Hybrid Metal Laminate (HML) Manufacturing Planning

Evaluation and Assessment Last Updated 2008/12/19

Charles Pitzer and Jenn Ming Yang Department of Materials Science and Engineering

University of California Los Angeles, 90095

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I. Introduction Even though European manufacturers have already incorporated glass-epoxy-aluminum

hybrid laminates into existing aircraft – namely, the Airbus A380 – a number of hurdles

impede the adoption of these advanced materials in the United States. The current US

industrial sector for producing this next-generation material is still in its infancy and is

capable of only limited production runs of secondary structure components. Few US

aircraft manufacturers are seriously considering use of hybrid laminates due to the

extensive research and development, capital investment, and FAA certification costs. In

short, the US is far behind European suppliers and aircraft manufacturers in

manufacturing capabilities, design standards, analysis techniques, QA testing standards,

in-service NDE testing, FAA qualification, and flight history with respect to hybrid fiber

metal laminates.

This paper will provide an assessment of the current industrial sector’s hybrid fiber metal

laminate manufacturing capability within the United States based on existing companies

with HML production histories. Next, we’ll evaluate the hurdles to transitioning the

technology to a production ready stage and the feasibility of doing so. Finally, we’ll

provide a future effort roadmap to the development of HML to enable it to be a viable

competitor to incumbent AL alloy technologies and other next-generation AL alloys

throughout the aircraft industry.

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FIGURE 1. Fatigue crack growth rates of aluminum 2024-T3 and some hybrid metal laminates. Reference [5].

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II. Background

Fiber-reinforced metal laminates (FML) are hybrid composites consisting of alternating

thin layers of metal sheets and fiber-reinforced epoxy prepreg. The most commonly used

metal for FML is aluminum, and the fibers can be Kevlar or glass. In Europe, FML with

glass fibers (tradename GLARE), and Kevlar fibers (tradename ARALL) have been

employed in new applications for aircraft structures. These laminates possess some of the

most desirable properties of both metal and fibrous composite materials. These hybrids

exhibit a number of major advantages over conventional AL alloy structural materials.

Most importantly, this family of fiber metal laminates (FML) results in an ability to

impede and arrest crack growth caused by cyclic loading that is significantly better than

all materials currently on the market - and even surpasses high-fatigue Li-Al alloys that

are currently in development.

Figure 1 shows the relative fatigue life performance of similar thickness specimens of

hybrid metal laminates and their monolithic aluminum counterparts. Most striking is the

almost linear crack growth progression of the hybrid metal laminate specimen versus the

exponential crack growth progression of the monolithic aluminum. The catastrophic rate

of crack growth in the monolithic AL alloy is what leads to the large burden of scheduled

inspections, airframe repair, and special inspections (currently ~37% of USAF man-hours

13). In 2006, the USAF spent 87 percent more on aircraft maintenance than it did in 1996,

while fleet availability has declined 13. Clearly, this is an accelerating problem. On the

other hand, even after a detectable crack is discovered in hybrid metal-glass fiber

laminates, there are a number of more opportunities to discover and repair this defect

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before catastrophic failure occurs - allowing for a significant increase in inspection

intervals and depot maintenance schedules.

Contrary to other "pure" laminated composites, the impact damage tolerance of hybrid

metal-glass fiber laminates are at least as good as monolithic aluminum alloys at low

velocities. For higher velocity impacts, hybrid metal-glass fiber laminates outperform

aluminum due to the strain rate effect and the increased strain hardening effect caused by

the glass fibers.

Corrosion resistance is enhanced by preventing through-thickness corrosion modes. The

prepreg layers act as moisture barriers between the various inner aluminum layers,

whereas the metal layers protect the fiber/epoxy layers from picking up moisture. The

aluminum layers also prevent UV radiation from degrading the glass-epoxy layer. Given

that these aluminum layers are much thinner than the conventional aluminum sheet that is

currently used, the quenching step in the tempering process occurs much faster - allowing

for less diffusion of alloying elements to the grain boundaries than in the thicker sheets -

thus resulting in even more corrosion resistance 8. Machine design article: 5% of all metal

products are lost due to corrosion each year. FMLs improve corrosion resistance.

