1-s2.0-1044580395000666-main

Lightweight J-l? Immarigeon, J. C. Beddoes*

ELSEVIER

Materials for Aircraft Applications R. T. Holt, A. K. Koul, L. Zhao, W. Wallace, and

NRC Institute for Aerospace Research, Ottawa KlA 0R6, Canada and *Department of Mechanical and Aerospace Engineering, Carleton University, Ottawa KIS SB6, Canada

Reducing structural weight is one of the major ways to improve aircraft performance. Lighter

and/or stronger materials allow greater range and speed and may also contribute to reduc-

ing operational costs. The article reviews some recent developments in lightweight mate-

rials for airframe components (aluminum alloys, composites, and hybrid materials) and

engine components (titanium aluminides and titanium-based composites). Emphasis is

placed on microstructural characterization and the relationship between the microstructure

and mechanical properties of specific materials systems being investigated at the National

Research Councils Institute for Aerospace Research.

INTRODUCTION

Improvements in the performance of aircraft have been closely linked to progress in materials. Breakthroughs such as those associated with the development of lightweight precipitationhardenable aluminum alloys, heat-treatable titanium alloys and, more recently, fiber-reinforced polymer (FRP) composites have made it possible to reduce aircraft structural weight significantly. This has enabled aircraft to fly farther and faster and has made commercial flying more econom- ical by allow-ing larger aircraft and by low- ering fuel consumption. The emergence of still lighter, stiffer, stronger, less fatigue- sensitive, more damage-tolerant, and/or more heat-resistant materials is expected to continue to play a significant role in the development of next-generation aircraft, both in terms of airframe and engine developments. For instance, weight reductions of 40-60% have been forecast for early-Zlst- century aircraft based on anticipated use of high-strength titanium alloys and metal-

matrix composites (MMCs) with oriented continuous fiber reinforcement [l] . Further- more, a growth in engine performance in excess of 50% is expected from the introduction of FRPs, WCs, and ceramic-matrix composites (CMCs) in engines and their containment structures [2], and low-density titanium aluminides have the potential to allow 50% weight reductions in some applications [3].

The most effective way to reduce aircraft structural weight is to reduce the density of structural materials. By comparison, increasing ultimate tensile strength, modulus of elasticity, or damage tolerance properties is 3-5 times less effective, with an increase in fracture properties being the least effective option, as explained by Ekvall et al. [4]. This makes FRPs particularly attractive for aircraft applications, because of the good combination of low density, high stiffness, and reasonably high strength that characterizes this class of materials. Over the last three decades, FRPs have been used in increasing quantities in aircraft structures, substituting

Presented at the International Metallographic Society Symposium on Microstructural Characterization of Light- weight Materials for Transportation, Montreal, July 24-2.5, 1994.

41 MATERIALS CHARACTERIZATION 35:41-67 (1995) 0 Elsevier Sciene Inc., 1995 655 Avenue of the Americas, New York, NY 10010

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42 J-l? lmmarigeon et al.

Table 1. Selection of Materials (Aluminum Alloys and Epoxy Composites) for the Canadair Global Express, Compared with the Canadair CL600 Challengera

Global Express CL600 Component (1994) Challenger (1977)

Fuselage skin C188-13 2024-T3 Lower wing cover 2024-T351 (bus) 7475-T7351 (ims) Upper wing cover 7150-T7751 (bus) 7050-T7651 (ims) Leading edge 6013-T6 2024-T3 Vertical stabilizer 2024-T3 All metal Horizontal stabilizer TGE camp. All metal Fairings TGE camp. Kevlar epoxy camp.

n Private communication, K. C. Overbuy, Bombardier Inc., Canadair, 1994. bus = built-up structure; ims = integrally milled skins; and TGE camp. = toughened graphite epoxy composite.

for metals in many secondary structures and, in a few isolated cases, as material for primary structures [5].

The first aircraft designed in Canada to use a polymeric composite for primary structure will be the new long-range business jet named Global Express. This aircraft is to be built by Bombardier Inc., Canadair, and will make extensive use of toughened graphite epoxy (TGE) which will be used for the horizontal stabilizer, the rudder, elevator ailerons, flaps, fairings, and so on. The Global Express will contain considerably more epoxy composite than the CL600 Challenger business jet which was designed by Canadair in the mid-1970s. It is interesting to compare structural materials in the CL600 with those expected to be used in the Global Express (K. C. Overbury, Bombardier Inc., Canada& Private Communication, 1994). As can be seen in Table 1, the fuselage skin and wing covers for the Global Express will still utilize aluminum alloys; a notable change in the Global Express is the proposal to use the new 2000 series aluminum alloy C188T3 for the fuselage skin. This new alloy developed by Alcoa takes advantage of proprietary processing to provide greatly improved fracture toughness and fatigue crack growth rates over the incumbent 2024-T3 without loss of strength or corrosion resistance. Most of the other aluminum alloys being specified for the Global Express are traditional alloys in traditional temper conditions. The alloy 6013-T6 is used in the bare polished condition on the wing leading edges because it has very good corrosion resistance.

Recently, a number of new metallic products and materials-processing technologies have also been developed which allow metallic structures to be highly competitive [6]. The new lighter metallic materials include low-density aluminum alloys, some capable of service temperatures well beyond the capabilities of polymer-based materials, aluminum-matrix composites (AMCs), and hybrid polymer-metal composites for airframes, as well as titanium aluminides and titanium-based metal-matrix composites (TMCs) for engine applications. New advanced synthesis and processing technologies [7] are also being developed to produce these new high-performance alloys and composites in a cost-effective manner. The techniques include advanced melting practices, mechanical alloying, powder metallurgy processing, spray forming, and vapor deposition, among others.

This article reviews the state of the art in lightweight metallic materials for aircraft applications. It also highlights some of the on- going developments in materials and process technologies that are being pursued in Canada by the National Research Councils (NRC) Institute for Aerospace Research in collaboration with others, for both airframe and engine applications.

ALUMINUM ALLOYS

Aluminum alloys have been the most widely used structural materials in aircraft for several decades. Three types of alloys make up

Aircraft Applications 43

the bulk of th.e aluminum found in modern aircraft. They are the 2000 series (Al-Cu-

Mg), the 6000 series (Al-Si-Mg), and the 7000 series alloys (Al-Zn-Mg-Cu). All are precipitation-hardenable alloys that rely on the precipitation of fine coherent precipitates and dispersoids for strengthening, the morphology and distribution of which dictate mechanical properties and environmental response of the materials. Their microstructure is controlled by heat treatment and they can be produced in a variety of microstructural conditicons, or temper, which allows specific design requirements to be met.

New alloys and engineered materials are emerging that have the potential to replace the conventional 2000,6000, and 7000 ingot metallurgy products. They are the low- density aluminum-lithium alloys, the powder- metallurgy-processed 7000 series alloys, the aluminum-based MMCs, and metal-polymer hybrid composites. Details on the properties, metallurgical. characteristics, and the use, or potential use, of both the conventional materials and the new products are provided below.

CONVENTIONAL 2000, 6000, AND 7000 SERIES ALLOYS

The most widely used alloy of the 2000 series is 2024-X$ which takes advantage of cold working followed by natural aging. The alloy has moderate yield strength but good damage tolerance (good resistance to fatigue crack growth and good fracture toughness). In the form of thick sheet, however, it is susceptible to exfoliation corrosion. The alloy is used mos,tly for fuselage skins, usually clad with a layer of pure aluminum for corrosion protection, and is found in most of the commercial and military transport aircraft built over the past 30 years. Strength is derived from the formation of a high- volume fraction of Guinier-Preston (GP) zones and coherent 0 (CuMgAlz) precipitate phase in the grain interiors, as well as by the presence of Al-Cu-Mn dispersoids. When alloyed with iron and nickel, the 2000 series alloys have reasonably high creep strength, as typified by alloy 2618 used in the Concorde, the Anglo-French supersonic aircraft. Addition of copper provides good

combinations of strength and ductility as well as good weldability, as typified by alloy 2219 (used for NASAs space shuttle external fuel tank and components of the shuttle re- mote manipulator system, Canadarm) .

The 6000 series alloys have much better corrosion resistance than the 2000 series alloys. The latest alloy of the 6000 series is 6013, which in the T6 temper offers a 12% strength advantage over alclad 2024-T3, with comparable toughness and resistance to fatigue crack growth and with the added advantage that it can be used bare, that is, without cladding. The 6000 series alloys also have excellent fabricability. However, they are not as widely used by the aircraft industries as the other alloys because they do not com- pete in terms of overall balance of properties.

