The CF6 engines three major maintenance cost factors can be PMR-I1-50 is one of the best high-temperature composites indicated: material cost (60 %); labour; and the severity of currently available. operation determined by flight length, start/stops and airport ambient conditions. • QacratignaLsgat jail; fuel consumption and costs V -CAP - 75 • N2 Pa originating from engine-caused delays. This percentage is retkr.ccpex" . ":144:4):::.44:14:14?-ar valid for long range civil aircraft and illustrates: that fuel costs rise dramatically with increased fuel price Polymer matrix it is of extreme importance to maintain or improve engine composites PMR-I1 50 efficiency to decrease fuel consumption For the short and medium range aircraft the contribution of fuel costs to operational costs is less significant. PMR-15 i, V

Environment , i In addition to safety and costs the environmental requirements 200 400 • 600 800 1000 1203 will become more stringent in the near future. Civil regulatory useful If a at 370°C (hours) authorities will require local airports to base their landing fees on noise and polluting emissions in addition to landing weight. Fig. 4 High-temperature capability of graphite-reinforced composites [Ref. 5]

3 TITANIUM Outer Inner form panel form panel Current alloys The combination of high strength, low density, moderate temperature Capability, excellent corrosion resistance and erosion resistance accounts for the wide use of titanium in the fan and compressor section. Titanium now represents as much as 25 % Honeycomb weight of the latest large turbofan engines. An overview of used Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1994/78873/V005T12A012/2405346/v005t12a012-94-gt-475.pdf by guest on 26 September 2021 titanium alloys is given in table I. Main applications include the fan and compressor discs and blades, ducts and casings [6, 7, 8]. One of the most significant recent developments has been the hollow wide chord fan blade introduced by Rolls Royce on the RB2I1-535E4. The blades, made from Ti-6A1-4V (IMI 318), are about 40 % wider in chord than the classic solid blades. Their shape makes them strong enough to resist flutter and fatigue without the support of stiffening features called 'snubbers' at mid- span. The blades have a cleaner aerodynamic shape which make the fan about 4 % more efficient compared to the fan on the 535C engine. This gives a specific fuel consumption 2.5 % lower than Section an equivalent solid blade fan. Fewer blades are needed to achieve through blade the same thrust, 22 compared with 33 on the 535C engine, which in turn make for a lighter fan disc. The fan blade construction, shown in figure 5, involves vacuum brazing the honeycomb core between two panels of titanium which have been formed and chemically etched to generate the correct airfoil shape. Subsequently the two panels are diffusion-bonded at the edges. Fig. 5 Wide chord fan construction (Ref. 6)

TABLE 1 CAPABILITIES OF SOME TITANIUM ALLOYS (6, 7, 8, 12, 131

Titanium,''. , Classification., Yiellstrength at Temperature capability ..0, Fracture '''..,' Density ec,, , M (War, - based oncreep proPertier Toughness lc , .1 (g/cm')

Currently applied titanium alloys Ti-6A1-4V a + g 950 350 55 4.47 (MI 318) Ti-6AI-2Sn-4Zr-6Mo a + 0 1100 400 25 4.65 Ti-4A1-4Mo-2Sn-0.5Si a + 0 950 450 60 4.60 (IMI 550) Ti-8A1-1Mo- I V near a 980 400 45 4.42 Ti-2.2A1-1 I Sn-5Zr-lMo-0.2Si near a 940 450 35 4.84 um 679) Ti-6A1-2Sn-4Zr-2Mo-0.1Si near a 990 520 50 4.54 Ti-6A1-5Zr-0.5Mo-0.25S1 near a 900 520 70 4.46 (IMI 685) Ti-5.5A1-3.5Sn-3Zr-0.3Mo- near a 860 550 60 4.55 INb-0.3Si (IMI 829) Ti-5.8A1-45n-3.5Zr-0.5Mo- near a 950 600 40 4.55 0.7Nb-0.35Si (IM! 834) Ti-15Mo-3Nb-3A1-0.2Si metastable 0 1300 595 - 4.93 (Timetal 21S) Materials under development Titanium MMC 35 vol% SCS-6 1600 600 low 4.5 fibre/Ti-6-4 MMC (L)