Even though some specialized tooling and processing with fiber metal laminates will be

necessary, many of the same material handling and construction techniques used in

monolithic metals can be applied. Milling, drilling, sawing, joining, etc. are similar to

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conventional practices for metals. Conventional lay-up methods for single-sided molds

for vacuum-forming are also directly applicable to HML lay-ups. Furthermore, this

material can be produced as sheet material, but it also can be cured in an autoclave as a

complete structure, e.g. a large curved panel with co-cured doublers and stiffening

elements. As a result, The HMLs are also attractive hybrid materials for light-weight,

fatigue critical large-scale structural applications.

Finally, lower weight leads to the so-called “snow ball” effect. If, for an equivalent

section, we can use a material that can even fractionally reduce weight, then less lift

needs to be generated, which reduces drag, which leads to less fuel consumption, thus

compounding further weight reductions. Conversely, existing aircraft can be retrofitted

to allow for larger payloads or longer flights – possibly increasing the capabilities of

existing aging aircraft to handle larger burdens.

Hybrid fiber metal laminates can be tailored to suit a variety of applications by varying

the fiber/resin system, the alloy type and thickness, stacking sequence, fiber orientation,

surface pretreatment technique, etc. The commercially available product forms of

GLARE laminates and their density are summarized in Table A. The density of the

GLARE laminates depends on the relative thickness of aluminum sheet and glass

fiber/epoxy layers, the number of layers in the laminate and the fiber volume fraction. In

all cases, the density of GLARE laminate is at least 8% lower than aluminum alloy.

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Reference [2]

A patent on Glare was filed on October 14, 1987 by AKZO with Roebroeks and

Vogelesang as inventors. A partnership between AKZO and ALCOA started to operate

in 1991 to produce and commercialize Glare. A formal agreement to form the Structural

Laminates Company (SLC), a joint venture of AKZO (1/3 owner) and ALCOA (2/3

owner), was signed on June 1, 1991. The agreement provided for production to be

concentrated in New Kensington, PA, while the research, development and marketing

would be done in Delft. For this reason, a separate company, a subsidiary of SLC with

the name Structural Laminates-bv was founded in the Netherlands, headed by Gunnink.

III. Current Industrial Sector Assessment Major hurdles do exist, however. In order to meet the market demand for a commercial

airliner, manufacturing capacity would need to be significantly expanded upon.

Processing of thin AL sheet, large scale production and handling of glass-epoxy prepreg,

automated layup, curing, and inspection processes would all need to be developed with a

significant capital investment. Perhaps more importantly, though, the material must be

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thoroughly qualified for aerospace service. This is an expensive and time consuming

endeavor that requires a compelling reason to proceed with.

For each flavor of glass-epoxy FML, hundreds of specimens from different production

runs will need to be tested for tensile strength, compression strength, shear strength,

fatigue strength, etc. Statistical variation must be characterized to anticipate any scatter

in performance. Each flavor of glass-epoxy FML can include variations in the

ingredients of metal layers (thickness, work hardening, temper, surface treatment), fibers,

adhesive, autoclave cycle, and the stretching process after the cure. All of these

variations must be carefully considered and controlled.

Since qualification of the material is started when it is clear that the material will be

applied in design, this leads to the “chicken before the egg” problem. Designers are not

inclined to design structures from materials that are not yet qualified. Nobody will

commit to qualify a material that has no impending design to propel the process. Plus,

there is no guarantee that the properties that have been reported for these materials can be

reproduced on a large commercial scale without significant opportunity for scatter in

those property values. Therefore, nothing happens.

The following discussion details the steps necessary to progress from secondary

structures, to doubler, repair, and reinforcement applications in primary structures, and,

finally, to full-scale production of primary structures such as new or replacement fuselage

and wing section applications. Although not ideally suited for all portions of the aircraft,

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there are a large of number applications that could benefit significantly from this next-

generation material.