The 7000 series alloys, of which 7075 has been the most widely used, have the highest strengths by far. The alloys are produced as either sheet, plate, forgings, or extrusions. They are used for fuselage skins, stringers, and bulkheads, as well as for wing skins, panels, and covers. Their strength is derived from the precipitation of n phase (coherent MgZn2) in the grain interiors and n phase (noncoherent MgZnz) along the grain boundaries. In the conventional peak age (high strength) condition (T6), the thick plates, forgings, and extrusions of the 7000 series alloys are highly susceptible to stress corrosion cracking (SCC) [8], particularly when stressed through the thickness, a shortcoming which has been well docu- mented for 7075-T6 [9]. Many theories have been developed to explain the susceptibility to SCC [lo]. Many of these regard hydrogen embrittlement to be an important factor and grain boundary precipitate size may also be important.

Considerable efforts have been expended over the years - and are still going on-to ad- dress the problem of SCC in 7000 series alloys. These efforts first led to the development of the overaged T73 temper for alloy 7075. With the T73 temper, the threshold stress for SCC of 7075 is increased by a factor of 6. The penalty is a 15% loss in yield strength [S, lo] relative to the T6 condition. Intermediate tempers (T74 and T76) have been developed to provide trade-offs in strength and SCC resistance between the

44 J-P. lmmarigeon et al.

Table 2. Effects of Impurity Content on. the Typical Mechanical Properties of Alloy 7049 Extrusions (After Petrak [20])

Alloy and Si temper max.

Fe max.

MTl max. % El

7049-T73511 0.25 0.35 0.20 504 552 11.6 26.2 7149-T73511 0.15 0.20 0.20 520 568 13.3 32.6 7249-T73511 0.10 0.12 0.10 529 579 13.3 37.2

YS = yield strength; UTS = ultimate tensile strength; El = elongation; Kt, = plane strain fracture toughness; and L = longitudinal.

T6 and T73 conditions. At about the same time, a technique known as retrogression and reaging (RRA) was also developed for reducing susceptibility to SCC in 7000 series alloys while maintaining the high strength of the T6 condition. The technique involves manipulation of grain boundary precipitate size in material initially in the T6 condition by multistage heat treatment. It was first applied to alloy 7075 by Cina [ll, 121 and has been extensively studied at the NRC from the late 1970s to the mid-1980s [13-191. Today, there is renewed interest in the use of RRA heat treatments to achieve corrosion- resistant tempers and particularly in treatments that could be applied retroactively to fully manufactured and service-exposed parts. More on this subject is provided in a separate section.

Demands for higher strength, coupled with good fracture toughness and SCC resistance, led to the development of several new 7075 derivative alloys (7175, forgings; 7475, mill products) and other alloys (7049, 7050). This progress resulted largely from tighter control over impurity levels but also from improvements in thermomechanical and heat-treatment practices such as T74, T76, and T77 tempers. Table 2 shows how reductions in the levels of silicon, iron, and manganese lead to significant gains in yield strength, tensile strength, and fracture toughness for alloy 7049-T7351 which, in effect, translates to chemical limits applied to alloys 7049,7149, and 7249 (D. Richardson, Lockheed Georgia, Private Communication, 1993). This information is derived from work by Petrak [20].

NEW 7000 SERIES ALLOYS

More recently, the drive for still higher- strength alloys without undue loss in SCC resistance or fracture toughness has led to further alloy developments at Alcoa (7150, 7055) with a new patented temper [21-231 registered as T77 by the Aluminum Associa- tion. Not much is said in the open literature about the proprietary process used to produce this temper for the new high-strength alloys. From the patent literature, the process is known to be a three-step aging treatment which may involve combinations of high and low aging temperatures, not unlike those associated with RRA processing, and possibly an intermediate mechanical working step. Table 3 compares the mechanical properties of the two new alloys in the T77 condition with the properties of 7075-T6. In the patented T77 temper, alloy 7055 offers

Table 3. Effect of Heat Treatment on the Typical Mechanical Properties

of 7000 Series Aluminum

Alloys

Alloy and temper

0.2% YS CL) UTS (L) Krc CL-T) (MPa) CMPa) % El (MPaJm)

7075-T6 475 555 11 25 7075-T73 451 510 13 30 7050-T76 503 541 14 - 7150-T77 553 585 11 - 7055-T77 592 603 11 22 7055-T76 537 569 13 32 7055-T74 467 523 17 40

Aircraft Ap,vlicr;rtions 45

in forgings the highest strength of any commercially available 7000 series alloys, with excellent combinations of SCC and fracture toughness capabilities [24]. The alloy is also available in the T74 and T76 tempers. The T76 temper -provides SCC resistance and fracture toughness comparable to 7175-T74 with at least !5% higher strength. In the T74 temper, the fracture toughness of the alloy is equivalent to that of 7050-T74 with a strength advantage of at least 6%. Finally, the data on fatigue properties obtained so far by Alcoa indicate that alloy 7055 in all three tempers will be at least at par with earlier high-strength 7000 series alloys [24].

The new 7000 series alloys are being considered for new and future generation aircraft as well as for the replacement of 7075-T6 components in existing aircraft. McDonnell Douglas has started using 7150-T77 plate for the upper wing skins and 7150-T77 extrusions for the upper stringers of the C-17 transport aircraft. Boeing is planning to use 7055-T77 for the upper wing skins and other components of the Boeing 777. Other potential aircraft applications for the high-strength 7055-T77 alloy include wing spars, web ribs, aircraft wheels, and landing-gear links [24].

PROCESSING OF 7000 SERIES ALLOY COMPONENTS FOR IMPROVED STRESS

CORROSION RESISTANCE

With the aging of commercial and military aircraft fleets, the susceptibility to SCC of the widely used aluminum alloy 7075-T6 has become a growing concern to users, manu- facturers, and regulatory bodies because there is still a large inventory of equipment in service containing this alloy. Experience has shown that aircraft components made from 7075-T6, tend to corrode rapidly, particularly in aiXri3ft operating in marine environ- ments. Excessively corroded parts must be replaced for obvious safety reasons and the replacement costs are high.

A numbe:r of options exist for replacing SCC-prone aircraft components to extend component lives and minimize operating costs [25]. When stress levels permit, the lower strength but significantly higher

corrosion-resistant 7075-T73 can be substituted for 7lY75-T6. Alternatively, an alloy having better corrosion resistance and mechanical properties similar to or better than 7CY75-T6, for instance, 7050-T74 or 7050-T7651, can be substituted for 7075-T6. Both of these options have been considered, and in some cases already implemented, by Lockheed, for instance, for the CC-130 Hercules transport aircraft (T. Ginter, Lockheed Aeronau- tical Systems Company, Private Communi- cation, 1992). A third and possibly more cost-effective approach might be to subject 7075-T6 components to ERA. This treatment could be applied to new parts from the existing stock of replacement parts or, as a refurbishment procedure, to components that have not yet corroded badly enough to war- rant replacement and are worth salvaging.

The RRA treatment was first described by Cina [ll, 121 as a process applicable to 7075 in the T6 condition to achieve stress corrosion resistance equivalent to the T73 temper while maintaining high strength characteristic of the T6 condition. The Cina process involves short time exposures to high temperature, typically l-100 s in the range from 200 to 280C (the retrogression) followed by reaging using the original T6 aging treatment (16-24 h at 120C). During retrogression, the hardness or yield strength falls to a minimum before increasing to a secondary peak, while continued treatment causes a further decrease in strength, as shown schematically in Fig. 1. At the same time, stress corrosion resistance increases as revealed by

FIG. 1. Schematic of the retrogression and reaging process showing the effect of retrogression on yield strength or hardness [13].


electrical conductivity measurements (ex- pressed as a percentage of the International Annealed Copper Standard [% IACS] [25]) and by SCC tests using bolt-loaded double- cantilever beam specimens [ 131. According to Cina, reaging from the initial minimum, that is, at the end of stage I (point A, Fig. l), restores strength and hardness in excess of the original T6 condition (point B, Fig. 1). Also during reaging, the conductivity and the stress corrosion resistance continue to increase toward T73 values. Since the retrogression times are very short, of the order of tens of seconds, the originally formulated Cina process is for all practical purposes limited to very thin section parts, typically l- to 6mm thick.

