Curing and Bonding of Composites using Electron Beam Processing
Daniel L. Goodman
Science Research Laboratory / Electron Solutions Inc., Somerville, MA 02143
Giuseppe R. Palmese
University of Delaware Center for Composite Materials, Newark, DE 19716
Drexel University Department of Chemical Engineering, Philadelphia, PA 19104
CONTENTS
1
2
3
4
5
6
7
8
Introduction
................................................................................................... 1
1.1
Advantages of EB processing........................................................................... 1
1.2
Current limitations to EB processing ............................................................... 3
1.3
Background ..................................................................................................... 4
Aerospace composite fabrication using EB curing and bonding ..................................... 6
2.1
Filament wound rocket motors......................................................................... 6
2.2
EB-cured aircraft components.......................................................................... 9
2.3
Integrated aircraft structures........................................................................... 12
2.4
Other EB-cured components .......................................................................... 14
Automotive and vehicle composite demonstrations and potential................................. 15
3.1
An EB-cured automotive frame ..................................................................... 15
3.2
EB bonding of the Composite Concept Vehicle............................................. 17
3.3
Other vehicle applications.............................................................................. 18
EB-curable resin and composite development and current status ................................ 20
4.1
EB curable resin chemistry............................................................................. 20
4.1.1 Free radical polymerization ........................................................................ 20
4.1.2 Cationic polymerization.............................................................................. 22
4.1.3 Interpenetrating networks ........................................................................... 23
4.2
EB-curable resin and composite properties.................................................... 24
4.2.1 Prepreg resins ............................................................................................. 24
4.2.2 VARTM systems ........................................................................................ 26
4.3
EB-curable resins: further development......................................................... 29
EB-curable adhesive development and current status ................................................... 31
Equipment and facilities for EB curing and bonding ..................................................... 34
Conclusions and future directions .................................................................................... 38
References
................................................................................................. 39
Preprint of a chapter to be published in the Handbook of Polymer Blends and Composites,
A. Kulshreshtha and C. Vasile, eds., Rapra Technology Ltd, publisher, Shropshire, UK
Curing and Bonding of Composites using Electron Beam Processing
Daniel L. Goodman
Science Research Laboratory / Electron Solutions Inc., Somerville, MA 02143
Giuseppe R. Palmese
University of Delaware Center for Composite Materials, Newark, DE 19716
Drexel University Department of Chemical Engineering, Philadelphia, PA 19104
1
1.1
Introduction
Advantages of EB processing
Fabrication of fiber-reinforced polymer matrix composites incorporates assembly and curing
steps. Many different processes have been used for assembly, such as hand lay-up, resin transfer
molding, filament winding, automated tape placement, etc. Although some resin matrices are designed to
cure near room temperature, most composites are processed at elevated temperature, while pressure is
applied for consolidation.
Electron Beam (EB) curing of fiber-reinforced polymer composites is a promising new curing
technology for fabricating aerospace and ground vehicle components. As a curing technology, it must be
combined with a method of compaction such as tape or tow placement, hand-layup with hot debulk,
vacuum assisted resin transfer molding, or pultrusion. The basic concepts of EB curing of composites and
adhesives are shown in Figures 1 and 2. The process reduces the time required to cross-link the polymer
matrix compared to conventional heat curing. This potential for high throughput is especially important
for automotive fabrication. For aerospace applications, the principal advantages of EB curing are:
•
Curing is done near room temperature, allowing the use of low cost, low temperature tooling such as
wood, plaster or foam.
•
Curing at low temperature can reduce residual thermal stresses.
•
Co-bonding and co-curing operations with EB-curable adhesive allow fabrication of large integrated
structure.
1
•
EB-curable resins and adhesives have long shelf lives and can be stored at room temperature. They
are typically one component and solvent-free.
•
For large parts that are inconvenient or impossible to fit in an autoclave, EB processing may be the
best low-temperature curing alternative. Using a portable electron beam system, one can bring the
curing equipment to the part, rather than visa-versa.
•
Recent cost comparisons of EB versus thermal fabrication have shown that EB processing can reduce
costs by 10-40% for production of a variety of aerospace parts. These studies include both recurring
and nonrecurring costs. Capital costs of EB curing systems (principally the electron accelerator and
concrete radiation shielding) are similar to large autoclave costs.
Figure 1: Composite manufacturing processes suitable for EB curing include (a) filament winding,
(b) pultrusion and (c) prepreg layup with vacuum bag. Other methods (not shown) include vacuum
assisted resin transfer molding and automated tape placement. Parts are transferred via cart or
conveyor to be cured under an EB scan horn. In-line or in-situ curing processes are under
development for pultrusion and automated tape placement. This is shown in (b) and discussed in
2
Section 6. Figure reprinted from Ref [1] with permission of T. Walton, Aeroplas International
Corp., Hollis, NH.
Adhesive bonding coupled with autoclave curing is widely used in the aerospace industry to
produce large composite parts with complex core structures. EB-curable adhesives have the potential to
replace thermally cured aerospace adhesives in bonding large integrated structures in fewer steps, at
reduced cost and without autoclave processing.
The automotive industry uses adhesives to bond composite parts and structural assemblies. For
ground vehicle fabrication, EB bonding offers high throughput and “command-cure” bonding near room
temperature. EB-curable adhesives can be used to bond composites to metals substructure. The process
avoiding de-bonding during the cool-down cycle that can occur due to differences in thermal expansion of
the metal and composite.
Figure 2: High energy electrons are produced in the accelerator (sometimes known as the EB gun),
transport through the air, and then penetrate deeply into materials. The depth of penetration is
proportional to the energy, measured in millions of volts (MeV). The dose is a measure of deposited
energy. Typical doses for curing composites or adhesives are in the range 50-200 kGy.
1.2
Current limitations to EB processing
Although electron beam curing and bonding of composites is an active area of research, the
technology has not yet been widely adopted in industry. This is partially due to the conservative nature of
the aerospace and automotive industries. Pre-qualification has begun as part of the industry-wide
Composite Affordability Initiative (CAI) program, but there will be many years between this work and
the use of the materials in production. This is because few new aircraft designs are planned, the design
cycles are extremely long and the amount of composite on aircraft may actually be shrinking. In the
3
automotive industry, work done by research groups on new technologies, though promising, are not
necessarily adopted for production. The reason that novel automotive concepts are not produced is often
due to market forces, rather than to technical issues.
In addition to overcoming the resistance to change of the risk-sensitive automotive and aircraft
industries, EB researchers still faces significant technical challenges, especially to meet demanding
aerospace requirements [2]. EB-cured composite properties have achieved those incorporating firstgeneration heat-curable resins, such as Cytec-Fiberite 3501-6. But composite properties, especially
compressive properties that require good fracture toughness, are not yet as good as those containing
toughened second-generation resins such as Hexcel 8552. Although preliminary material qualification has
begun, the EB community does not have a reputation for "stable" materials with consistent properties.
Suppliers and users have not developed the processing specifications and acceptance standards that are
the hallmarks of a mature technology.
EB-curable adhesives have made significant strides in the last few years. Their properties meet
the requirements of automotive applications. The strength of the best EB-curable paste adhesive is close
to that of industry-standard aerospace systems. But these adhesives cannot be used in the hottest or
coldest aerospace environments, and are just beginning to be used on secondary structure. In some cases,
aerospace companies find it more convenient to have their adhesive supplied as a film rather than as a
paste.
However, EB-curable film adhesives do not currently possess the strength needed for any
aerospace application.
There are several reasons why EB-curable composites and adhesives properties do not match
those of state-of-the-art aerospace systems. One reason is that the resin chemistry (either free radical or
methacrylated and cationic epoxy-based) produces different polymer structures than heat-curable systems.
Another oft-cited reason for reduced composite properties is reduced fiber-matrix interface strength due
to incompatibility between EB-curable resins and fiber finish. These issues and current research to
improve material are discussed in Section 4.
1.3
Background
Electron Beam composite curing and bonding is a subset of the field of radiation processing of
polymers. Electron Beam irradiation is a widely used commercial process. In the electronics industry, EB
processing is used to crosslink wire and cable insulation, to produce heat-shrinkable tubing and to make
polymer devices such as resettable fuses for automotive and portable electronics [3,4]. The Radtech
organization, founded in 1988, promotes the technology of inks and coatings curing by ultraviolet light or
EB. Radtech sponsors biannual conferences in the US and Europe [5]. Electron beam irradiation is also
4
used to sterilize medical products, increase the strength of supermarket plastic bags and crosslink rubber
in tires.
