ArticlePDF Available

Processing and performance of out-of-autoclave bismaleimide composite sandwich structures


Abstract and Figures

Bismaleimides (BMI) are thermosetting polymers that are widely used in the aerospace industry due to their good physical properties at elevated temperatures and humid environments. BMI-based composites are used as a replacement for conventional epoxy resins at higher service temperatures. Out-of-Autoclave (OOA) processing of BMI composites is similar to that of epoxies but requires higher cure temperatures. Polymer properties such as degree of cure and crosslink density are dependent on the cure cycle used. These properties affect mechanical strength as well as glass transition temperature of the composite. In the current research, carbon fiber/BMI composite laminates were manufactured by OOA processing. The void content was measured using acid digestion techniques. The influence of cure cycle variations on glass transition temperature and mechanical strength was investigated. Properties of manufactured specimens were compared with that of conventional autoclave cured BMI composites. Laminates fabricated via OOA processing exhibited properties comparable to that of autoclave cured composites.
Content may be subject to copyright.
Copyright 2015. Used by CAMX The Composites and Advanced Materials Expo with permission
CAMX Conference Proceedings. Dallas, TX, October 26-29, 2015. CAMX The Composites and Advanced Materials Expo.
S. Anandan, G. S. Dhaliwal, and K. Chandrashekhara
Department of Mechanical and Aerospace Engineering
Missouri University of Science and Technology, Rolla, MO 65409
T. R. Berkel
Boeing Research & Technology, St. Louis, MO 63166
D. Pfitzinger
GKN Aerospace, St. Louis, MO 63042
Composite sandwich structures offer several advantages over conventional structural materials
such as lightweight, high bending and torsional stiffness, superior thermal insulation and
excellent acoustic damping. In the aerospace industry, sandwich composites are commonly
manufactured using the autoclave process which is associated with high operating cost. Out-of-
autoclave (OOA) manufacturing has been shown to be capable of producing low cost and high
performance composites. Unlike the autoclave process, OOA processing avoids the issue of core-
crushing due to high pressure. However, Bismaleimide (BMI) prepregs require high cure and
post-cure temperatures which can lead to high internal core pressures, core-to-facesheet
disbonding and voids. In the current work, OOA sandwich composite panels are manufactured
using aluminum honeycomb core, BMI adhesive film and carbon/BMI prepregs. Two vacuum
levels were used during OOA processing, full vacuum (100 kPa) and partial vacuum (80 kPa).
Adhesive bond quality was evaluated using flatwise tensile and fracture toughness tests.
Mechanical performance was evaluated using edgewise compression. It was observed that
vacuum level variation during processing had no significant effect on mechanical properties of
manufactured laminates. Tests are performed at room temperature and 177 °C (350 °F). All
manufactured laminates exhibited room temperature flatwise tensile strengths comparable to
those of aerospace grade epoxy adhesives. Sandwich mechanical properties reduced when test
temperature was increased.
High performance composite materials are conventionally processed in autoclaves which are
associated with high capital requirements and operating costs. Part size is also limited by the
autoclave chamber volume. Out-of autoclave (OOA) processing has emerged as an alternative to
autoclave processing [1, 2]. OOA involves curing at atmospheric pressures which obviates the
need for an autoclave and provides greater design flexibility for production of large and complex
shaped composite components. It also reduces capital and operating costs. In an OOA prepreg
process, removal of entrapped air from the prepreg layup is critical which can otherwise lead to
voids which degrade mechanical properties of the composite. Sandwich components can be co-
cured, a process where adhesive and the prepreg are cured simultaneously, resulting in time and
cost reduction. Manufacturing co-cured sandwich panels out-of-autoclave, eliminates core
crushing which can occur at autoclave pressures [3] and enables the use of lighter cores.
In a composite sandwich structure, the facesheet absorbs in-plane loads while the core is resistant
to shear loading. The composite facesheet is bonded to the core using a film adhesive. The
adhesive bond between core and facesheet is of great importance in the design of sandwich
structures. The relation between the adhesive fillet and sandwich properties was described by
Grimes [4]. The mechanical performance of a sandwich composite depends on the quality of the
core-to-facesheet adhesive bond. In the case of honeycomb cores, the cell walls provide a
relatively small area for bonding. Structural strength depends on the presence of a well-formed
fillet at the interface of the core cell wall and facesheet. Larger adhesive fillets absorb a higher
amount of energy upon fracture, due to which they have high fracture toughness [5]. Grove et al.
[6] showed that the higher debonding energy was obtained with larger, regular-shaped adhesive
fillets between the honeycomb cell walls and the skin. Rion et al. [7] showed the failure occurs in
the adhesive meniscus when low weight adhesive is used. Butukuri et al. [8] evaluated the effect
of the reticulation technique on properties of sandwich constructions. Core reticulation resulted
in improved fillet geometry and flatwise tensile strength. Hou et al. [9] reported that thermal pre-
treatments can reduce adhesive foaming in OOA cured composite sandwich structures.
