Aerospace Struct Mater

Acta Materialia 51 (2003) 5775–5799www.actamat-journals.com

Progress in structural materials for aerospace systems�

James C. Williamsa,∗, Edgar A. Starke, Jr.b

a Materials Science & Engineering Department, The Ohio State University, 142 Hitchcock Hall, 2070 Neil Ave, Columbus,OH 43210, USA

b Materials Science & Engineering Department, The University of Virginia, Thornton Hall, Room B102A, Charlottesville,VA 22903, USA

Accepted 31 August 2003

Abstract

This paper examines the progress in aircraft and aircraft engines from the standpoint of the role that better materialsand processing has played. Such progress includes the relatively recent transformation of the aircraft industry frompurely performance driven products to products that are driven by customer value. It is demonstrated that advances inmaterials and processing technology and understanding has enabled much of the progress that has been made sincethe inception of manned, heavier than air flight. The recent constraints of cost, as determined by customer value, havechanged the way new materials are introduced and these trends appear to be the new paradigm for the aircraft andaircraft engine industry.

While the focus of this paper is aircraft and aircraft engines, the broader focus is on the role of materials in creatinglightweight structures. There are examples used in this paper that are relevant to automotive applications once theyare adjusted for cost. This matter is briefly discussed at the end of the paper. 2003 Acta Materialia Inc. Published by Elsevier Ltd. All rights reserved.

Keywords: Solidification; Aluminum alloys; Titanium alloys; Ni-base alloys; Mechanical properties

1. Introduction

In the 21st century the key characteristics of suc-cessful aerospace and automotive products aregood customer value and minimal environmentalimpact. Value is most simply thought of as service

∗ Corresponding author. Tel.:+1-614-292-2836; fax:+1-614-292-3244.

E-mail address: [email protected] (J.C. Williams).� The Golden Jubilee Issue—Selected topics in Materials

Science and Engineering: Past, Present and Future, edited byS. Suresh.

1359-6454/$30.00 2003 Acta Materialia Inc. Published by Elsevier Ltd. All rights reserved.doi:10.1016/j.actamat.2003.08.023

received from a product for money paid, includingboth initial cost and cost of ownership. Here, own-ership cost includes the cost of capital as well asthe cost of fuel and maintenance. Historically,improved structural materials have been a majorenabler of better (higher value) products. Half acentury ago better was measured largely in termsof performance. Today “Better” includes additionalmetrics as already mentioned. When these metricsare disaggregated into detailed design require-ments, materials are still a critical enabler, but thecharacteristics that lead to the choice of a materialfor a new product are more complex and the

5776 J.C. Williams, E.A. Starke, Jr. / Acta Materialia 51 (2003) 5775–5799

choices available to the designer are more numer-ous. In this paper we will attempt to describe, inqualitative terms, the evolution of structuralmaterials for application in aerospace systems andautomotive products and the role of materials increating customer value. Environmental impactalso has several associated dimensions, includingemissions of the product, effluent created duringproduct manufacture and disposal/recycling capa-bility.

A recent trend in the transportation industry isto improve customer value through creation of pro-ducts that incorporate advanced structural materialsand benefit from new manufacturing technologies.The objective is to create value by improving per-formance, reducing ownership costs, extending thesystem life and reducing environmental impact.Improved performance, as determined bymaterials, typically translates into higher structuralefficiency, resulting in reduced product weight.Structural efficiency is the combined result ofmaterials capability and design methodology. Forexample a stiffness limited component may incor-porate a higher modulus material using the samedesign or it may use a new design that increases thesection modulus or both. Honeycomb constructionused in aircraft is an example of an increased sec-tion modulus design. The decision to introducesuch a design impacts material selection becausenot all materials can be fabricated into honeycomb.Moreover, in recent years, honeycomb constructionhas fallen from favor because of the propensity forcorrosion in the core when moisture is allowed toenter. This simple example illustrates the interac-tion between materials selection, manufactur-ability, design methods and ultimate productacceptance. Performance also must be normalizedby damage tolerance because of the need forreliability. Improved reliability adds value throughincreased availability of the product to perform itsintended function. Due to the increased number ofdesign constraints driven by the growing numberof product requirements and the greater range ofstructural materials now available, the designer isfaced with complex choices for selecting a materialto meet the requirements for a particular system.The outcome of competition between variousclasses of materials may also be associated with

availability of the material in the appropriate pro-duct form, e.g. castings, forgings, rolled productsand/or extrusions, and how amenable they are tonet shape forming methods such as, superplasticforming, stretch aging, welding, casting and othermanufacturing methods. These latter operationshave historically been thought of as materials pro-cessing, but have seldom been treated as designconstraints. While this characterization is stillappropriate, it is no longer appropriate to treat pro-cessing as a separate consideration frommaterials selection.

Lightweight construction, consistent with meet-ing design intent, has become a universal require-ment for all transportation systems but the com-plexity of current design and manufacturingmethods now requires structural materials tosatisfy a much wider variety of properties. This hasled to many improvements in materials perform-ance during the past few decades, which are mostlyassociated with an increase in our understanding ofthe relationships among composition, processing,microstructure and properties. This understandinghas come from the technical literature, which pro-vides guidelines for these relationships, and fromindustrial experiments, which provide detailed, butoften empirical, information on how processingaffects microstructure and properties. Integratingthe information obtained from both sources hashelped replace the ‘ trial and error’ approach toalloy development and has reduced the timerequired to obtain the desired result. Furtherreductions are essential, however, because the pro-duct design cycle continues to get shorter andmaterials readiness at time of product launchbecomes a “ take it or leave it” opportunity win-dow. We assert that the most obvious path to short-ening the materials readiness cycle is through thedevelopment and use of more accurate and robustcomputational methods. These methods will, nodoubt include both modeling and simulation. Thisis a major challenge for the materials community,but one that must be addressed and conquered ifthe continued contribution to competitive productsattributable to materials is to be realized in futuregeneration products.

5777J.C. Williams, E.A. Starke, Jr. / Acta Materialia 51 (2003) 5775–5799

2. Materials for aerospace systems

Since the first days of powered flight, aircraftdesigners have focused on achieving minimumweight, both in airframes and in propulsion sys-tems. For example, the original Wright Flyeremployed an engine with an aluminum block,which was virtually unheard of at the turn of the20th century. From 1903 to about 1930 absoluteminimum weight was necessary for practical flight,due in large part to the limited capability of theavailable propulsion systems. Consequently, thestrength/weight ratio was the prime driver formaterials selection for both engines and aircraft.While this consideration continues to be of firstorder importance, light weight is now necessarybut not sufficient. Current design criteria are muchmore complicated, as mentioned in the Introduc-tion, and winning products require new designmethods as well as improved materials and pro-cessing methods.

The transition from internal combustion, pistonengines to turbines represented a major “gamechanger” in aircraft. Beginning in the early 1940’swith the German Messerschmitt Me-262 fighteraircraft, turbines became the preferred type of pro-pulsion system. The performance of the early tur-bine engines was severely limited by materialscapability, especially in terms of operating tem-perature. Jet powered fighters were in operation inthe late 40’s but large jet aircraft, e.g. bombers andtransports did not become feasible until nearly adecade later, largely because of the increased dif-ficulty of producing large turbine engines withadequate reliability. As will be described later, tur-bine technology has evolved and, today, propulsiontechnology is not the limiting factor it once was.

2.1. The evolution of aircraft and the role ofmaterials

The evolution of aircraft that fly farther and fas-ter has been a consistent goal of aircraft designers.Achieving this goal has required new materialswith other attributes in addition to strength andlight weight. For example, aircraft that fly at higherspeeds require higher temperature capabilitymaterials due to frictional heating. As a conse-

quence, skin materials have progressed from woodand fabric used in the early aircraft to advancedalloys of aluminum, titanium, and polymer matrixcomposites containing high-strength carbon fibers.Perhaps the most sophisticated aircraft ever builtin the western world is the SR-71 Blackbird (Fig.1) which had an all Ti alloy skin. This militaryaircraft was used for high altitude reconnaissancemissions and was capable of speeds in excess ofmach 3. The lessons learned from military oper-ation of the SR-71 showed what practical limi-tations existed for commercial supersonic flight,especially above mach 2, where the skin tempera-tures exceed the capability of Al alloys. As aconsequence, there has not yet been a large super-sonic vehicle produced and entered into scheduledservice that flies at these speeds. The Soviet Tupo-lev design bureau did produce several all-Ti Tu-144 aircraft but even in their economic system, itwas not considered practical. The Concorde oper-ates at mach 1.8 and is an Al airplane. Even so, ithas not been economical to operate and is beingretired from revenue service despite its technicalsuccess.

Damage tolerance became recognized as animportant issue in 1954 when three Comet jet air-planes, manufactured with 7075-type aluminum,crashed [1]. The cause of the crashes was attributedto premature fatigue failure of the pressurizedfuselage associated with stress concentrations atwindows and hatches. Today, fracture toughnessand fatigue crack growth have been incorporatedas a primary design criterion in many products in

Fig. 1. Front view of the all titanium SR-71 Blackbird.


the same manner as strength was used 35 yearsago. In fact, newer, higher toughness Al alloys forfuselage skins have enabled significant weightreductions through removal of some of the circum-ferential frames that serve, among other things, toarrest a running crack. This is discussed in detaillater and is mentioned here as an example.