The French firm Aerospatiale was the first company to use electron beam composite curing in a
production environment. Aerospatiale fabricated large filament-wound rocket motor cases and cured the
resin with EB [6]. They reduced the time to cure a rocket motor from about a week to less than eight
hours, while producing lower residual stresses in their part. This application is described in Section 2.1.
Encouraged by this success, the US Defense Advanced Research Project Agency (DARPA)
decided to fund work on the use of electron beam curing to lower the cost of manufacturing aircraft
composites. The two largest DARPA-sponsored programs were the Affordable Polymer Composite
Systems (APCS) program led by Northrop Grumman during 1994-1998 [7,8] and the Integrated Airframe
Technology for Affordability (IATA) program led by Lockheed Martin from 1995-1997 [9]. Both
programs produced demonstration aircraft parts using EB curing and bonding. The APCS program found
significant cost savings for the use of EB in a variety of aircraft parts. The APCS program and its results
are described in Section 2.2. The IATA program designers were encouraged to be forward-looking and
futuristic. They asked how novel integrated designs using EB curing and bonding as an enabling
technology could radically reduce fabrication costs. Their results are summarized in Section 2.3.
The most recent large aerospace program to examine EB composite curing and bonding is the
Composite Affordability Initiative (CAI), a cooperative program between the Air Force, Navy, Lockheed
Martin Corporation, Boeing and Northrop Grumman. One of the CAI demonstration articles is an EBcured and bonded keel-duct interface. The part is based on an IATA design, similar to that shown in
Section 2.3. CAI has also tested many of the commercially available EB-curable resins and adhesives and
is creating benchmarks, comparing the systems to thermal systems that have flown or are planned to fly.
The CAI data was not available for this review, but will soon be available from the companies
participating in the testing [10]. A material selector CD with processing data is also planned for release.
The high production rate achievable with EB processing is an important advantage for automotive
production. As early as 1992, researchers at Chrysler were considering the possibility of a completely EBcured and bonded automotive body [11]. This concept is described in Section 3.1. They decided that such
extensive use of an unproven technology was too large a technological leap. The Composite Concept
Vehicle (CCV) was instead produced from resin injection molded thermoplastic [12]. One possibility for
bonding this car is with EB-curable adhesive, as described in Section 3.2. Other applications of EB curing
to ground vehicle structures are described in Section 3.3.
EB-curable resins and adhesives for aerospace and automotive applications cure by either cationic
or free-radical polymerization. Aerospace programs have primarily adopted cationic resins, because of
5
their higher use temperature and low shrinkage. Oak Ridge National Laboratory has coordinated the
largest development effort in cationic resins and has licensed their materials to several commercial
suppliers [13]. Free radical resins and adhesives have an advantage that they cure more quickly than
cationics and are not inhibited by nucleophilic (alkaline) materials on fibers or surfaces [11]. Cationic,
free-radical and interpenetrating network EB-curable resins for composite matrices are described in
Section 4.1. Section 4.2 describes the current status and the requirements for EB-curable composites,
primarily in the aerospace industries. Additional planned research to improve EB-curable resins are
summarized in Section 4.3. The current status of EB-curable adhesive development is described in
Section 5.
Several new facilities devoted exclusively to EB composite curing and bonding have recently
begun operation. The capabilities and cost of these facilities are described in Section 6. Section 7
contains our conclusions and expected future directions in EB composite curing and bonding.
2
2.1
Aerospace composite fabrication using EB curing and bonding
Filament wound rocket motors
The French company Aerospatiale has manufactured solid propellant rocket motors for more than
20 years. The cases were traditionally made from filament wound carbon fiber with heat-cured epoxy
resin. Beginning in 1979, they looked for ways to increase production speed and to reduce thermal
stresses during cure and cool-down. Researchers found that a combination of EB and X-ray curing
reduced their curing time from four days to less than eight hours while keeping the structures near room
temperature [6]. Because appropriate EB-curable resins did not exist, they developed an acrylated
formulation with the low viscosity needed for filament winding [14].
Aerospatiale was the first to realize the need for new fiber sizings to improve the adhesion of EBcurable resins to carbon fibers. They developed specialized coupling agents that improved the fibermatrix interface by creating covalent bonds between the fiber and the resin [15]. The hydroxyl groups on
the carbon fibers react with an isocyanate group on the sizing, and the hydrogen bonds on the sizing's
acrylate group are free to crosslink with the resin during EB curing. This improved the burst strength of
the system to exceed their heat curable baseline.
Aerospatiale installed a 10 MeV, 20 kW electron accelerator to cure their rocket motors. In their
dedicated concrete bunker, the accelerator could move linearly and the composite rotated to be able to
cure parts up to four meters in diameter and 10 meters long. The system is shown in Figure 3.
6
Figure 3: EB curing of a filament wound rocket motor at Aerospatiale. (Photo copyright 1997
Aerospatiale Matra Lanceurs, Cedex , France, used with permission.)
Following Aerospatiale’s lead, several groups have recently used electron beam curing to
fabricate filament wound aerospace components.
A US Army-sponsored program investigated the
feasibility of developing faster, more economical manufacturing techniques for filament wound
composite motor cases for small diameter tactical rocket motors [19]. They produced thin-wall pressure
vessels from a variety of EB-curable resins and demonstrated the use of low cost foam tooling. They
achieved burst pressures that compare favorably with their baseline thermally cured pressure vessels.
Figure 4(a) shows their pressure vessels during EB curing. Figure 4(b) shows one of their pressure
vessels after burst testing.
A group in Italy patented an in-situ curing process that combines layup and EB curing in one
system and has developed a system containing a low energy electron beam accelerator coupled to a twoaxis filament winder for in-situ curing of pressure vessels [16, 17]. Their layer-by-layer curing process is
applicable to a variety of composite production processes and is capable of producing parts of arbitrary
size. Using their layer-by-layer process, the group was able to achieve interlaminar shear strengths equal
to that obtained when curing their structures all at once using a high energy (10 MeV) electron
accelerator. Using a similar process, a US group has coupled a low energy electron beam gun to an
7
automated tape placement machine for in-situ curing [18]. Both systems are shown and described in
more detail in Section 6.
Figure 4: (a) Filament wound pressure vessels during EB curing. (b) After burst testing. Reprinted
from Ref. [19] with permission of R. Foedinger, DE Technologies, King of Prussia, PA.
A 1998 program at the NASA Marshall Space Flight Center (MSFC) examined the use of EB
curing to correct manufacturing problems with their 60,000 lb.-thrust "FastTrack" rocket motor. They
were experiencing difficulties due to a thermal expansion mismatch between the filament wound graphite
and the silica-phenolic tapewrap at the throat of this one-piece composite nozzle. They showed that EB
curing was a practical alternative, as long as the dose rate (and therefore heating due to EB dose and resin
exotherm) was kept low. This part is shown in Figure 5.
Figure 5: Filament wound nozzle cured by EB near room temperatue to avoid cracking due to
thermal expansion mismatch between the silica (center) and graphite overwrap. Photo courtesy of
William McMahon, NASA Marshall Space Flight Center, Huntsville, AL.
8
2.2
EB-cured aircraft components
The DARPA-sponsored Affordable Polymer Composite Structures (APCS) program examined
many of the key issues facing the aerospace industry as it begins to adopt EB curing and bonding. The
overall goal of the APCS program was to see if EB curing could significantly reduce the recurring and
non-recurring cost of aerospace composites.
Northrop Grumman personnel evaluated EB-curable
materials, developed new processes, accessed EB cost benefits, and produced demonstration assemblies
including a full-scale article. Although new material development was outside the program scope, resin
suppliers produced novel systems to meet Northrop requirements.
The EB-curable resins evaluated by Northrop included acrylated and methacrylated epoxies,
cationic-curing epoxies, a modified bismaleimide, a modified epoxy novolac and a modified epoxy.