Nagarajan et al. [10] studied the effect of processing variables on OOA sandwich panels with
carbon/epoxy facesheets. Presence of dissolved moisture in the adhesive layer was found to
increase adhesive foaming.
Internal core pressure in the cells of a honeycomb core has been shown to affect adhesive
foaming and fracture toughness in sandwich composites with carbon/epoxy facesheets [11]. Core
pressure fluctuations were observed when internal core pressures during cure equaled
atmospheric pressure. Increased adhesive permeability due to perforation resulted in enhanced
fracture toughness of the adhesive layer. The optimal range of honeycomb core pressure during
processing was estimated to be 40-70 kPa, to reduce adhesive outgassing and ensure good fillet
formation [12].
OOA epoxy prepreg systems have previously been used in the Boeing wing spar, Advanced
Composite Cargo Aircraft (ACCA) and the Ares V interstage and payload shroud [13, 14]. OOA
Bismaliemide (BMI) systems were evaluated as a part of NASA CoEx (Composites for
Exploration) project [15]. Due to tooling limitations, the cure temperature was fixed at 177 °C
(350 °F) followed by a freestanding post-cure at elevated temperatures. BMI prepreg systems
from Renegade, Stratton Composites and Tencate were evaluated. Two adhesive systems namely
Cytec FM 2550 and Renegade RM 3011 were evaluated. Sandwich structures exhibited no
warping on post-cure, and flatwise tensile strength obtained was comparable to aerospace grade
epoxy adhesives.
In the current work, sandwich composites were manufactured using aluminum honeycomb,
carbon/BMI facesheets and a toughened BMI adhesive system. The sandwich composite was
cured via the OOA prepreg process. One set of panels was fabricated under full vacuum pressure
(100 kPa). Remaining panels were manufactured using a reduced vacuum pressure of 80 kPa.
This level was chosen to ensure that the internal core pressure does not exceed atmospheric
pressure during the high temperature cure. The effect of varying vacuum pressure on mechanical
properties of the sandwich composite was evaluated. Adhesive bond quality was evaluated using
flatwise tensile tests and adhesive bond fracture toughness tests. Load bearing capability
structure was evaluated using edgewise compression tests. Tests were performed at room
temperature as well as an elevated temperature of 350 °F (177 °C).
The sandwich facesheets were fabricated using unidirectional IM7/AR4550 prepreg, a toughened
carbon/BMI system suitable for OOA curing. It has a resin content of 35% and a prepreg areal
weight of 304.8 gsm. Cytec Metlbond 2550 film adhesive was used to bond the facesheet to the
honeycomb core. Metlbond 2550 is a modified BMI film adhesive with good high temperature
properties. The core material was aluminum honeycomb. Materials are listed in Table 1.
Table 1. Materials used
Aldila Composite Materials
Recommended cure cycle:
143.33 °C (290 °F) for 1 hour
followed by 190.5 °C (375 °F)
for two hours. Post cure at 232
°C (450 °F).
BMI film adhesive
(Metlbond 2550)
Cytec Engineered Materials
Recommended cure cycle:
177 °C (350 °F) for 4 hours. Post
cure at 226.6 °C (440 °F).
Density: 8.1 pcf (129.7 kg/m3)
Thickness: 1 in. (25.4 mm)
Cell size: 1/8 in. (3.175 mm)
Honeycomb sandwich composite panels were manufactured using an Out-of-Autoclave (OOA)
manufacturing process. One set of panels of size 304.8 mm. x 304.8 mm (12 in. x 12 in.) and
facesheet configuration [0 °/90 °]s were manufactured, which yielded samples for flatwise tensile
and edgewise compression tests. A second set of sandwich panels 304.8 mm. x 152.4 mm (12 in.
x 6 in.) and facesheet configuration [0 °/90 °]2s was manufactured for adhesive bond fracture
toughness evaluation.
An aluminum mold was cleaned of surface defects using sand paper, and a non-porous Ethelyne
tetrafluoro ethelyne (ETFE) release film was placed. The prepreg layers were cut into required
dimensions and laid up according to the required facesheet configuration. Rollers along with
hand pressure were used to press the prepregs, starting from the center of the layup and moving
progressively towards the edges. This process was repeated for all the prepreg layers to remove
entrapped air as well as folds or wrinkles. The tool-side and bag-side facesheet layups were
debulked at room temperature under full vacuum for 60 minutes to ensure good compaction.
The schematic of the bagging procedure employed for the OOA process is shown in Figure 1.
The compacted tool side facesheet layup was laid on the aluminum mold which was covered
with the ETFE release film. A layer of film adhesive was placed on the prepreg layup. The
honeycomb core was placed on the adhesive layer followed by the bag-side adhesive layer and
facesheet stack. Edge breathing using EB1590 edge bleeder was used to evacuate entrapped air
from the sandwich layup. A layer of ETFE release film was placed on the sandwich layup
followed by an aluminum caul plate. Airweave N10 breather was used to distribute the vacuum
pressure evenly. The entire layup was sealed using a double vacuum bagging scheme.