Beginning in the ‘80s, jet powered flight hadbecome “a given” and other considerations for air-craft such as fuel costs, the revenue opportunitiesassociated with increasing range and payload, andreducing landing weight fees again returned thetechnical focus to weight reduction but withoutsacrificing life. Then, in the ‘90s, the realization ofthe benefits of extending the life of the aging air-craft fleet resulted in a technical focus on improveddamage tolerance and improved corrosion resist-ance. One response to this requirement was thepossibility of retrofitting existing aircraft withnewer and advanced alloys. This was done exten-sively in the case of the B-52 bomber and othermilitary aircraft. In the case of commercial pro-ducts, derivative aircraft models and the emergenceof new, large twin engine aircraft was more rep-resentative of product trends.

The currently used airframe design methodshave evolved over many years and the practicesemployed today incorporate both the benefit of thispast experience and availability of better analyticalmethods. Static strength has been an important firstorder consideration since the beginning of flightand aircraft are usually designed to withstand amaximum operating load plus a safety factor,which is typically 1.5. In modern aircraft, staticstrength is necessary but not nearly sufficient, larg-ely because of the protracted service lives expectedof aircraft. The effect of fatigue on aircraft integ-rity began to be considered as long as 70 years ago.Design methods that consider fatigue as a con-straint were introduced. Today, there are 2 concep-tual approaches to calculating fatigue limited struc-tural life: safe life and fail safe. Safe life designrequires that no fatigue failure occurs in N life-times, where N always is greater than one and typi-cally on the order of four. Safe life design wasintroduced in the 1930s and 1940s and reliesstrongly on detailed knowledge of service experi-ence and requires rigorous product testing. Fail

safe design was introduced in the 1950s andassures that catastrophic failure is not probable asthe result of a failure of a single structural element.To achieve this result, fail-safe designs incorporateredundant crack pathways in the design. Finally,damage tolerance was introduced in the late 1960sand early 1970s, and couples crack growth analysiswith periodic inspections to detect cracks andremove cracked load bearing members from ser-vice in situations where these would have a highprobability of failure prior to the next scheduledinspection.

The material property requirements for use inairframes vary depending on the particular compo-nent under consideration [2]. The fuselage is asemi-monocoque structure that is made up of skinto carry cabin pressure (tension) and shear loads,longitudinal stringers or longerons to carry thelongitudinal tension and compression loads due tobending, circumferential frames to maintain thefuselage shape and redistribute loads into the skin,and bulkheads to carry concentrated loads includ-ing those associated with pressurization of thefuselage. Strength, Young’s modulus, fatigue crackinitiation resistance, fatigue crack growth rate,fracture toughness and corrosion resistance all areimportant, but fracture toughness is often the limit-ing design consideration. The wing is essentially abeam that is loaded in bending during flight. Thewing supports both the static weight of the aircraftand any additional loads encountered in servicedue to maneuvering or turbulence. Additional wingloads also come from the landing gear during taxi,take-off and landing and from the leading and trai-ling edge the flaps and slats during that aredeployed during take-off and landing to createadditional low speed lift. The upper surface of thewing is primarily loaded in compression becauseof the upward bending moment during flight butcan be loaded in tension while taxiing. The stresseson lower part of the wing are just the opposite.Compressive yield strength and modulus of elas-ticity in compression are the static material proper-ties that influence design and, due to alternatingloads caused during flight, fatigue resistance is alsoimportant. The tail of the airplane, also called theempennage, consists of a horizontal stabilizer, avertical stabilizer or fin, and control surfaces e.g.


elevators and rudders. Structural design of both thehorizontal and vertical stabilizers is essentially thesame as for the wing. Both the upper and lowersurfaces of the horizontal stabilizer are often criti-cal in compression loading due to bending and,therefore, the modulus of elasticity in compressionis the most important property.

Improved structural materials can contribute toimproved performance and reduced operating costsfor aircraft and spacecraft as noted in the schematicof Fig. 2. Since weight is probably the most sig-nificant contributor to performance and operatingcosts associated with fuel efficiency, the designernormally considers how a certain property willimpact weight savings. In order to select the cor-rect material for a component of a new system,aerospace companies have developed computerprograms that perform trade-off studies. Theproperties of materials are placed in a database andcomparisons are made between baseline materialsand newer advanced materials for a particularcomponent and failure modes as schematicallyillustrated in Fig. 3. By using this type of programa designer can determine the real potential weightsavings. In addition, the cost of materials forweight reduction should not exceed the costs savedfrom reduced fuel burn, maintenance and landingfees [3]. Consequently, aerospace companies con-duct a cost/benefit analysis for new candidatealloys. The total life-cycle cost consists of acqui-sition, operation and support costs [3]. The non-recurring development costs for design allowables,

Fig. 2. Plot showing the relative contribution of various tech-nology advances to improved fuel efficiency of aircraft.

Fig. 3. Flow chart showing the order of events in computeraided materials selection.

etc., are typically divided by the number of air-planes to be built using the candidate material, andthus adds a fixed amount to the cost per poundsaved [4].

2.1.1. Evolutionary improvement of aluminumalloys for aircraft

Improved structural materials are usually classi-fied as either a revolutionary product, e.g. rapidlysolidified and mechanically alloyed P/M alloys,discontinuous and continuous fiber reinforced met-als and polymers, and structural laminates, or evol-utionary, i.e. derivative alloys and tempers. Revol-utionary products have had limited success inincorporation into aircraft due to cost of manufac-ture, qualification and certification and possiblymodification of existing material production infra-structure. Although polymer matrix composites arebeing used in modern commercial aircraft, e.g. forthe fin of the Airbus A310, the horizontal stabilizerof the Airbus A340 and the Boeing 777, aluminumalloys have remained the materials of choice forthe airframe of most commercial aircraft. Evol-utionary improved aluminum alloys have had amuch more rapid insertion due to their lowermanufacturing and/or life cycle costs, low substi-tution risk and the use of an existing material pro-duction infrastructure.

A summary of various microstructure/propertyrelationships for aluminum alloys is given in Table


1. The basic metallurgical concepts for maximizingproperties of age-hardenable aluminum alloys arewell known [2,5]; the challenge is putting theminto practice in alloy design. For example, theimpurities Fe and Si form coarse constituents in2XXX 7XXX, and 8XXX aluminum alloys andresult in lower fracture toughness [2,6] and havean adverse effect on both fatigue crack initiationand fatigue crack growth resistance [7]. Coarse pri-mary phases formed when solubility limits areexceeded at the solution heat treatment temperature(or those formed during processing and not re-dis-solved during subsequent heat treatment) have asimilar effect [8]. Consequently, low Fe and Si lev-els, a good knowledge of complex phase diagramsand tight controls on composition can be used toproduce a good balance of strength, fracture tough-ness and fatigue crack growth resistance. As men-tioned, the cost of implementing these concepts isoften considered when selecting an alloy for aparticular application, and the cost issue has oftentimes delayed their implementation. Incrementalimprovement in yield strength of aluminum alloysis shown schematically in Fig. 4, in specific stiff-ness in Fig. 5.

The two major passenger aircraft manufacturers(Airbus and Boeing) have recently introduced air-craft, e.g. the Boeing 777 and the Airbus A340,

Table 1Property-microstructure relationships in aluminum alloys

Property Desired microstructural feature(s) Function of feature(s)

Fine grain size with a uniform dispersionStrength of small, hard particles Inhibit dislocation motion

Fine structure with clean grain boundariesand no large particle or shearable Encourage plasticity and work hardening,

Ductility & toughness precipitates inhibit void formation and growthThermally stable particles within the Inhibit grain boundary sliding and coarse

Creep resistance matrix and on the grain boundaries microstructureFine grain size with no shearable particles Prevent strain localization, stress

Fatigue crack initiation resistance and no surface defects concentrations, and surface slip stepsLarge grain size with shearable particles Encourage crack closure, branching,

Fatigue crack propagation resistance and no anodic phases or hydrogen traps deflection and slip reversibilityPrevent preferential dissolution of second

Pitting No anodic phases phase particlesHomogenize slip and prevent crack

Stress corrosion cracking & hydrogen Hard particles and no anodic phases or propagation due to anodic dissolution orembrittlement interconnected hydrogen traps HE

Fig. 4. Plot of yield strength for new Al alloys as a functionof the year of introduction.

which utilized evolutionary improvements of oldermaterials. The advanced materials on the Boeing777 include a number of improved aluminum andtitanium alloys and polymer matrix composites aswell as laminates and lightweight sealants. Someexamples of the aluminum alloys are shown in Fig.6, and include 7150-T77 that has higher strengthand damage tolerance when compared with 7050-T76, alloy 7055-T77 that has higher strength than7150-T6 along with similar fracture toughness andfatigue crack growth resistance and alloy 2524-T3that has approximately 15–20% improvement in


Fig. 5. Plot of the density normalized modulus of Al alloysas a function of the year of introduction.