Ti-25A1-10Nb-3V-1Mo a2 1000 600 14 4.1 (Ti-aluminide) Ti-24A1-17Nb-IMo a2 1000 600 17 4.2 (Ti-aluminide) Ti-22AI-27Nb 0 + do 1100-1200 650 28 4.4

The most widely used material in engines (and airframes) superior thermal fatigue strength and hot corrosion resistance over continues to be the alloy Ti-6A1-4V. Applications include fan the Ni-base alloys. Ni-base alloys are stronger at low and discs and blades in the RB211 series, the General Electric CF6, intermediate temperatures and have better oxidation resistance and the Pratt & Whitney JT9D, 4000 and 2027 engines. It is also when Al is added as an alloying element. The combustor and used for lower temperature applications in the compressor stages afterburner are currently protected by a ceramic coating These of military engines such as the RBI99. Other alpha-beta alloys thermal barrier coatings (TBCs) typically consist of an oxidation used include IMI 550 (Pegasus and Olympus 593), Ti-6A1-25n- resistant bondcoat and a thermal insulating topcoat, both of which 4Zr-6Mo (Pratt & Whitney FI00) and Ti-17 (Ti-5AI-2Sn-2Zr- are applied by plasma arc spraying. Advanced TBC-systems 4Cr-4Mo) used by GE in the F404. In Europe the most widely consist of a dense oxide-free Ni(Co)CrAlY bond coating and a used creep resistant alloy is IMI 685 for discs and blades in the porous 7-8 wt% yttria-stabilized zirconia Zr0 2/Y203 top coating. Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1994/78873/V005T12A012/2405346/v005t12a012-94-gt-475.pdf by guest on 26 September 2021 R.8211, R8199, Adour, SNECMA M53 and Larzac engines. Ongoing research will lead to improved oxidation resistance of bondcoats and better adherence between substrate and coating [14, Newer developments 15]. Inconel 718 is the most widely used sheet metal for the More recent alloy developments, since the mid-1970s, resulted cooler parts in the combustor and exhaust because of its high in the alloy IMI 829 which has a temperature capability of about strength up to 650°C, good workability and weldability. 550 °C. This alloy is a successor to IMI 685 and is used in the Rolls-Royce RB211-535E4 high pressure (HP) compressor, where Turbine blades and vanes it saves 540 kg over an alternative design in nickel alloy. IMI's An overview of the turbine materials used in the Tay 650, latest alloy, IMI 834, is being introduced into engines now. IMI powering the Fokker 100, is given in figure 6 [16]. In table 2 an 834 has also been seriously considered for use in the low pressure overview is given of the chemical composition and high (LP) sections of the turbine. IMI 834 is stronger than IMI 829 and is capable of long term operation to 600 °C. Alloy Timetal 2IS will be used by Boeing to reduce weight of the engine Tay 610/612/620 Tay 650 exhaust system of its new 777 jumbo jet [9]. The new alloy is ..A Sake ...„."1 C1023 heat-and corrosion resistant and can compete with nickel alloys at 11P1 NGV single impartment -ix9., lain compartment service temperatures up to 595 °C. potty cooling For several years, titanium matrix composites and titanium aluminides have been of great interest to aircraft engine builders 14108 leaped DS cast Mal M002 because of their low density and high-temperature strength HP1 blade sine feed, dual teed multi-pass relative to conventional titanium alloys. Important limitations, singe pass coding however, have been brittleness at room temperature and or low HP1 disc Stainless steel Waspaloy fracture toughness (Table 1). The latest titanium aluminides alloyed with niobium are stronger and tougher than earlier C1023 ____4 sea) Mar M002 developed a 2 (TOD material [10] These alloys have been Pain compartment HP2 NGV mulfrpass cooling studied at General Electric. The new materials have two-phase cooing microstructures consisting of ordered orthorhombic (0) and kz,.. single aystal cast ordered beta (13.). The new Ti-aluminides are potential weight- %,.. equiax Mar M002 saving alternatives to nickel-base superalloys in relatively low HP2 blade uncooled SFIR99 uncooled temperature (650 °C) applications. Examples are: exhaust-nozzle structures and compressor components. The density of the HP2 disc Stainless steel Waspaloy Ti2AINb-based alloys is less than two-thirds that of the widely used Inconel 718. The alloys are not limited by high temperature creep resistance. Rather they are sensitive to oxygen penetration Fig. 6 Tay 650 - Improved HP turbine materials and cooling and oxidation [111. These barriers must be overcome through (Ref. 16] further alloy and coating development before widespread use is expected. In fact, it is unlikely that titanium aluminides will see service much before 2010. The incentive to develop new high temperature titanium alloys 1800 and titanium aluminides comes mainly from their much lower 1700 density compared to superalloys: 4.1-4.8 Went', table 1, versus 1600 7.7-8.9 gran', table 2. 15C0 Turbine 14 00 inlet temp. SUPERALLOYS (°C) 1300 1203 Combustor end exhaust 1103 Both, Ni- and Co-base alloys are used in gas turbine low combustors and exhaust segments. For the hottest parts like 900 combustion chambers and afterburner liners with metal temperatures up to 1100 °C, the Co-base alloy Haynes 188 has 1960 1970 1980 1993 2000 2010 proven its outstanding performance in service for several years. Year of iritroduction The widely used Co-alloy is a solid solutioning strengthened alloy containing 22 16 chromium. In general Co-base alloys have Fig 7 Evolution in turbine blade cooling technology (Pet 17)