Subsequent work spearheaded by Wallace and co-workers [ 13-151 demonstrated that lower temperatures and longer retrogression times (up to point C, Fig. 1) could be used to produce more effective combinations of strength and stress corrosion resistance in thicker sections. Retrogression times em- ployed were typically several minutes at 220C or up to 2 h at 180C. After reaging, tensile properties equivalent to 7075-T651 properties (point D, Fig. 1) and SCC growth rates similar to those for the overaged 7075- T73 were achieved for 25mm (lin.) plate. The work was subsequently extended to 25mm (lin.)-thick specimens cut from63mm (2.5in.) 7475 plate material. RRA processing gave tensile strength equivalent to 7475-T6 and corrosion properties much better than T6 and almost as good as 7475-T7351 [16, 171. In other work [l&19], it was shown that RRA processing of 7075 has little effect on fatigue or corrosion-fatigue crack growth rates. However, material in the T6RRA or T7351 tempers appears to be marginally more resistant to fatigue crack growth in the threshold region at low frequencies in corrosive solutions. This is indicative of good corrosion- fatigue resistance. The yield strength, the electrical conductivity, and the SCC crack growth rate (CGR) for alloy 7475 extrusions in various heat-treat conditions are compared in Table 4. Each of the RRA conditions (3 h at 180C and 5 minutes at 220C) resulted in higher yield stress than the T651

Table 4. Properties of Aluminum Alloy 7475 Extruded Bar in the T651, T7351, and T6RRA (3 h at 180C and

5 minutes at 220C) Conditions [17]

Temper

T651

T7351

T6RRA/180-180

T6RRA/5-220

0.2% Conduc- YS tivity SCC CGR

(MPa) (% IACS) (10-4mm h-l)

490.5 33.6 773.7

405.8 42.2 5.0

498.6 40.9 -

505 39.5 18.6

condition. The conductivity provides an in- dication of corrosion resistance: the higher the conductivity, the greater the corrosion resistance [ 171.

The effects of RRA processing on 7O75-T6 microstructure have been investigated by transmission electron microscopy [ 14,15,26, 271. Results from these studies indicate that RRA processing is primarily a coarsening treatment for the grain boundary precipitates. Retrogression, through stages I and II (see Fig. 1), leads to full or partial disso- lution of the finer 1 precipitates (MgZnr) while the size and volume fraction of the coarser n and r\ precipitates increase slightly [26, 271. At the same time, retrogression in- duces rapid coarsening of the grain boundary precipitates (mainly q). Reaging from stage II, or slightly beyond the end of stage II, leads to precipitation of fme n in the grain interiors, minor growth of the partially dis- solved II, and agglomeration of the grain boundary precipitates into closely spaced coarse particles of n [14]. Figure 2 compares the microstructures of 7075 plate in the T6RRA, T651, and T7351 conditions and shows that the RRA processing produces a microstructure comprised of fine matrix TJ + n precipitates (MgZnr), characteristic of the T651 temper, with coarse grain boundary precipitates of n representative of the T7351 temper. These features are believed to be responsible for the combination of high strength and SCC resistance of the T6RRA condition.

It has been suggested that the coarse grain boundary precipitates, characteristic of the


FIG. 2. Transmksion electron micrographs of 7075 plate in the (a) T651, I:b) T7351, and (c) T651RRA conditions [14, 191.

T6RRA and T73 conditions, reduce susceptibility to SCC by acting as trapping sites for hydrogen [ 141. There is supporting evidence to show that hydrogen, entering the metal at the crack tip by hydrolysis, tends to con- dense at the trapping sites into molecular gas bubbles [28,29]. This in turn would reduce the concentration of atomic hydrogen along grain boundaries ahead of the crack

tip. Atomic hydrogen is believed to cause embrittlement by forming weak hydrogen- metal bonds along the grain boundaries, and the condensation of hydrogen into molecular gas bubbles would, therefore, reduce the number of such bonds. Metallographic evidence of hydrogen gas bubbles attached to grain boundary precipitates in transmission electron microscopy thin foils of 7075- T6RRA has been reported [14, 291.

The beneficial effects of RIG&type treatments on resistance to SCC and exfoliation of 7000 series alloys have been confirmed in numerous studies [ll, 12,30-331. As just noted, the technique is particularly well suited to the treatment of thin section parts. In thick section parts, only the surface microstructure would be modified by processing, to a depth of up to 25mm, depending on geometry. This may be sufficient to con- fer improved SCC resistance where it is most needed, that is, close to the surface, while the bulk of the part would remain in the high-strength T6 condition.

ALUMINUM-LITHIUM ALLOYS

Among the new aircraft materials, aluminum- lithium alloys are particularly attractive because of their weight-saving potential. When aluminum is alloyed with lithium, for every 1% addition of lithium, there is approxi- mately a 3% reduction in alloy density and an increase in stiffness of about 6%. The commercial alloys typified by 2090, 2091, 8090, and 8091 contain from 1.9-2.7% lithium. Therefore, they offer up to about 10% density advantage over the 2000 and 7000 series alloys. They also have correspondingly higher stiffness and offer a 25% advantage in specific stiffness. With aluminum-lithium alloys, weight saving in aircraft structures of up to 10% is possible in strength-critical structures and of up to 18% in stiffness- critical structures [34]. In addition to being light and stiff, the alloys are strong, damage tolerant, and corrosion resistant. However, their properties are strongly sensitive to processing conditions and, therefore, product quality is more difficult to control than for conventional alloys. Other shortcomings in-

48

elude high anisotropy of unrecrystallized products caused by the strong crystallographic textures developed during processing, low short-transverse properties of thick plates, lack of thermal stability of some products, limited experience with manufacturing requirements, and limited amounts of design data [35]. Because of these limitations, together with the very high cost of material and the added expense for the recycling of scrap, the impact of Al-Li alloys on the aircraft industry has fallen short of initial expectations.

Aluminum-lithium alloys were introduced more than 30 years ago by Alcoa as alloy 2020 for use on the RA-5C Vigilante military aircraft. Outside of the U.S.S.R., where several alloys were developed in the 196Os, the technology appeared to lay dormant until the mid-1980s, when Alcoa, Alcan, and Pechiney introduced alloys 2090, 8091, and 2091, respectively, and Alcan and Pechiney jointly introduced alloy 8090. During the late

J-l? lmmarigeon et al.

198Os, a mechanically alloyed powder metallurgy product known as Al-905XL (formerly IN 9052) was also introduced by IncoMAP and, more recently, a cast Al-Li-Cu alloy known as Weldalite 049 (now registered as alloy 2095) was developed by Martin Marietta Laboratories [36]. The latter material has excellent weldability, superior to that of the 2000 series alloys, including alloy 2219, and is a strong contender as fuel tank material for NASA& space shuttle because of the materials excellent cryogenic properties.

The 2000 and 8000 series Al-U alloys are available commercially in a variety of forms and tempers which can be selected to meet the specific design requirements of either high strength (e.g., 2090-TSX, 809%T8), medium strength combined with corrosion resistance and damage tolerance (e.g., 8090- l%XXX, 2091-T8X), or high damage tolerance (e.g., 209%T8XXX) [35]. The commercial alloys normally have strongly developed textures resulting in strong anisotropy of

Table 5. Minimum Properties (AMS Standards) of 2090 and 8090 Al-U Alloys Compared to Conventional Sheet and Plate Aluminum Alloys

FO?Yn Alloy

Tensile 0.2% AMS strength YS EXCO &CT-L) Density

number Orientation (MPa) (MPa) % El min. (MPadm) (g/cm3)

Sheet 2024-T3 up to 6.25mm 2024-T861

7475-T61

4037 4193 4084

L, LT 441 290 15 L, LT 490 455 4 L 517 455 9 LT 517 441 9 L 490 420 9 LT 490 415 9 L 517 483 4 LT 503 455 5 L, LT 396 303 4

L, LT 441 290 12 L, LT 483 441 4 L, LT 455 393 6 L 530 475 10 LT 540 460 10 L 490 414 10 LT 490 414 9 L 517 483 4 LT 517 469 3

NS NS

NS

NS

EB

NS 2.77 NS 2.77 66 2.80

7475-T761 4085 88 2.80

2090-T83 4251 EB 2.60

8090-T6 4259 EB NS 2.53

KICK-T)

NS NS

26.4 33

Rate 2024-T351 4037 up to 12.7mm 2024-T861 4193

2124-T851 4101 7475-T651 4090

NS NS

NS NS

2.77

2.77

2.77

2.80

7475-T7351 4202 NS 42 2.80

2090-T81 4303 4303

EB 27.5 2.60

NS = not specified; EXCO = exfoliation corrosion; EB = rating B for the EXCO test.

Aircraft Applications

strength and fracture properties. Strength and fracture are also strongly influenced by grain size and structure [37]. Strength is derived from precipitation of 6 (AlsLi), Ti (Al;?CuLi), S (A12CuMg), 8 (Al#Zu), and other phases. Volume fraction, distribution, and morphology of the precipitates depend on alloy composition and processing conditions (extent of hot and cold working and heat treatment). The presence of coherent shearable 6 precipitates is conducive to in- homogeneity of slip resulting in severe strain localization during deformation. This local- ized slip planarity strongly influences fracture properties, being beneficial in terms of resistance to fatigue crack growth rate (FCGR) but detrimental to toughness [37]. It has also been suggested that toughness is strongly -influenced by the level of alkali metal impurities and the occurrence of liquid metal embrittlement caused by these impurities [38].

