Experimental studies of thin-ply laminated composites

COMPOSITES SCIENCE AND TECHNOLOGY Composites Science and Technology 67 (2007) 996–1008 www.elsevier.com/locate/compscitech

Experimental studies of thin-ply laminated composites Sangwook Sihn a

a,*

, Ran Y. Kim a, Kazumasa Kawabe b, Stephen W. Tsai

c

Nonmetallic Division, University of Dayton Research Institute, 300 College Park, Dayton, OH 45469-0168, USA b Industrial Technology Center of Fukui Prefecture, 61, Kawaiwashiduka-cho, Fukui-city, Japan c Think Composites, 101 Alma Street #703, Palo Alto, CA 94301, USA Received 2 February 2006; received in revised form 24 May 2006; accepted 14 June 2006 Available online 24 August 2006

Abstract A new processing method was developed for spreading fiber tows to make thin-ply laminated composites. The present method with a constant airflow through sagged fiber filaments can efficiently spread the thick tows without damaging any fibers. This method is robust and easy compared with other available thin-ply methods. The thin plies of thickness less than one-third of the conventional plies can easily be made with the tow-spreading technology. Experiments were performed to evaluate the performance of tow-spread, thin-ply laminated composites. To study the thickness effect of the laminated composites, the test specimens were made with the same material and the same spread tows, but with dispersed and grouped laminations of the plies. Uniaxial tension static and fatigue loadings were applied on both unnotched and open-hole specimens. Impact and compression-after-impact tests were also conducted. From stress–strain curves, acoustic emission counts, X-ray photos, c-scan images and observation of damage modes of failed specimens, it was observed that the thin-ply laminated composites can suppress the microcracking, delamination and splitting damage for static, fatigue and impact loadings without special resin and/or 3-D reinforcements. Therefore, the laminate design can be simplified by using higher strain allowable without a progressive failure analysis. Ó 2006 Elsevier Ltd. All rights reserved. Keywords: Tow-spreading technology; C. Transverse cracking and delamination; B. Static and fatigue; C. Notch; B. Impact behavior

1. Introduction Fiber-reinforced laminated composites have been used in many structural applications such as airplanes, ships and sporting goods because of their superior specific properties compared with metal materials. The composites structures can be damaged under mechanical and thermal loadings. The typical damage behavior in the laminated composites is transverse microcracking, fiber-breakage and delamination. Typically, the transverse microcracking through the thickness of the ply occurs as the first-ply failure, and the delamination damage follows. The fiber breakage usually happens at the last stage of the failure. However, a catastrophic failure can occur only with the microcracking and delamination damage without the fiber *

Corresponding author. Tel.: +1 937 255 9324; fax: +1 937 258 8075. E-mail address: [email protected] (S. Sihn).

0266-3538/$ - see front matter Ó 2006 Elsevier Ltd. All rights reserved. doi:10.1016/j.compscitech.2006.06.008

breakage. The failure behavior in the laminated composites is usually complicated and highly dependent on the properties of the constituent materials, fiber orientation, stacking sequence, nature of loading, etc. The delamination damage is known to happen because of excessive interlaminar normal and shear stresses at the ply boundaries. The interlaminar stresses can be concentrated relatively easily near the free edge of the laminated composites. Delamination along the straight free edge of the laminated composites under in-plane uniaxial loading has been studied since the early 1970s [1,2]. Since then, a great amount of work has been reported on the free-edge problem in the laminated composites, indicating that free-edge delamination is attributed to the existence of interlaminar stresses which are highly localized in the neighborhood of a free edge under the in-plane loading [3,4]. The nature of the interlaminar stresses with regard to their magnitude and sign can be accurately calculated

S. Sihn et al. / Composites Science and Technology 67 (2007) 996–1008

using the analytical model developed by Pagano and Soni [5,6]. The magnitude and distribution of the interlaminar stress components are widely varying and depend upon the laminate layup, stacking sequence, properties of the constituent materials and nature of loading. Among the many components, a thickness effect on the onset of delamination damage was studied by Kim and Soni [7]. Experiments were performed with the laminated composites having the same stacking sequence and volume fraction of each ply orientation, but a different number of layers. Results were reported for various combinations of ply orientations, stacking sequences and number of repeated layers, including the layups of [±30n/90]s, [±30n/ 90n]s, [0n/±45n/90n]s, [0/±45/90n]s, [0/90n/±45]s, [0n/90n/ ±45n]s, etc. [8], where n is the number of repeated layers. The results show that for all cases, the onset of the delamination stress decreases as n increases, meaning although the stacking sequence and volume fraction of each ply orientation remain unchanged, the delamination threshold decreases with increasing numbers of grouped layers. The thickness effect on the onset of the microcracking damage cannot be adequately explained by the stress analysis alone. Since the stress analysis will determine identical stress magnitudes for different thicknesses of laminates under the same applied axial stress, a criterion based on the stress magnitude would necessarily predict the same onset stress for the microcracking damage. Rodini and Eisenmann [9] used a probabilistic argument that the laminate with thick plies contains statistically more defects than laminates with thin plies. Consequently, the thicker ply is likely to fail at a lower stress level. Wang, et al. [10,11] used an energy release rate approach to describe the fracture growth event in a series of specific laminates and found that the onset stresses for transverse cracking and delamination damage are seen to be inversely proportional to the square-root of the thickness of the 90° layer in [±25/90n]s, n = 1, 2, 3. More experiment studies with a series of graphite-epoxy laminates by Rodini and Eisenmann [9] on the edge delamination in a laminate of [±45n/0n/90n]s, n = 1, 2, 3, and by Crossman et al. [11] on initiation and growth of transverse cracks and edge delamination in a laminate of [±25/90n]s, n = 1, 2, 3 illustrated this thickness effect convincingly in