Figure 1. Schematic of bagging layup
The layup was debulked at 49 °C (120 °F) for two hours. The sample was cured according to the
cure cycle shown in Figure 2, and allowed to cool down to room temperature. This was followed
by a free-standing post-cure at 232 °C (450 °F) for two hours. The prepreg cure cycle was used
to manufacture the sandwich panel. The samples did not exhibit warping following post cure.
Figure 2. Cure Cycle
4.1 Thermogravimetric Analysis
Outgassing of prepreg and adhesive during cure can lead to voids in facesheets and adhesive
fillets. These volatiles need to be removed during cure. The evolution of volatiles from the
prepreg and adhesive was studied using Thermogravimetric analysis (TGA). The adhesive was
removed from storage and allowed to warm up to room temperature. A small sample was
removed and the prepreg cure cycle was simulated in a TGA. Air at a flow rate of 40 mL/min.
was used as the sample purge gas and the weight loss was recorded.
4.2 Flatwise Tensile Test
Flatwise tensile (FWT) strength primarily serves as a quality control parameter for bonded
sandwich panels. This test produces information on the quality and strength of the core-to-facing
adhesive bond. The flatwise tensile test for sandwich samples was done according to ASTM
C297. Sandwich samples had a facesheet configuration of [0°/90°]s. This test method consists of
subjecting a honeycomb sandwich samples to a uniaxial tensile force normal to the plane of the
sample. Honeycomb sandwich samples of 50.4 mm x 50.4 mm (2 in. x 2 in.) were cut from the
manufactured panel. The sample edges were polished progressively using 80-220 grit sandpaper.
The facings of the samples were lightly roughened using 80 grit sand paper. For RT testing,
aluminum blocks of 50.4 mm x 50.4 mm x 50.4 mm (2 in. x 2 in. x 2 in.) were bonded to the
sample facings using a high strength epoxy, 3M Scotch-Weld DP-460NS. The adhesive was
allowed to cure at room temperature for 24 hrs and was post cured at 121 °C (250 °F) for two
hours. Samples for elevated temperature testing at 177 °C (350 °F) were bonded to aluminum
blocks using Metlbond 2550 film adhesive. Test samples were conditioned at 177°C (350°F) in
the thermal testing chamber for 2 hours prior to testing. Average relative humidity was
maintained at 50%. Sample loading rate was 0.5 mm/minute. Each test was replicated four times.
4.3 Edgewise Compression Test
Edgewise compression test is used to measure the compressive properties of a sandwich
specimen in a direction parallel to the facing plane. This test is a measure of the load bearing
capacity of the sandwich structure. The edgewise compression test for sandwich samples was
performed according to the ASTM C364. Sandwich samples had a facesheet configuration of [
90°]s Sandwich specimens were cut into 50.4 mm x 127 mm (2 in. x 5 in.) sizes. Tests were
conducted at room temperature as well as 177 °C (350 °F). Test samples were conditioned at 177
°C (350 °F) in the environmental testing chamber for 2 hours prior to testing. Average relative
humidity was maintained at 50%. A loading rate of 3 mm/min. was used. Mechanical load was
applied along the core ribbon direction. Each test was replicated three times.
4.4 Adhesive Bond Fracture Toughness Test
The interfacial fracture toughness was measured using a modified cracked sandwich beam (CSB)
specimen [16]. Sandwich composites were manufactured with [0° 90°]2s facesheet orientation,
using the manufacturing procedure mentioned in section 3. Specimens of size 25.4 mm x 254
mm (1 in. x 10 in.) were extracted from the manufactured panel. An initial crack of 25.4 mm (1
in.) was cut as shown in Figure 3. Aluminum blocks, with hinges, were bonded onto the sample
as shown in Figure 4. A roller support was added at the free end to prevent large rotations.
Predetermined crack lengths were marked on the specimen and mechanical load was applied
using an Instron 5985 universal testing machine. This procedure was repeated for multiple
loading/unloading cycles. Crack growth was along the core ribbon direction.
Figure 3. Schematic of modified CSB test
(a) Test setup (b) Crack propagation on mechanical loading
Figure 4. Crack propagation in a modified CSB test
5.1 Thermogravimetric Analysis
Results of TGA are shown in Figure 5. The temperature was increased to 143.3 °C (290 °F) and
held for 60 minutes. This isothermal hold is required to maximize the mobility of reacting
groups. An initial mass loss of 1.8 % was seen during the first isothermal hold. The temperature
was then increased to the cure temperature of 190.5 °C (375 °F) at 10 °C/min. An adhesive mass
loss of 1.7 % was observed during this period. Mass loss during cure is due to adhesive
outgassing which can lead to foaming defects in the adhesive layer.