5 TABLE 2 NOMINAL COMPOSITIONS OF CAST NICKEL-BASE SUPERALLOYS USED IN TURBINE BLADES [17, 19, 20, 21]

Alloy Composition; wt% Approximate year Temperature i Density, , of introduction capability, C g/cm Cr Co AI Ti Mow Nb Ta Zr B Other

Conventionally ca t Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1994/78873/V005T12A012/2405346/v005t12a012-94-gt-475.pdf by guest on 26 September 2021 1N-718 0.05 19 - 0.5 1.0 3.0 - 5.0 0.01 0.005 18Fe 1965 700 8.2 IN-713C 0.05 12 5.9 0.6 4.5 2.0 0.1 0.010 - 1955 985 7.9 IN-100 0.18 10 15 5.5 4.7 3.0 0.05 0.015 1.0V 1958 1000 7.7 Rene 80 0.17 14 9.5 3.05.0 4.0 4.0 0.03 0.015 1965 1000 8.2 IN-738LC2 0.11 16 8.5 3.4 3.4 1.7 2.6 0.9 1.7 0.05 0.010 1970 980 8.1 IN-9392 0.15 22.5 19 1.9 3.7 2.0 1.0 1.4 1.1 0.100 1973 970 8.2 IN-620P 0.03 20 20 2.4 3.6 0.5 2.3 1.0 1.5 0.05 0.800 1978 1010 8.2 MAR-M 246 0.15 9.0 10 5.5 1.5 2.5 10 1.5 0.05 0.015 1966 1025 8.6 Direct. onall solidified MAR-M200 Hf 0.15 9.0 10 5.0 2.0 12.5 1.0 0.05 0.02 2.0Hf 1970 1040 8.6 MAR-M002 DS 0.15 9.0 10 5.5 1.5 10 2.5 0.05 0.015 1.5H1 1975 1045 8.6 IN-62032 0.15 22 19 2.3 3.5 2.0 0.8 1.1 01 0.01 0.75Hf 1981 1020 8.2 Single crystal PWA 1480 10 5.0 5.0 1.5 4.0 12 1980 1060 8.7 CM SX-2 73 4.7 Si 1.0 0.6 8.0 6.0' 1980 1070 8.6 SRR-99 0.015 8.5 5.0 5.5 2.2 9.5 2.75 1980 1080 8.5 AM 1 7.5 6.5 5.2 1.2 2.0 15 8.6 1985 1090 8.6 PWA 1484 5.0 10 5.6 2.0 6.0 8.7 - 3 Re 0.1Hf 1986 1100 8.9 CM SX-40 6.2 9.5 Si 1.0 0.6 6.5 6.5' 2.9Re 0.1Hf 1986 1110 8.8 MC 2 - 8.0 5.0 5.0 1.5 2.0 8.0 6.0 1990 1125 8.6 SC 162 16 3.5 3.0 3.5 1990 1030 8.2