Minimum mechanical properties for design purposes, taken from Aerospace Mate- rials Specifications (AMS) standards for two commercial Al-U alloys 2090-T8 and 8090-T6, are compared to conventional alloys 2024 and 7475 in Table 5, which shows that strength levels of 2090-T8x sheet and plate are almost #equivalent to those of 7475-T~xx, with about a 7% saving in weight. How- ever, 7475-1651 plate has better guaranteed plane strain fracture toughness (Kt,) than does 209OT81 plate. Note that, for sheet, plane stress fracture toughness K, is quoted, and K, is always greater than Kr,. The two aluminum.-lithium alloys also exhibit much higher resistance to the growth of fatigue cracks when compared to conventional alloys. This is illustrated in Fig. 3, which shows that their FCGRs in the L-T orientation (i.e., in compact tension specimens loaded in the rolling direction with cracks propagating in the transverse direction) are 1 order of magnitude lower than the FCGRs of conventional alloys [39]. The crack growth rate data in Fig. 3 are compensated for material differences in yield strength and work-hardening coefficient, as well as for differences in the R ratios used to collect the data. This is done to separate the effects of texture from those

: A

103- i

: v _ 0

102: +

2024-T351 221 g-T651 7075-T651 7475-T761 8 2090-T61 OV

6090-T6771 n -

FIG. 3. Fatigue crack growth rates in conventional high- strength aluminum alloys and the aluminum-lithium alloys 2090-T81 and 8090-T8771 in the L-T orientation 1391.

arising from differences in materials properties and experimental techniques. The compensated data reveal that the large differ- ence in FCGR is due primarily to texture [39]. Alloy 8090-T8771 possesses a strong (110) brass-type texture with (110) crystallographic planes lying in the rolling plane and [112] poles in the rolling direction.

The superior FCGR resistance of Al-Li alloys can be traced to a tortuous crack path resulting from their highly planar slip characteristics and the presence of texture. In alloy 8090-I8771 tested in the L-T orientation, cracks are found to follow well-defined slip planes and to meander about the T direction over a significant fraction of the crack path [39]. This tortuosity increases the energy required to propagate the crack and increases crack closure and crack tip shielding. Both of these factors impede crack growth. It also results in well-defined slip band facets on the fracture surface, as revealed by frac- tography. Metallography of sections taken through the specimens indicate that the angle of intersection between the facets de-

J-l? Immarigeon et al.

FIG. 4. Fatigue crack profiles in a compact tension specimen of L-T orientation: (a) sections parallel to the rolling plane and (b) sections across the thickness of the plate, normal to the transverse direction (see Fig. 5 for schematic view) [39].

pends on the orientation of metallographic sectioning. For sections parallel to the rolling plane, the facets are about 109.5 apart, whereas for sections across the thickness of the plate, normal to the transverse direction, the facets are about 60 apart, as shown in Fig. 4. These observations can be rationalized in terms of simple crystallographic consid- erations, from which it can be shown that the angles of intersection between the (111) favorable slip planes in sections parallel to the rolling plane and across the thickness of the plate, normal to the transverse direction, agree with the measured angles (Fig. 5)

[391. Applications of Al-U alloys are not wide-

spread to date. Alloy 8090-T83 is used in limited quantities by Airbus Industries, for the D-nose skins of the leading edge of the A330/

FIG. 5. Three-dimensional schematic showing the rela-

tive orientation of favorable slip planes in a compact tension specimen in the L-T orientation [39].

340 aircraft wing. Alloys 2090-T83 and 2090- T62 are used by McDonnell Douglas for some flooring sections in the C-17 airlifter craft. The new Boeing 777 aircraft makes only limited use of Al-U alloys [40]. In con- trast, Westland-Agusta, U.K. /Italy is unique in making extensive use of 8090 forgings and sheets and 2090 and 2091 sheets for the EHlOl helicopter. The alloys are also being tested for a variety of new applications, including lower wing skins and fuselage applications (panels and doors) [35].

NEW ALUMINUM ALLOYS

Over the last 10 years, new processing techniques have led to the introduction of new aluminum alloys with greatly enhanced properties over conventionally processed ingot metallurgy (IM) products. The techniques include powder metallurgy (P/M) processing in conjunction with rapid solidification (P/M-RS) or mechanical alloying (P/M-MA), spray forming, vapor deposition [7], and XD processing.

Powder-Processed (P/M-RS and P/M-MA)

Materials

Aluminum has a limited ability to alloy with more than a handful of elements. These elements are those found in the conventional cast alloys (essentially Mg, Zn, Cu, Si, and Li) just discussed. RS more than triples that


number, which vastly extends the possibil- ities for eithler solid solution, precipitation, or dispersion hardening [41]. Additional benefits derived from RS are finer grain sizes and more uniform microstructures. Mechanical alloying is an alternative to RS processing for extending solubility limits, refining microstructure, and producing non- equilibrium phases. The powders obtained by RS or MA are consolidated into usable products by extrusion (or hot isosatic pressing) and this may be followed by rolling or forging, depending on the desired product. Both P/M-F!S and P/M-MA have been explored as alternative processing routes for 7000 series alloys, as well as to produce lower-density Al-Li alloys and new alloys with improved high-temperature properties [7].

The first-generation wrought P/M alloys derived from the conventional 7000 series alloys were Alcoas 7090 and 7091 and Kaisers P/M61 and 64 alloys, which were introduced in the early 1980s. Alcoa has since developed a new highI-strength P/M alloy designated X7093 (formerly CW67) with notably improved fracture toughness over early-generation P/M alloys. As shown in Table 6, X7093-T7E92 die-forging material has excellent strength and fracture toughness and the alloy has also been reported to have good fatigue and corrosion properties [42]. Alloy X7093 is available commercially as both extrusions and forgings and with SCC resistance equivalent to 7075-T73; and it is being promoted as a replacement for 7075-T6 die forgings. De- tailed information on the properties of this new material, including standardized properties, is currently being developed.

The lithium content of IM Al-Li alloys is limited to ;!.7% in commercial alloys. Above about 3% Li, serious loss in toughness occurs in wrough.t IM products. By virtue of RS effects on solid solubility limits, powder processing allows the Li content to be increased up to 3.6% of Li. The result is reduced density, much better corrosion resistance, and mechanical properties that compare favorably with conventionally processed 7075 products [7]. Mechanical alloying adds the benefits of dispersion strengthening by

Table 6. Mechanical Properties of Some New Aluminum Alloys Including Powder Metallurgy (P/M) Alloys

Alloy and temper

0.2% YS UTS % KI

Ref. CMPa) (MPa) El (MPaJm)

Weldalite 049-T8 36 692 712 5 NA P/M X7093-T7E92 42 615 637 9 47 P/M 5091 (Into) 25 455 595 9 25 X8019 (Al-Fe-Ce) 43 393 414 NA NA

NA = not available

oxides. The surface oxide film present on the aluminum powder feed stock is broken up and incorporated into the interior of the MA powder as a fine dispersion. Corrosion properties very much superior to those of 7075-T73 material have been reported for experimental MA Al-Li alloys [7]. Into alloy 5091 (Al-Mg-Li) is a non-heat-treatable mechanically alloyed material containing 1.2- 1.4% of Li, exhibiting a respectable balance of properties roughly equivalent to 7075-T73 [25]. The properties of this alloy, as well as those of Weldalite 049 just described, are also shown in Table 6.

P/M-RS has also been taken advantage of to produce alloys with good high-temperature strength and thermal stability. This is achieved by alloying with transition metals (e.g., Fe, MO, or V) and rare earth elements (e.g., Ce). Alloying with these elements is not possible with conventional IM because of these elements propensity to form coarse brittle intermetallic phases during ingot solidification. P/M-RS promotes the formation of a high-volume fraction of these phases in a uniformly distributed sub-micron-scale morphology which remains stable at high temperatures and confers excellent high- temperature mechanical properties to the materials [41]. Two of the most promising alloy systems that have been developed for aircraft applications are the Al-Fe-Ce-based system developed by Alcoa and the Al-Fe-V- Si-based system developed by Allied Signal [43]. Both systems are available in all the conventional product forms: sheet, plate, extrusions, and forgings. In sheet, specific

52 1-P. lmmarigeon et al.

strength is typically equivalent to 2024-T8 up to about 150C. Above this temperature, however, the two alloy systems provide a significant strength advantage over 2024-T8 (50% at 260C after 100 h exposure) [43]. Plane strain fracture toughness, Kr,, is of the order of 22MPadm in thick-section extrusions and forgings and FCGR characteristics are similar to those of 7075-T6. Their resistance to surface corrosion is equivalent to that of 6061-T6 and their SCC threshold is of the same order of magnitude as that of 7050-T7451. Alloy X8019 (Al-8Fe-4Ce) is the latest of the high-temperature alloys from Alcoa designed for applications in and around engines, both reciprocating and tur- bines [42]. These high-temperature aluminum alloys could be used as substitutes for titanium alloys in moderate-temperature applications, up to 300C and possibly beyond.