Fig. 6. Plot showing the improvements in strength-toughnesscombinations of some newer Al alloys.

fracture toughness and twice the fatigue crackgrowth resistance when compared with 2024-T3.The improved properties of 2524, compared with2024 were obtained using the principles describedin the previous paragraph. The higher toughnessand greater resistance to fatigue crack growth of2524-T3 helped in the elimination of tear straps ina weight-efficient manner on the Boeing 777. TheT77 temper, developed by Alcoa, is based on athree step aging treatment that produces a higherstrength with durability and damage tolerancecharacteristics matching or exceeding those of7050-T76. The improved fracture toughness of7150-T77 products is attributed to the controlledvolume fraction of coarse intermetallics particlesand unrecrystallized grain structure, while the com-bination of strength and corrosion characteristics is

attributed to the size and spatial distribution andthe copper content of the strengthening precipitates[2]. Phase diagrams and metallurgical simulatorsare currently being used to optimize compositionsand fabrication schedules for a wide variety ofaluminum alloys. Alloys 7449, used for upper wingskins, and alloy 7040, used for spars, weredeveloped by Pechiney using this technology [9].The relatively slow quench rate of thick plate wasimportant in selecting the composition and low lev-els of Mg and Cu reduced heterogeneous precipi-tation significantly, ensuring good fracture tough-ness without sacrificing strength. The incrementalimprovements in usable damage tolerance are illus-trated schematically in Fig. 7.

The 6XXX alloys have been considered toreplace 2024 on a number of US Navy programsand alloy 6013 is being used on the Boeing 777.The major problems with 2XXX alloys for fusel-age skins is that they must be clad since they canbe susceptible to intergranular corrosion. Inaddition, the 2XXX alloys cannot be fusionwelded, a process that is being considered forreducing weight and costs of manufacturing. The6XXX alloys are weldable and cheaper than 2XXXalloys, however, Cu-rich 6XXX, e.g. 6013-T6 and6056-T6 are also susceptible to intergranular cor-rosion. This susceptibility is associated with theformation of precipitate free zones at grain bound-aries, which are developed during artificial aging,are depleted in Si and Cu and are anodic withrespect to the grains [10]. A new temper has beendeveloped, designated T78 that has a controlled

Fig. 7. Plot comparing yield strength-toughness improvementsof 7000 and 2000 series Al alloys.


degree of overaging that desensitizes 6056 to inter-granular corrosion, keeping the yield strength at anacceptable level with respect to the T6 temper.Other experimental tempers have been developedthat prevent intergranular corrosion without anyloss in strength [11]. An alternative to changingalloys to facilitate joining, is the use of a relativelynew joining method known as friction stir welding(FSW). In this method, the alloys are not melted,but instead are joined in the solid state by mechan-ical working. In a sense, they simply are “kneadedtogether” with a rotating tool. The use of FSW cre-ates the opportunity to join a wide range of Alalloys that cannot be fusion welded.

Superplastic forming is also a way to reduce partcount and manufacturing cost since it lends itselfto the manufacture of very complicated parts. Forexample, a single sheet can be formed into a com-plex arrangement of ribs and stiffeners and hencereplace an assembly of parts and fasteners at areduced manufacturing cost and at a reduced rate.Superplastic 7XXX alloys have been available forsome time, however there have been difficulitieswith developing superplastic 6XXX alloys. A ther-momechanical process has recently been developedfor 6013-6011 alloys that produces a microstruc-ture amenable to superplastic forming [12]. Theprocess involves obtaining a fine, uniform distri-bution of one micron-sized particles for subsequentparticle stimulated nucleation of recrystallizationin order to obtain a fine (9.5 micron), equiaxed,recrystallized grain structure and a random texture.The alloys are then superplastic above 500 °C.Uniaxial tests indicated a strain rate of 0.5 at 540°C and elongations of 375% using a stress of4.7MPa.

2.1.2. Lighter weight, higher stiffness materialsfor aircraft

Aluminum-lithium alloys are attractive for aero-space applications because they have lower densityand higher modulus than conventional aluminumaerospace alloys. Each weight percent of lithiumlowers the density of aluminum by approximately3% and increases the modulus by approximately6%. The second generation of Al-Li alloys (thefirst being the Alcoa alloy 2020) was developed inthe 1970s (alloy 1420 in Russia) and the 1980s

(alloys 2090, 2091 and 8090). The Al-Mg-Li alloy1420 and the Al-Li-Cu-X alloys 2090 and 8090 arenow in service in the MIG 29 and the EH1 helicop-ter. Alloy 1420 has only moderate strength and theAl-Li-Cu alloys that contain approximately 2%lithium (2090, 2091 and 8090) have a number oftechnical problems, which include excessive ani-sotropy of mechanical properties, crack deviations,a low stress-corrosion threshold and less thandesirable ductility and fracture toughness. NewerAl-Li alloys have been developed with lower lith-ium concentrations than 8090, 2090 and 2091.These alloys do not appear to suffer from the sametechnical problems.

The first of the newer generation Al-Li alloyswas Weldalite 049 (2094) which can attain ayield strength as high as 700 MPa and an associa-ted tensile elongation of 10%. A refinement of theoriginal alloy, 2195, which has a lower copper con-tent, is now being used for the U.S. Space ShuttleSuper-Light Weight Tank. Alloy 2195 replaced2219 and, along with a new structural design,enabled a 7500 pound weight reduction on the60,000-pound tank. This allows an increased pay-load for the Shuttle and reduces the number offlights necessary for the construction of the Inter-national Space Station, thus saving millions of dol-lars. Three other recent derivatives of the third gen-eration of Al-Li alloys are 2096, 2097 and 2197.They contain lower copper and slightly higher lith-ium content. Alloys 2097 and 2197 contain a verylow Mg content to improve SCC resistance and Mnto prevent strain localization normally associatedwith the shearable Al3Li present in the higher Li-containing alloys. Alloy 2097/2197 was recentlyselected to replace 2124, which had inadequatefatigue strength, for bulkheads on the F16. Alloy2097 has a 5% density advantage over 2124 andat least three times better spectrum fatiguebehavior or approximately 15% higher spectrumfatigue stress allowable. Although Al-Li alloys aremore expensive than conventional aluminumalloys, the replacement of 2124 by 2097 for theBL 19 Longeron of the F16 doubles the servicelife of the part, saving over twenty-one million dol-lars for the 850 aircraft fleet operated by the USAF.Al-Li alloys, due to their better fatigue life, are alsoreplacing engine access cover stiffeners, currently


made from 2124. The Al-Li alloy 2098 has beensuccessfully demonstrated/flight tested for F16skins and has showed a 6X life improvement over2024 skins.

There are a number of new aluminum materialsthat are being developed for use in commercialtransport aircraft. Laminated hybrids of aluminumsheet with aramid-fiber-reinforced (ARALL) orglass-fiber-reinforced (GLARE) composites havehigh fatigue resistance and the potential for sig-nificant weight savings in aircraft. These materialsalso have resistance to burn through in the eventof a fire and can potentially substitute for titaniumin firewalls. On the negative side are the very highmaterial costs, typically 7–10 times that of mono-lithic aluminum sheet. The cargo door of the C-17is fabricated from ARALL. Laminates also offerthe potential for adding extra functionality, e.g.load monitoring, damage detection, etc. Combiningsuch functionality would make them more costeffective in aerospace structures. Other aluminumalloys under development include Al-Mg-Scalloys. Although potentially very expensivebecause of the presence of scandium, they appearto have excellent corrosion resistance, and as abody skin they may not need to be clad or paintedwhich would reduce maintenance costs. Higher-strength forgings, age-formable alloys, less-quench-sensitive alloys, and rivet alloys withimproved formability are all being examined anddeveloped for future use in subsonic aircraft.

Since a significant portion of airframe design isstiffness driven, there are opportunities for newmetallic materials providing high specific stiffnessrelative to currently available materials. Forexample, aluminum beryllium alloys may providean affordable sheet metal alternative to resin matrixcomposites in stiffness critical airframe structure.A comparison of the specific modulus and specificstrength of AlBe (62%Be, 38%Al) withboron/epoxy, Ti, and Carbon/epoxy is shown inFig. 8. Aluminum beryllium alloys are currentlyused for secondary structures such as equipmentshelves and support structure due to low density,high stiffness, and good vibration damping charac-teristics. Application of these materials in primarystructure requires the establishment of a databasefor design and manufacturing, safe manufacturing

Fig. 8. Plot of density normalized stiffness vs. strength for Al-Be alloys and Boron and Carbon fiber reinforced polymermatrix composites.

practices that avoid exposure to beryllium dust andparticulates, and corrosion protection methods forcomponents in the field. Also appropriate safetyprotocols to control exposure to beryllium in fieldmaintenance must be applied.

Metallic sandwich panels with periodic, open-cell cores are being developed with novel fabri-cation and design tools [13]. They can be fabri-cated using protocols based both on sheet formingof trusses and textile assembly. Analysis, testingand optimization have revealed that sandwichpanels constructed with these cores sustain loadsat weights greatly superior to stochastic forms andcompetitive with the lightest known honeycombcore systems. The benefits of the truss/textile coresover honeycombs reside in their higher specificstrength at low relative density, and lower manu-facturing costs.