/ 100 hr to rupture at 140 Mpa. 7) High-chromium alloy suitable for land and marine-based gas turbines. 3) Indicates combines Nb + Ta.

50 temperature creep performance of superalloys used for turbine blades and vanes. The 100 hr/I40 Mpa creep rupture strength for superalloys has increased at an average rate of 10 °C per year 40 from about 1940 to 1980. Specific aeroengine fuel consumption .621A, has been halved, thrust has been increacPd 50 fold, and thrust-to- PW2037 1 1 '92500 r 30 weight ratios have increased by a factor of ten during this period, #M,L;S; owing largely to the increases in alloy performance. The alloy .:eJT9D-7R4 PW4000 performance improvements are significantly pushed by Specific fuel ;Pr consumpfion 20 improvements in casting techniques, coatings and internal cooling improvement design (Figs. 7, 8). (7 ) Turbine inlet temperatures in military aircraft engines are close 10 to 1580 °C currently, and there is hope for 1700 °C before the turn of this century [171 Consequently, turbine airfoil alloys in baseline modern gas turbines must satisfy multiple property requirements. advanced The airfoil section of a turbine blade operating in the hot gas single blade crystal stream must have favourable combinations of creep strength, 0/20 thermal fatigue resistance and oxidation and hot corrosion r- 1960 1970 1980 1990 resistance; whereas the root attachment and extended neck Certification date regions, protected from the gas stream by the platform, run considerably cooler but with higher centrifugal stresses. These Fig. 8 The improvement in fuel efficiency with engine model and cooler regions of the blade must possess favourable combinations the contributions made by cast superalloy technology of monotonic and cyclic properties, with emphasis on tensile yield [Ref. 18] strength and notched low-cycle fatigue properties. To meet the hot