These were compared with a heat-curable Cytek-Fiberite 3501-6 baseline. The cationic epoxy CAT-B
resin had the best overall properties [8]. CAT-B demonstrated a dry glass transition temperature (Tg) of
180°C, close to the 190°C baseline. Composite lamina and laminate properties were measured on woven
and unidirectional material forms. The CAT-B properties approached those of 3501-6 except in
compression and shear strength. Northrop attributed this to poor resin-fiber adhesive, as indicated by dry
fibers in the failed region of the specimens. They believe that a fiber sizing or a coupling agent
compatible with the EB-curable resin can improve the interface.
The APCS process development concentrated on ways to reduce porosity in prepreg composite
laminates, which were initially high, in excess of 5%. Hot debulk prior to EB cure significantly reduced
porosity, and application of external pressure during hot debulk increased the fiber volume in the
laminates. For resin transfer molding (RTM) and vacuum-assisted RTM, they found that laminate
porosity could be controlled by proper degassing of the resin before injection and by maintaining vacuum
integrity throughout the infusion and cure cycle. Northrop Grumman combined a wide variety of
fabrication processes with EB curing including automated fiber placement, RTM, VARTM and manual
lay-up together with cocuring, cobonding and secondary adhesive bonding. They fabricated many parts
including skins, covers, doors and various substructures.
The culmination of the APCS program was the fabrication of a full-scale demonstration article.
Northrop Grumman selected the aft center fuselage assembly of the F/A-18E/F fighter plane containing
engine inlet ducts, frames, fuel cells, keels, dust-to-side skin webs and side skins. A wide range of part
forms, resins and fabrication processes were used to demonstrate all the EB curing and bonding processes
examined during the program. Materials and processes used included:
•
Woven AS4/Cat-B prepreg and IM7 /RB-47 tape with hand layup.
•
Tow placed AS4 Tape.
9
•
Resin transfer molding of woven AS4 with VEB-2 and CAT-B resins.
•
Selective area curing for later assembly to "wet" flanges.
•
Co-bonding and EB adhesive bonding.
•
Use of syncore for buckling resistance.
Northrop Grumman’s materials and assembly methods for the EB-cured fuselage assembly are shown in
Figure 6. Figure 7 shows the assembly undergoing EB irradiation of adhesive bond lines during final
curing steps. After assembly, the fit of the frames to the duct showed a maximum gap of less than 1 mm.
Figure 6: The aft center fuselage assembly of the F/A-18E/F fighter was selected by Northrop
Grumman to demonstrate a variety of EB curing processes and materials. EB-curable materials
used included CAT-B cationic epoxy prepreg and Loctite 334 adhesive. Figure courtesy of R.
Vastava, Northrop-Grumman Co. El Segundo, CA.
Cost studies constituted a significant portion of the APCS program. Northrop conducted cost
trade studies for composite monolithic details, sandwich and integrally stiffened structures, and various
assemblies. They found that the cost savings potential of EB processing varied with the structure,
complexity and manufacturing process, and ranged from ~10% in case of simple details with direct
process substitution to up to 60% for a complex assembly.
The cost estimates included the cost of EB accelerator systems, buildings with the required
features, part handling systems, operating, maintenance and repair. Typical EB facilities costs are
described in Section 6. Capacity analysis for each facility and return on investment (ROI) analysis were
also performed. The pay-back period naturally depends on the volume of composite parts processed.
Northrop found a 2-4 year payback period for the assumptions of a single program using EB, a seven year
10
equipment life, an equipment acquisition value of $5M and a total of 54,000 to 90,000 lbs of composite
processed. Typical ROI curves are shown in Figure 8.
Figure 7: Duct assembly undergoing EB curing as part of the APCS program. The EB gun (at
right) is directed along the adhesive bondline to cure the duct splices.
11
Figure 8: Return on investment as a function of savings for an electron beam aerospace composite
processing facility. (Based on data collected by the APCS program [8].)
2.3
Integrated aircraft structures
The DARPA-sponsored Integrated Airframe Technology for Affordability (IATA) program at
Lockheed Martin examined cost reduction from a design and manufacturing approach [20]. From the
IATA perspective, the largest cost reductions come from advanced integrated designs that make use of
advances in composite material and manufacturing process. A large portion of the IATA program
focused on EB-curable composites and adhesives as a means to create large integrated structures. IATA
selected for design review the wing/fuselage section of the JAST/ASTOVL aircraft. because it is the most
costly portion of the airframe and the most structurally challenging for composites.
The IATA program strategy reflects where production dollars are spent during manufacture of
high performance aircraft. For prototypes or short production runs, cost is dominated by the tooling (up
to 70% of final costs), which is related to the number and complexity of the parts required to produce the
final structure. For larger production runs, cost is dominated by the recurring costs for fabrication and
12
assembly (up to 65% of total costs). The primary fabrication costs elements are material placement and
processing, which can be 40-70% and 25-35% of fabrication costs, respectively, depending on the number
of parts and their complexity. Unlike metallic parts that can be fit-up and assembled economically using
mechanical fasteners, the mechanical assembly of composite structures is expensive, and is the largest
contributor to scrapped parts.
To reduce costs, the IATA program examined integration of parts for reduced part count and
combinations of manufacturing processes including advanced fiber placement, EB-cured prepreg and
resin transfer technologies (RTM, VARTM). They also examined co-processing technologies and unique
methods for assembly that eliminates the extensive use of fasteners.
Among the results of the program were many detailed designs, several demonstration
manufacturing parts and preliminary material specifications for EB-curable resins and adhesives. One of
the IATA fuselage section designs is shown in Figure 9. A demonstration part, shown in Figure 10, was
fabricated to verify the concept design and provide a check on cost estimates.
Figure 9: An Integrated Airframe Technology for Affordability design utilizing EB curing and
bonding to reduce costs. (a) The VARTM tri-resin keelson is EB-cured and rotated into place. (b)
An integrated EB-cured tape-placed upper skin is lowered into place. EB-curable adhesive bond
lines are also shown. IATA figures courtesy of D. Sidwell, Lockheed Martin Corp, Palmdale, CA.
13
Figure 10: Demonstration EB-cured bulkhead section of JAST/ASTOVL airframe. (a) As displayed
at the SAMPE Symposium in Long Beach CA, 1997. (b) The position of the structure, next to the
inlet duct. IATA figures courtesy of D. Sidwell, Lockheed Martin Corp, Palmdale, CA.
2.4
Other EB-cured components
A windshield frame and arch for the T-38 Talon is one of the few EB-cured aircraft components
that has flown. The T-38 Talon is the US Air Force's primary supersonic jet training aircraft. Two
windshield assemblies were fabricated, substituting EB-cured materials for their thermally cured
analogues [21].
As conventionally fabricated, most of the windshield frame is composed of 121°C-cured
epoxy/6781 S-2 woven fiberglass prepreg, thin strips of 301 half-hardened stainless steel bonded together
with high-strength epoxy film adhesive, and 6061 aluminum for structural attachment points. The
windshield is rated to withstand a 1.8 kg bird impact at 740 km/hr.
The windshield frame is made from four components that are assembled into one windshield
frame and one windshield fairing. The four components are the fairing, innerskin, bulkhead doubler and
outerskin/arch. The first three components are sufficiently thin (< 5mm) to be easily penetrated by
electrons with energy of <10 MeV. The outerskin/arch is quite thick, and required the deeper penetrating
ability of X-rays to cure the adhesive through the stainless steel. The prepreg frame after lay-up and
14
during cure is shown in Figure 11. The frame survived bird impact testing with damage comparable to
frames fabricated using conventional processing [22].
The T-38 is one of several projects that required deeper penetration than is possible using direct
electron irradiation. Researchers at Aerospatiale, Oak Ridge National Laboratory and at NASA Marshall
Space Flight Center have also used X-rays to cure thick filament-wound structures. The effects of X-rays
and of electrons on radiation-curable materials are very similar, since both cause polymerization via
secondary electrons. However, X-ray curing is much slower than direct EB curing.
Placing a thin, high-density water-cooled metal target in front of the electron beam generates the
X-rays. Up to five percent of the EB energy can be converted to X-rays in this manner. The low
conversion efficiency together with X-ray absorption throughout a large volume is the reason why X-ray
curing is a slow (and expensive) process.
Figure 11: An EB-cured and bonded windshield frame and arch for the US Air Force T38
supersonic trainer. (a) Prepreg layup, (b) During EB cure. From Ref. [21], used with permission.