2.1.3. Improved Al base materials for newersystems

The Airbus A380 will be the largest commercialaircraft ever built and requires the introduction ofadvanced and new materials—combined with newmanufacturing technologies. Hinrichsen has writ-ten a series of papers that describes Airbus’ selec-tion process for new materials and processes forthis advanced aircraft [14–16]. The first step intheir selection process was to examine “ the lessonslearned” from their operations of existing Airbusfleets, including in-service experience concerningcorrosion protection, inspections and repairs for


Fig. 9. Schematic depicting materials chosen for use in thenew, very large Airbus A-380.

crack growth, etc. Each of their design solutionsand material applications envisaged had to getapproval from A380 customers with respect toinspections and repairs. Workshops with the air-lines were regarded as a key element of the “ tech-nology down-selection process.” The finalmaterials selection is displayed in Fig. 9 andincludes advanced aluminum alloys, carbon fiberreinforced plastics (CFRP), fiber metal laminates(GLARE) and glass thermoplastics. Note that evenfor this advanced aircraft, the majority of theweight of the airframe is still made of aluminum,Fig. 10. New materials, combined with new manu-facturing processes, deliver the smallest portion ofadded value for the launch version. Hinrichsenpoints out that this fact is in contradiction to theperception that a new aircraft family would pioneer

Fig. 10. Materials distribution in the A-380 by per cent ofempty weight.

new technology and should take full benefit fromday one. Airbus considers that a more evolutionaryapproach is preferred since past experience hasshown that almost every new technology has someinitial technical problems; water-ingress withArmid fibers and de-bonded longitudinal lap jointsof metal fuselage skins were two that were men-tioned by Hinrichsen. However, new materialsdevelopment will continue to be monitored forpossible mid-term candidates for continuousimprovement with respect to weight savings, poss-ibly replacing 2024 and 2524. Fig. 11 comparesstandard 2024-T3 with sheet metal material withadvanced alloys and new developments withrespect to fracture toughness and yield strength.Laser-beam-welding (LBW) technology haspushed the development of 6056 at Pechiney andof 6013 at ALCOA since they offer improved yieldstrength relative to 2024 and 2524 and can beLBW. This process replaces riveting forstringer/skin assembly in fuselage panels and addsvalue by reducing manufacturing costs. Alumi-num-magnesium scandium alloys show materialsperformance similar to 2024-T3 but with much bet-ter corrosion resistance with the added benefit thatthey don’ t require complex heat treatments. TheC68-type alloys are being judged as mid-term can-didates due to their higher strength with respect to2024 and 2524.

There has long been interest among the majoraircraft manufacturers in producing a second-gen-

Fig. 11. Fracture toughness—yield strength domain with sev-eral alloys positioned to show their utility for deigns that arestatic strength or damage tolerance limited.


eration supersonic transport to operate in the long-range international market. Although there is nocurrent active program for such an aircraft, futuredemands may regenerate production activity. Thechoice of speed for such a vehicle will be the majordeterminant for the materials selection for thefuselage and may range from Mach 1.6 to Mach2.4. Aluminum alloys would be viable candidatesup to Mach 2 while more expensive titanium alloysand polymer composites would be candidates forhigher Mach numbers. Alloy 2618 (CM001) wasselected for the primary structure of the Concorde,a Mach 2.0 airplane. The variant of 2618 that wasused is a specially processed clad sheet and wasselected over other candidates, such as 2024-T8,7075-T6 and 2014, because of its static strength,fatigue strength, and especially because of its creepstrength. Recent studies have shown that in orderto meet the desired economy and range goals of anew SST, a weight reduction of approximately 30–33% is required compared to the projected weightof an airframe similar to that used for the Concordebut enlarged to meet the passenger requirements[17]. Following the research of Polmear [18],recent studies on a NASA program showed thatsmall additions of Ag and Mg to 2519 couldincrease the peak aged tensile yield strength by10%. The addition of small amounts of Ag and Mgstimulate the precipitation of a plate precipitate on{111}, designated �, in addition to the �’ precipi-tates that form on {100} of the matrix. Using com-puter simulations, Zhu et al [19,20] have shownthat an optimum balance of {111} and {100} pre-cipitates can increase the strength compared toalloys that contain similar volume fractions of onlyone type of precipitate. A number of alloys basedon 2519, but with variations in Cu, Mg, and Mnand with 0.5%Ag and 0.13%Zr were examined onthe NASA program. Two alloys looked particularlyattractive with respect to mechanical properties andwere designated C415 and C416. Tensile yieldstrengths and creep resistance of these alloys aresignificantly better than 2519-T87 and 2618-T61,Figs. 12 and 13.

2.1.4. New methods for Al alloy designThe aerospace industry, both air framers and

engine manufacturers, have indicated the need for

Fig. 12. Plot of temperature dependence of yield strength ofseveral Al alloys including two new alloys.

Fig. 13. Creep strain vs. time plots for three Al alloys. Testsconducted at 107°C with a load of 207 MPa.

new high temperature aluminum alloys, which canoperate successfully at temperatures to 150 °C fora supersonic airframe or for components of theengine. Due to economic considerations, alloydevelopment can no longer be performed usingpurely empirical, trial and error approach that hasbeen dominate over the past 100 years. Modelingand simulation, in conjunction with experiments,can be employed to improve the efficiency of alloydesign, optimizing processing and manufacturingoperations [21]. This approach can greatly reducethe time associated with alloy development and thegeneration of the knowledge base required forinsertion of new materials into aerospace systems.One approach to streamline alloy design is shownschematically in Fig. 14. The first step is to select asystem that offers promise for obtaining the desired


Fig. 14. Flow diagram showing main elements of a typical alloy development campaign.

microstructure and properties to meet the statedobjective and this is accomplished by an extensiveliterature search and evaluation of available data.After selecting the system, one needs to determinethe phase space that offers the greatest potential tomeet the program goals.

Since phase diagrams represent the state of analloy as a function of temperature, pressure, andalloy concentration they can be used for identifyingthe ideal phase field. Often times the phase dia-grams for a complex alloy system have not beendetermined and must be calculated using programssuch as CALPHAD (CALculation of PHAseDiagrams). This method is an approach to estab-lishing equilibria among phases, through thermo-dynamic modeling of the individual phases basedon the idea of phase competition in a system. Thephases are modeled according to the phase stab-ilities, measured thermodynamic properties, andthe measured transus points (including liquidus,solidus, eutectic, etc). An article by Kattner [22]presents an excellent overview of the use of ther-modynamic functions for the calculation of phasediagrams. Such a calculation reduces the effortrequired to determine equilibrium conditions in amulti-component system. Dubost [23] has illus-trated the industrial application and determinationof equilibrium phase diagrams for light alloys, e.g.aluminum. Materials design can now be viewed as

the best application of available models for the pre-diction of synthesis of alloys with desired proper-ties. Of course, all theoretical calculations must beverified experimentally, but the experiments shouldbe based on sound theory and not on a trial anderror approach.

Once the alloy system and phase field have beenidentified, the desired optimum precipitate struc-ture and morphology for maximum strength maybe determined from a dislocation simulationmethod [19]. The knowledge gained through thesimulation method has been verified in somealuminum alloys [19]. For alloys for high tempera-ture application, the thermal stability of the precipi-tate structure is important and one must considerboth phase coarsening and phase competition.Coarsening of a single second phase is thermodyn-amically attributed to the tendency for decreasingthe total interfacial energy between the secondphase and the aluminum matrix. The interfacialenergy consists of the chemical energy and theassociated strain energy. First principle calcu-lations using programs, such as the Vienna ab initiosimulation program (VASP) [24] may be used tocalculate interfacial energies and changes in totaland interfacial energy by the addition of selectedtrace elements to either the interface or the internalprecipitate structure. Such calculations will aid inthe selection of trace elements that, hopefully, will


reduce the coarsening of the precipitates withoutusing a trial and error method. This program mayalso be used to verify crystal structures and to cal-culate elastic constants. Quantitative data of theelastic constants for the second phases are signifi-cant and fundamental for determining the strength-ening mechanisms. These parameters also play arole in determining the habit plane and the evol-ution of equilibrium shapes of the second phaseparticles since they determine the elastic strainenergy associated with lattice misfit. Once candi-date alloys (and microstructures) have been ident-ified the methodology must be evaluated withmechanical property tests, etc. but such anapproach should lead to streamlined alloy designand therefore aid in the early insertion of new highperformance materials.

2.1.5. Ti alloy usage in aircraftTi alloys have been used for special purpose

applications in both military and commercial air-craft for several decades. In addition to the highspeed military airplane, the SR-71, described earl-ier, there are a number of applications of Ti alloysfor heavily loaded structure such as bulkheads infighter aircraft, the wingbox on the B1-B bomberand the landing gear beam in the Boeing 747.These components are illustrated in Figs. 15–17. In

Fig. 15. Photos of a forged and machined Ti alloy bulkheadfor a fighter aircraft.

Fig. 16. Photo of the wing box of the B1-B bomber aircraftmade by diffusion bonding.