a

such a way that the occurrence of the transverse microcracks and ply delamination might be suppressed by a decrease in the ply thickness. Consequently, in order to make better laminates resistant to the microcracking and delamination damage, it is desirable to disperse the plies of the same orientations rather than to group the plies, so that the thickness of the plies of the same kind is thin. For example, the quasi-isotropic laminates [(0/±45/90)n]s will be better damage-resistant laminates than [0n/±45n/ 90n]s. It will be even more desirable to use thinner plies whose thickness is a fraction of the conventional plies. There have been a few efforts to reduce the ply thickness below that of the conventional ply of 0.125 mm (5 mil) thickness. However, it required costly and slow processes. Sometimes, the fibers in the ply can be damaged during the process. Recently, a new novel process was developed to make the thin plies by using a tow-spreading technology [12,13]. This new process is low-cost and robust and does not damage the fibers. This paper will begin with the introduction of a new processing method to make the thin-ply laminated composites. This part will explain a novel tow-spreading technology and its principle to make the thin plies. The latter part of the paper will show experimental results to evaluate the performance of tow-spread thin-ply laminated composites. This part will compare thin- vs. thick-ply laminated composites and will show the thickness effect on suppression of microcracking and delamination damage. 2. Tow-spreading methodology There are several known technologies to make a thin ply. Among them, a promising and cost-effective way is to use a tow-spreading technology that was developed by Industrial Technology Center in Fukui Prefecture. The method uses the conventional thick tow such as 12 K filament tow (see Fig. 1). The tow passes though a spreading machine that is equipped with an air duct and a vacuum that sucks the air downward through the air duct. The air duct is located between these guide rolls. As the air flows through the air duct with the help of the vacuum, the tow sags downward toward the air flow direction so that it loses tension and results in a tension-free state

b Guide Roll

Original Tow (Thick) CF 12K

997

w Spread Tow

Spread Tow

Air Duct

(Thin) Original Tow

Spread Width 16-32mm

Fig. 1. Schematic of tow-spreading method with a pneumatic method.

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S. Sihn et al. / Composites Science and Technology 67 (2007) 996–1008

momentarily. With the uniform airflow continuously operating on the tow, the tow can be spread continuously and stably. Furthermore, the airflow does not usually cause any damage to the fiber filaments during this processing because the airflow velocity is fairly low. Therefore, with this present method, the tow can easily be spread without damaging the filament fibers. The thin-ply laminate can be made with the spread tows. Fig. 2 is a schematic that shows how the present towspreading processing method works with the help of airflow. When the air flows around both sides of the tow, the difference in the velocity of the airflow, near and away from the tow, results in the difference in the pressure at these locations. As shown in the leftmost figure, the pressure near the tow becomes greater than that away from the tow. The pressure difference creates an aerodynamic force that helps the filament fibers to lose the tension momentarily. As the tow-spreading begins, the air flows between the filaments of the tow to help the tow to spread more and faster, as shown in the middle and rightmost figures. Fig. 3 shows samples of the original unspread 12 K tow and the two different widths of the spread tows. The wider the tow is spread, the thinner the tow thickness becomes.

Initial state

After spreading

Air flow Filament

Fig. 2. Schematic of tow-spreading process with the help of airflow.

Fig. 3. CF12K tow and its spread tow: (a) original tow (width: 6 mm), (b) spread tow (width: 12 mm) and (c) spread tow (width: 20 mm).