3
3.1
Automotive and vehicle composite demonstrations and potential
An EB-cured automotive frame
Although the majority of research funding for EB curing and bonding has been for aerospace
applications, the technology offers significant advantages for ground vehicle fabrication. A primary
advantage of EB curing for automotive applications is cure speed. An entire composite automotive frame
can be cured in less than two minutes using currently available high-power electron accelerators [11].
However, curing is only a small portion of the technical challenge of automotive assembly with EB.
15
Molding, tooling and assembly issues for EB-curable automotive parts are major challenges, but have
received little attention to date.
The first substantial work in this area was in 1991 by Billiu [23,24]. Billiu devised a method to
form a hollow section fiber-reinforced composite by injecting a two-part resin that reacts to form a
thermoplastic. Using supersonic tack welding, the sections can be fusion welded together and then the
"B-staged" composite can be cured to a hard thermoset with electron beams. Billiu brought his ideas to
the attention of the Chrysler Corporation's advanced technology group, who were investigating ways to
make an all-composite car body. In theory, an entire car can be EB cured, even including the adhesive
and the paint. Chrysler studied the idea and designed an automotive production facility that included
injection molding and EB curing. A result of that study is shown in Figure 12.
Figure 12: An automotive body production facility using EB curing. Car frames awaiting transport
are shown at the top right, resin transfer molding equipment is at the bottom right and an EB
curing system with car frames transported through a radiation maze is at the top middle-left.
Figure courtesy of C. Billiu, Ticom Co., Warren, MI.
The Chrysler advanced technology group determined that there are too many technical risks to
this approach, and instead decided to produce an all-composite thermoplastic car frame using injection
molding [12]. The result is the composite concept vehicle (CCV) shown in Figure 13. The CCV body is
composed of six large injection-molded pieces made of glass-reinforced polyethylene terephthalate
(PET). Four of these sections are shown in Figure 14. The CCV inner- and outer-frames are bonded with
adhesive to themselves and to a steel chassis. EB curable adhesive can be used to bond these parts
together. This is described in the next section.
16
Figure 13: The Daimler-Chrysler Composite Concept Vehicle (CCV) features a thermoplastic body.
3.2
EB bonding of the Composite Concept Vehicle
The baseline materials for bonding together automotive body panels are heat-curable adhesives.
Conventional heat-curing adhesives require many minutes to bond. Total bonding time includes the time
to heat the material at the bond-line plus the time to cure the adhesive. Plastic auto bodies such as the
CCV are effective insulators, preventing heat from reaching the bond-line. Oven or forced hot-air
temperatures must be limited to avoid deforming or melting the thermoplastic. These effects combine to
give hot-air bonding times of many minutes, whereas auto manufacturers prefer cycle times under twominutes.
Some of the bonds in the CCV join the plastic to a steel frame chassis. After joining with heatcurable adhesives, some sections of these bonds and near-by composite can crack when cooled to room
temperature. The cracking is due to the thermal expansion mismatch, caused by the much larger
contraction of the plastic than the steel during cooling.
To avoid these problems with heat-curable adhesives, some manufacturers have formulated
adhesives that cure at room temperature. Once mixed, these systems have very short working life before
they harden. Such systems are well suited to bonding of small areas, but are very difficult to use on large
areas such as auto bodies. Designers would prefer a room temperature “command cure,” which provides
17
extremely rapid curing once the adhesive, parts and tooling are aligned in place. Curing of adhesive with
electron beams provides this processing flexibility.
Figure 14: Inner and outer frames are bonded together to form the rigid CCV body. One method to
bond the CCV frame to itself and to the steel frame is with EB-curable adhesive. Other thermoplastic components (not shown) include the roof and undercarriage.
Automotive adhesives for structural bonding are much more flexible (although not nearly as
strong) as those used to bond aerospace structures. It is difficult to maintain tight dimensional tolerances
in molded parts, so the bond-line thickness vary considerably in the CCV (and in automotive applications
generally.) The adhesive for the CCV thus has the additional roles of shock absorber and of caulk or
sealant. Researchers at Chrysler and SRL found that for the CCV, EB-curable adhesive requirements
included a lap shear strength of 3.5-5 MPa, an elongation to failure of >100%, and an immediate cure
strength >50% of final strength.
They developed adhesives that met these specifications and
demonstrated bonding of CCV panels [25, 26]. Although the CCV is not being produced (due to market
issues and difficulties in part molding), Daimler-Chrysler is considering EB curing for other automotive
parts [27].
3.3
Other vehicle applications
The US Army has sponsored several projects to determine whether EB curing has advantages for
military vehicle fabrication. In one of these programs, side skirt armored panels for the Composite
Armored Vehicle (CAV) Advanced Technology Demonstrator were fabricated using EB curing [26]. The
18
CAV is the first armored combat vehicle to use a fully integrated armor and structural composite hull
[28]. Figure 15 shows a picture of the CAV with the location of the EB-cured armored panel indicated.
The demonstration showed that multiple layers, including base composite, ceramic tile armor and
signature protection, could be cured together at low temperature using EB.
The EB-cured demonstrator panels performed as well in all respects as side skirts fabricated using
conventional resins and heat curing at United Defense L.P., the prime contractor on the CAV-ATD.
Figure 16 shows a panel after ballistic testing.
Deflection
limiting
bumper
Armor
panel
Composite
Armored Vehicle
Flexible
skirt
Seven-layer
armor panel
Figure 15: Side skirts for the Composite Armored Vehicle (CAV) were fabricated using EB
composite curing. The seven-layer panels consist of composite, rubber, adhesive, ceramic tile, EMI
shielding and signature protection.
19
Figure 16: EB-cured CAV panels after ballistic testing. The hexagonal pattern on the right is
"shine-through" of electrons at ceramic tile edges [26].
4
4.1
EB-curable resin and composite development and current status
EB curable resin chemistry
High energy electrons generate ionic species, free radicals, and molecules in excited states
capable of initiating and sustaining polymerization. Figure 17 summarizes the pathways for induction of
polymerization by EB irradiation. Depending on the chemistry of the resin system being irradiated,
polymerization can occur by free radical as well as ionic mechanisms. The polymerization of
acrylic/methacrylic systems, maleic and fumaric polyester resins, maleimides, and thiol-ene systems
proceeds via free-radical mechanisms without initiators. EB-induced polymerization of nitroethylene
proceeds via anionic mechanisms, and epoxies are polymerized cationically with the appropriate
photoinitiator under EB irradiation. Of these systems, free-radical-cured systems based on acrylate and
methacrylate functionality and cationically cured epoxies catalyzed using diaryliodonium or
triarylsulfonium salts have shown the most promise for composites applications.
4.1.1 Free radical polymerization
The polymerization of monomers containing double bonds is often conducted thermally by free
radical chain polymerization. This consists of three major steps: initiation, propagation and termination.
Initiation is usually aided by the use of peroxides. When heated, these homolytically decompose to form
free radicals that add to a monomer thus transferring the active site. Such initiators are not needed for EB
initiation since the beam generates free radicals directly.
20
The rate of EB initiation can be expressed as Ri=IeGF, where Ie is the electron beam intensity or
current, and G and f are efficiency factors for the production of initiating species and their initiating
efficiency. The G value depends on the chemical structure of the irradiated material. Unsaturated
materials have G values hundreds of times larger than saturated moieties.
The formation of radical initiating species occurs readily for vinyl-containing monomers such as
acrylates and methacrylates. Polymerization proceeds via propagation steps that occur by the successive
addition of vinyl-containing monomers to the active radical center. Termination occurs as with thermal
systems by several processes including combination, disproportionation, chain transfer, and radical
trapping. The rate of initiation and polymerization in EB-cured systems is determined by the dose rate.
The polymerization induced by high dose rate EB curing can differ significantly from polymerizations
initiated via thermal decomposition of initiators, because the high concentrations of initiating radicals in
EB curing changes the balance between propagation and termination.
High Energy Electrons
Ions and
Ion Radicals
Excited
Molecular States
Free Radicals
Ionic
Polymerization
Free Radical
Polymerization
Figure 17: High energy electrons in EB-curable resins and adhesives generate ionic species, free
radicals, and molecules in excited states that initiate and sustain polymerization.
EB-curable composites with acrylate/methacrylated matrices have been studied extensively [29,
30]. These systems provide high reactivity, and they have good stiffness, good control over processing
viscosity, and very long shelf lives. Among the shortcomings associated with such systems are high cure
shrinkage (8–20%), the potential for oxygen inhibition, low Tg relative to high-temperature thermally
cured epoxies, and low fracture toughness.