Fig. 17. Photo of the Ti alloys forged and machined landinggear beam for the Boeing 747.

each of these cases, Ti alloys were chosen because,compared to Al alloys, they have equal or betterdensity corrected strength, good damage toleranceand excellent corrosion resistance. The alloyschosen for each of these three applications is theold, but still widely used a + b alloy, Ti-6Al-4V(Ti-6-4). Higher strength a + b and b Ti alloys areavailable, but the damage tolerance, both tough-ness and crack growth, typically decreases in thehigher strength alloys. For fracture critical struc-tures, i.e. those that are sized for static strength,the critical flaw size prior to the onset of unstablecrack extension is proportional to the parameter(KIc/YS)2. Thus incorporation of higher strengthalloys without a commensurate increase in tough-ness increases the risk of catastrophic failure whichis generally unacceptable. In addition to the intrin-


sic mechanical properties of Ti alloys, the cor-rosion resistance makes them attractive, especiallyfor components that are embedded in the aircraftand, as a consequence, are very difficult to inspectfor corrosion attack. The better resistance to gen-eral corrosion and essential immunity to exfoliationcorrosion compared to high strength Al alloys isan advantage for Ti alloys. Ti alloys are also usedin circumstances where their higher strength allowsthe same load to be carried by a physically smallerstructural member, even though there is no weightadvantage because of the higher density of Tialloys. An example of the latter situation is thelanding gear beam shown in Fig. 17. From a strictstructural efficiency standpoint, Al alloys would becomparable and considerably less expensive. How-ever, had an Al alloy been selected, the landinggear beam would not have fit within the availablespace in the airframe fuselage, thereby creating anaerodynamic penalty.

Ti alloys are also a very good alternative to highstrength steel, even though at the highest strengthlevels, the structural efficiency achievable withsteel is considerably higher. The issue here is thatthe susceptibility of steel to hydrogen embrittle-ment becomes much higher at strength levels�1250 MPa. Using steel at strengths higher thanthis requires use of a protective coating such aspaint or plating with Cd or Cr. The former of theseapproaches requires periodic maintenance to repairscratches and chips and the latter of these is beingphased out for environmental and health hazardreasons. The landing gear of most commercial air-craft has traditionally been made of high strengthsteel such as AISI 4340 heat treated to strengthsof 1800–1900 MPa. Service history has shown thatnumerous hydrogen embrittlement failures havebeen experienced, despite the use of stringentmaintenance procedures. Recently, a high strengthb alloy, Ti-10V-2Fe-3Al (Ti-10-2-3), has beenselected for the landing gear of the Boeing 777.For this application, Ti-10-2-3 is heat treated to the1250 MPa strength level. The decision to use Ti-10-2-3 is driven by the ability to achieve weightreduction and eliminate the risk of hydrogenembrittlement failure, albeit at an increase in cost.This is a good example of added customer valuedirectly attributable to materials. The B-777 land-

ing gear is shown in Fig. 18. The new, very largeAirbus aircraft (A-380) mentioned earlier appearsto be committed to a Ti alloy landing gear also,although the choice of alloy has not been finalized.The driver for this selection is also weightreduction and, perhaps, available space if Al alloysare being considered. This relatively new appli-cation of Ti alloys for landing gear is made poss-ible by the advancement in b alloy understandingand production capability.

Another relatively new application of b alloys inlarge aircraft is for springs. These Ti alloy springsreplace steel springs at a significant weightreduction and also eliminate the need for protectionby painting. Because springs are generally loadedin torsion, fracture is a lesser concern. Conse-quently b alloys with very high strengths but lowtensile ductility can be used safely. The alloys arecold drawn or rolled, coiled into springs and agedto achieve strengths in excess of 1400 MPa. Therelatively low modulus of b alloys (60–100 GPa)coupled with the high yield strength permits a verylarge elastic displacement range for the spring,

Fig. 18. Photo the landing gear assembly for the Boeing 777.The horizontal forging is Ti-10-2-3 and is the largest b alloyforging in use today.


which also can be beneficial. Examples of severalb alloy springs are shown in Fig. 19. Several b Tialloys are commonly used for springs including Ti-3Al-8V-6Cr-4Mo-4Zr (Beta C) and Ti-15V-3Cr-3Al-3Sn.

Ti alloys are also being used in aircraft undercircumstances where the corrosion resistance is theprime consideration. Ti alloys exhibit good gal-vanic compatibility when placed in contact withcarbon fiber composites, whereas Al alloys do not.Consequently, as the use of composites increaseswith each generation of aircraft, the need to use Tialloys for fittings and attachments for mitigation ofgalvanic corrosion also has increased. These fit-tings and attachments often are not heavily loadedand, therefore represent prime opportunities to useTi castings which are less costly than machinedforgings or fabricated components. Ti casting tech-nology has progressed in the past 15–20 years tothe point that complex net shapes are producible.Such castings have been in use in aircraft enginesfor a number of years but the aircraft designershave been slow to adopt this technology. Anexample of a casting that is used on a militarycargo aircraft is shown in Fig. 20. This single piececasting replaces a fabricated part made up of 22pieces resulting in major cost savings. Ti alloycastings that have been hot isostatically pressed(HIP’d) exhibit no porosity and therefore havefatigue strength that is comparable to wrought pro-ducts with the same microstructure. This is in con-

Fig. 19. Photo showing three types of β alloy springs used inBoeing aircraft.

Fig. 20. Photo of a Ti alloy casting for use in a large militarytransport aircraft.

trast to the situation for Al alloys where gasinduced porosity is common and such porosity can-not be closed by HIP processing. This has led tothe use of a casting factor for Al castings, whichserves to penalize their structural efficiency. Thereis no comparable need in the case of HIP’d Ti alloycastings, making the design process more straightforward and yielding more attractive results interms of structural efficiency. This point also willbe mentioned again in connection with Ti usage inaircraft engines.

Another alternative to machining forgings ormachining plate to create a rib-on-plate geometryis a new process known as laser additive manufac-turing. In this process, powder is fed into a laserbeam which fuses it and deposits the molten metalon a flat substrate, creating an upstanding rib. Thisrib is not necessarily straight and can take on acurved shape because the work piece moves underthe laser and is guided by a computerized numeri-cally controlled servo mechanism in the same wayas a machine tool would be in a CNC machine.This process, while still under development, exhib-its promise for making net shapes where alternativeprocesses would require removal of as much as80% of the starting weight in some cases. Thereare unanswered questions about this processincluding the cost of powder and the ability to con-trol porosity, but it is an attractive means ofincreasing the utilization of input material,especially for designs where there are deep pocketsthat would have to be created by machining. An


Fig. 21. Ti alloy shape with a deep pocket made by laser addi-tive manufacturing.

example of a part made by laser additive manufac-turing is shown in Fig. 21.

Finally, the use of high strength Ti alloys forhelicopter rotors has enabled the operation of theaircraft at higher gross weights than was originallyintended. In one example, the rotor material of aWestland helicopter was changed from Ti-6-4 tothe higher strength Ti-10-2-3 b alloy which permit-ted the aircraft gross weight to be increased from~3860 to 5585 kg. This provided major customervalue by avoiding development of an entirely newsystem. These rotors consist of three individualforgings as is illustrated in Fig. 22. The fatiguesensitivity of helicopter rotors is extreme, thus theavailability of high strength b alloys was the keyelement in this system modification.

In summary, Ti alloy usage in aircraft, expressedas a percentage of empty weight, is increasing rela-

Fig. 22. Drawing of a helicopter rotor assembly showing threeparts that were changed from Ti-6-4 to Ti-10-2-3 to enable agross weight increase.

tive to Al alloys and steel with each new productgeneration. For example, the original Boeing 747-100 contained about 2.6% Ti whereas the Boeing777 contains about 8.3%. This usually adds to theinitial aircraft cost which creates the need to dem-onstrate a commensurate contribution to customervalue. The decision to use Ti alloys, therefore,must be determined by the trade-off between costand customer value. Weight reduction, lower main-tenance cost and improved reliability all areaspects of customer value that may be used to jus-tify the use of Ti alloys.

2.2. Evolution of propulsion systems and the roleof materials

In a similar fashion to the evolution of thedesign methods for airframes, design methods foraircraft engines also have evolved and improved.The improvements are based both on lessonslearned from field experience and on the avail-ability of better computer based design tools.Materials with improved capability also arerequired to enable realization of the full potentialof these designs. In the case of engine materials,there are several discrete materials requirementsthat drive changes in materials or processingmethods: higher strength, better damage tolerance,absence of material defects and, in some cases,higher temperature capability. In terms of percentby weight, the principal materials of constructionfor gas turbine engines are Ti alloys and Ni basesuperalloys. High strength steel is used for themain shaft and for bearings and there also are afew applications of polymer matrix-carbon fibercomposites. Due to length limitations, the focus ofthis paper will be Ti and Ni alloys.

A modern subsonic aircraft engine for a passen-ger or transport aircraft consists of discrete sectionsor modules: fan, low pressure compressor (LPC),high pressure compressor (HPC), combustor, highpressure turbine (HPT) and low pressure turbine(LPT). Generally, the fan, LPC and about 2/3 ofthe HPC are made from Ti alloys, whereas the bal-ance of the HPC, all of the combustor and both theHPT and the LPT are made from Ni base alloys.This transition from Ti to Ni alloys is dictated bythe respective operating temperature capability of


these materials. An example of such an engine isshown in Fig. 23. These engines can be very large;the fan diameter of the engine on the Boeing 777-300LR is �3 m and the engine weighs �7300 kg.