Current usage in industry for fiber metal laminates includes applications such as ARALL

in the C-17 cargo door skins, GLARE 5 in the Boeing 777 cargo floor, HML in Fokker

F50 lower wing access panels, HML in T38 crown fuselage panels, ARALL in the C-130

flap skins, HML in De Havilland DHC-8 flap skins, and, significantly, GLARE in the

Airbus A380 horizontal and vertical tail plane leading edges in addition to GLARE in the

upper fuselage skin of the A380. Leading edges have been found to be appropriate

applications of GLARE (glass-epoxy fiber metal laminates) due to the favorable strain

rate effect during impacts. Southwest Airlines, in cooperation with Boeing, is currently

flight-testing GLARE flap skins on a 737.

A sampling of the current suppliers in the US industrial sector for producing Hybrid

Metal Laminates is given in Table B. Some infrastructure has been developed for short

production run retrofits and components – such as the examples given, above. The

preliminary list, below, is based on the current licensees of the GLARE brand products

and their limited production to date.

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MATERIAL/PROCESS SUPPLIER Per Spec

FM 94 Adhesive Cytec MIL-A-25463

FM 906 Adhesive Not commercially available

S-2 Glass Fibers Cytec

FM 94/S-2 Glass Prepreg Cytec AMS 3906/2

FM 906/S-2 Glass Prepreg Not commercially available

2024-T3 AL Sheet Alcoa AMS 4037 AMS-QQ-A-250/4A

7475-T761 AL Sheet Alcoa AMS 4085B

Phosphoric Anodize Metal Improvement (Pa) Aviation Equip. Corp (Ca)

ASTM D 3933

Chemical Conversion Coating Extensive List of Suppliers

MIL-DTL-5541 MIL-DTL-81706

Primer Application BR 127 or BR 6747-1

Extensive List of Suppliers

AMS 3107/2 MIL-F-18264

Secondary bonding AAR Composites (FL) GKN Westland (AL)

MIL-A-83377B

Repair/Rework/Retrofit GKN Westland (AL) Vought (TX) CTL Aerospace (OH) GD-ATP (VA) Alliant Tech Sys (UT) Hexcel Struct Prod (PA)

MIL-P-9400C MIL-HDBK-337

NDI Inspection NDE, Inc (TX) US Inspection Ser (OH) West-Pro (OR)

AMS-STD-2154 AMS 3920

Table B. Preliminary list of suppliers in the industrial segment of HMLs

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IV. Manufacturing Processes of GLARE GLARE laminates are manufactured by bonding together unclad metal sheets with fiber

composite prepreg using either a press or preferably an autoclave. Prior to bonding lay-

up, the metal layer surfaces are pretreated to improve the adhesion to the prepreg. After

the lay-up procedure, the laminate is cured in an autoclave. The adhesive system in

which the fibers are embedded contributes significantly to the performance of the

laminates. It also determines the bond strength between the fiber layers and metal layers.

The adhesive system that is used in GLARE is epoxy FM 94. Several GLARE variants

with different laminate lay-up and fiber orientations have been developed. For example,

a typical 2/1 lay-up consists of two layers of metal bonded by one layer of prepreg.

Thicker laminates are achieved by adding more layers of each constituent to form a “3/2”

or “4/3” lay-up. The laminate can be produced as semi-finished sheet material. Post cure

operations can be performed. These operations include milling, drilling, riveting, bolting,

and bending. Bending perpendicular to the glass fiber direction requires very generous

bend radii. Bending parallel to the fiber direction results in bend radii and spring back

angles comparable to the aluminum sheets used in the product. As such, hybrid metal

laminates can be press-formed into shapes as long as spring back is taken into account.

But these post-cure formed components will be pre-stressed and correspondingly

weakened. More desirable is for hybrid metal laminates to be cured in an autoclave into

a complete structure, i.e. a large curved panel with co-cured doublers and stiffening

elements. The development of “splicing concept” also allows the fabrication of a larger

panel size compared to conventional aluminum structures. In the spliced laminate, thin

aluminum sheets are laminated with a very narrow seam in between. The seams in the

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various aluminum sheets are at different locations such that they are bridged by both the

fiber layers and the un-spliced aluminum layers. Several concepts are presented in

Appendix A and one detailed manufacturing process is presented in Figure 9. The fiber

layers between the metal sheets bridge across the gaps, providing load transfer.