3. Experimental results Experiments were conducted to evaluate the performance of tow-spread thin-ply laminated composites. The composites test specimens were made with Toray’s carbon fibers (T800SC-24K) and Bryte’s epoxy film resin (BT250E-1) as the fiber and matrix material in the composites, respectively. The pristine tow was 6 mm wide and 0.14 mm thick. This fiber tow was spread with the present tow-spreading method and wound to a spool. The dimensions of the spread tow were 20 mm in width and 0.04 mm (1.58 mil) in thickness. An areal weight is 40 g/m2. The spread tows in eight spools were wound off and placed side by side to make a dry fiber sheet of 320 mm width. The spread tow sheet was then placed on the BT250E-1 film resin sheet. By the tackiness of epoxy resin film, the thin sheet adhered to the epoxy resin film. At this stage, epoxy resin did not impregnate the fiber tow yet. The thin sheet adhered to the epoxy resin film was converted to a prepreg sheet by an in-house prepregging device developed by Industrial Technology Center in Fukui Prefecture. The device is equipped with a sheet press system and a sheet transfer system. The prepreg sheet was produced by repeating the press and transfer with the following processing conditions: press pressure = 1 MPa, press time = 3 s, press temperature = 90 °C, sheet transfer length in one cycle = 90 mm for 2 s, processing speed = about 1 m/min. The prepreg sheets were laid up on a tool with desired layup orientations. The laminates were cured in an autoclave at a temperature of 125 °C and a pressure of 0.5 MPa for 1 h [14]. The fiber volume fraction was approximately 60%. The thickness of the spread thin ply is 0.04 mm (1.58 mil), which is less than one third of the conventional ply of 0.125 mm (5 mil) thick. To study the thickness effect of the laminated composites, test specimens were made with the same material (T800SC-24K/BT250E-1) and the same spread tows, but with two different laminations, which are designated as THIN and THICK laminations. Both THIN and THICK specimens were made with the spread thin plies of 0.04 mm thickness. In the THIN lamination, the thin plies of the same fiber orientation were dispersed in the thickness direction as shown in Fig. 4a, whereas in the THICK lamination, five thin plies of the same fiber orientations were grouped together to form a single thick ply, as shown in Fig. 4b. Therefore, the thickness of the single thick ply in the THICK specimen is 0.20 mm, which is five times thicker than that of the single thin ply in the THIN specimen. Note that the total thicknesses of THIN and THICK specimens are the same since the same numbers of thin plies were used. Uniaxial tension static and fatigue loadings were applied on both unnotched and open-hole specimens. We recorded stress–strain curves and acoustic emission (AE) counts and took X-ray photos to observe damage modes of failed specimens.

S. Sihn et al. / Composites Science and Technology 67 (2007) 996–1008

999

Fig. 4. Micrographs of test specimens made by tow-spread thin plies with (a) distributed lamination (THIN) and (b) grouped lamination (THICK). The thickness of a grouped THICK ply is five times thicker than that of a THIN one.

Table 1 Longitudinal Young’s modulus (EL), transverse Young’s modulus (ET), longitudinal tensile strength (X), transverse tensile strength (Y) and Poisson’s ratio (m) of unnotched unidirectional laminates

3.1. Unidirectional laminates Before comparing the THIN and THICK laminates, we measured the basic laminate properties such as longitudinal Young’s modulus (EL), transverse Young’s modulus (ET), longitudinal tensile strength (X), transverse tensile strength (Y) and Poisson’s ratio (m). We used unidirectional [0]48 and [90]48 laminates for measuring longitudinal and transverse properties, respectively. Fig. 5 shows the stress–strain curves of longitudinal and transverse unidirectional laminates under static loading, and Table 1 lists these properties of the unnotched unidirectional laminates. Based on the property values and the standard deviation, it can be considered that the laminate specimens fabricated from the tow-spread thin plies were of good quality.

SD (%) EL (initial) EL (Overall) ET X Y m

168 GPa (24.3 msi) 185 GPa (26.8 msi) 8.40 GPa (1.22 msi) 2740 MPa (397 ksi) 66.0 MPa (9.62 ksi) 0.33

50.8 mm (2")

178 mm (7")

3.2. Unnotched tension

a

Fig. 6. Test specimen under unnotched tension loading.

strains were measured in the central gauge location. Three duplicate specimens were tested for each THIN and THICK laminate. Fig. 7 shows the stress–strain curves under the UNT loading. Strain measured in the middle of the specimen in longitudinal and transverse directions are shown in the figure. The dotted and solid lines are for the THIN and THICK specimens, respectively. The initial slopes of

b

3000

EL [MPa]

1500

ET [MPa] 8.4 8.4 8.4 8.3 8.3 8.1

60

Stress [MPa]

Stress [MPa]

2000

80 70

170 172 166 164 170 170 161

2500

50.8 mm (2")

w=25.4 mm (1")

strain gauges

The first laminate test considered was the unnotched tension (UNT). Fig. 6 shows the configuration of the UNT specimen. The width of the specimen is 25.4 mm (1.000 ). We tested the THIN and THICK specimens of a quasi-isotropic (QI) laminate (25/50/25), where three numbers in parentheses represent the percentage of the [0], [±45] and [90] plies, respectively. Strain gages were attached in the middle of the specimen to measure the central far-field strain away from the free edge, and 1.27 mm (0.0500 ) away from the free edge to measure the edge strain. Both longitudinal and transverse

2.3 2.3 2.2 4.8 10.4 2.7

1000

50 40 30 20

500

10

0

0 0

0.5

1

Strain [%]

1.5

2

0

0.2

0.4

0.6

0.8

1

1.2

Strain [%]

Fig. 5. Stress–strain curves of (a) longitudinal ([0]48) and (b) transverse ([90]48) unidirectional laminates under static loading.