While weight is important for aircraft engines,the options available to reduce weight are fewercompared to aircraft because of the high loadsplaced on the major components during service andbecause of the high operating temperatures. Themost common weight-related figure of merit forengines is the ratio of thrust to weight, also calledspecific thrust. Therefore, higher specific thrustachieved, for example, by increasing maximumoperating stresses or temperatures often is con-sidered as important as weight reduction per se.Particularly in commercial engines, reliability anddurability have become key product characteristics.The trend in large commercial aircraft favors twinengine designs, such as the Boeing 767, the Airbus310, the Airbus 330 and the Boeing 777. This cre-ates a requirement for a high degree of enginereliability because these aircraft all operate on longover-water routes. The US Federal AviationAgency and the European Joint Aviation Authorityhave explicit criteria for certifying anairplane/engine combination for such routes. Therating is called Extended Twin-engine Operations(ETOPS). The ETOPS rating is expressed inminutes and higher ETOPS ratings enable a longerthe over-water portion of the total flight. HigherETOPS ratings make more direct routes to beflown. This impacts total fuel consumptions andflying time, both of which are important to the air-

Fig. 23. Large commercial turbofan engine showing the sixmodules.

lines. In-flight shutdowns of engines negativelyimpact the ETOPS rating. Therefore, materialsrelated in-flight shutdowns are unacceptable. Thiscreates the requirement for a high degree of confi-dence before a new material is introduced intoan engine.

Durability, especially in the turbine section, dic-tates the number of hours an engine remains in ser-vice before it must be removed for maintenance.Over the past 30 years, the durability of aircraftengines has improved dramatically, especially inthe case of commercial products. Improved dura-bility means better utilization of the aircraft andtypically lower operating costs, both of which areimportant to the operator. When the Boeing 707first entered service in the 1950’s the engines typi-cally were removed for maintenance after about500 hours of operation. Much of the need forremoval was related to performance deteriorationof the HPT. Today, a Boeing 747 class engineremains on wing for more than 20,000 hours. Thisremarkable improvement is in part due to morerobust designs and in part due to better materials.When it also is considered that engine operatingtemperatures have increased significantly toimprove fuel economy, the impact of bettermaterials becomes even clearer. Finally, it is essen-tial that engine materials, especially those used inthe rotating parts of the engine be free of anydefects from melting or forging. From a practicalstandpoint, it is physically impossible to design anengine that can contain a burst rotor. Therefore theintegrity of the rotors becomes a matter of seriousconcern with regard to flight safety. Better damagetolerance coupled with intense efforts to eliminateboth intrinsic material defects and manufacturinginduced defects has led to major improvements inrotor integrity over the past 20 years. The efforts toeliminate materials defects represent contributionsfrom many specialties ranging from melting andforging to non-destructive evaluation. Improve-ments in damage tolerance has come from animproved understanding of microstructure-pro-perty relationships in Ti and Ni alloys much in thesame way as has been described previously forAl alloys.


2.2.1. Evolutionary improvement in Ti alloys foraircraft engines

Because of excellent specific strength, gooddamage tolerant properties, temperature capabilityup to 600 °C and well-established manufacturingprocess capability, Ti alloys make excellentmaterials for the cooler portions of aircraft engines.Especially in the fan, where the stresses on the diskare very high, it is essential to use premium qualityTi alloys to minimize the incidence of materialdefects. There are several types of materials; thosethat are melt related and those that are caused dur-ing forging. Among these, the most serious are themelt related defects, especially the interstitial stabi-lized inclusions, called hard a or Type I defects.These defects are essentially hard, brittle inclusionscontaining as much as ~10wt% nitrogen, typicallyin the form of TiN. Because of the brittle natureof the inclusions and the high operating stresses inthe rotors, hard a is essentially an incipient crackwhich begin to propagate starting with the firstcycle on the engine. A great deal of effort has beenexerted to minimize the occurrence of these defectsand the rate of occurrence now is on the order of1 per million kg of rotor grade material produced.This represents a 5–10× reduction over 35 yearsago.

Ti alloys also are costly and therefore it is com-mon to re-use turnings and chips generated whenforgings are machined to create the final Ti alloycomponents. This practice, which is common in theindustry and which is an economic necessity, cre-ates the possibility of accidentally incorporatingWC inclusions from broken machine tools into thefinal material. These WC, melt-related defects arecalled high density inclusions (HDI). Strict man-agement of the turnings and chips, including 100%radiographic inspection, has reduced the incidenceof HDIs to relatively low levels also.

Driven by concern over the consequences ofmaterials defects, production methods for premiumquality or rotor grade Ti alloys has evolved overthe past 40 years through a number of incrementalchanges. These changes include sponge quality,handling of scrap, VAR electrode preparation, hottop practice during melting and melt rate control.The attendant increase in material reliability haspermitted designers to use higher operating stresses

than were initially the case. In the past 5 years amajor shift in Ti alloy melting practice hasoccurred. Historically rotor grade Ti alloys weremelted by vacuum arc re-melting (VAR) and thenre-melted two additional times to produce sound,homogeneous ingots. The product of this process iscalled triple melt VAR materials. Today, a processknown as hearth melting is being used to producemuch of the rotor grade material for jet engines.Hearth melting is done in a furnace that uses awater-cooled copper hearth and either plasmatorches or electron beam guns as the source of heatfor melting. The heat extraction from the hearth iscarefully managed to ensure that a thin layer ofsolid Ti alloy (called a skull) remains and is indirect contact with the hearth. The skull separatesthe molten Ti alloy and the copper hearth. Thusthe Ti alloy being hearth melted only contacts Tialloy during the time it is in a molten state. Thematerial to be melted is fed into one end of thehearth, is melted by the heat sources, and flowsinto a water cooled mold at the other end of thehearth. When turnings are used, any WC particlesink and are trapped in the skull and cannot beincluded in the final ingot. Hearth melting andVAR melting including schematic illustrations aredescribed in more detail elsewhere [25].

Other types of material defects also are observedin forged Ti alloy products, either billet or finalforgings include b flecks, strain induced porosityand blocky a. Each of these has the potential toreduce the fatigue life of the material and none areallowed in rotor grade material. Among these, bflecks are caused by solute segregation during sol-idification of the ingot, strain induced porosity iscaused by improper mill practice when the ingotis converted to billet and blocky α is the result ofimproper reheat practice during the final forgingoperation. The defects just described can occur inessentially any of the commonly used a + b Tialloys, although the propensity for b flecks is morepronounced in those alloys containing b eutectoidalloying elements such as Cr, Fe or Cu. The useof modeling of ingot solidification and billet con-version processes has been helpful in identifyingprocess parameters that eliminate these defects. Ifthe process windows that have been defined bymodeling are broad to be practical and are carefully


followed, these defects can be eliminated. This rep-resents an important accomplishment in improvingmaterials reliability for jet engines and other com-parably demanding applications.

Jet engine rotors are manufactured from forgingsor rolled rings. The fan disk is typically a large,single-piece forging. An example of a fan diskafter final machining is shown in Fig. 24. From aweight standpoint, it is beneficial to use a higherstrength alloy for the fan disk and the two mostcommon alloys (in addition to Ti-6-4) are Ti-5Al-2Zr-2Sn-4Cr-4Mo (Ti-17) or Ti-6Al-2Sn-4Zr-6Mo(Ti-6-2-4-6). The limiting properties for the fandisk (at a given strength) are low cycle fatigue(LCF) and fatigue crack growth. Attention to pro-cessing can optimize these particular properties andb forging is now commonly used to create a Wid-manstatten microstructure. This structure reducescrack growth without an unacceptable penalty inLCF life. Ten years ago it was not possible to dothis because the control of the forging process wasnot adequate to provide reproducible structures. Incontrast, the compressor rotor, also called a spool,is a multi-stage component frequently made fromseveral forgings or rolled rings and joined togetherbefore machining. The compressor spool can incor-porate as many as 7 compressor stages into a single

Fig. 24. Photo of a forged and machined Ti alloy fan disk fora large commercial aero-engine.

component, as shown in Fig. 25. The compressorspool is typically Ti-6-4 in the first 5 stages withthe last 2 stages being a higher creep strength alloysuch as Ti-6Al-2Sn-4Zr-2Mo+Si (Ti-6-2-4-2S).The front stages of the spool are LCF limited andthe final 2 stages are creep limited. Here, the Ti-6-4 stages are a + b forged and the Ti-6-2-4-2Sstages are b forged and the pieces are joined byinertia (friction) welding. The benefit of the spoolconstruction is the absence of bolted joints betweenstages with a resulting lower risk of fatigue crackinitiation at bolt holes. The final stages of the HPcompressor rotor are made from Ni base alloysbecause the operating temperatures exceed thecapability of Ti alloys. The use of Ni base alloysin jet engines will be discussed in a later section.