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Figure 2. Current manufacturing process for glass-epoxy fiber-metal laminates

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V. Current Status of Manufacturing Capacity in the US

The current industrial sector can provide a limited number of units of glass-aluminum

hybrid metal laminates every month based on short production runs by Aviation

Equipment Corp. Figure 2 illustrates the notable steps in the process. Each process step

is discussed and bottlenecks are identified and addressed.

Step 1: Consolidation of the S-2 glass fibers and the FM 94 adhesive into 0.127 mm

nominally thick sheets is currently automated. The FM 94 adhesive system cures at

120 °C. The S-2 glass fibers are approximately 10 µm in diameter. US production of

glass-epoxy prepreg is of 0.127 mm (.005 inch) thick sheet already relatively well-

established. Figure 3 roughly illustrates the prepreg process from the constituent FM 94

epoxy and S-2 glass fibers.

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Figure 3. Prepreg manufacturing process. After fabrication and rolling, glass-epoxy prepreg must be stored at or below -18 °C for the FM 94 adhesive system. Reference [7].

Figure 4. Current process for preprocessing aluminum sheet for bonding HMLs

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Step 2: Aviation equipment can process components and stock materials in etching and

anodizing tanks that measure 14’ long, 3’ wide, and 7’ deep. Only a limited number of

60in by 120in aluminum sheets can be processed in one day. This could be considered

one the limiting constraints (bottleneck), currently, on the production rate.

Step 3: Single-sided layup tools for producing the hybrid-metal laminate shapes use

similar techniques as conventional one-sided vacuum forming molds for other composite

layups. The technology is well-understood.

Step 4: Manufacturing of laminate panels of AL2024-T3 sheet alternating with S-2

glass/FM 94 epoxy prepreg sheet is a highly manual process. The operator is required to

meticulously lay the prepreg onto the aluminum layers – one layer at a time. Usually this

is done while the operator is laying on their stomach suspended over the tooling. This is

another limiting constraint in the process.

Step 5: For curing, the sub-assembly is then bagged and attached to a vacuum system in

the autoclave. Using the FM 94 adhesive system, the sub-assembly is heated to 250 °F

(120 °C) at a rate of 3-5 °F (1.7-2.8 °C) per minute. 40 psi of vacuum pressure is then

applied to the vacuum bagged setup 2. Then it is held for 60 minutes at 250 °F (120 °C)2.

The total processing time for this step is approximately 210 minutes.

Step 6: For large, flat panels, ultrasonic C-scan inspection is an accurate and appropriate

way for detecting cracks and delaminations1. However, it becomes more difficult to use

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this method for detecting defects in contoured panels. Smaller hand-held eddy current

testers, Fokker Bondtesters, and ultrasonic probes must be used ibid. The detectable flaw

size is directly proportional to the probe size – leading to very time consuming

inspections for large contoured panels.

Step 7: Panels of hybrid metal laminates are first cured into a contoured mold. After

curing and inspection operations are complete, the panel is then machined to the

necessary outer profile in addition to the fastener and cut-out holes.

Step 10: Many of the conventional methods for inspecting composite and aluminum

structures directly apply to inspecting hybrid metal laminates. Practical eddy current

methods for detecting sub-surface cracks using sliding probes such as the Nortec SPO-

2181 have shown a maximum reliable crack detection depth of 9-10 mm (12-13 AL

layers) 1.

At fastener row joints (lap joints and butt joints), where the maximum stress occurs in the

faying aluminum layers due to secondary bending, cracks initiate at the mating surfaces.

The low-frequency eddy-current technique can detect cracks ≥ 2 mm at a depth of 3 to 4

mm. At a depth of 5 to 6 mm, cracks of length ≥ 6 mm can be reliably detected. Greater

than a depth of 6 mm, cracks can no longer be reliably detected 1.