6 section turbine demands, alloy compositions have become more strength and ductility usually are the life-limiting criteria for complex. Matrix strengthening elements such as tungsten and polycrystalline turbine blades, the removal of this constraint made molybdenum are now added. It was discovered that the latter it possible to use very strong compositions with a high volume element made it possible to increase the amount of the 7'-forming fraction of 71 . elements, aluminium and titanium. The 7'-forming elements will The single crystal airfoils were first cast by Pratt & Whitney in give high volumes of strengthening intermetallic phases resulting the mid-1960s. Several nickel-based superalloys were especially in improved creep strength. designed for single crystal casting, such as PWA 1480 and Directional solidification (DS), pioneered by Pratt & Whitney, CMSX-2 (USA), 5RR99, RR 2000 (GB), AM1 and AM3 produced a further jump in the strength and temperature (France). The use of such single crystal alloys has led to capabilities of superalloy castings. This improvement is attributed temperature capabilities of about 80 °C higher than that of Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1994/78873/V005T12A012/2405346/v005t12a012-94-gt-475.pdf by guest on 26 September 2021 to two important characteristics of a directionally solidified conventionally cast polycrystal superalloys such as IN 100. The microstructure: favourable properties of single crystal alloys in terms of life are - the elimination of transverse grain boundaries given in figure 10. Single crystal PWA 1480 turbine blades first - the achievement of preferred, rather than random, grain entered service in the Tf9D-7R4 engine in 1982, powering the orientation such that the direction of maximum lattice strength Boeing 767 and Airbus A310, and in military engine applications coincides with the stress axis of the blade or vane. (18]. The alloying elements in PWA 1480 such as Al, Ti, W and Further improvements in the grain boundaries were made possible Ta were added to obtain 60 vol. % of 7'. PWA 1480 obtains its by the addition of hafnium. excellent oxidation resistance from its high levels of aluminium For directionally solidified blades (and by analogy single and tantalum in combination with a coating. Among the developed crystal substrates) Young's modulus is reduced by about 40 % in alloys in France the industrial gas turbine alloy SC 16, aimed to the <001> direction (spanwise for blades). Thermal fatigue is replace IN-738LC, and the aerospace alloy AM I alloy are worth the result of strains introduced by thermal expansion and mentioning [19, 20]. The AM1 alloy is now being used in the contraction. This is readily seen from the relationship do = SNECMA M 88 military engine which powers the RAFALE a.E.ST, where a is the thermal expansion coefficient. However, fighter plane [17]. the a in the <001> direction is greater than the average a. The significantly lower E and a slightly larger a give a significant reduction in plastic strain during an engine thermal fatigue cycle. 9x NO p olycrysal This results in a marked improvement of six- to tenfold in thermal '; Ara columnar-crystal fatigue resistance [18]. single crystal The development of the single crystal casting process (Fig. 9) 7x Relative made it possible to eliminate grain boundaries from parts and to We use simplified alloy compositions in which grain boundary 5x 1 strengthening elements were unnecessary. Because grain boundary 1

SS V 1 induction t lumace molten metal lx I. • creep strength thermal fa gue resistance corrosion resistance crucible Fig. 10 Comparative properties of polycrystal, columnar and single crystal superalloys [Ref. 18]

More recently, second generation nickel based single crystal superalloys have been introduced [21]. The higher strength single crystal alloys, such as CMSX-4 and PWA 1484, were developed furnace in the USA These two alloys contain 3% rhenium and have a 30 °C use-temperature advantage over first generation single crystal alloys such as PWA 1480. The addition of 3 % rhenium provides solid solution strengthening and permits higher Al+Ti contents, which in turn produce a higher volume fraction of 7'. Rhenium also reduces the coarsening rate of 7' by decreasing the fusion kinetics at the matrix/y' interface. The strongest non- heated hold rhenium containing single crystal alloy MC2 [17], developed by single crystal selector ONERA in France, competes with the USA-developed alloys PWA 1484 and CMSX-4. It should be noted that besides developments in blade materials, water cooled plate leading to the latest single-crystal alloys, there have been developments in blade cooling techniques. Most recently there have been developments in single crystal casting to enable the withdrawal transpiration cooling technique to be used for blades (this technique has been known for more than 20 years, but only for Fig. 9 A process schematic of a single crystal casting Met 18) less advanced materials and applications).

Turbine discs overview about nonmetallics written by Sims [22]. He uses in this Besides blades and vanes, the introduction of powder article the superalloy toughness, 16 to 22 MPa.m 1/4 , as the metallurgy disc materials has required a combination of alloy benchmark acceptable toughness for hot-stage use of ceramics and design and process development technologies. Alloy developments other nonmetallic materials. for conventionally cast-wrought nickel-base auperalloy discs are restricted by the excessive chemical segregation and forging difficulties associated with the levels of alloying additions needed