Figure 5. Mass loss of adhesive during TGA
5.2 Flatwise Tensile Test
Results of the FWT tests are shown in Table 2. All samples failed in a combination of cohesive
and adhesive failure mode. Values depicted in Table 2 are comparable to previously reported
FWT properties of BMI sandwich structures [15]. Samples cured at full vacuum pressure and at
partial vacuum exhibited slightly higher FWT strengths. The adhesive bond strength reduced by
23% (full vacuum) and 21% (partial vacuum) when samples were tested at 177 °C (350 °F). The
room temperature FWT strength values compare well with the FWT strength of 6167.7 kPa
reported by Butukuri et al. [8] for OOA cured epoxy sandwich systems using Cycom 5320
facesheets (Cytec), Cycom FM 300-2U epoxy adhesive (Cytec) and aluminum honeycomb core
(1/8 in. cell size, 12 pcf). Adhesive fillets are distributed on either side of the cell wall (Figure
Table 2. FWT test results
Flatwise tensile strength
Average (kPa)
S.D. (%)
Figure 6. Adhesive fillet formation in sandwich composites manufactured under full vacuum
Optical photographs of fracture surfaces are shown in Figures 7 and 8. A few voids are visible on
the facesheet side fracture surface of all samples (Figures 7 (a), 7 (b),8 (a) and 8 (b)). This can be
a result of adhesive outgassing during cure. Voids and air packets are seen along the fillet
regions (Figures 7(c), 7 (d),8 (c) and 8 (d)). These defects may be a result of low internal core
pressures which have been linked with adhesive foaming in previous studies [11]. The presence
of these defects did not have a significant effect on the FWT strength. While cohesive mode of
failure dominates in all specimens, regions of adhesive failure are also seen (Figure 8 (b)).
(a) Facesheet side (b) Facesheet side
(c) Core side (d) Core side
Figure 7. Representative images of fracture surfaces of samples cured under full vacuum at
different locations
(a) Facesheet side (b) Facesheet side
(c) Core side (d) Core side
Figure 8. Representative images of fracture surfaces of samples cured under partial vacuum at
different locations
5.3 Edgewise Compression Test
In the present study, edgewise compression test was conducted under room temperature and 177
°C (350 °F). Results of edgewise compressive test are shown in Table 3. Similar compressive
strengths were observed in samples cured under full vacuum and partial vacuum. When test
temperature was raised to 177 °C, the edgewise compression strengths dropped by 20.8 % and
32.2 % respectively. This can be an effect of reduced core stiffness under elevated temperatures.
Table 3. Edgewise compression test results
Vacuum pressure
Edgewise compressive strength
Average (MPa)
S.D. (%)
Full vacuum (100)
Full vacuum (100)
Partial vacuum (80)
Partial vacuum (80)
A representative load-displacement plot of edgewise compression test is shown in Figure 9. The
load increases gradually until it reaches a peak value. Progressive failure initiates at this point.
The exact location of facesheet failure was unknown. Continued loading resulted in stable end
crushing of the sandwich specimen (Figure 10).
Figure 9. Load vs Extension in edgewise compression test at 177 °C (350 °F ) of sandwich
samples cured under full vacuum
Damage initiation in the sandwich composites was in the form of facesheet failure, because the
failure strain of the core is much higher than that of the facesheets. The composite facesheets can
exhibit various failure modes such as elastic microbuckling, plastic microbuckling, fiber
crushing, splitting, buckle delamination and shear band formation. On continued loading after
failure initiation, sandwich structures can exhibit various failure modes such as (1) Unstable
buckling of the sandwich column; (2) Facesheet delamination followed by unstable buckling in
opposite directions; and (3) Stable end crushing of sandwich panels [17, 18].The unstable failure
modes are seen when the bending strength of the facesheets is high, compared to the strength of
the fracture interface. The sandwich falls apart rapidly as the crack in the interface propagates
along the length of the sandwich specimen. When the interface strength is high compared to the
bending strength of the facesheets, crack propagation is more stable. The crack arrests after a
short propagation distance and other failure modes such as facesheet bending, delamination and
core crushing begins to dominate.
(a) Front view (b) Isometric view
Figure 10 (a) and (b). Stable end crushing failure
5.4 Adhesive Bond Fracture Toughness Test
A sample load/displacement curve is shown in Figure 11. All tests were conducted at room
temperature. The load increases until it reaches a critical value and crack begins to propagate. As
the crack propagates, load reduces. Critical strain energy release rate was calculated from three
loading/unloading cycles for each specimen.
Figure 11. Loadingunloading cycle in modified CSB test of a sample cured under partial
The critical strain energy release rate, , can be expressed as,
= 
where, P is the external load;  is the displacement; B is the width of specimen and  isthe
incremental crack length during the process of test. The energy required for crack growth, ,
can be calculated by calculating the area enclosed by the loading-unloading path. The area under
the loading/unloading path is obtained by integration using the trapezoid rule. The average
interface fracture energy release rate for samples fabricated under full vacuum and partial
vacuum is shown in Table 4. About 10% variation was noticed in samples fabricated under full
vacuum and partial vacuum.