Spray-Formed and Vapor-Deposited Alloys

Spray forming is typified by the Osprey spray-deposition process. In spray forming, molten metal is fed through an atomizing nozzle into a retort filled with inert gas. At the exit of the nozzle, the molten metal is atomized by a high-velocity inert gas jet into a fine spray of metal droplets which is di- rected onto a collecting surface [44]. The process is amenable to the production of solid billet shapes for further processing by extru-

sion, hot isostatic pressing, rolling, or forging. It can also be used to produce powders. Spray forming has been used by Alcan Cospray Products Division [45] to produce monolithic alloys for standard compositions (2618,7075, and 8090) and AMCs based on silicon carbide reinforcement (2014 MMC, 2618 MMC, 8090 MMC, 6061 MMC, 7075 MMC, and 7049 MMC). By virtue of the rapid solidification rates associated with splat quenching of the molten droplets, materials of standard compositions are produced with much refined microstructures and more uniform second-phase distribu- tions. Their tensile (longitudinal direction) and fracture toughness (L-T and T-L orien- tations) properties are accordingly superior to those of the conventionally processed alloys, as shown in Table 7. Note that, for the Cospray MMC materials, neither the reinforcement nor the volume fraction of reinforcement is identified in the company literature [45]. Spray forming has also been used to produce Al-Li alloys containing 4.2% of lithium (alloy UL40), a level which is well above the limits permitted by conventional IM techniques.

Recently, a process known as REFORM (for reactive spray fomzing) has been developed by Perma, a research and development company in Montreal, for producing alloys or composites from molten metal and reactive gases. In the REFORM process, a vari- ant of the spray-forming process, atomiza-

Table 7. Tensile and Fracture Toughness Properties of Spray-Deposited Aluminum Alloys and AMCs [45]

Material 0.2% YS

CMPd

UTS

(MPaJ % El E

(GPaJ

KI, ~MPdnd

L-T T-L

7049-T6 576 614 11.3 69.9 - - 7049-T6 MMC 598 643 2.8 90.1 -

7075-T651 572 617 10.5 71.7 38.9 29.4 7075-T651MMC 556 601 3.7 94.9 - -

7075-T7351 458 529 11.7 71.8 45.2 29.3 7075-T73 MMC 357 458 6.5 92.9 - -

8090-T651 512 550 5.1 79.2 46.4 17.8 8090-T651MMC 499 547 3.0 100.8 14.1 -


tion takes place in the high-velocity flame of a p1asm.a torch within which the molten metal is made to react with a gas to produce an alloy for spraying. Preliminary work has shown that it is possible to produce aluminum composites reinforced with a fine dispersion of TiAls by using aluminum powder and Tic& gas as reaction precursors [46]. Al-TiAls colmposites made by P/M-MA have been shown by others [47] to exhibit at high temperatures very attractive mechanical properties that are significantly superior to those of P/M-RS Al-Fe-X high-temperature alloys.

High-rate electron-beam evaporation is another method for producing bulk materials with novel microstructures by vapor quenching onto a temperature-controlled collector. The concept of bulk vapor deposition has been used to produce new high-temperature alloys comaining normally insoluble elements, such as Fe or Cr. Alloy RAE 72 (Al-7.5Cr-3_.2Fe) is one such material which exhibits exceptional specific strength, higher than that of Ti-6Al-4V, at temperatures up to 300C. ljtrength is derived from solution strengthening by chromium, precipitation hardening by fine (3-5nm) particles of Fe3A1, and an ultrafme grain size arising from rapid deposition from the vapor onto the collector [48].

53

XD-Processed Materials

XD processsing is a patented technology developed by Martin Marietta Laboratories for the production of particle reinforced materials. The process relies on the intrinsic thermodynamic stability in metals of ceramic particles, and their known ability to increase strength and stiffness when present in sufficiently small size (20%), and homogeneous distribution [49]. With this technology, it is possible to produce composites composed of a wide variety of matrix materials (e.g., Al, Cu, Ni, Ti, and intermetallics) and second-phase particles (e.g., borides, carbides, or nitrides). Because the dispersoids are produced in situ, the metal-particle interface is clean, which is beneficial to toughness. XD process-

ing is an inexpensive and generic technology which has been used to produce a variety of particulate-reinforced metal systems, including aluminum-based metal-matrix composites (AMCs) as discussed next.

Aluminum-Based Metal-Matrix Composites

A number of AMCs have been recently developed for aerospace applications and include products reinforced by discrete ceramic particles or whiskers, known as DRA composites (for discontinuously reinforced aluminum) as well as products reinforced by continuous ceramic fibers (known as FRA for fiber-reinforced aluminum) [50]. The fabrication methods for DRA and FRA composites are varied. Alcoa, for instance, has recently introduced a series of DRA composites reinforced by low-cost silicon carbide (Sic) particles, produced by P/M techniques [51]. Some specific Alcoa DRA products are X208O/SiC/15p (where the powder alloy X2080 is reinforced with 15 vol.% silicon carbide particulate), X7093/SiC/15p, 61131 SiC/15p, 6113/SiC/2Op, and X80191SiC112.5p. The latter has excellent high-temperature properties.

Advanced Composite Materials Corpora- tion (ACMC) has also been producing a class of powder-processed DRA composites known as SXA@ composites. These composites are available in all the common product forms for alloys 2009, 6061, 6013, and 7475 reinforced with up to 30 vol.% of Sic whiskers or up to 55 vol.% of Sic particulate [52]. Extruded SXA 2009/SiC/15w-T8 is being field tested on C-141 aircraft as a main landing- gear actuator strut material to improve fatigue life and corrosion resistance over the current material, 7075-T6 [53]. Alloy 2009 is registered with the Aluminum Association specifically for use in composites [51]. At 1992 prices, 2009 composite sheet material has been said to be more cost effective than polymer-based composites for stiffness- critical applications [53].

DRA composites have also been produced by techniques such as stir casting [54], squeeze casting [55], spray forming [44,56, 571, and XD processing [49]. A great attrac-


tion of the latter process is that XD composites, once engineered, can be recast, extruded, or forged as dictated by component requirements. This makes the approach highly cost effective. Similarly, as regards the fabrication of FRA composites, several processing routes have been demonstrated. FRA composites can be produced by invest- ment casting or diffusion bonding using fiber prepregs in the form of green tapes, plasma-sprayed aluminum tapes, or woven fabric [50]. The AVCO hot-molding technique is one approach which is best described by drawing a parallel with autoclave processing of graphite epoxy laminates. Alu- minum prepregs consisting of a single layer of fibers that have been plasma sprayed with aluminum are staked against a ceramic tool of the desired shape, in the desired orientation sequence. Laminate consolidation is achieved by vacuum bagging using a metallic vacuum bag. The technique has been used by AVCO to produce Zee stiffen- ers for aircraft panels [50].

The properties of DRA and FRA composites are strongly influenced by the matrix characteristics and the matrix alloy can be selected to meet specific application needs. The 2000 series alloys offer strength and damage tolerance, the 7000 series alloys offer higher strength potential, and the 6000 series alloys are conducive to good corrosion resistance and improved machinability, whereas the Al-Fe-X (8XxX) alloys provide opportunities for high-temperature performance [58]. Meanwhile, a matrix based on Al-Li provides a unique combination of high stiffness and low density. Composite properties are also strongly influenced by the type of reinforcing medium. Fibers provide the highest stiffness, strength, and toughness combination . Particulate reinforcement is often used for wear-resistance applications and offers high stiffness but only low strength and low toughness, whereas whisker reinforcement offers high stiffness, medium strength, and low toughness [50]. Several reinforcing mediums have been used for AMCs, including alumina, carbon, and Sialon fibers, but Sic is the most common reinforcing medium.

0 200 400 600 BOO loo0 1200 1402 lea0

FIG. 6. Effect of volume fraction and type of reinforcement on the mechanical properties of 6061 composite material [59].

The extent by which reinforcement with Sic influences properties of 6061 aluminum alloy is shown in Fig. 6 as a function of the nature and volume fraction of the reinforcement medium (particulate, whisker, and fiber reinforcement) [59]. The trends are typical of AMCs. Stiffness and tensile strength are greatly increased and toughness is improved but transverse properties can be low in FRAs because these properties are dominated by matrix properties and are similar to those of the unreinforced alloy. Cross- plying may be applied to achieve multiaxial reinforcement for more in plane isotropy or properties tailored to special design requirements.