For testing delaminations, the Fokker Bondtester is preferred for detecting defects on the

order of 0.25 inch diameter at a depth up to 6 mm (~ 7 layers) and 1.25 inch diameter

defects at a depth of at least 25 mm (~ 32 layers). Ultrasonic testing allowed for

comparable results to the Fokker Bondtester method 1.

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VI. Possible Future Process of Manufacturing Hybrid Metal Laminates in the US

Figure 5. Potential future manufacturing process for glass-epoxy fiber-metal laminates

Figure 6. CNC Tape Laminate Machine. Reference [3].

Figure 7. Proposed aluminum sheet preprocessing continuous rolling system.

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Step 1: Based on current production rates, only limited opportunities for improving

capabilities of the US prepreg manufacturing sector exist. However, the FM 94 adhesive

could potentially have some limitations on service temperature. Hagenbeek reported a

glass transition temperature of as low as 67 °C 8. With a glass transition, a significant

departure in material mechanical properties is observed – which is strongly advised

against in any aerospace application. This falls well within the common service

temperature of a commercial airliner of -55 °C to 80 °C. However, it must be mentioned

that other sources report the glass transition and service temperatures of 80 °C (85%

RH)9, 103 °C 9, 104 °C 11, and 107 °C 10, with some numbers showing a dependence on

relative humidity. This would suggest that future work would include testing for the Tg

of FM 94 using differential thermal analysis (DTA), studying specific heats, CTEs,

specific volumes, or viscosities to determine a valid reportable number. Possible

replacements for the FM 94 adhesive system include FM 906 which has a 180 °C curing

temperature and a greater glass transition temperature. Other future work could include

investigation of ways to modify the epoxy systems to improve Tg performance.

Step 2: Figure 7 illustrates a proposed aluminum sheet preprocessing continuous rolling

system. Process control will be implemented by changing roller heights within the

respective processing tanks to reduce or increase exposure times.

Step 3: Existing technologies for single-sided vacuum molding will require little

transitional effort to apply to hybrid metal laminates. However, work-holding for

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performing precision operations such as laser welding or highly loaded operations such as

friction stir welding may require some limited low-level development.

Step 4: Existing tools for automated composite layups can be directly applied to hybrid

metal laminates. Prepreg tape laying machines can currently feed 3”, 6”, or 12” wide

tape onto a panel in X-Y-Z space – allowing for complex contouring of aerospace

structures. Different platforms are available in either high or low gantry configurations

allowing for dispensing rates of up to 2000 IPM – a theoretical maximum of

approximately 166 sq ft per minute. Figure 6 illustrates one potential automated CNC

tape laying system on a high gantry configuration.

Step 5: Machining contouring will take place using conventional 5-axis routers on a

gantry-type system. Figure 8 illustrates a potential system produced by Thermwood.

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Figure 8. Five-axis router for machine contouring of hybrid metal laminate formed panels. Thermwood C67-510DT 5-axis CNC router

Figure 9. One proposed manufacturing method for building and curing hybrid metal laminate panels. Doubler bonded to pre- or post-curing using an adhesive system such as FM73 adhesive film. Illustration of steps 4, 5, 6, and 7.

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Step 6: Splicing is of special concern since AL sheets can currently only be

manufactured with widths up to 1524mm (60in) to the required accuracy and with the

necessary nominal thicknesses between 0.3 and 0.5mm 6. This limitation would imply

the necessity of applying many costly mechanical longitudinal or circumferential joints in

an aircraft fuselage. Intersplicing would help avoid this disadvantage. Obviously, the

splices might cause a reduction in performance. According to Roebroeks, however, the

actual reduction in performance is quite small when loaded transversely to the splice line

5. Surprisingly, if the spliced panels are loaded in the direction of the splice lines (e.g.

circumferential hoop stress in a fuselage), the splices actually serve as crack stoppers in

both residual and fatigue tests ibid. It effect, it acts like a DCF (damage control feature).