Table 4. Adhesive bond fracture toughness test results
Vacuum pressure
Strain energy release rate
Average (J/m2)
S.D. (%)
Full vacuum (100)
Partial vacuum (80)
High temperature sandwich composite panels were manufactured using the OOA prepreg
process. Two levels of vacuum pressure were studied, full vacuum (100 kPa) and partial vacuum
(80 kPa). The prepreg cure temperature were used to manufacture the panels. No warping was
visible after post cure at 232 °C (450 °F), which implies that the panels had sufficient green
strength to enable a free-standing post-cure. Outgassing of adhesive during the cure cycle was
studied using TGA. A mass drop of 1.7 % was observed during the ramp and cure isotherm.
These volatiles need to be removed during cure to avoid foaming issues. Adhesive bond strength
was evaluated using FWT tests and adhesive bond fracture toughness tests. The samples
manufactured using full vacuum pressure during cure exhibited in a 6-8.5 % increase in FWT
strength compared to those manufactured at a partial vacuum pressure. FWT strength reduced
when samples were tested at 177 °C (350 °F). Cohesive failure dominated in all samples while
some regions of adhesive failure were also present. Sandwich samples with [0°/90°]2s facesheet
orientation were fabricated and a modified CSB test was conducted at room temperature.
Samples exhibited a strain energy release rate of 335.19 J/m2 (full vacuum) and 301.80 J/m2
(partial vacuum). Edgewise compression tests were conducted on samples with facesheet
orientation [0°/90°]s. Vacuum pressure levels used in this study, did not have a significant effect
on edgewise compressive strength. Increase in test temperature to 177 °C (350 °F) resulted in a
reduction in edgewise compressive strength. This can be a result of reduced core stiffness at
increased temperatures. Measured FWT strengths at room temperature were comparable to those
mentioned in previous studies involving aerospace grade epoxy adhesives.
This research is sponsored by the Industrial Consortium of the Center for Aerospace
Manufacturing Technologies (CAMT) at Missouri University of Science and Technology. The
authors would like to thank Stratton Composite Solutions and Cytec Engineered Materials for the
materials supplied.
[1] T. Centea, L. Gunenfelder and S. Nutt, "A Review of Out-of-Autoclave Prepregs
Material Properties, Process Phenomena, and Manufacturing Considerations,"
Composites Part A: Applied Science and Manufacturing, vol. 70, pp. 132-154, 2015.
[2] C. Ridgard, "Next Generation Out-of-Autoclave Systems," Proceedings of the
International SAMPE Symposium and Exhibition, pp. 1-18, Seattle, WA, May 17-20,
[3] H. Hsiao, S. Lee and R. Buyby, "Core Crush Problem in Manufacturing of Composite
Sandwich Structures: Mechanisms and Solutions," AIAA Journal, vol. 44, no. 4, pp. 901-
907, 2006.
[4] G. Grimes, "The Adhesive-Honeycomb Relationship," Applied Polymer Symposium, vol.
3, pp. 157-190, 1996.
[5] R. Okada and M. Kortschot, "The Role of Resin Fillet in the Delamination of the
Honeycomb Sandwich Structures," Composites Science and Technology, vol. 62, pp.
1811-1819, 2002.
[6] S. Grove, E. Popham and M. Miles, "An Investigation of the Skin/Core Bond in
Honeycomb Sandwich Structures using Statistical Experimentation Techniques,"
Composites Part A: Applied Science and Manufacturing, vol. 37, no. 5, pp. 804-812,
[7] J. Rion, Y. Letterrier and J.E. Manson, "Prediction of the Adhesive Fillet Size for Skin to
Honeycomb Core Bonding in Ultra-Light Sandwich Structures," Composites: Part A, vol.
39, pp. 1547-1555, 2008.
[8] R. Butukuri, V. Bheemreddy, K. Chandrashekhara, T. Berkel and K. Rupel, "Evaluation
of Skin-Core Adhesion Bond of Out-of-Autoclave Honeycomb Sandwich Structures,"
Journal of Reinforced Plastics and Composites, vol. 31, no. 5, pp. 331-339, 2012.
[9] T. Hou, J. Baughman, T. Zimmerman, J. Sutter and J. Gardner, "Evaluation of Sandwich
Structure Bonding in Out-of-Autoclave Processing," Proceedings of the International
SAMPE Technical Conference, pp. 1-12, Salt Lake City, UT, October 11-14, 2010.
[10] S. Nagarajan, V. Menta, K. Chandrashekhara, T. Berkel, J. Sha, P. Wu and D. Pfitzinger,
"Out-of-Autoclave Sandwich Structure: Processing Study," SAMPE Journal, vol. 48, no.