Work is under way at Ceramics Kingston Ceramiques Inc. (CKCI) of Napanee, On- tario, to optimize a low-cost Sic whisker production technique for composites application. Whiskers produced with this proprietary technique have been studied by metallography [60, 611. The whiskers have an aspect ratio as large as 50 and transmission electron microscopy evaluation reveals they are l3 Sic single crystals with a large number of twinning faults in the (111) plane, which is a characteristic of this type of product (Fig. 7). Tests performed on 208O/SiC/lO-12, composites containing the CKCI whiskers have demonstrated the potential of these whiskers as a reinforcement agent for aluminum alloys. The test material for this exploratory work


800 , ,120

55

FIG. 7. Transmission electron micrograph of a sili-

con carbide whisker produced by CKCI [61].

was obtained by a proprietary process involving blending of the whiskers and alloy powder in an aqueous solution and slip casting the mixture into billet form, followed by drying, hot pressing, and extrusion to rod shapes [62]. Tensile tests were performed in air at room temperature and the results revealed quite attractive mechanical properties for the composites with ultimate tensile strength up to 690MPa, yield strength up to 525MIa, 4.5% elongation, and with a modulus increase over the base alloy of 25% to about 1100GPa. Figure 8 shows that the tensile properties of the CKCI composites (solid symbols) compare favorably with those of commercial AMCs (open symbols), as discuss,ed elsewhere [62].

HYBRID COMPOSITES

Hybrid composites are FRP-metal sandwich laminates consisting of alternating layers of high-strength aluminum alloys and fiber- reinforced. epoxy adhesive. This hybrid structural material, illustrated schematically in Fig. 9, was developed in the late 1970s at

$ 110 g 600

tat 100

E 2

3 L 400 80 f

g Y

f 80 z

v) 200 P

Yield (pwtlcutate) 3

Z! 70 4

> n E B

0 60 0 5 10 15 20 25

Whisker or particulate content (volume %)

FIG. 8. Tensile properties of aluminum alloy 208O/SiClxxw-T4 and 208O/SiC/xxp-T6 composites as a function of whisker or particulate content [62]. UTS = ultimate tensile strength and E = modulus of elasticity.

Delft University in the Netherlands and Fokker Aircraft and was later commercialized in collaboration with Alcoa and Akzo [63]. Two categories of hybrid composites are available commercially today, the ARALL@ and GLARE@ laminates, which differ in the type of fiber used for reinforcement [64]. ARALL laminates (for aramid reinforced aluminum Iaminate) use 50% fiber volume of adhesive prepreg of high-modulus aramid fibers. GLARE laminates (for glass reinforcement) are unidirectionally or biaxially reinforced with 60% fiber volume of high- strength glass fibers. GLARE laminates are a more recent development, complementing the original ARALL product through provi- sion of higher compression strength.

Both ARALL and GLARE laminates come in different configurations ranging from two layers of aluminum with one FRP layer in between, to five layers of aluminum with four interspace FRP layers. In GLARE laminates, the glass fibers can be layed up in a cross-ply configuration. Also, both ARALL and GLARE laminates can be fabricated with different aluminum alloys. This allows laminate properties to be closely tailored to component design requirements. The laminates are produced by curing in a heated platen press. After curing, ARALL laminates can be stretched to eliminate undesirable resid- ual stresses. Stretching greatly increases resistance to fatigue crack growth [64].

J-l? lmmarigeon et al.

FIG. 9. Schematic representation of a 312 FRP-metal sandwich laminate [64].

FRP-metal laminates have the ability to impede and self-arrest fatigue crack growth, which makes the materials highly damage tolerant. As cracks develop in the aluminum face sheets, fiber bridging across a propagating crack causes the unbroken fibers to carry increasing portions of the load, which may decrease the stress intensity at the crack tip to the point where the cracks cease to grow. This makes the material particularly well suited to applications requiring good fatigue resistance. ARALL laminates are best suited to tension-dominated fatigue applications. The GLARE laminates, because of their higher compression strength, can also be used for tension-compression applications. High strength and high stiffness are other attractive attributes of FRP-metal laminates. The specific strength and specific stiffness of four different types of ARALL laminates and other aircraft materials, including 2024-T3, 7075-T6, Al-Li alloy 2090, and carbon-fiber-reinforced polymer (CFRP) are compared in Fig. 10. The FRP-metal laminates properties fall well above those for the alloys but below those for the CFRPs. The spread in ARALL properties arises from differences in the properties of alloys used in the different types of ARALL laminates. ARALL 1, 2, 3, and 4 are based on alloys 7475-T6, 2024-T3, 7475-T76, and 2024-T& respectively. The stronger GLARE laminates also based on 7475-T6 or 2024T3 offer weight savings that are competitive with CFRP composites but at lower production cost than for CFRPs [64].

ARALL was introduced to service in 1988

by the Douglas Aircraft Company for the aft cargo door of the C-17 airlifter. ARALL laminates have since been test flown in the form of underwing inspection hatches on a Fokker 50, as a laminate-skinned inboard flap on the DeHavilland Dash .8 aircraft, and as retrofit materials for C-130 wing-flap lower skins and T-38 dorsal fuselage panels. Areas where the fiber laminates can be used are fatigue-critical and fracture-critical locations where static strength, stiffness, fatigue, damage tolerance, and fracture toughness are required, such as lower wing skins and fuselage skins. GLARE laminates are particularly well suited for frrewall applications because of a high burn-through resistance.

TITANIUM ALLOYS

Titanium-based alloys are another group of lightweight materials which can be used to

0 2 3 4 5

5peclfic Modulus (E I p, 106 m)

FIG. 10. Comparison of the specific strength and specific stiffness of ARRAL laminates and other lightweight aircraft materials [64].


increase the strength-to-weight ratio in structures and provide heat resistance with weight savings in engines. Major thrusts in titanium alloy developments have been to increase thle strength and temperature capabilities of the alloys so that they could com- pete with the much heavier steels in airframe structures and much heavier superalloys in engines. Having a lower density and modulus than steel, titanium alloys enable weight savings of up to 70% and potential volume savings of 50% [65]. Therefore, it is not sur- prising that titanium usage in commercial aircraft ha:s increased dramatically, from less than 1% of total weight in the Boeing 707 of the mid-1950s to roughly 10% of total weight in the new Boeing 777 [66, 671. The titanium content of McDonnell Douglass commercial and military aircraft, including the C-17 airlifter and the MD-11, averages about 9%, which compares with a figure of 70% for aluminum [67]. In some military aircraft, titanium usage has reached considerably higher levels, including a stand-alone 95 wt.% in the recently retired SR-71 Black- bird spy plane [65]. The SR-71, which was developed in the 196Os, was made almost entirely from titanium. Meanwhile, the temperature capabilities of basic titanium alloys have also been increased significantly, from around 400C in the early 1960s [68] to about 62OC today [69], allowing significant improvements in engine thrust-to-weight ratios. Emerging low-density titanium aluminides are promising even greater opportunities for growth in engine performance.

BASIC ALLOYS

Three classes of titanium alloys are used in aircraft: t.he near-alpha alloys (e.g., Ti-6242 and IMI-834), the alpha-beta alloys (e.g., Ti- 6-4 and Ti-6-22-22S), and the beta alloys (e.g., Ti-10-2-3 and Timetal21S). Here, alpha and beta refer to the low-temperature hexagonal- close-packed and high-temperature body- centered-cubic phases of titanium, which are present in titanium alloys in proportions and morphologies that depend on alloy composition and heat treatment. Control of the alpha-to-beta proportions is obtained through additions of alloying elements that

FIG. 11. Typical microstructures in near-alpha Ti-6242S alloy: (a) fme-grained duplex alpha-beta worked and heat treated, and (b) coarse-grained lamellar beta transformed.

preferentially stabilize either one or the other of the two phases. Meanwhile, phase morphology (and to some extent volume fraction) is controlled through heat treatment by phase transformation and precipitation reactions. Most structural alloys are two- phase materials which can be produced in a variety of microstructural conditions. Be- cause properties are strongly influenced by microstructure, a broad range of properties can be achieved by heat treatment.

Near-alpha and alpha-beta alloys can be produced in two basic microstructural conditions: a beta-transformed Widmanstatten microstructure containing lenticular alpha in a beta matrix, obtained by rapid cooling from above the beta transus, and a duplex microstructure consisting of equiaxed alpha and transformed beta grains obtained by alpha-beta working and heat treatment (Fig. 11). The beta-transformed microstructure provides a good combination of creep strength, fracture toughness, and FCGR resistance whereas the equiaxed duplex micro-


structure provides good tensile ductility and low cycle fatigue (LCF) resistance. To achieve a good balance between creep strength and LCF resistance, an alpha-beta-worked alloy can be heat treated just below the beta transus. This maximizes the amount of beta- transformed material for good creep strength, while a sufficiently fine duplex grain size is retained for good LCF properties [68].