Future work could include the development of design rules to properly place these splice

lines in their most beneficial configuration.

Figure 9 diagrams one potential process for a hybrid metal laminate panel build-up. First,

step 6a and 6b would rigidly and precisely position the sheets relative to one another with

a fixture for either a friction stir weld operation or a laser weld operation. Post-

machining of the weld bead may be necessary in order to ensure that the glass-epoxy

prepreg lays flush on the surface (step 6c). As seen in Step 4 for flat or contoured single

aluminum panel hybrid metal laminates, a CNC tape machine can be used for larger

panels that require splices. Step 6d would also require that the aluminum sheet be rigidly

fixtured for the routing operation. A small piece of sacrificial material would need to be

slipped under the edge of the sheet to be machined to prevent any damage to the glass

fibers – glass fibers have exceptional toughness due to their large strain to failure, but that

is significantly degraded when scratched. Step 6e would likely require an epoxy bond (or

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none at all) instead of anything hot enough to cause a fusion weld in the AL sheet since

the prepreg will be close by. However, one concept would be to control weld processing

temperatures tightly enough to avoid degradation of the underlying prepreg. Figure 10

shows the friction stir weld process. Relative to glass fibers, AL alloys have a relatively

low melting point (~520 °C). By ensuring that the heat affected zone (HAZ) is limited to

only the top fractional distance of the sheet, the aluminum, with a high thermal

conductivity, would adequately distribute the weld heat to the entire sheet without

affecting the underlying prepreg. Steps 6f through 6h would repeat the process for the

remaining layers of the requisite product. All along, the minimum splice spacing

distance would have to be maintained and staggered throughout the panel for minimum

effect on the longitudinal or transverse properties.

Production of 3.6m (914.4in) wide sheets is planned for the near future and will

significantly reduce the number of required splices.

This process would be similar for either new construction or for aging aircraft retrofits –

although retrofit patches likely wouldn’t exceed one panel in width. Retrofits can employ

portable cure equipment for performing the final step of the vacuum-bagged retrofit

process in situ. Typical cure temperatures for FM 73 adhesive film is 120 °C.

Alternative ideas for splice joints are presented in Appendix A.

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Figure 10. Friction Stir welding process for AL sheet. Reference [4].

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Step 12: Ultrasonic Inspection techniques could be improved by developing specialized

transducers to minimize the complicated interference patterns present when used on a

layered structure such as GLARE or other hybrid metal laminate. Even though the

Fokker Bondtester is reasonably reliable at all depths, the ultrasonic method for detecting

delaminations is especially promising due to its layer-by-layer accuracy for finding

defects.

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VII. Possible Future Process of Retrofitting Aging Aircraft with Hybrid Metal

Laminates in the US

Figure 11. Potential retrofit manufacturing process for glass-epoxy fiber-metal laminates

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Given in Figure 11 is a potential process for retrofitting existing aircraft that exhibit

fatigue cracking with patch plates composed of hybrid metal laminates. The smaller

scale of the physical components would enable an incremental increase in the

infrastructure capabilities of existing US manufacturers without a tremendous investment

in capital equipment and the increased liability of a new material without flight history.

Step 1: Manufacture of retrofit components could begin with the existing process for

manufacturing hybrid metal laminate composites detailed in Figure 2. Then

manufacturing infrastructure could gradually shift capabilities to the more advanced

capabilities and processes detailed in Figure 5 Potential future manufacturing process for

glass-epoxy fiber-metal laminates.

Step 2: Aluminum surface preparation begins with paint removal, cleaning, and abrading

of the surface seal and underlying fasteners per conventional processes used for existing

repair operations. Surface treatment then proceeds with a phosphoric acid anodize to

which a BR127 primer is applied and cured. BR127 is a chromate epoxy-phenolic primer

manufactured by Cytec.

Steps 3-6: The patch panels and retrofit components are then bonded onto the aircraft

using a structural film adhesive such as FM 73. The adhesive is supported by a polyester

knit fabric scrim that controls bondline thickness and flow during cure. Curing of this

bond is carried out at 115-125 °C with a bond line pressure of 240-310 kPa using

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portable vacuum bag and curing equipment 12. Doublers and stringers can be co-cured or

installed in a separate operation.