1000

S. Sihn et al. / Composites Science and Technology 67 (2007) 996–1008 Table 2 Young’s modulus (Ex), Poisson’s ratio (m) and longitudinal tensile strength (X) of unnotched QI THIN and THICK specimens

1200

THIN (tply=0.04 mm)

Transverse strains

1000

Longitudinal Longitudinal strains strains

Stress [MPa]

800 600

THICK (5*tply)

400

Dotted lines: THIN

200

Solid lines: THICK 0

-1

-0.5

0

0.5

1

1.5

2

Strain [%]

Fig. 7. Stress–strain curves of QI THIN and THICK laminates under UNT static loading. Strains were measured in the middle of specimen.

stress–strain curves in both longitudinal and transverse directions are almost identical for the THIN and THICK specimens. While the stress–strain curves of the THIN specimens retain linearity nearly up to the ultimate failure stress, the curves of the THICK specimens show a sudden increase of strains near 690 MPa (100 ksi). The sudden big increase of the strain indicates unrecoverable severe damage due to microcracks and delamination in the THICK specimen. The averages and standard deviations of longitudinal Young’s modulus, Poisson’s ratio and longitudinal tensile strength are listed in Table 2. The ultimate strength of the THIN specimens is approximately 10% higher than that of the THICK ones. We also tested the UNT specimens under cyclic loading. The fatigue condition was tension–tension fatigue with 2 Hz of frequency and 0.1 of stress ratio. The maximum stress level was 483 MPa (76 ksi), which is approximately 60% of the ultimate strength of the QI THIN UNT specimens. We monitored the increase of strains during the fatigue loading. Firstly, a static loading was applied before the fatigue loading to measure the initial stress–strain curve (0 cycle). Then the fatigue loading was stopped after 1000, 10,000 and 50,000 cycles to measure the strains near the free edge. Fig. 8 shows the stress–strain behaviors of the THIN and THICK specimens. The edge strain near the free edge increased as the fatigue cycles increased.

X

SD

m

2.1%

0.94 GPa (136 ksi)

5.1%

0.38

THICK 58.7 GPa (8.5 msi)

1.2%

0.85 GPa (123 ksi)

4.8%

0.38

The THIN specimen shows little increase in the free-edge strain after 50,000 cycles, while the THICK specimen shows a significant increase only after 1,000 cycles. Fig. 9 shows the X-ray images for the QI THIN and THICK specimens taken after 50,000 cycles of the fatigue loadings. The THICK specimen shows many microcracks and delamination damage starting from the free edge at the sides of the specimen because of high interlaminar stresses, while the THIN specimen shows little or no visible damage. After 50,000 fatigue cycles at a stress level of 483 MPa (76 ksi), the static loading was applied to measure the residual stiffness and strength left in the fatigued specimens. Fig. 10 shows the residual stress–strain curves under the

200

0 cycle 1,000 cycles cycles

100

(tply=0.04 mm)

b

600

THICK (5*tply=0.20 mm)

500

300

THIN THIN

(5*tply=0.20 mm)

Fig. 9. X-ray images of UNT QI: (a) THICK and (b) THIN specimens after 50,000 cycles of fatigue loading at 60% stress level of ultimate static strength.

THIN (tply=0.04 mm)

400

THICK THICK

a

b

600 500

Stress [MPa]

SD

Stress [MPa]

a

Ex THIN 61.2 GPa (8.9 msi)

400 300 200

0 cycle 1,000 cycles cycles 10,000cycles cycles 50,000cycles cycles

100

10,000 cycles cycles 50,000cycles cycles

0

0 0

0.2

0.4

0.6

Strain [%]

0.8

1

1.2

0

0.2

0.4

0.6

0.8

1

1.2

Strain [%]

Fig. 8. Stress–strain curves of UNT QI: (a) THIN and (b) THICK laminates after fatigue loading. Strains were measured in the middle of specimen.