for significantly improved tensile, creep and low cycle fatigue Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1994/78873/V005T12A012/2405346/v005t12a012-94-gt-475.pdf by guest on 26 September 2021 strength. Powder processing has overcome these problems and has Ititanium super. led to the development of Astroloy, Rene' 95 and IN 100 having 100 aluminium alloys alloys 00 proof strengths some 50 % higher than earlier superalloys. alloys 50 stainless alloy steels SO Intermetallic compounds (IMCs1 Intermetallic compounds being investigated, some of which have near-metallic properties, include nickel and niobium -1 )in'iun,1 / //, 0 aluminides (NiAl, Ni 3A1 and NbAI), dicobalt niobide (Co2Nb) and Fracture 10 - SIC spinet molybdenum disilicide (MoSi2) [22]. NiAl is receiving toughness (Kt ). toughened b 02 considerable attention, but still has brittleness problems. If NiAl ksi ko/2 5 SiC rn 513 N4 becomes commercially successful (which is not assured), it will Si3 N4 M2 03 be limited to a use temperature of about 1200 °C. Co2Nb is competitive with superalloys in terms of tensile properties, and SIC recently the compound MoSi 2 has been alloyed with SiC, yielding interesting results.

0.5 - CERAMICS - 0.5 The evaluation of ceramics for use in gas turbines started 20 glass years ago. Extensive research and an expenditure of hundreds of 5 10 millions of dollars have demonstrated that the high temperature Density (gicm3 ) structural (load bearing) ceramics cannot be incorporated into aircraft or industrial gas turbines. The key reason is a lack of Fig 11 Fracture toughness between 16 and 22 MPa . m 1/2 is ductility for the ceramics, see table 3. Figure 11 illustrates how required in materiels for use in gas turbine engines; ceramics and ceramic composites fall short in toughness compared ceramics and ceramic-matrix composites currently fall with superalloys. For this section, many data are used from an short of this requirement [Ref 221

TABLE 3 OVERVIEW OF CERAMICS AND CERAMIC MATRIX COMPOSITES 122, 231

, Material Strengtkat:RT Strength at 1200 °C ' Fracture toughness K„ at , Density Remarks (MPa) - - (MPa) RI , ' (g/cm) (MPa.m a) Monolithic ceramics SVI, 300-400 (tensile) 100 4-5 3.2 (stntered bonded) 800-1000 (flexural) 700 SiC 450 (flexural) 450 3-4 3.1 (sintered a) SiA1ON 5-7 3.1 MoSi2 300-700 (flexural?) 200-600 3 6.3 Crystallographic texture influences properties significantly Zr02 . 8 6.0 (Toughenedl• Ceramic Matrix Composites (CMCs) 20Sic/MoSi, 400-600 (flexural) 400 (flexural) 8 16 Fracture toughness K„ at 1400 °C is 12 MPam" 30SiCKSCS6)/Si,N4 500-600 (tensile) 12 Sic/A1,03 200 (tensile) 100 3.8 400-500 (flexural)