Near-alpha and alpha-beta alloys have been developed for engine applications. Of the alpha-beta alloys, Ti-6-4 is by far the most widely used, accounting for almost half of all titanium used in aircraft [67]. They are used for rotating and static components, mostly as fan blade and compressor disc, blade, and vane material, but also for compressor cases and in temperature-sensitive areas of airframes close to engines. Engines contain a significant amount of Ti-6-4 as well as the stronger and more creep-resistant near-alpha Ti-6242 and Ti-6242s. The latter is a silicon-modified version of Ti-6242 with a 30C advantage in operating temperature limit over the base Ti-6242 alloy. Silicon increases the creep resistance of near-alpha and alpha-beta alloys [70] and is now added to most of the modern alloys, including the state-of-the-art IMI-834 (IMI Ltd., U.K.) and Ti-1100 (or Tin-ret 1100, Timet, U.S.A.). Some of the physical and mechanical properties of IMI-834 and Ti-1100 are compared in Table 8 [71] with properties of the older alloys Ti-6-4 and Ti-6242. The table also provides the temperature capabilities of each of the alloys, from which it can be seen that IMI-834 and Ti-1100 offer a 300C advantage over the old standby alloy Ti-6-4. Thus, the temperature capability of titanium alloys has improved considerably over the 30-year time span sep- arating introduction of Ti-6-4 in the mid- 1950s from the introduction of IMI-834 and Ti-1100 in the middle to late 1980s. Emerg- ing titanium aluminides may provide as significant a gain in temperature capability over these near-alpha alloys, as discussed next.

Beta alloys are, for the most part, meta- stable alloys which precipitate a second phase, usually alpha, when aged at high temperatures. Thus, beta alloys are age hardenable. They are normally used in the

Table 8. Typical Properties of Some Titanium Alloys Used in Engines (After Seagle and Wood [71])

0.2% E YS UTS

A&is &m3) (GPa) (MPa) (MPa)

% Creep

El K,.x, C)

Ti-64 4.43 110 965 1035 8 300

Ti-6242 4.54 120 990 1010 3 450

Ti-1100 4.52 114 895 1000 10 600

&II-834 4.55 117 905 1035 10 600

Alpha 2 4.84 125 620 725 4 730

Gamma 4.0 169 480 585 2 900

solution-treated and peak-aged condition but can also be used in the overaged condition for increased stability at high temperatures [65]. Some of the beta alloys are subject to strain-induced transformation to an orthorhombic martensite during quenching. Subtransus solutioning can be used to avoid the undesirable martensite and retain as much single-phase beta as possible for further aging [66].

The beta alloys have higher strength but are less creep resistant than the near-alpha and alpha-beta alloys. Beta alloys also provide the best combination of strength and toughness and have better cold formability than the other titanium alloys. Another plus is their excellent forgeability. On the minus side, they are 10% heavier than alpha-beta alloys because of the large amount of heavy elements required to stabilize the beta phase. Yet, they offer the highest strength-to-weight ratio of all titanium alloys. These characteristics make the beta alloys most suitable for airframe applications. However, they are also used for intermediate-temperature applications in engine containment structures and ducting. Properties of some of the beta alloys developed for airframe applications are compared in Table 9 with properties of the high- performance alpha-beta Ti-6-22-22S, an alloy that provides significant improvements in damage tolerance with respect to strength relative to Ti-6-4.

As compared to the near-alpha and alpha- beta alloys, the beta alloys have not been used to any great extent by the aerospace


Table 9. Typical Properties of Some Titanium Alloys Used in Airframes (After Seagle and Wood [71])

Alloy type

Alpha-beta Alpha-beta P-rich a-B Beta Beta Beta

Alloys

Ti-6-4 Ti-6-22-22s Ti-17 Ti-10-2-3 Ti-15-3 Timetal21S

E 0.2% YS UTS KZC &3) CGPaJ (MPa) (MPa) % El (MPah)

4.43 110 965 1035 8 42.9 4.57 121 1000 1100 10 82 4.65 116 1125 1205 10 - 4.65 107 1070 1170 8 50-80 4.76 107 1070 1135 10 - 4.93 116 1170 1240 6 -

industry. Their high cost has been a deterrent, as has been the cost of qualifying the materials [66]. Beside some military applications (Lockeeds SR-71, Rockwells B-1B bomber, and McDonnell Douglas C-U), beta alloys have been used by Fairchild for engine access doors on the Fairchild-Saab FS- 340 commuter aircraft (Ti-15-3) and by Gen- eral Electric Aircraft Engine for compressor discs (Ti-17, strictly speaking a beta-rich alpha-beta alloy). The beta alloys will be used extensively in the new Boeing 777, including alloy Ti-U-2-3 for the large forgings of the aircraft main landing-gear structure, an application traditionally assigned to high-strength steels in earlier-generation aircraft. Alloy Ti- 10-2-3, when properly processed, offers the best combination of strength, toughness, and resistance to high cycle fatigue of any titanium alloy [66]. It is interesting to note that the Eioeing 777 will be the first aircraft since the SR-71 in which Ti-6-4 will not be the dominant titanium alloy [65]. Other beta alloy applications on the Boeing 777 aircraft will include ducts (rolled Ti-15-3), engine nacelle components (rolled Timetal21S), and cargo-handling components (cast Ti-15-3). Alloy Ti-10-2-3 is also being used by the helicopter industry for rotor components. New beta alloys, including low-cost Timetal@ LCB and Beta CEZ (a beta-rich alpha-beta alloy), could provide further opportunities for weight reductions in structures and are being appraised by designers. The basic titanium alloys are available in all product forms and can be produced by casting or as wrought IM products. They can also be produced by P/M techniques either via the low-cost

blended elemental approach or the prealloyed powder approach [68].

TITANIUM ALUMINIDES

When mixed in the right proportions, aluminum and titanium form several ordered intermetallic compounds, two of which are of particular engineering relevance to applications in aero engines. The compounds are T&Al (hexagonal alpha-two phase) and TiAl (face-centered tetragonal gamma phase). Materials of interest to aero engines are alloys containing mixtures of alpha-two and beta phases, called alpha-two alloys, and mixtures of alpha-two and gamma phases, called gamma alloys [72]. Properties of alpha- two alloys and gamma alloys are compared with properties of conventional titanium alloys in Table 8.

With a density close to half that of superalloys, and temperature limits for creep and oxidation approaching 850 and 95OC, respectively, the gamma alloys are good candi- dates for replacing superalloys operating at intermediate temperatures in aero engines. Gamma alloys also have relatively high specific modulus (relative to conventional Ti alloys) and reasonable oxidation resistance, as well as high-temperature strength retention and good fire resistance. Their major shortcoming is low room-temperature ductility which arises from a limited number of active slip systems in the constituent phases, at room temperature [73]. The alpha- two alloys have properties that lie some- where between those for conventional titanium alloys and the gamma alloys. Their

60 I-?? lmmarigeon et al.

room temperature ductility is also limited, but not as much as in gamma alloys.

Both the alpha-two and gamma alloys are candidate materials for compressor applications as compressor discs and spacers, as well as impellers. Because of their higher temperature capabilities, the gamma alloys could also be used for aft-compressor cases, low-pressure turbine (LPI) blades, and turbine exhaust components, as well as after- burner components and support attach- ments. Until recently, it appeared that alpha- two alloys might reach application status before gamma alloys, partly because of their superior room-temperature ductility and more advanced state of development. How- ever, interest in alpha-two alloys appears to have diminished greatly over the last 2 years, judging from the number of publications on the subject. Papers at recent symposia on intermetallics and titanium aluminides have dealt almost exclusively with the gamma alloys. This loss of interest in alpha-two alloys could be due to the marginal performance advantage offered by these alloys over conventional titanium alloys, as compared to the gamma alloys.

Interest in gamma alloys is likely to grow even further in light of recent reports from General Electric Aircraft Engines (GEAE) that the first engine test involving LPT blades fabricated from a gamma alloy had been suc- cessfully completed, without the blades showing any evidence of abnormal damage. The test engine was a CF6-80C2 and the blades were part of the fifth-stage turbine rotor. The test lasted over 3 months and in- volved 1032 engine cycles simulating flight conditions (T. J. Kelly, GEAE, Private Com- munication, 1994). A close-up of the rotor after completion of the test is shown in Fig. 12. The rotor contains 98 blades. The solid blades were produced by Howmet using conventional foundry practice. Each blade is about 31cm long and weighs about 22Og, which is close to half the weight of the superalloy blades used at present in the engine. It has been estimated that, if used in the low-pressure section of the new GE 90 engine as a blade material, the gamma alloy could cut engine weight by more than 136kg

FIG. 12. Low-pressure turbine rotor with solid airfoils fabricated from gamma titanium aluminide after engine testing (T. J. Kelly, GEAE, Private Communication, 1994).