S. Sihn et al. / Composites Science and Technology 67 (2007) 996–1008

1001

1200

THIN (tply=0.04 mm)

Stress [MPa]

1000 800

THICK THICK(5*t (5*tplyply)

600 400

THIN - Static THICK -- Static THICK Static THIN - Residual Residual THICK -- Residual THICK Residual

200 0 0

0.5

1

1.5

2

Strain [%] Fig. 10. Residual stress–strain curves of UNT QI THIN and THICK laminates for static loading and after 50,000 cycles of fatigue loadings at a stress level of 483 MPa (76.0 ksi).

UNT loading. The original static stress–strain curves taken from Fig. 7 are also shown in this figure for comparison. The static and residual curves for the THIN specimens are nearly identical. However, the residual curve of the THICK specimen shows degradation of both stiffness and strength properties after fatigue loading. Table 3 lists averages and standard deviations of Young’s modulus and tensile strength under static and residual loadings. After 50,000 cycles, the averages of Young’s moduli of THIN and THICK specimens decreased by 2.8 and 17.7%, respectively. The average strength of THIN specimens increased slightly by 4.4% while that of THICK specimens decreased by 16.8%. Therefore, it can be concluded that the THIN specimen is more durable than the THICK one under the UNT fatigue loading. 3.3. Open-hole tension The next test considered was the open-hole tension (OHT) (see Fig. 11). The width of the specimen is 38.1 mm (1.500 ), and the diameter of the hole is 6.35 mm (0.2500 ); thus the ratio of the width to diameter of the hole is 6. We tested the THIN and THICK specimens of (1) the quasi-isotropic (QI) laminate (25/50/25) and (2) the hard laminate (50/40/ 10), where three numbers in parentheses represent the percentage of the [0], [±45] and [90] plies, respectively. 3.3.1. QI laminate The layup sequence of the QI laminates is [45/0/45/ 90]ns, where subscripts n and s represent the number of

50.8 mm (2")

178 mm (7")

50.8 mm (2")

D=6.35 mm (1/4")

strain gauges

w=38.1 mm (1.5")

Fig. 11. Test specimen under open-hole tension loading.

repeats and the symmetric layup, respectively. The values of n for the QI THIN and THICK specimens are 10 and 2, respectively. Therefore, the grouped ply thickness of the THICK specimen is 0.2 mm (7.9 mil), which is five times thicker than that of the THIN one. The total number of tow-spread thin plies used for the QI laminates is 80. The total thickness of the QI specimen is 3.2 mm (0.12600 ), which is same for both THIN and THICK specimens. Two strain gages were attached. One was attached in the middle of the specimen between the hole and a grip to measure the far-field strain away from the hole edge and the free edge, and the other was attached 1.27 mm (0.0500 ) away from the hole edge to measure the edge strain. Two duplicate specimens were tested for each THIN and THICK laminate. Fig. 12 shows the stress–strain curves under the OHT loading. Strain measurement by both the far-field and hole-edge strain gages are shown in the figure. The dotted and solid lines are for the THIN and THICK specimens, respectively. The initial slopes of the far-field stress–strain curves are almost identical for the THIN and THICK

Table 3 Young’s modulus (Ex) and tensile strength (X) of unnotched QI THIN and THICK specimens under static and residual loading Static

SD

Residual

Degradation

Ex THIN THICK

61.2 GPa (8.9 msi) 58.7 GPa (8.5 msi)

2.1% 1.2%

59.5 GPa (8.6 msi) 48.3 GPa (7.0 msi)

2.8% 17.7%

X THIN THICK

0.94 GPa (136 ksi) 0.85 GPa (123 ksi)

5.1% 4.8%

0.98 GPa (142 ksi) 0.70 GPa (102 ksi)

4.4% 16.8%

1002

S. Sihn et al. / Composites Science and Technology 67 (2007) 996–1008 600

Far-field Far-field

500

Stress [MPa]

THIN (tply)

400

THICK (5*tply)

300

Edge strain Edge strain

200

Thin - Edge Thin - Farfield Farfield Thick - Edge Edge Thick - Farfield Farfield

100 0 0

0.5

1

1.5

2

2.5

Strain [%] Fig. 12. Stress–strain curves of QI THIN and THICK laminates after OHT static loading.

specimens. However, the hole-edge stress–strain curves of the THIN are more linear than those of the THICK. The averages of the ultimate OHT strength of the QI THIN and THICK specimens are 492 MPa (71.3 ksi) and 547 MPa (79.4 ksi), respectively. Unlike the UNT specimen without the hole, the ultimate strength of the THIN specimens is approximately 10% lower than that of the THICK

b

600

500

500

400

400

Counts

Counts

a 600

one. One of the possible reasons for the higher ultimate strength with the THICK laminate is due to the stress relaxation near the hole edge after the initial failure. Because of the stress concentration at the edge of the hole, the initial failure tends to occur near the hole. This initial failure relaxes the stress near the hole, which leads to lower the stress concentration. The THIN laminate suppresses the initial damage, so that the stresses are concentrated until catastrophic failure occurs. Fig. 13 shows AE counts of the QI THIN and THICK specimens monitored during the OHT loading. The AE counts of the THIN appear to be less in count number and amplitude than those of the THICK. The THIN case shows almost no count at low load and big counts near the ultimate load. The THICK case shows many big counts throughout the loadings up to the ultimate failure. For one test specimen, the loading was increased up to 57.8 kN (13,000 lbf), which corresponds to a far-field stress of 448 MPa (65 ksi), the loading was stopped, and an X-ray was taken to see the damage up to this loading. Fig. 14 shows the X-ray images of the THIN and THICK specimens. Because of the stress concentration at the hole edge as well as at the free edges on the sides of the specimen, one