8 Monolithic nonoxide ceramics Metal-matrix composites (IVIMCs) Extensively investigated ceramics are silicon nitride (Si 3 N,), Metal matrix composites (NIMCs) are not new materials. SiC and Si-A1-0-N. Many process variations of Si,N, were Oxide-dispersion strengthened (ODS) alloys consisting of a studied, and entire small-turbine hot stages were built. However, dispersion of oxides, such as thoria and yttria, in superalloy brittle failure of these turbines was more the rule than the matrices have been in existence for • some time. Greater exception. Some successful applications of Si 3N4 have been temperature capability is sought in MMCs reinforced with small- achieved in turbochargers of automobiles (Nissan Motor Co.) diameter SiC, A1 203 and IMC whiskers and fibres. The two major using small (6 to 8 cm in diameter) monolithic Si 3N4 . The rotors problems associated with this approach are: reach maximum speeds over 100,000 RPM. This application is a - coefficient of thermal expansion (CTE) mismatch landmark achievement but the operating temperature is limited to - chemical reactivity between fibres and matrix. Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1994/78873/V005T12A012/2405346/v005t12a012-94-gt-475.pdf by guest on 26 September 2021 only about 600 °C. A ceramic rotor is still considered a luxury A drawback of the [MC whiskers and fibres is again their lack of item in terms of costs. Daimler-Benz AG has operated a Si 3N4 toughness. turbine rotor at a temperature of 1100 °C in short-term tests (less It can be concluded that the potential of nonmetallics to replace than 100 hours) and up to 50,000 RPM (12.7 cm in diameter). 'heavy' superalloys and to provide higher operating temperatures Despite these development successes Si 3N4 is not used in load- is far from being realised. Oxide ceramics could be an exception, bearing parts in aircraft or industrial turbines. The reasons are: but will require long-term development. - Si3N4 loses its tensile and creep strength at temperatures above 1200 °C. - Its brittleness makes it susceptible to catastrophic failure. REVOLUTIONARY STEPS Most importantly, its inherent protective Si0 2 film is Materials are the key to increasing the performance of aircraft ineffective above 1500 °C; this operating temperature is gas turbine engines. Estimates have been made that 50 % of the readily achievable using moderately cooled commercial improvement in performance will come from improved materials superalloys. and processes. Reduced leakage contributing another 25 % will Oxygen contamination of Si3N4 during service further also rely heavily on better materials [24]. In the USA, the embrittles the material. Integrated High Performance Turbine Engine Technology It is worth noting here that the performance problems of Si,N, are (IHPTET) initiative was launched in the 1980s with the overall characteristic of many other silicon containing ceramics. goal of doubling turbine propulsion capability, that is a 20:1 thrust-to-weight ratio, by the year 2000. In the UK there is a Monolithic oxide ceramics programme with a similar target. The programmes include: The best known oxide ceramics are A1 203 and Zr02 . Both - increasing turbine inlet temperature to over 2000 ° C provide satisfactory service in all types of gas turbines as the - reducing the density of materials used in the hot section from common protective films on superalloy pans: A1 203 as a 8 glee to 5 gice protective film on blades and vanes, and Zr0 2 as a thermal eliminating component cooling. bather coating (TBC). The protective oxide films operate at temperatures ranging from 1200 to 1400 °C, and the oxide IHPTET and other programmes include material capability tests ceramics show superior high temperature surface stability. The in advanced demonstration engines. GE presented results last major hurdle to overcome now is to improve strength and year; however, details are classified. GE reported that meeting toughness of oxide ceramics for structural applications. the demonstration engine requirements needs revolutionary materials [25], such as: Ceramic-matrix comoosites (CIVICS! - MMC discs SIC fibre-reinforced Si 3/41 and A1 203 began to appear after it - CMC turbine blades and exhaust parts became apparent that monolithic ceramics had little potential for - lightweight, high-temperature intermetallics use in full-size turbine applications at very high temperatures. - 370°C use-temperature PMC casings and static structures Again the presence of Si in most of these composites means that - TiAl composite blades protection against oxidation depends on Si0 2. This in turn - ceramic bearings and dry lubricants. restricts service to temperatures below 1500 ° C. Capability and potential market estimates are frequently The toughening mechanism involves fibre pullout from the unrealistic, particularly in the field of advanced ceramics. The matrix. This pullout functions as a crack stopper. However, fibres application of ceramics for critical hot section components could unbonded to the matrix (which is generally the case) significantly require very long-term development. Continuation of such reduce the potential strength of the composite as a whole. Test research must be reconsidered if material costs become data generally show that CMCs are not nearly as strong as excessively high. The task to develop an oxide ceramic having superalloys, or other nonmetallic materials. They seem to be thermal stability at temperatures to 1650°C for use as a TBC unacceptable for use in turbine applications. could be far more realistic than to develop a ceramic-matrix composite for use as a load-bearing material in turbine engines. Recent research claims that SiC,/MoSi 2 composites are strong, ductile and more oxidation resistant than the current best high- New materials will remain the centrepiece to more powerful, temperature materials at temperatures up to 1800 °C. Data lighter and more fuel efficient aero engines. Although many more comparing this composite with Ni- and Co-base superalloys years of research and testing are needed, the aero engines will see appear to indicate a 300 °C use-temperature advantage [23]. intermetallics, MMCs and CMCs substituting partly for today's However, too little data are currently available to make a fair titanium and nickel alloys. comparison.

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