21
4.1.2
Cationic polymerization
Cationically cured epoxies require a photoinitiator to enable polymerization. Commercially
available photoinitiators are diaryliodonium and triarylsulfonium salts of weak bases such as
diphenyliodonium hexalfluoroantimonate or triarylsulfonium hexalfluoroantimonate [31, 32]. Based on
such initiators, a consortium led by Oak Ridge National Laboratory has formulated and evaluated a great
number of epoxy resin systems for EB curing [13]. Such systems offer low shrinkage (<6%) exhibit Tg
significantly higher than thermally cured systems, are not inhibited by oxygen, and do not require curing
agents, as do their thermally cured counterparts [33]. On the other hand, cationic systems tend to cure
more slowly than acrylate/methacrylate systems, the photoinitiators are easily poisoned by nucleophilic
contaminants, and they require additives to improve fracture toughness [34].
The mechanisms of EB-induced cationic chain polymerization are not well understood.
However, investigations conducted on the thermal curing of epoxy resins by cationic mechanisms form a
basis for understanding cationic polymerization of epoxies by EB irradiation.
Several authors have investigated the thermal curing of epoxy resins by cationic polymerization
[35, 36]. Such polymerizations are comprised of initiation, propagation and termination steps. Chabanne
et al. have investigated the mechanisms associated with cationic homopolymerization of epoxy initiated
by BF3-amine catalysts [37, 38]. The salt R'NH3+BF4- is thought to activate the epoxy monomer (M) by
forming an oxonium ion active center. Propagation has been reported to proceed via two competing
mechanisms known as the activated chain end (ACE) or activated monomer (AM) mechanism. In the
ACE mechanism, growing chain cyclic tertiary oxonium ions are formed, and polymerization proceeds by
chain addition of monomer. On the other hand, the AM mechanism proceeds via the addition of
molecules containing hydroxyl groups to the activated monomer. This is accompanied by a charge
transfer to regenerate an activated monomer that continues to react. During the initial stages of reaction
and in systems with low initial concentration of hydroxyl groups, the ACE mechanism is predominant.
However, the presence of hydroxyl groups, i.e., water, polymer chain ends, and comonomer, can favor the
AM mechanism. Other reactions include the chain transfer reactions in which an activated chain end
reacts with a hydroxyl moiety, water, or alcohol. This transfer reaction via an OH-containing compound
results in an inactive oligomer.
Initiation of epoxy cationic polymerization by irradiation is effected using onium salts, like
diaryliodonium hexafluroantimonate, that undergo photolysis and release a powerful Bronsted acid. In
subsequent steps, this strong protonic acid efficiently initiates the polymerization of epoxy-containing
monomer. This mechanism is well understood for UV irradiation [39, 40], but considerable debate exists
regarding the mechanism of initiation by EB irradiation. The mechanism is complex, since many radical,
22
radical ion and ionic species produced during irradiation can lead to the generation of a cationic active
center. Among these is the H+ ion that reacts with the epoxy group of the monomer and produces an
active species capable of polymerization. Some work suggests that the propagation steps in cationic EB
polymerization are analogous to those that occur in thermally cured systems [41].
Figure 18: The formation of an interpenetrating network: The first network forms at room
temperature, and EB irradiation forms a second network, increasing Tg.
4.1.3
Interpenetrating networks
A third type of EB-curable polymer resin is based on Interpenetrating Polymer Networks (IPNs).
Developed for liquid molding applications, these systems can be C-staged at room temperature to form
handleable solids, that, when irradiated, achieve a high degree of conversion with Tg approaching 200°C.
[42]. An interpenetrating network is produced in which the EB-crosslinked portion develops around a
previously formed network structure based on epoxy-amine or isocyanate-alcohol step growth
polymerization [43]. A schematic representation of epoxy-based systems is shown in Figure 18. These
IPNs are comprised of epoxy monomers crosslinked by amine curing agents in the presence of
(meth)acrylate reactive diluents that react by free radical polymerization when subjected to EB
irradiation. Network coupling is achieved by selecting monomers with dual functionality.
There are several advantages to this type of resin.
First, the cure shrinkage shortcoming
associated with traditional acrylate/methacrylate systems is reduced via the physical constraints imposed
by the interpenetrating network and a lower concentration of double bonds. These systems have a total
shrinkage of 5–7%, with most of this shrinkage occurring during the C-staging process. Second, the
systems maintain good processability via viscosity control by manipulation of reactive diluent
concentration and type. Third, C-staging allows for EB irradiation after removal of the part from its mold
or tooling. Finally, the IPN structure results in materials with inherently higher fracture toughness. Some
23
of the drawbacks associated with these systems include relatively low Tg compared to aerospace-grade
epoxies and the need to control the temperature during the staged curing process.
4.2
EB-curable resin and composite properties
Most EB-cured composite parts have been fabricated from prepreg or by Vacuum Assisted Resin
Transfer Molding (VARTM). Resins for these two processes are described in this section.
4.2.1 Prepreg resins
The mechanical properties of three industry standard prepreg resin systems are given in Table 1.
The amine-cured epoxy resin 3501-6 was developed to operate at temperatures up to 177°C and is
recommended for general-purpose structural applications. Cytec-Fiberite’s 977-3 amine-cured epoxy has
a 177°C dry service capability and a 132°C wet capability. It is a toughened system designed for use in
primary and secondary aircraft structures or any application which requires impact resistance and good
hot/wet performance. Hexcel’s 8552 is an amine-cured toughened epoxy resin system recommended for
structural applications requiring high strength and damage tolerance. It was developed to operate in
environments up to 121°C.
Also included in Table 1 are the properties of some representative EB-cured systems. CAT-B is
an untoughened cationically-cured epoxy resin representative of cationic systems developed by the Oak
Ridge National Laboratory-led consortium [13] and evaluated under the APCS program [7, 8]. The resin
designated RB-47 is a developmental free-radical-cured methacrylate prepreg system supplied by
Aerospatiale that was also evaluated during the APCS program.
Data in Table 1 shows that the Tg of the methacrylate system is lower than that of the other
systems and that the tensile strength of RB-47 is significantly lower. CAT-B has a dry Tg roughly
equivalent to values for the thermally cured systems. CAT-B also possesses equivalent tensile strength
and tensile modulus to the untoughened 3501-6. The major shortcoming associated with this cationically
cured system, however, is low fracture toughness.
A comparison of composite properties for autoclave-cured versus EB-cured prepreg carbon-fiber
composites is given in Table 2. This table summarizes data for AS4-G' sized carbon fibers and IM-7
carbon-fiber prepregs. The thermally cured systems are unidirectional and 5HS AS4 fiber-based systems
impregnated with 3501-6 and unidirectional IM-7 prepregs made with toughened 977-3 and 8552 resin.
The EB-cured systems are based on the CAT-B and RB-47 resins described earlier and on EB4 and EB6,
two other cationically cured epoxy systems [44].
24
Table 1: Comparison of thermally and EB-cured prepreg resin systems. Thermal cure data are
taken from manufacturers published product information. EB cure data are from Refs [7, 8]
except data marked *, obtained from testing on an equivalent system.
Thermal Cure
EB Cure
Resin
3501-6
977-3
8552
CAT-B
RB47
Epoxyamine
toughened
epoxyamine
toughened
epoxyamine
cationic
epoxy
free-radical
methacrylate
E"
190
—
—
180
160
tanδ
210
190/240
200
—
—
Tensile strength ksi
10
—
17.5
9.7
7.4
Tensile modulus ksi
640
—
680
570
680
Elongation %
1.7
—
4
2
1
Flexural strength ksi
—
21
—
—
—
Flexural modulus ksi
—
550
—
—
—
GIC J/m2
130
220
680
70*
—
KIC mpa • m1/2
—
0.9
1.6
0.4*
—
Description
Tg (°C)
These systems illustrate the high Tg potential of cationically cured epoxies. The unidirectional
AS4 fiber composites made with RB-47 perform poorly, with reductions in tensile, compressive, and
interlaminar shear strength compared to the 3501-6 baseline system. CAT-B, on the other hand, performs
better than the methacrylate system but also exhibits reduction of compressive and interlaminar shear
strength relative to the 3501-6 baseline.