300

300

200

200

100

100

0 0

10000

20000

30000

40000

50000

60000

70000

Load[N]

0 0

10000

20000

30000

40000

50000

60000

Load[N]

Fig. 13. Acoustic emission counts of QI (a) THIN and (b) THICK specimens after OHT static loading.

THICK (ply thickness: 5*tply=0.20 mm) Total thickness = 3.2 mm (0.126").

Near hole

THIN (ply thickness: tply=0.04 mm)

Fig. 14. X-ray images of QI (a) THICK and (b) THIN specimens after OHT static loading at 448 MPa (65 ksi).

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S. Sihn et al. / Composites Science and Technology 67 (2007) 996–1008

can expect the initial failure such as microcracking to occur near these edges. The THICK specimen shows many microcracks near the hole edge as well as the outer free edges than the THIN one does. A closer view near the hole edge on the right-hand side of the figure shows the microcracks in the [0], [90] and [±45] directions, which is very typical in QI laminated composites. However, the THIN specimen shows no visible damage at the hole edge and the free edges. Therefore, the difference of the ply thickness results in this phenomenal difference in the damage behavior. Fig. 15 shows failed THIN and THICK specimens. The THICK specimen shows a pull-out failure mode with the failed sections at 45° and 45° angles as well as many delaminations from the side view. Meanwhile, the THIN specimen shows a brittle type of net-section failure mode whose failed section is perpendicular to the applied tension direction. The side view of the THIN specimen shows little delamination. We also tested the OHT specimens under cyclic loading. The fatigue condition was tension–tension fatigue with 10 Hz frequency and 0.1 stress ratio. The maximum stress

1003

level was 354 MPa (51.3 ksi), which is approximately 70% of the ultimate strength of the QI OHT THIN specimens. The increase of strains was monitored during the fatigue loading. Firstly, a static test was conducted before the fatigue test to measure the initial stress–strain curve. Then the fatigue loading was stopped after 1, 20,000 and 100,000 cycles to measure both far-field and hole-edge strains. Fig. 16 shows the stress–strain behaviors of the THIN and THICK specimens. Similar to the static case, the change of far-field strain in both THIN and THICK specimens was negligible during the fatigue loading. However, the edge strain near the open hole increased as the fatigue cycles increased. The THIN specimen shows little increase in the hole-edge strain after 100,000 cycles, while the THICK specimen shows a significant increase (nearly 100% increase) after only 20,000 cycles. Photos and X-ray images were taken of the fatigue specimens after 100,000 cycles of fatigue loading at 70% stress level of the ultimate static strength (see Fig. 17). Side and front views were taken with the digital camera and X-ray, respectively. The fine line blurry areas in white color in the X-ray image indicate microcracking and delamina-

a THICK

After ultimate failure

Near hole

Side view

b THIN

Fig. 15. QI (a) THICK and (b) THIN specimens after ultimate failure after OHT static loading.

a

b

400

400

Far-field

Edge strain

350

350

Far-field 300

Stress [MPa]

Stress [MPa]

300 250

Edge strain

200 150

0 cycle 1 cycle 20k cycles cycles 100k cycles cycles

100 50

250 200 150 100

0 cycle 1 cycle 20k cycles cycles

50

0

0

0

0.2

0.4

0.6

Strain [%]

0.8

1

1.2

0

0.5

1

1.5

2

Strain [%]

Fig. 16. Stress–strain curves of QI (a) THIN and (b) THICK laminates after OHT fatigue loading.

2.5

3

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S. Sihn et al. / Composites Science and Technology 67 (2007) 996–1008

a

THIN THIN

b

THICK THICK (5*tply=0.20 =0.20 mm) (5*t mm)

=0.04 mm) (t(tply=0.04 mm)

Side view

Front view

Side view

Front view

Fig. 17. X-ray images of QI (a) THIN and (b) THICK specimens after 100,000 cycles of OHT fatigue loading at 70% stress level of ultimate static strength.

tion damage, respectively. The difference in the severity of the damage in the THIN and the THICK specimens can easily be seen from the figure.