[74]. Rotor centrifugal loads are reduced by lighter blades and a thinner rotor disc could be used, which explains the large weight reduction.

The gamma alloy tested by GE was a second-generation cast alloy developed by GEs Corporate R and D (research and development) Group. Third-generation alloys are being developed at GE and elsewhere with modified chemistries and controlled microstructures to provide improved high- temperature capabilities, The new alloys are either cast or wrought and, in both cases, are solution strengthened with refractory metals (e.g., W) and other elements (e.g., Si) to improve creep resistance [75]. The on-going developments are focused on achieving more balanced properties between strength, ductility, creep resistance, oxidation resistance, and toughness. This is not a simple task, since alloying elements that improve ductility, such as Cr or Mn, reduce oxidation resistance, while elements that improve oxidation resistance, such as Nb or W, reduce ductility. Similarly, microstructural changes that improve tensile strength tend to lower toughness and creep resistance [76]. Also, with gamma alloys, toughness and ductility are inversely related, which is a limiting factor, but novel processing routes are being explored to produce microstructures that would provide both toughness and ductility [75].

Aircrafl Applications 61

Gamma alloys can be produced by conventional fabrication methods including in- vestment casting and mechanical hot working of ingolts. Casting is usually followed by HIPing to remove porosity. The alloys are hot worked by isothermal forging, hot-die forging, or extrusion. Gamma alloys can also be processed by powder metallurgy techniques using either blended elemental, prealloyed, or mechanically alloyed powders as starting materials [71]. A number of other P/M-processing routes, such as reactive powder processing [77] or thermochemical powder processing [71], have also been explored. The consolidation of powders into usable structural materials is normally achieved through HIPing, hot pressing, extrusion, or forging.

The effects of mechanical hot working on texture development, microstructural evolu- tion, and properties in wrought IM or powder-pmcessed gamma alloys are not well understood. Final microstructure is normally dictated by heat treatment, following thermomechanical processing. However, microstructure can also be controlled through multistep hot working and, in the case of HIPing, by incorporating the heat treatment with the HIP cycle [78, 791. In addition, controlled cooling rates in the HIP vessel from the peak temperature soak can be used to produce interlocking or serrated grain boundaries to improve the creep properties of compacts [80, 811.

Investigations have been carried out at NRC both on binary (Ti-48Al) and ternary (Ti-48Al-2W and Ti-48Al-3Cr) P/M alloys to assess the effects of mechanical attrition, hot isostatic pressing, isothermal forging, and heat treatment on alloy microstructure and properties. The four basic microstructures which can be produced in wrought IM alloys [75,76] can also be produced in powder processed alloys [80]. These microstructures are the so-called near-gamma, duplex, nearly lamellar, and fully lamellar microstructures, examples of which are shown in Fig. 13 for the Ti-48Al-3Cr alloy. Details of the HIP cycles and heat treatments used to produce the microstructures are discussed elsewhere [80]. Mechanical attrition of preaIIoyed pow-

der drastically reduces the grain size of compacts produced by hot isostatic pressing. For the two NRC ternary alloys, compacts produced from as-atomized powder had grain sizes from 5 to 1Opm while compacts produced from attrited powder had grain sizes of about 3OOnm, as shown in Fig. 14. Evi- dence of beta phase is more prominent in the W-containing alloy. The NRC investigations have also shown that forgeability of ternary powder compacts varies with alloy composition and processing conditions [82, 831. Processing windows exist within which the compacts exhibit superplastic behavior. In Ti-48AL2W, the beta phase at grain boundaries promotes good forgeability by prevent- ing grain growth during forging [83].

Gamma alloys in the near-gamma or duplex microstructural conditions are generally stronger and more ductile, whereas in the nearly lamellar or fully lamellar conditions, the same alloys are tougher and more creep resistant [76]. Creep properties are also strongly influenced by alloy composition. This is shown in Fig. 15, which compares the creep curves for the two NRC ternary alloys. The Ti-48Al-2W alloy has much better creep resistance than the Ti-48Al-3Cr alloy. Also, both ternary alloys are more creep resistant in the fully lamellar microstructural condition [BO] .

TITANIUM-MATRIX COMPOSITES

The stiffness and strength of titanium alloys and titanium ahnninides can be significantly increased by reinforcement with continuous ceramic fibers. Different types of fiber rein- forcements have been considered to date, although Sic is by far the preferred choice over C, B, or A1203 fibers because of its long-term stability at temperatures greater than 480C [84,85]. The major driving forces for TMC development in the United States have been the NASP (National Aerospace Plane), the HSCT (High-Speed Civil Trans- port), and the IHPTET (Integrated High- Performance Turbine Engine Technology) programs. Bladed compressor rings, shafts, ducts, fan components, structural rods, and so on, are expected to be made out of TMCs

J-?? lmmarigeon et al.

FIG. 13. Typical microstructures in powder-processed gamma titanium aluminide: (a) fine-grained near-gamma, (b) tine-grained duplex, (c) coarse-grained nearly lamellar, and (d) coarse-grained fully lamellar. The material was produced by hot isostatic pressing from Ti-4&41-3Cr prealloyed powder.

in the HSCT. Similarly, the TMCs are being considered for aircraft skin, internal structures, and medium-temperature-range engine parts in the case of NASI? In traditional airframe and aeroengine structures, however, the TMCs have only been considered for limited use in landing gears and as inserts in engine discs and impellers or fan-blade airfoils. Efforts are under way worldwide for developing techniques for reinforcing ma- trices of either conventional titanium alloys, such as Ti&Al+W, IMT-844, Ti-15-3, Timetal 21S, or recently developed Ti&l and TiAl aluminides with 30-5096 of Sic fibers.

Experimental TMC materials as well as prototype components have been produced by textron using the foil-fiber-foil (FFF) technique, as shown in Fig. 16 [84]. For ring- shaped inserts for discs and impellers, layers

of metallic foils and circumferentially wound fibers are stacked together and consolidated by hot pressing or hot isostatic pressing. There are, however, some disadvantages associated with this process: the cost of the foils is high and debuckling problems, that is, displacement of the fibers, can occur during HIPing. Another promising technique for manufacturing TMCs is the matrix-coated fiber (MCF) process, which relies on coating the Sic fibers with the matrix material using electron-beam deposition or other spraying techniques [86]. These coated fibers can then be easily wound on a mandrel to produce ring-shaped components either through hot pressing or I-Wing.

Two sources of Sic fibers are Textron Spe- cialty Materials from the United States, producing SCS-X fibers, and British Petroleum,

Aircraff Applications 63

FIG. 14. Optical and transmission electron micrographs of ternary gamma alloys produced by hot isostatic pressing from attrited prealloyed powders: (a and b) Ti-4&U-3Cr, and (c and d) Ti-Ml-2W.

producing Sigma fiber. The latter is an un- coated W core Sic fiber which is stiffer and about 25%~ more dense than the SCS-6 fiber. The success in producing TMCs with either fiber type will depend upon the processing conditions used to produce the component. Any process or fiber type which will sup- press the deleterious interface reactions and minimize other fiber-matrix interface defects is likely to be preferred.

C:ONCLUDINC REMARKS

Quite a number of new materials and processing concepts have emerged in the last few years, as described in this article, which offer opportunities for further reductions in aircraft weight or further growth in engine performance. While some of these new ma-

terials and processing concepts have been developed to the point where they can be evaluated in structures, many are at a very early stage of development. Much work will

16, I

76OoC - 276 MPa I I

0 100 200 300 400 500 TIME IN HOURS

FIG. 15. Creep curves for two ternary gamma alloys produced by hot isostatic pressing from prealloyed powders in either the duplex (as HIPed) or fully la- me&r microstructural conditions [So].

1-P lmmarigeon et al.

Blade / l-l

References

Center Line ________________

Fabric

Foil

Fabric

Foil

Fabric

FIG. 16. Schematic of the textron fiber-foil-fiber fabri-

cation method for the production of turbine rotor reinforcing rings [84].

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Some of the most innovative concepts re- late to the development of metal-matrix composites for either airframe and engine applications. Forecasts indicate that titanium metal-matrix composites could displace steels for landing-gear components in airframes and the conventional titanium alloys in engines by the year 2000. However, the cost to qualify new materials can be very high, depending on criticality of the application. With diminishing defence expen- ditures everywhere, this could prove to be a deterrent for commercial applications. Whether the new materials can be produced and qualified at affordable costs will ulti- mately dictate their rate of introduction in commercial aircraft.

This work was performed under NRC-IAR Proj- ect JHROO with financial assistance from the De- partment of National Defence under CRAD- DREP FA 1446 92NRC. Theauthors thank Ms. Michelle Gagnon and Mr. Vinko Totic for their assistance in preparing the manuscript.

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