4, pp. 24-31, 2012.
[11] S. Tavares, N. Callet-Bois, V. Michaud and J.-A. Manson, "Vacuum-bag Processing of
Sandwich Structures: Role of Honeycomb Pressure," Composites Science and
Technology, vol. 70, pp. 797-803, 2010.
[12] J. Kratz and P. Hubert, "Processing Out-of-Autoclave Honeycomb Structures: Internal
Core Pressure Measurements," Composites Part A: Applied Science and Manufacturing,
vol. 42, no. 8, pp. 1060-1065, 2011.
[13] G. Marsh, "De-autoclaving Prepreg Processing," Reinforced Plastics, vol. 56, pp. 20-25,
[14] J. Sutter, W. Kenner, L. Pelham, S. Miller, D. Polis, C. Nailadi, T. Hou, D. Quade, B.
Lerch, R. Lort, T. Zimmerman, J. Walker and J. Fikes, "Comparison of autoclave and
Out-of-Autoclave Composites," Proceedings of the SAMPE Fall Technical Conference,
pp. 1-15, Salt Lake City, UT, October 11-14, 2010.
[15] T. Hou, S. Miller, T. Williams and J. Sutter, "Out-of-Autoclave Processing and Properties
of Bismaleimide Composites," Journal of Reinforced Plastics and Composites, vol. 33,
no. 2, pp. 137-149, 2014.
[16] S. Smith and K. Shivakumar, "Modified Mode-I Cracked Sandwich Beam (CSB)
Fracture Test," Proceedings of the 19th AIAA Applied Aerodynamics Conference, pp. 1-
18, Anaheim, CA, 11-14 June, 2001.
[17] A. Mamalis, D. Manolakos, M. Loannidis and D. Papapostolou, "On the Crushing
Response of Composite Sandwich Panels Subjected to Edgewise Compression:
Experimental," Composite Structures, vol. 71, no. 2, pp. 246-257, 2005.
[18] A. Lindström, In-Plane Compressive Response of Sandwich Panels, Doctoral
Dissertation: KTH Engineering Sciences, Stockholm, Sweden, 2009.
... Fiber-reinforced polymers are lightweight materials that can be tailored for modern high-performance structural applications allowing for efficient engineering solutions to severe and varying operating conditions [1,2]. Carbon fiber reinforced polymer (CFRP) composites have gained significant attention in recent decades and are widely used in the aerospace industry to replace metals due to their excellent mechanical and electrical properties [1,[3][4][5][6][7]. Autoclave manufacturing is the dominant manufacturing process used to produce high-performance composites, especially for aerospace applications. ...
... The temperature and pressure requirements add demand to tooling materials used during this manufacturing process. As a result, alternative curing methods such as the out-of-autoclave (OOA) prepreg process have been investigated to manufacture high-performance composites [5,7,8]. ...
Full-text available
Carbon fiber-reinforced polymer composites have been increasingly used by the aerospace, automobile, and other industries due to their high specific stiffness and strength properties. Manufacturing polymer composites using conventional prepreg curing is time-consuming and energy-intensive. Microwave curing of composites is an attractive alternative technology to manufacture composites with exceptional processing speed and energy savings. In this work, a traditional microwave oven was modified and used to manufacture composite panels. The void content, short beam strength properties, flexural properties, and impact response of the microwave cured panels were evaluated. Composite panels (control) were manufactured using carbon/epoxy prepreg cured in walk-in oven. The performance of microwave cured composite specimens was compared against the performance of conventional walk-in oven cured composite specimens. Results showed that the short beam strength and flexural strength of panels cured in the microwave oven remained approximately the same as the control specimen, however, the flexural modulus, the impact peak force, and the rebound energy increased by 12%, 6% and 18% respectively for the panels manufactured through microwave curing.
This paper successfully achieves the in-situ monitoring on core pressure of honeycomb sandwich composite during hot-pressing process by using a self-designed pressure sensing system. The core pressure variation and surface quality of honeycomb sandwich composites were investigated under open or sealed conditions, where the edges of sandwich composites were exposed to air or sealed with sealant tape, respectively. The results show that the core pressure increases with the increase in temperature and then declines at the temperature platform and cooling stage under open condition. The core pressure at the central area is bigger than that at the edge areas, and the dimpling defects are most pronounced on the thin facesheets at the edge areas under open condition during hot-press process. Under sealed condition, the core pressure is greatly enhanced and maintains at a higher pressure during the whole manufacturing process, and thus the dimpling defects are minimized. The maximum core pressure at the edge area and central area of 25 cm × 25 cm sandwich composite may reach 4 kPa and 13 kPa, respectively. Furthermore, by using a self-designed device, the air permeability of the whole sandwich structure is measured, and a two-dimensional model is built to simulate the air flow and core pressure variation inside sandwich composites during hot-pressing process. The simulated core pressure variation shows the same tendency with experimental results, which explains the mechanism of core pressure variation and guides the control of process quality.