Edge strain 1000

Far-field strain

Stress [MPa]

3.3.2. Hard laminate The hard laminate is made with 50% of [0], 40% of [±45] and 10% of [90] degree plies. This layup is often used in axial-loading dominant load-carrying structures such as wings in an airplane. The layup sequence of the hard laminates is [45/02/45/90/45/02/45/0]ns. The values of n for the hard THIN and THICK specimens are 5 and 1, respectively. Therefore, the thickness of grouped plies in the THICK specimen is 0.2 mm (7.9 mil), which is five times thicker than that of the THIN one. Total numbers of tow-spread thin plies used for the hard laminates are 100. The total thickness of the hard specimen is 4.0 mm (0.15700 ). Far-field and hole-edge strain gages were attached in the hard laminates. Two duplicate specimens were tested for each THIN and THICK laminate. Fig. 18 shows the stress–strain curves under the OHT loading. Strain measurement by both the far-field and hole-edge strain gages are shown in the figure. The initial slopes of the far-field stress–strain curves are almost identical for the THIN and THICK specimens. The hole-edge stress–strain curves of the THIN specimens are more linear than those of the THICK ones until the ultimate failure stress. The averages of the ultimate OHT strength of the hard THIN and THICK specimens are 720 MPa (104 ksi) and 967 MPa (140 ksi), respectively. Comparing with the QI OHT case, the ultimate strengths of the hard THIN specimens are even much lower than those of the THICK ones. It is approximately 34% lower in this case. The reason why the OHT ultimate strength of the THIN QI specimens falls below that of the THICK ones can also be applied to this hard laminate case. Because of the even higher stress concentration at the hole edge, the difference becomes more pronounced in this hard laminate case than the QI case.

1200

800 600 400 Thin - Edge Thin - Farfield Farfield Thick - Edge Edge Thick - Farfield Farfield

200 0 0

0.5

1

1.5

2

Strain [%] Fig. 18. Stress–strain curves of hard THIN and THICK laminates under OHT static loading.

Fig. 19 shows failed hard THIN and THICK specimens. Similar to the QI laminates, the THIN specimen shows the brittle type of net-section failure mode whose failed section is perpendicular to the applied tension direction. Meanwhile, the THICK specimen shows a pull-out failure mode with the failed sections at 45° and 45° angles. The side views show that the THICK specimen has much more delamination damage than the THIN specimen. The AE counts of the hard THIN and THICK specimens were monitored during the OHT loading. The AE counts were cumulated with the load increase and plotted in Fig. 20. The figure clearly shows that the THIN specimens have much less damage events than the THICK ones throughout the load increase up to the ultimate failure. We also tested the hard OHT specimens under cyclic loading. The fatigue condition was tension–tension fatigue at 5 Hz frequency and 0.1 stress ratio. The maximum stress level was 483 MPa (70 ksi), which is approximately 70% of the ultimate strength of the hard OHT THIN specimens. Fig. 21 shows X-ray images for the hard THIN and

S. Sihn et al. / Composites Science and Technology 67 (2007) 996–1008

1005

a THIN THIN

After ultimate failure

b

Side view

THICK THICK

Fig. 19. Hard (a) THIN and (b) THICK specimens after ultimate failure under OHT static loading.

microcracking and delamination damage. As shown in the figure, a big difference can be observed in the amount of the damage after 73,000 fatigue cycles in this layup.

10000

Cumulative Counts

THICK 8000

6000

3.4. Impact

4000

If the thin-ply laminates suppress the onset of the microcracking and delamination damage, one can expect that the similar superiority can be observed under impact loadings. We conducted impact testing of the THIN and THICK specimens and compared c-scan images that show the size and depth of the delamination damage due to the impact loading. We also performed compression-after-impact (CAI) tests to study the severity of the delamination damage. Panel specimens were prepared for the CAI tests with width and length dimensions of 102 mm (400 ) by 152 mm (600 ), respectively. The panels were clamped at 600 sides in the length direction and free at 400 sides in the width direction [15].

2000

THIN 0 0

30000

60000

90000

120000

150000

Load [N]

Fig. 20. Cumulative acoustic emission counts of hard THIN and THICK specimens after OHT static loading.

THICK specimens taken after 73,000 cycles of the OHT fatigue loadings. Because of the high percentage of the 0° ply, split damage behavior is expected in addition to the

a

THI THIN =0.04 mm) (t(tply mm) ply=0.04

b

THICK THICK (5*tply=0.20 =0.20 mm) (5*t mm)

Fig. 21. X-ray images of hard (a) THIN and (b) THICK specimens after 73,000 cycles of OHT fatigue loading at 70% stress level of ultimate static strength.