The properties of unidirectional IM7 composites based on the cationic EB-cured resin systems
EB-4 and E-B6 show similar reduction of properties compared to 977-3 and 8552 systems. Compressive
and interlaminar shear strength are significantly deficient in these systems. Thus there is a well-supported
trend that carbon composites based on cationically cured epoxies possess poor compressive and
interlaminar shear strength. This is perhaps a result of the brittle nature of EB-cured cationic epoxy
systems and also the existence of a weak interphase between the carbon fiber and the EB-cured resin
system..
25
Table 2: Comparison of thermal and EB-cured carbon-fiber composites fabricated from prepreg.
Thermal cure data are taken from manufacturers published product information. EB cure data
are from Refs [7, 8] and [44, 45] except data marked *, obtained from testing on an equivalent
system. All data is normalized to 62% fiber volume fraction.
AS4 G' sizing
Reinforcement
Tg
(°C)
Unidirectional
IM7
12K 5HS
Unidirectional
9773
Resin
3501-6
RB47
3501-6
CAT-B
Cure
Thermal
EB Cure
Thermal
EB Cure
E'
190
160
190
180
—
—
—
—
tanδ
210
—
210
—
190/
240
210
212
237/40
0
310
240
120
100
364
373
319
328
20
—
—
—
23.5
23.7
22.8
23.8
Tensile
strength ksi
Tensile
modulus msi
Compressive
strength ksi
Compressive
modulus msi
Flexural
strength ksi
Flexural
modulus msi
ILSS ksi
8552
Thermal
EB4
EB6
EB Cure
250
175
100
90
244
245
207
221
18
—
—
—
21.6
21.7
—
—
250
—
150
127
256
274
256
—
18
—
—
21.7
18.4
22.3
—
18.5
8
10
18.5
19.9
12.9
11.2
8.5
4.2.2 VARTM systems
VARTM is a liquid molding process that is driven by vacuum pressure. The reinforcing preform
is typically formed around one-sided tooling, and a vacuum bagging technique is used to seal the system.
Impregnation is conducted using low- permeability media to rapidly direct flow over large surface areas.
This technique can be used for manufacturing large structures such as roadway bridge decks and large ondeck ship structures like mast enclosures. This fabrication technique is also under evaluation for cost
reduction in the manufacture of aircraft structures, as part of the Air Force-led Composite Affordability
26
Initiative program. A major requirement of such systems is that they possess a viscosity below 500 cps
with adequate working time for processing conducted at ambient temperature.
Table 3: Comparison of thermally cured and EB-cured VARTM resins. Thermal cure data are
taken from manufacturers published product information. EB-cured data are from Ref [46].
Thermal Cure
Resin
Description
EB Cure
Ciba
Resinfusion
8611
Shell
862/Lindride
6k
API SC-15
VCCM1
VCCM4
EpoxyAmine
EpoxyAnhydride
Toughened
EpoxyAmine
IPN EpoxyMethacrylate
Toughened
IPN
2500 (25°C)
680 (25°C)
350 (25°C)
240 (25°C)
320 (25°C)
Viscosity cps
170 (52°C)
Tg (°C) E"
136
120
105
132
129
Tensile strength ksi
5.9
12
8.7
—
—
Tensile modulus ksi
450
480
380
—
—
Elongation %
—
4.9
4.4
—
—
12
19
—
18
19
410
490
—
530
560
GIC J/m2
—
110
990
250
650
KIC mpa • m1/2
—
1.9
1.2
1.5
Flexural strength
ksi
Flexural modulus
ksi
The properties of several thermally commercially available VARTM resins are given in Table 3.
The VARTM resin systems being considered for this comparison are materials having a service
temperature around 100°C. The three thermally-cured reins are all epoxy-based systems. The Resinfusion
8611 system manufactured by Ciba Geigy and the Shell 863/Lindride 6k systems are amine- and
anhydride-cured systems respectively. Both need to be mildly heated for optimal viscosity and therefore
require processing temperature 20–30°C above ambient conditions. SC-15 is a two-phase toughened
amine-cured epoxy supplied by Applied Poleramics Inc. that possesses a sufficiently low viscosity to
allow for room temperature VARTM processing.
27
The EB-curable systems were developed under the US Army Research Laboratory program,
“Non-Polluting Composites Repair and Remanufacturing for Military Applications” and are based on the
epoxy-methacrylate IPN systems described earlier. VCCM4 is a two-phase toughened system based on
VCCM1.
The higher fracture toughness of the IPN systems compared to the Shell 862/Lindide system.is
apparent from the data in Table 3. Toughening using second-phase tougheners results in a 2.5 fold GIC
increase for VCCM4 compared to VCCM1. Toughening is accompanied by an acceptable viscosity rise,
practically unchanged Tg, and slightly improved flexural properties.
The thermally cured toughened SC-15 has the highest values of GIC and KIC but lower Tg and
modulus values. In general, epoxy-based VARTM resins do not perform as well as the systems
formulated for prepreg applications. This is a direct result of the requirement for low viscosity at ambient
temperature. Thus, the IPN-based EB-curable systems outperform many of the commercially available
thermally cured VARTM systems and possess a very good balance of properties. These include control
over viscosity, fracture toughness, Tg, stiffness, and strength. IPN based EB-cured toughened VARTM
resin systems with higher Tg (180°C) are currently under development.
A comparison of composite properties for thermally cured versus EB-cured VARTM-processed
carbon-fiber composites is given in Table 4. The composites were manufactured using 5 harness satin
(5HS) fabric woven from 12k carbon tows.
The data for 3501-6 are included for comparison to
autoclave-processed prepreg materials standard to the aerospace industry. The property data for the
thermally cured Shell 862/Lindride 6k system were obtained under the DARPA-sponsored Bridge
Infrastructure Renewal Program for the design and construction of a carbon-fiber Composite Army
Bridge.
Generally, the composite properties of the EB-cured materials are significantly better than those
of the Shell 862/Lindride 6k system. The EB-cured systems have much better compressive properties 98-110 ksi for VCCM resins compared with 80 ksi for the thermal system. This is perhaps a result of the
greater fracture toughness of the IPN EB-cured resins. In addition, the interlaminar shear strength for the
EB systems is slightly greater than that of the thermally cured system.
Carbon- fiber composites
fabricated by VARTM using VCCM1 and VCCM4 also compare favorably with autoclave-processed
3501-6 based systems.
28
Table 4: Comparison of thermally-cured and EB-cured composites fabricated with Vacuum
Assisted Resin Transfer Molding (VARTM). All data is normalized to 62% fiber volume fraction.
Thermal cure data from Manufacturer's specifications and Ref. [47]. EB cure data from Ref. [46,
48].
Thermal Cure
EB Cure
Resin
Epon
862/Lindride 6
3501-6
Description
VARTM
Autoclave
Fiber
12k 5HS T-300
Tg (°C) E"
Tensile strength
ksi
Tensile modulus
msi
Compressive
strength ksi
Compressive
modulus msi
Flexural strength
ksi
Flexural modulus
msi
ILSS
4.3
VCCM1
VCCM2
12k 5HS AS4 G'
12k 5HS AS4 G'
12k 5HS AS4 G'
120
190
132
129
115
120
121
125
10.3
10.4
10.6
10.8
80
100
98
110
9.5
10.4
10.0
10.8
—
150
132
135
—
—
10.4
10.4
8.0
10.0
8.4
8.8
EB-curable resins: further development
Much of the early materials development work to date has taken what could be termed an
Edisonian approach since initially key factors that affect the performance of EB-cured resin systems and
to some extent what properties they influence were not well understood. The limitations of today’s EBcured resin systems and composites are clear. The cationically-cured epoxy systems generally do not
possess sufficient fracture toughness and composite properties such as compressive strength that are
29
sensitive to resin performance suffer accordingly. In addition, interlaminar shear strength of EB-cured
composites is much lower than is acceptable for standard thermally cured epoxies. This is due not only to
the low fracture toughness of the EB system but also to poor interfacial strength between carbon fiber
surfaces that have been optimized for thermal cure of epoxy-amine systems and not EB-cured epoxies and
acrylate/methacrylate composites.
The EB curing community has recognized these limitations, and several research and
development programs are underway to address these issues. Two examples are the Army Research
Laboratory-led SERDP program and a second phase of the Oak Ridge National Laboratory-led
consortium. The ongoing ARL program is working to improve fracture toughness, a key barrier to the
widespread insertion of environmentally-friendly EB resin systems in current and future military
composite applications. The ORNL-led consortium has been established to address the issues relating to
poor interfacial performance.