In addition to the highly stressed rotors, the fanblades and at least 6–8 stages (depending on theparticular engine) of the compressor air foils alsotypically are made from Ti alloys. These compo-nents are typically life limited by high cycle fatigue(HCF), although resistance to impact damage fromforeign objects also is important. The HCF strengthof Ti alloys usually scales with the yield stress,so the use of higher strength alloys than Ti-6-4 isattractive in principle. In practice, the difficulty of

Fig. 25. Photo of a Ti alloy seven stage HPC spool after forg-ing, inertia welding and final machining.


manufacturing net shape air foils from higherstrength alloys is significantly greater and thedecision to use a higher strength alloy representsan economic trade-off that is ultimately made onthe basis of customer value. Thus most of the fanand compressor air foils in service in commercialengines today are made from Ti-6-4. There aresome exceptions driven by severe fatigue sensi-tivity or by the occurrence of aero elastic exci-tations. In the former case, the higher strength alloyTi-4Al-4Mo-2Sn (known as IMI-550) is used andin the latter case, the higher modulus alloy Ti-8Al-1Mo-1V (Ti-8-1-1) is used. Increasing Al contentraises Young’s modulus of α+β Ti alloys and thischanges the natural resonant frequency of the airfoil. The latest generation of very large engines forthe Boeing 777 (cf Fig. 23) has such large fans thatweight becomes an issue. Consequently, these fanblades are either a hollow Ti alloy part made bysuper plastic forming and diffusion bonding or aremade from solid polymer matrix carbon fiber com-posites. Both of these solutions are expensiveresponses to design requirements, but are the onlyavailable means of creating very large diameterturbofan engines.

Ti alloys are also used for static components injet engines. Included are frames, casings, mani-folds, ducts and tubes. The largest use today is castframes because they can be produced in near netshape and can replace a fabricated componentcomprised of dozens of individual parts. Eventhough the castings are relatively expensive, theyrepresent a significant cost reduction compared tothe fabricated part. A typical cast Ti alloy frontframe is shown in Fig. 26. As was discussed in theairframe case, these castings are HIPd to eliminateany porosity. In many instances, the engine mountis an integral part of these frames so having repro-ducible property values is important to assureintegrity of the design.

2.2.2. Improved Ti alloys for new propulsionsystems

By comparison to Al alloys, there has been lesseffort devoted to developing and commercializingnew Ti alloys and very little success in introducingnew alloys into aircraft engines. There are severalreasons for this. First, the volume of Ti alloys used

Fig. 26. Photos of two cast Ti alloy engine frames showingthe detail that can be achieved by the casting process.

is small compared to that of Al alloys, so amortiz-ing development cost is more difficult. Second, thequalification cost for a new rotor alloy is very highbecause of the inherent risk of rotor failure. Thiscost–risk combination becomes an unavoidabledeterrent to introduction of a new alloy. Third, theapplication requirements typically preclude realiz-ation of any benefit from improvement of a singleproperty. A possible exception to this situationwould be the availability of an alloy with signifi-cantly higher temperature capability. Such an alloycould displace one or more Ni alloy stages fromthe rear of the compressor with a significant weightbenefit. A relatively new alloy Ti-5.8Al-4Sn-3.5Zr-0.5Mo-0.7Nb-0.35Si-0.06C (IMI 834) with a 50 °Cincreased temperature capability has beendeveloped in the UK. This alloy is attractive tech-nically but there is only one IMI834 producerworld-wide which makes it commercially unattrac-tive. Consequently, it has not been widely adopteddespite its attractive properties.

In addition to conventional a + b Ti alloys, alot of effort has been devoted to developing Ti-based intermetallic compounds for high tempera-ture applications. These compounds are based onthe phases Ti3Al, Ti2AlNb and TiAl. While eachof these has the potential for higher temperatureuse, there are issues associated with low ductility,environmental sensitivity and cost that have pre-vented their application until now. This situation is


analogous to the revolutionary Al alloy technologydescribed earlier. There is reason to believe thatlimited applications of TiAl based materials willoccur when the need for higher temperature, lowdensity alloys is great enough. Market pull andcustomer value will be the determinant here, notmaterials technology.

2.2.3. Evolutionary improvement in Ni alloys foraircraft engines

At temperatures above ~550 °C, Ni base alloysare used instead of Ti alloys. This means the rearof the compressor, the combustor and the entireturbine sections are made of Ni base alloys. Earlyin the development of the jet engine, the tempera-ture capability of the early generation Ni basealloys was the main limitation in achieving higherperformance through increased operating tempera-tures. The current alloys have minimized this limi-tation although newer designs could benefit fromhigher temperature capability than is availabletoday, especially at the rear of the HPC. It is com-mon in rotor design to restrict the operatingstresses to levels that eliminate creep as a life limit-ing consideration. Consequently, LCF life andfatigue crack growth become the limiting proper-ties just as in the case of Ti alloy rotors. Over thepast 25 years, Ni base rotor alloys have evolved interms of temperature capability. This is shown inFig. 27. This figure compares the creep strength oftwo classes of Ni base alloys; those produced byingot metallurgy (IM) and those that are producedby powder metallurgy (PM).

The IM alloys generally contain lower total con-

Fig. 27. Comparison of the creep strength of five Ni base diskalloys. Time is for 0.2% strain at 650 °C at a stress of 800 MPa.

centrations of alloying additions and have lowercreep strength than PM alloys. They also are sig-nificantly less expensive to make and fabricate. Aswas the case for Ti alloys, melting technology foringot metallurgy Ni base alloys also has evolvedover the past ~35 years. Today triple melting iscommonly practiced to ensure composition andinclusion control and homogeneity of the final pro-duct. In the case of Ni base alloys, triple melt prac-tice starts with a vacuum induction first melt, fol-lowed by an electroslag second remelt and a finalVAR melt. This practice has led to a majorreduction of melt related defects such as frecklesand white spots and large inclusions that were thecause of rotor failures during the early days of Nibase alloy rotor development. As a consequenceof these improvements, with the exception of theinherent temperature limitations of IM alloys, thesematerials are excellent in terms of reproducibilityand reliability.

The use of powder metallurgy (PM) technologyto make turbine disks also has evolved since itsinception ~25 years ago. Alloys are made intopowder if they contain such high concentrations ofsolute that large (�30 cm diameter) ingots cannotbe cast without having unacceptable levels offreezing segregation. The use of PM methods forproducing rotor alloys solves this problem butrequires the utmost care. For example, a highlydisciplined powder handling process is essential toavoid accidental incorporation of unwanted con-taminants that can have deleterious effects onfatigue behavior. A single, large inclusion in a highstress region of a rotor can cause a rotor burst withdisastrous consequences.

A typical processing sequence is as follows:

1. Inert gas atomization to form powder;2. Screening of the powder to a predeterminedmaximum powder particle size, typically �270mesh (44 µm);3. Vacuum degassing of the powder4. Placing the powder in an extrusion can5. Extruding the powder into a billet using anextrusion ratio that assures full density;6. Forging the billet into a final forged part, oftenby an isothermal, hot die forging process.


One reason to include this process outline is toemphasize the cost of using PM disks.

Ni base alloys contain reactive elements such asAl and refractory element additions for solid sol-ution and precipitation strengthening. Thesereactive elements can form small oxide particlesduring inert gas atomization, even under the mosttightly controlled circumstances. These small oxideparticles can serve as premature fatigue crackinitiation sites. Because they are not uniform insize and are not uniformly distributed throughoutthe volume of the disk, a different methodology isrequired to calculate the cyclic life as a functionof stress distribution. This requirement has led todevelopment of an elaborate probabilistic approachto accounting for the effect of inclusions on fatiguelife. Use of this method is necessary but also cre-ates an added cost, even after the initial data basehas been generated. The remaining aspect of Nibase rotor alloy performance is the effect of grainsize on critical properties including tensilestrength, creep strength, LCF life and fatigue crackgrowth rate. The impact of grain size on each ofthese properties is shown in Fig. 28, which is aschematic derived from actual data. From this dia-gram it can be seen that LCF life and strength haveopposite grain size dependence to creep strengthand fatigue crack growth. Thus a processingmethod must be chosen that suits the particular lifelimiting property for the design under consider-ation. This ability to tailor properties is a potentialbenefit and the current level of understanding of

Fig. 28. Schematic drawing based on real data showing effectof grain size on creep, low cycle fatigue life, fatigue crackgrowth rate and tensile strength.

property trade-offs is the cumulative result ofnumerous studies by many investigators.

Finally, the HP turbine air foils are perhaps themost critical component in a modern gas turbineengine, in terms of both time between overhaul andperformance. Creep strength and oxidation resist-ance usually are the life limiting properties of tur-bine air foils. Turbine air foil production tech-nology also has evolved. The early airfoils wereforged but the creep strength of the forged materialwas a severe limitation. Today, these air foils arecast Ni base alloys made by a highly sophisticatedcontrolled solidification process. The processingevolution for turbine air foils can be traced fromequiaxed fine grained forgings to coarse grainedequiaxed castings to a columnar grain structureproduced by directional solidification and finally toa monocrystal structure which is essential free ofhigh angle grain boundaries. The creep mechanismof grain boundary sliding limits the temperaturecapability of the forged and equiaxed cast products.This mechanism is minimized in the columnargrain structure but still occurs. Grain boundarycracking is an issue in Ni base alloys but can bemitigated with alloying additions such as B, Zr andHf. These elements are necessary but are not help-ful to the creep strength. In the monocrystal struc-ture, there are no large angle grain boundaries toundergo sliding or to crack so the rate controllingmechanism shifts back to intragranular flow byclimb and glide of dislocations. Monocrystal alloysalso require minimal or no grain boundarystrengthening additions, which also is beneficial tocreep strength. The macrostructures of the threetypes of cast air foils are Fig. 29.