Sandwich composite structures are comprised of a low-density core (commonly honeycomb) and facesheets. They are typically used in applications that require lightweight for efficient design, such as in the marine and aerospace industries. This work investigates the feasibility of adopting triply periodic minimal surface (TPMS) cellular structures as the core for sandwich composites. Sandwich structures were manufactured using a carbon fiber-reinforced polymer (CFRP) facesheet and three different 304 L stainless steel core structures (honeycomb, gyroid TPMS, and diamond TPMS). Three mechanical tests, namely edgewise compression, three-point bend, and impact test, were carried out to evaluate the performance of each sandwich configuration. The experimental results of the non-traditional sandwich configurations were compared against those of a honeycomb core sandwich composite. The edgewise compression test showed that the ultimate edgewise compressive strength increased by 7% when the honeycomb core was replaced by the gyroid core and reduced by 2% when the diamond core replaced the honeycomb core. The three-point bend test showed that the traditional honeycomb core sandwich configuration had a higher shear yield stress when compared to the non-traditional sandwich structures. The shear yield stress was reduced by 54% when non-traditional sandwich cores were used. The shear ultimate stress was reduced by 41% and 37% when the honeycomb core was replaced by the gyroid and diamond structure, respectively. Impact test results, on the other hand, showed that the peak force recorded during the impact event was reduced, while the absorbed energy was increased when non-traditional cores were used. Peak force was reduced by 28% and 39%, while the absorbed energy was increased by 9% and 16% when the honeycomb core was replaced by the gyroid and diamond cores, respectively.
Full-text available
Composite sandwich structures offer several advantages over conventional structural materials such as lightweight, high bending and torsional stiffness, superior thermal insulation and excellent acoustic damping. One failure mechanism in a composite sandwich structure is the debonding of the composite facesheets from the core structure. A well-formed adhesive fillet at the interface of the honeycomb core cell walls and the laminate is a significant factor in preventing bond failure. In the present work, honeycomb composite sandwich panels are manufactured using a low-cost vacuum-bag-pressure-only out-of-autoclave manufacturing process. CYCOM®5320 out-of autoclave prepreg is used for the facesheet laminates and FM® 300-2U film adhesive is used for the facesheet-to-core bond. The manufactured composite sandwich panels are of aerospace quality with a facesheet laminate void content of around 1%. In this study, adhesive fillet formation and adhesive mechanical strength are evaluated as a function of several different sandwich construction design variables. Both aluminum and aramid Nomex® honeycomb core materials are considered to study the effect of core cell size and core material. The effect of film adhesive thickness is studied. A process for reticulation of the adhesive is applied and studied. A quantitative investigation of the adhesive fillet geometry is carried out for all the panels. Manufactured panels are evaluated for flatwise tensile strength in accordance with test method ASTM C297. Optimized combinations of core material, core density, cell size and adhesive thickness are identified. Results show that the reticulation process improves fillet formation and increases flatwise tensile properties.
Full-text available
Manufacturing of composite honeycomb sandwich structures is significantly impacted by poor production yields caused by the core crush problem that occurs during the autoclave curing process. It is a major manufacturing defect that leads to costly part rejects because the defects are nonrepairable. This failure mechanism also constrains aircraft engineers, limiting the design range of core density and core thickness in attempts to mitigate core crush. Recent studies that have led to basic understanding of core crush mechanism are discussed. It was found that the prepreg frictional resistance is the key factor in controlling core crush. Research in the scientific community has mainly focused on resin effects in core crush. However, studies conducted show that core crush can also be significantly reduced by controlling construction of the fiber tow shape and the fabric architecture. Rounder fiber tow or more open fabric produces rougher prepreg surface, which results in a higher prepreg frictional resistance, reducing the effects of core crush. Experimental results indicate that a developed core crush resistant prepreg increases the prepreg frictional resistance and effectively reduces core crush, without changing the resin system.
Composite sandwich structures offer several advantages such as reduced weight, high bending/torsion stiffness, superior thermal insulation and excellent acoustic damping over conventional structural materials. One failure mechanism in a composite sandwich structure is the debonding of the composite facesheets from the core structure. A well-formed adhesive fillet at the interface of the honeycomb core cell walls and the laminate is a significant factor in preventing bond failure. Improved adhesion strength with a well-formed adhesive fillet at the interface of the honeycomb core walls and laminate is the aim of sandwich composite manufacturing, and is one area of investigation in this work. Out-of-autoclave (OOA) processing of composites offers several key benefits compared to autoclave processing such as lower manufacturing cost resulting from a lower capital cost and lower energy consumption. In the present work, the effects of moisturized adhesive film, core type, material system and variable levels of vacuum on adhesive fillet formation are evaluated. Aluminum honeycomb and vented aluminum honeycomb are used as core materials in this study. Two material systems of prepregs and adhesive films, Cycom 5320 prepreg with FM 309-1M adhesive, and MTM45-1 prepreg with MTA-241/PK13 adhesive are used to manufacture the panels and evaluate the performance. A quantitative investigation of the adhesive fillet geometry is carried out for all the panels. Manufactured panels are evaluated for peel strength and flatwise tensile strength in accordance with ASTM standards. Various combinations of core material, moisture content, material system and vacuum level are evaluated for effects on adhesive strength and adhesive foaming.