The ARL program has taken advantage of the C-staging capacity of interpenetrating network
systems to control phase separation of a toughening additive prior to irradiation. The goal is to develop a
microstructure that provides toughness.
The short polymerization time for traditional cationic and
acrylate/methacrylate systems makes it difficult to control phase separation and microstructure during EB
processing. Further improvements of toughness for these materials will require a more complete
understanding of EB cure mechanisms and the associated kinetics and thermodynamics of
polymerization.
Historically, much work was conducted to improve the interfacial strength between carbon fibers
and thermally cured epoxy systems. Eventually the proper surface treatments and sizing materials were
developed to optimize this critical aspect of composite performance. The solutions relied initially on trial
and error and more recently have been based on a fundamental understanding of the chemical and
physical processes that occur in the interphase region. It is well known that fibers have the potential to
perturb polymer network structure in the vicinity of the surface. New mechanisms influencing interfacial
behavior may be associated with EB cure. These range from chemical incompatibility to physical
phenomena linked only to irradiation. Investigations by the ORNL-led consortium are addressing these
issues by developing sizings and coupling agents to improve interfacial performance with cationically
cured systems. The success of much of this and similar work will depend on attaining a more complete
understanding of the processing-structure-property relationships of EB-cured epoxies.
30
5
EB-curable adhesive development and current status
Adhesives that cure using electron beams are a subset of radiation-curable adhesives, which also
include systems that cure by ultraviolet light (UV). Until about 1995, all commercially available EB
curing adhesives were developed for UV. During the past 5 years, there has been considerable effort to
develop EB-curable adhesives for aerospace and automotive applications. The aerospace work is in
response to the results of the APCS and IATA programs that showed that new EB-curable adhesives are
needed to reduce fabrication and assembly costs and to increase design freedoms. There has also been
demand from the Composite Affordability Initiative (CAI) program, which has tested and ranked many
commercially available EB-curable adhesives.
The first experiments on EB-curable adhesives began 20 years ago.
Campbell et al. [49]
recognized the advantages to include fast curing, reduced air pollution due to 100% reactive formulations,
reduced energy consumption and reduced residual thermal stress. His group evaluated both acrylate and
bismaleimide formulations for bonding 7075-T6 aluminum to itself. Lap shear strengths up to 14 MPa
were obtained for the acrylate-based adhesives, obtained from the Loctite Company. They found up to
90% strength retention after immersion in water, jet fuel and acetone at room temperature for 500 hours
and after 500 hours in air at 350°F.
Campbell and Brenner also studied cross-link density effects in epoxy acrylate adhesive
formulations and found that a difunctional reactive diluent resulted in a 20% higher lap shear strength
than the same epoxy acrylate diluted with a monofunctional reactive diluent [50].
The EB-cured
formulations have up to about 5% greater strength than the same formulations cured with benxoyl
peroxide at 80°C.
Other groups have studied electron beam curing of adhesives more recently, with results
indicating that commercially available UV-curable acrylate adhesives bonded to both thermally and EB
cured carbon fiber composites are insufficient for aerospace applications. Lopata et al.[51] studied some
of the same Loctite adhesives as Campbell et al., as well as more advanced Loctite adhesives. They
achieved lap shear strengths up to 21 MPa on sanded or chromic acid-treated composites. Scanning
electron microscopy results suggest that a surface primer used on carbon fiber composites can improve
the adhesive bond strength.
Farmer et al.[52] studied EB cured epoxy paste adhesives for bonding the Air Force T-38
supersonic jet trainer’s composite windshield frame, composed of 36 adhesive-bonded sheets of 301 hard,
0.41mm thick stainless steel. The materials were all cationic epoxies developed originally for composite
matrices by the Oak Ridge National Laboratory-led CRADA [13]. Their results, summarized in Figure
31
19, indicated that the EB curable epoxies came within 13% of the thermally cured adhesive and met the
minimum requirements for this application.
Figure 19: (a) Lap shear strength of cationic adhesives on stainless steel substrate for bonding of
the T-38 windshield. (Application described in Section 2.4.) (b) Lap shear strength with various
scrim cloths and Adhesive 11L. The T-38 windshield has a minimal acceptable bonding strength of
21 MPa. Adhesive labeled EA is the Hysol EA9628 thermally-cured baseline. (From ref. [52], with
permission.)
However, there are several difficulties with using cationic epoxies as adhesives. A lap shear result
of 21 MPa (3100 psi) is frequently not sufficient for aerospace applications. Even if this value is
acceptable at ambient temperature, it leaves no margin for reduction in strength that typically occurs as
the temperature increases. Cationic epoxies are inhibited by nucleophilic compounds often added to
adhesives to improve properties as well as found on the surfaces of cured composite resins and primers.
This makes it difficult to improve their adhesive properties and restricts their range of application. Two
free-radical and eight cationic adhesives were recently tested by the Composite Affordability Initiative
32
(CAI) program. The average cationic epoxy lost 70% of its strength when bonding a (nucleophilic) 977-3
composite substrate, whereas the free-radical systems maintained full strength.
Table 5. Property comparison of aerospace and EB-curable adhesives. The Cytec adhesives are
widely used higher-temperature curing systems. Hysol EA 9396 is the heat curable baseline for this
work. Loctite 334 was the strongest commercially available EB-curable adhesive as of 1996 [8]. The
most recent EB-curable acrylate adhesives meet the 9394/6 baseline [53]. EB-curable film
adhesives require considerable improvement to meet aerospace structural requirements. Test
Conditions: aFloating Roller Peel(in-lb/in) bClimbing Drum Peel Test(in-lb/in) cBell-Peel Test(lb/in)
d
T-Peel Test(lb/in) Loctite 334 data is on grit-blasted steel substrate. Other data is for primed
aluminum substrate.
Adhesive
Cytec
Cytec
FM – 300 BR - 100
Film
Paste
(1)
(2)
Hysol
EA 9394/6
Paste
(2)
Loctite
334
Paste
(1)
Type and
number of
Components
Cure Cycle 175°C/2h 93°C/1h 25°C/5 days EB/Room
Temp
65°C/1hr
Cure
40
40
40
15
Pressure
(psi)
Ambient
Ambient
Storage
-18°C
5°C
Temperature
Shelf Life
6
6
Long
>12
(months)
Tensile
5080
2120
3300
No data
Shear (psi):
-55°C
5145
3770
3500
2600 - 2800
24°C
3995
2810
2300
1100 -1300
120°C
2910
2320
1800
420 - 600
150°C
Peel
29a
12b
20 - 25c
15 - 25d
Strength
(lb/in)
EB-cure
Acrylate
Paste
(1or 2)
EB-cure
Film
Film
(1)
EB/Room
Temp
15
EB/Room
Temp
15
Ambient
Ambient
>12
>12
2800-3300
No data
4000-5700
2000-3500
No data
13-25
1200-1700
No data
No data
No data
As of 1996, the Loctite 334 adhesive was considered the best commercially available EB-curable
paste adhesive based on work by Lopata et al. [51] and unpublished Northrop-Grumman IR&D data.
Loctite 334 was used to bond a variety of EB-cured composites in the Affordable Polymer Composite
Structures Program, including co-cured subassemblies and a full-scale composite fuselage assembly [7].
The properties of Loctite 334 are listed in Table 5. Table 1 also shows the properties of commercially
available adhesives including Hysol 9394/6, a heat-curable system that has been qualified by Lockheed
Martin for bonding large cryogenic tanks [20, 54, 55].
A recent US Army-sponsored program has led to the first EB-curable adhesives with properties
needed for aerospace applications. The adhesives developed by Science Research Laboratory Inc., the
33
Army Research Laboratory and the University of Delaware are of three types: one-component paste
adhesives, two-component paste adhesives and film adhesives.
adhesives are based on free radical chemistry.
The one-part adhesives and film
The two-part dual curing adhesives have an
interpenetrating network due to both thermal reaction and EB-crosslinking. Such a hybrid material is
similar in concept to the B-stageable EB-curable IPN resins described in Section 4.1.3 [42] and the dualcuring adhesive used for bonding of the Chrysler Composite Concept Vehicle described in Section 3.2
[25].