Alloy composition also has a major influence onboth creep and oxidation behavior. The combinedeffects of alloy composition and casting method ontemperature capability are shown in Fig. 30. In thisfigure, each generation of alloy, e.g., N4 to N5 con-tains a higher concentration of slow diffusingrefractory alloying elements which reduces thecreep rate, especially in monocrystals. An interest-ing comparison in the figure is N4 vs. DS R142,which shows that composition effects can be aspowerful as grain structure effects. Directionallysolidification process is more economical than thatused to create monocrystals, therefore DS airfoils


Fig. 29. Photos of three cast Ni base turbine blades macro-etched to showing the effect of the solidification process on thegrain structure. (Courtesy Howmet Corporation).

Fig. 30. Comparison of the creep strength of eight Ni baseturbine blade alloys. Temperature is for constant rupture timeat the same stress.

are still used in less demanding application suchas the second stage of the HP turbine and in theLP turbine.

These cast air foils are often air cooled usingbleed air from the compressor. It is now becomingcommonplace to apply a 125 µm thick ceramiccoating on the surface of the air foil. This coatingis called a thermal barrier coating (TBC) andreduces the heat flux through the air foil wallwhich permits greater gas-metal temperature dif-ferences to be maintained during engine operation.Modern, high performance gas turbines typicallyoperate at maximum gas temperatures at the tur-

bine entrance of 1450–1500 °C. This is consider-ably higher than the metal capability but is possiblebecause of the use of air cooling (Fig. 31).

3. Closure

The availability of improved materials hasenabled the continuous improvement in capabilityof aircraft and aircraft propulsion. In combination,enhanced materials capability, improved materialsprocessing methods and more efficient designmethods account for the existence of modern air-craft that can fly more than 20,000Km non-stop.Concurrently, the noise and emissions from currentgeneration commercial aircraft are considerablylower than ever before. As the capability of mod-ern aircraft has been enhanced, the simultaneousimprovement in reliability of propulsion systemshas evolved. The current high levels of reliabilityin current propulsion systems has enabled twinengine aircraft to fly long over water routes thatwere previously reserved for three and four engineaircraft. This has added considerable customervalue.

Improvements in materials will continue to bemade possible by new computation methods,advances in modeling and simulation both for alloyand process design, new sensing devices for pro-cess control, and the ability to tailor materials tospecific applications. There will also be changes instructural concepts directed towards lowering thecosts of manufacture and saving weight. Theseinclude integral stiffening/iso and ortho grid con-

Fig. 31. Cross section of a turbine blade with the TBC on itshowing the columnar structure of the TBC, the bond coat andthe diffusion zone in the bond coat base metal interface.


struction, precision structural castings, superplas-tically formed parts, and welded structures. Newmanufacturing methods may require modificationin existing materials or the development of newmaterials that would be conducive to these newtechnologies. Economic considerations haveresulted in current aircraft operating well beyondtheir original design life bringing to the forefrontissues of materials stability, corrosion resistance,fatigue behavior, and maintenance procedures.Evaluation of new materials, either for retrofittingolder aircraft or for new systems, must include thecosts associated with their qualification, as well asan assessment of performance and operating costsover the entire life cycle of the aircraft, from fabri-cation to maintenance.

While the focus of this paper is on aircraft andaircraft engines, many of the messages also pertainto automobiles in a general way. The need forimproved fuel economy, driven by government andby competitive pressures, has created a heightenedinterest in lightweight automotive structures.Clearly, any experience from the aircraft industrywill need to be viewed in light of the differencesin requirements of the automotive industry. Centralto these differences are expected product lifetimeand cost. The added cost to reduce the weight ofan airplane is amortized over more than a 30 yearlife whereas for an auto it is more like 5–10 years.The possible exception to these significant differ-ences, and the product which has the closest paral-lel to aircraft requirements, is over large highwaytrucks. In the case of trucks, the differences com-pared to aircraft are smaller because of the longerproduct life, higher duty cycle and the fixedmaximum weight, at least in the US, which pro-vides a direct economic incentive to pay more forlighter vehicles. A similar situation could evolvefor autos if the government mandated fuel econ-omy standards become more stringent. Fuel pricesare increasing but, unto itself, this has proven tobe insufficient motivation to pay much more fora lighter auto. Speculation about the influence ofregulatory pressures on product trends has beenhistorically ineffective, largely because the regulat-ory trends have not been introduced on the righttimescale and because they have not always exhib-ited constancy of purpose.

Nevertheless, the wealth of experience availablefrom the aircraft industry can be valuable asdemands for lighter structures in a variety of pro-ducts becomes more important.

Acknowledgements

One of us (EAS) acknowledges support of theAir Force Office of Scientific Research underGrant S49620-01-1-0090. the other of us (JCW)acknowledges support from NASA under theURETI for Propulsion and Power.

References

[1] Smith RK. The quest for excellence. In: Greenwood JT,editor. Milestones of Aviation. New York: CrescentBooks; 1991. p. 222–96.

[2] Starke Jr. EA, Staley JT. Application of modern aluminumalloys to aircraft. Progress in Aerospace Sciences1996;32:131–72.

[3] Hyatt MV, Caton R, Lovell D. Advanced materials devel-opment in commercial aircraft, paper presented at theAIAA/AHS/ASEE Aircraft Design, Systems and Oper-ations Conference, Seattle, WA, AIAA, WashingtonD.C., 1989.

[4] Hyatt MV, Axter SE. Aluminum alloy development forsubsonic and supersonic aircraft. In: Herano K, OikawaH, Ikeda K, editors. Proceedings of the International Con-ference on Recent Advances in Science and Engineeringof Light Metals. Tokyo: Institute of Light Metals; 1991.p. 273–80.

[5] Hornbogen E, Starke Jr. EA. Theory assisted design ofhigh strength low alloy aluminum. Acta Metall Mater1993;41:1–16.

[6] Quist WE. Effect of composition on the fracture propertiesof aluminum alloy 7178, MSc Thesis, University of Wash-ington, 1963.

[7] Bucci RJ, Nordmark G, Starke EA, Jr. ASM Handbook,Vol. 19, Fatigue and Fracture; 1966, p. 771.

[8] Hyatt MV. Program to improve the fracture toughness andfatigue resistance of aluminum sheet and plate for airframeapplications, Boeing Commercial Airplane Company,Technical Report AFML-TR-73-224, Wright-PattersonAir Force Base, 1973.

[9] Warner TJ, Shahani RA, Lassince P, Raynaud GM,Aluminum alloy developments for affordable airframestructures, Paper presented at the ASM International Con-ference, Paris, France, June, 1997.

[10] Dif R, Bechet D, Warner T, Ribes H. 6056-T78: A cor-rosion resistant copper-rich 6XXX alloy for aerospaceapplications. In: Sato T, Kumai S, Kobayashi T, Murak-


ami Y, editors. Aluminum Alloys, Their Physican andMechanical Properties—Proceedings ICAA6. Tokyo,Japan: The Japan Institute of Light Metals; 1998.

[11] Behaghel P-F, Starke EA, Jr., The effects of processingand microstructure on two aluminum alloys for aerospaceapplications, Journal of Advanced Materials, (in press).

[12] Troeger LP, Starke Jr. EA. Microstructural and mechan-ical characterization of a superplastic 6XXX aluminumalloy. Materials Science and Engineering2000;A277:102–13.

[13] Wadley HNG, Fleck NA, Evans AGI. Fabrication andstructural performance of periodic cellular metal sandwichstructures, Composites Science and Technology, (inpress).

[14] Hinrichsen J. Airbus A3XX: Design features and struc-tural technologies review, Presented at the InternationalSchool of Mathematics, G. Stampacchia, 28th Workshop:Advanced Design Problems in Aerospace Engineering,Galileo Galilei Celebrations, Eric-Sicily, Italy, 11–18July 1999.

[15] Hinrichsen J. The material down-selection process forA3XX, Key Note Address at the 19th European Confer-ence on Materials for Aerospace Applications: Confeder-ation of European Aerospace Societies (CEAS) Munchen,Germany, 6–8 December 2000.

[16] Hinrichsen J. Airbus A380: Requirements for the selectionof materials and manufacturing technologies, Presented atAluminum in the Aerospace Industries, Gosselies,Belgium, 8 March 2001.

[17] Starke Jr. EA. Aerospace materials for future technologies.Journal of the Japan Institute of Light Metals2001;51:77–85.

[18] Polmear IJ, Couper MJ. Design and development of anexperimental wrought aluminum alloy for use at elevatedtemperatures. Metall Trans 1988;19A:1027.

[19] Zhu AW, Csontos A, Starke Jr. EA. Computer experimenton superposition of strengthening effects of different par-ticles. Acta Materialia 1999;47:1713–21.

[20] Zhu AW, Starke Jr. EA. Strengthening effect of unshear-able particles of finite size: A computer experimentalstudy. Acta Materialia 1999;47:3263–9.

[21] Zhu AW, Gable BM, Shiflet GJ, Starke Jr. EA. The intelli-gent design of age hardenable wrought aluminum alloys.Advanced Engineering Materials 2002;4:839–46.

[22] Kattner UR. Journal of Metals 1997;12:14.[23] Dubost B. Rev Metall 1993:195.[24] Kresse G, Furthmueller J. J. Comput. Mater. Sci. Vol-

ume 1996;6:15.[25] Lutjering G, Williams JC. Titanium. Berlin: Springer-Ver-

lag, 2003.

Aerospace Struct Mater

Documents

Transcript of Aerospace Struct Mater