Steps 7-8: Many of the conventional methods for inspecting composite and aluminum

structures directly apply to inspecting hybrid metal laminates. Practical eddy current

methods for detecting sub-surface cracks using sliding probes such as the Nortec SPO-

2181 have shown a maximum reliable crack detection depth of 9-10 mm (12-13 AL

layers) 1.

Step 10: Seal and paint using conventional paints and sealants.

VIII. Developing Infrastructure for Using Hybrid Metal Laminates in Aging

Aircraft

In order to facilitate the accumulation of flight history with this material and improve its

acceptance and familiarity within the design community, a first possible step in

employing this material is to introduce it extensively for retrofitting aging military

aircraft. Implicit is that there will not be time for developing splice technology and that

manual lay-ups of the prepreg with aluminum sheets must be employed rather than using

automated CNC tape laying machines and 5-axis CNC routers. Much of the existing

infrastructure would be utilized to manufacture the components. However, in order to

instill confidence in the design community, a concerted effort to fully characterize the

properties of hybrid metal laminates along with thoroughly developed design guidelines,

FEA analysis techniques, NDI techniques, thoroughly proven fatigue crack growth

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models, and in-situ bonding techniques must be developed. Figure 12 is a timeline of the

major activities that must take place in order to gain acceptance into the MIL-HDBK-5

for Design Allowables. Subsequently, the community would then petition the FAA for

testing and certification.

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Figure 12. Aggressive future effort roadmap for readying the industry for retrofits using hybrid metal laminates.

30

IX. Developing Infrastructure for Using Hybrid Metal Laminates in the

Development of New Aircraft (e.g. the KC-135 tanker primary structure)

Once hybrid metal laminates have gained wide-spread acceptance in the design

community, and approval by the FAA for primary structures, the material will then be

considered during trade studies when developing new aircraft. A likely candidate would

be the KC-135 that will be developed by either Boeing or the Northrop Grumman-Airbus

partnership. Once a commitment is made to use this material for a large-scale program,

resources can be devoted to developing the infrastructure for producing hybrid metal

laminates for large production runs. Figure 13 details the proposed process that would

have evolved from present capabilities. Investments will need to be made in developing

splice technologies in conjunction with CNC tape laying machines, CNC 5-axis routing

of hybrid metal laminates, advanced welding (FSW and laser), and continuous rolling

aluminum sheet preprocessing systems. Additionally, investments will need to be made

in developing analysis and design guidelines to accompany the new benefits and

limitations of this material.

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Figure 13. Future effort roadmap for readying the industry for new design using hybrid metal laminates.

32

X. Conclusion The research and development activities to date have covered a variety of important

aspects pertaining to mechanical properties of GLARE. However, there is still little and

insufficient information available about mechanical behavior of GLARE in published

literature. A lot of areas are open to future investigation, especially for the cross-ply

configuration of GLARE, and some areas still remains to be further verified by more

detailed testing. More research and testing of the basic mechanical behavior such as in-

plane shear strength, bearing strength and tensile/compressive behavior in different

environments, estimation of fatigue lives and crack growth rates, notched sensitivity,

impact behavior, delamination and damage characterization are necessary to generate

adequate data to facilitate greater utilization of GLARE in future aircraft structures.

The manufacturing capabilities of hybrid metal laminates have been evaluated. Figure 14

highlights the major discussion points in this paper. However, much is left on the future

effort roadmap in order to fully understand what the requirements are for design, analysis,

and NDE maturity. The most direct way to facilitate the development of this very

promising new material is to start to commit to applications within industry. Funding for

research and development will always be limited unless a return on that investment can

be directly related to production of flight components.

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Figure 14. Preliminary evaluation of the current capabilities, transition possibilities, and future efforts has been proposed for manufacturing

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APPENDIX A: Alternative Splice Suggestions for Limited Aluminum Sheet Widths

Reference [16] and [17].

35

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