The properties of these systems are shown in Figure 20 and are comparable to the baseline
thermally cured system (Hysol 9394).
40
Strength (MPa)
35
30
25
Acrylate
20
Epoxy Acrylate
15
Hysol 9394
10
5
0
Ambient
Hot
Cold
Hot-W et
Thermal
Cycle
Figure 20: Lap-shear strength of one-part (Acrylate) and two-part (Epoxy Acrylate) EB-curable
adhesives on aluminum substrate, compared with a heat-curable baseline. Testing parameters are
Hot: 82°C, Cold: -55°C, Hot-Wet: Water immersion at 90°C for 24 hours, Thermal Cycling from
-194°C to 140°C [53].
6
Equipment and facilities for EB curing and bonding
Electron beam equipment has been used for the past 50 years for radiation processing of
materials. The many types of EB accelerators and their economics for polymer crosslinking are described
in Refs [56, 57]. With the exception of Aerospatiale’s system for curing rocket motors, EB facilities used
for composite curing have been designed for other uses; i.e. irradiation of wire, tubing and film or
34
sterilization of medical products. Figure 21 shows such a system. This 10 MeV, 50 kW EB facility in
Cranbury, NJ was used for several of the programs described in this review.
Figure 21: EB curing research has primarily used equipment designed for other industrial
processes. The Impella 10 MeV accelerator was designed for sterilization of medical products, but
has been extensively used for composite curing experiments. (a) EB curing of composite tanks. (b) A
cart-on-track system brings product into the x-ray shielded concrete maze. Photos courtesy of T.
Knobel, Ebeam Services, Cranbury, NJ.
Several of the features characteristic of high-energy electron beam processing facilities can be
seen in Figure 21. The scan horn in Figure 21(a) sweeps a “curtain” of electrons back and forth, while the
transport system moves the part under the horn. This facility uses a cart-on-track system, whiles others
use conveyor belts to move product through the concrete radiation-shielding maze. The concrete wall
thickness scales with beam energy, and is typically 1.5-3 meters thick for beam energies in the range of 310 MeV.
Safety interlocks prevent personnel from entering the shielded vault while the machine is in
operation. Ozone gas is generated when electrons pass through air and then exhausted from the curing
area by high-speed blowers.
Researchers have proposed and demonstrated other facilities for EB curing and bonding beside
the traditional closed maze shown in Figure 21. Because narrow concrete walls and a fixed scan horn
place limitations on the size and shape of composites to be cured, Aerospatiale and Science Research
Laboratory have built large shielded rooms with movable accelerators. These facilities are accessible to
large composite parts through sliding doors or removable concrete blocks.
35
Figure 22: Sketch and photo of a filament winding device coupled with a low energy electron
accelerator. Reprinted from Ref [16], by permission from the Society for the Advancement of
Material and Process Engineering (SAMPE).
Cart or conveyor handling systems do not provide precise positional control, and make it difficult
to attach sensors to parts during cure. A new facility dedicated to composite curing research uses an
accurately positioned table that moves back-and-forth under the beam to overcome these problems [58].
This will allow the use of in-situ cure monitoring and other in-line measurements.
Constructing new EB curing facilities is very expensive. Cost analysis performed under the APCS
program estimated $5 million for a facility capable of meeting their production needs, with the majority
of that cost due to items other than the electron beam system [8]. One way to lower this cost is to install
low energy (0.2-0.3 MeV) EB systems into existing facilities and use in-situ or layer-by-layer curing. Xray shielding requirements at these energies are 2-6 cm of lead or 10-30 cm of concrete blocks. This
approach is being used for filament winding [16] and for automated tape placement [18].
Figure 22 shows a sketch and a photo of a filament winding system coupled with a low energy
(0.25 MeV) electron accelerator. The system is enclosed in lead, allowing personnel to stand next to the
machine during operation. Figure 23 shows an Automated Tape Placement (ATP) system coupled with a
portable, low energy (0.25 MeV) electron gun. The EB gun is attached to the ATP head, and the entire
assembly is attached to a four-axis gantry system. The ATP system is shielded with a wall of portable
concrete blocks that can be easily installed or removed. The wall can be seen in the background.
Although the in-situ EB curing systems shown in Figure 23 are easier to install and cost less than
high-energy systems, obtaining good composite properties are more difficult. To obtain good
consolidation, researchers often use a warm debulk and vacuum bag pressure prior to high energy EB
curing. The ATP head applies heat and pressure for only a very short time, proving little room for error.
Good tape quality with full fiber wetting is critical to producing in-situ cured composites without voids.
36
This is because unlike conventional autoclave curing, there is no time for resin flow to fill tape voids, and
thus flaws must be avoided during tape fabrication.
Voids can also occur if the top layer springs back
due to insufficient “memory” before it is cured. The degree of spring-back as a function of EB cure dose,
as well as the tack, drape and memory of the uncured tape are determined by the visco-elastic properties
of the EB-curable resin.
Figure 23: In-situ or layer-by-layer EB curing allows use of lower energy electron beam systems. It
is thus possible to add the EB gun and localized x-ray shielding to existing composite facilities An
in-situ curing system at NASA Marshall Space Flight Center in Huntsville, AL laying and curing
tape is shown. The EB gun is on the left and the ATP head is in the foreground. A concrete block xray shielding wall is in the background.
Despite these challenges, initial results from in-situ curing systems are encouraging. Researchers
using the filament winding system shown in Figure 22 report interlaminar shear strength equal to highenergy EB-cured systems [16]. Researchers using the ATP system have yet to report their results using
the ATP system, but found good interlaminar strengths using an hand-operated ATP simulator [18].
37
7
Conclusions and future directions
Electron beam curing and bonding is a promising new technology that offers technical and cost
advantages for fabricating aerospace and automotive structures. Technical advantages for aerospace
structures include curing at room temperature using low-cost tooling, and the ability to fabricate large
integrated structure including structures too large to fit inside autoclaves. Studies by aerospace companies
have shown potential cost savings of 10-60% by using EB curing and bonding of integrated structures.
For automotive fabrication, advantages of EB processing include high production rates and near-room
temperature curing and bonding.
Researchers have spent a significant amount of effort during the past decade to develop this
technology. The majority of this work has been to develop processes for manufacturing aerospace
structures. Some of this work combines EB curing with prepreg and VARTM composite fabrication
techniques. These methods have been used to demonstrate large integrated structures such as fuselages,
and wing skins. Other demonstrations have combined filament winding with EB curing to produce rocket
motors and large filament-wound pressure vessels.
Considerable work has also been spent to formulate EB-curable materials. EB-curable resins and
adhesives have been developed with properties close to those needed for aerospace applications. EBcurable adhesives with the properties needed for automotive bonding are now available.
Several companies have also developed specialized electron beam equipment for curing and
bonding of composites. For example, in-situ EB curing systems using low-energy electron guns with a
minimal amount of X-ray shielding are being added to existing composite facilities.
Despite these successes, EB curing and bonding has not yet been widely adopted in industry, and
the technology remains primarily an area of research. For example, research areas include toughening of
cationic resin systems, improvement of interfacial properties, and extending use temperature ranges.
In 1997, the Air Force technical monitors for the APCS and IATA programs evaluated the
potential of EB for aerospace manufacturing [2]. Their conclusions provide a roadmap for the continued
development and commercialization of EB processing. Although there has been some additional work in
the last several years, especially in the areas of material testing, pre-qualification and novel EB facility
development, their conclusions remain valid:
"From studies of successful transitions of technologies from the laboratory to structural
applications, some common factors have been identified. These factors include the
following that are important to bear in mind when working with materials and processes:
stabilized material and/or material processes; characterized mechanical properties; and
predictability of structural performance. Although this technology [EB curing and bonding]
is still quite new to the aerospace composites community, it is important to recognize the
development path necessary for this technology to be considered for use on military weapon
systems. Stabilized materials and processes are essential for designers of systems.
38
Preliminary material qualification, processing specification and acceptance standards are
sought. Material suppliers capable of supplying the material in appropriate quantities and
forms are needed. Manufacturing processes compatible with the shop environments are
required. Structural properties must be sufficiently defined such that usable weight estimates
are possible. Properties must be evaluated for material performance at acceptable extremes
of material and process parameters. And finally, for the materials and process to reach a
stage where designers can confidently incorporate it, statistically significant data must be
generated. Such a development program represents a significant investment."
8
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