The emergence of bismaleimide composites has fulfilled some of the increasing demand for higher temperature performance aeronautics and space exploration vehicles. This study examines and evaluates three bismaleimide matrix resins and two bismaleimide adhesives and reports on the processing properties of these resins and composites by out-of-autoclave-vacuum-bag-only oven processing. Measurements were conducted under various cure cycles to characterize and correlate thermal and viscoelastic properties of the materials. Specimens of all three aged matrix resins exhibited an out-time life up to 30 days when stored at room temperature. Solid and sandwich panels were fabricated with the out-of-autoclave-vacuum-bag-only process. Because of tooling limitations in industry practices, composite fabrication of these bismaleimides was restricted to a maximum 177? curing, followed by a free-standing postcuring at elevated temperatures in an oven. The adhesive foaming characteristics, composite resin/void content, flat wise tensile strength, and fracture surface features were evaluated. Due to the unique temperature limitations of this work, the resulting panel properties were not necessarily representative of manufacturer specifications or recommendations.
While prepregs have speeded up the composite process and improved quality and consistency, acquiring and operating large autoclaves is very costly and their finite capacities create process bottlenecks. Hence out-of-autoclave prepreg processes able to speed up fabrication and reduce costs whilst also equalling autoclave quality have become a focus of development. George Marsh reports.
A strong desire to reduce manufacturing costs in the aerospace industry, coupled with recent advancements in out-of-autoclave prepreg materials, has created the potential to manufacture composites with only vacuum pressure. As a step towards better manufacturing of sandwich structures, we focus on the behaviour of air inside a honeycomb core during vacuum bag processing. The pressure of the air inside the core was measured in a lab-scale test fixture to simulate full-scale manufacturing. The lab-scale tests revealed pressure fluctuations if the internal core pressure equalled atmospheric pressure during processing. Furthermore, an increasing number of pressure fluctuations were observed for increasing cure temperatures. Sandwich panels cured at higher temperatures had a higher void content and a worse bond between the skin and core.
In this study, the mechanism of peel fracture (or core/skin delamination) in honeycomb-core sandwich panels was investigated, and the role of resin fillets on energy absorption was elucidated. Fillet deformation and fracture were shown to control the delamination resistance of sandwich panels, to a great extent. In fact the area fraction of fillets on the fracture surface proved to be very important. The delamination resistance data for a variety of panels, made with various materials, cores etc., collapsed to a single master curve when the peel energy was normalized by the delamination energy of the skin material, and this ratio was plotted against the residual area fraction of the resin remaining on the core side after the peel test. A detailed stick/slip mechanism for crack advance was proposed, and its existence was supported by both the load traces and by fractography.
In the present work the compressive properties, collapse modes and crushing characteristics of various types of composite sandwich panels were investigated in a series of edgewise compression tests. The tested sandwich panels were constructed trying four types of polymer foam core (more specifically PMI foam, two grades of linear PVC foam and polyurethane foam) and two types of FRP faceplate laminates made of glass fibre reinforcements impregnated in modified acrylic resin in eight different material combinations. Three modes of collapse were recorded in the compressive tests, one of which being progressive end-crushing of the sandwich panel featured by significant crash energy absorption, feature that was a highly desired since the tested hybrid composites were candidate materials for the manufacture of parts of transportation vehicles. The influence of the most important material properties of the faceplate laminates and foam core and the sandwich construction geometry on the compressive response and the crushing characteristics of the tested sandwich panels such as the peak load, crash energy absorption and collapse modes is extensively examined and analysed. Particular attention is paid on the analysis of the mechanics of progressive deformation and crumpling of the sandwich panels in each of the three collapse modes and especially of the end-crushing mode, emphasizing on the mechanisms related to the crash energy absorption during the edgewise compression of the sandwich panels.
Experimental design has been used to investigate the effect of processing parameters on the peel strength of Nomex honeycomb core/carbon fibre-epoxy prepreg skin sandwich panels, manufactured by vacuum bagging. The four processing parameters considered were: (i) consolidation pressure; (ii) temperature cure cycle; (iii) temperature ramp rate and (iv) vacuum pressure application time. Each parameter was considered at three levels. It was found that (ii) was the dominant parameter. Panels manufactured under the optimum combination of processing parameters gave a mean peel strength more than twice that of the worst panels. Qualitative microscopy has associated the higher peel strengths with larger, regular shaped adhesive fillets between honeycomb cell walls and skin laminate.