S. Sihn et al. / Composites Science and Technology 67 (2007) 996–1008

Fig. 22 shows in-depth c-scan images after the impact loading. Each round shape in a different color indicates the delamination damage in each layer through the thickness of the panel. It appears that the overall sizes of the delamination damage are similar for THIN and THICK specimens. The impacted specimens were tested in compression using the CAI test. Fig. 23 shows the test configuration for the CAI test. Two strain gages were attached in the middle of the panel on both sides of the surfaces to monitor the instability under the compression loading. Fig. 24 shows the stress–strain curves of the THIN and THICK CAI specimens. We tested two duplicates of each kind. The bifurcations of the curves toward left and right sides indicate the two strain gage measurements in front and back faces of the panel. Initially, both strain gages on the panel were under compression. As the load increased, the panel buckled to cause the bifurcation of the strain measurement. The figure clearly shows that the THICK specimens bifurcate earlier than the THIN ones. The early bifurcation at the lower stress level in the THICK specimens indicates more severe delamination damage than the THIN ones. This result is preliminary and more study is needed for better understanding of the ply thickness effect under impact loading. 4. Summary A new processing method was developed for spreading fiber tows to make thin-ply laminated composites. The present method with a constant airflow through sagged fiber filaments can efficiently spread the thick tows without damaging any fibers. This method is robust and easy compared with other available thin-ply methods. The thin plies of thickness less than one-third of the conventional plies can easily be made with the tow-spreading technology.

Fig. 23. Test setup for compression after impact test.

1600

THIN 1400 1200 1000

Stress [MPa]

1006

800

THICK

600 400

Thin #1 Thin #1 Thin #2 Thin #2

200

Thick #1 Thick #1 Thick #2 Thick #2

0 -1

-0.8

-0.6

-0.4

-0.2

0

0.2

0.4

Strain [%] Fig. 24. Stress–strain curves of QI THIN and THICK laminates after CAI loading.

Fig. 22. In-depth c-scan images of QI (a) THIN and (b) THICK specimens after impact loading.

S. Sihn et al. / Composites Science and Technology 67 (2007) 996–1008

We measured basic elastic and strength properties of the thin-ply composites. Then the thickness effect was studied by comparing THIN and THICK specimens of QI and hard laminates with and without the hole defect under static, fatigue and impact loadings. Experimental observations from stress–strain curves, AE counts, X-ray and c-scan images and observation of damage modes in failed specimens indicate suppression and/or delay of microcracking, delamination and splitting damage with the thin-ply laminated composites with and without the hole defect. Under UNT static loading, the stress–strain curves of the THIN specimens retain linearity nearly up to the ultimate failure stress, while the THICK specimens showed a big increase in the strain, which indicates unrecoverable severe damage due to microcracks and delamination in the THICK specimen. The ultimate strength of the THIN specimens is approximately 10% higher than that of the THICK ones. Under UNT fatigue loading, THIN and THICK specimens show phenomenal difference in both stress–strain curves and X-ray images. The residual stiffness and strength of the THIN laminates degrade negligibly after 50,000 fatigue cycles at a 60% stress level of the ultimate static strength compared with the THICK laminates. With the open-hole defect, we observed a larger increase of the strain and more AE counts near the hole edge with the THICK specimens than with the THIN ones. The X-ray images and failed specimens also show a phenomenal difference between THIN and THICK specimens with both QI and hard laminates. However, the ultimate strength and strain-to-failure of the OHT THICK specimens are higher than that of the THIN ones with both QI and hard laminates due to a possible reason of the stress relaxation near the hole edge after the initial failure, which leads to a lower stress concentration in the THICK specimen. Under fatigue loading, the change of hole-edge strain in the THIN specimens was negligible after 100,000 cycles of fatigue loading, whereas that in the THICK specimen significantly increased after 20,000 cycles. Finally, the in-depth c-scan images after the impact loading indicate that the overall sizes of the delamination damage are similar for THIN and THICK specimens. However, the stress–strain behavior after the CAI tests shows a delay of the bifurcations of the curves with the THIN specimens indicating less severe delamination damage in the THIN laminates than in the THICK ones. 5. Concluding remarks We observed from many experimental results that the thin-ply laminated composites can suppress the microcracking and delamination damage without special resin and/or 3-D reinforcements. Therefore, the laminate design can be simplified by using higher strain allowable without the need for a progressive failure analysis. The thinner ply thickness will also provide more choice in optimizing the laminate composites structures. The ply drop will be smoother than in conventional thick-ply laminates.

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Thin plies can further be used in a hybrid lamination. The conventional thick plies can still be used in the fiberdominant mode such as in the [0] ply, where the dominant failure mode is fiber breakage. The hybrid usage of thin and thick plies has the potential for highly damageresistant composites structures without much increasing layup costs. The present thin spread tow can also be used in the textile composites. Thin tows in textile composites such as woven and braided composites result in less undulation of the woven and braided tows than the conventional thick tows. Less undulation means higher modulus and less resin-rich area between the tow. Therefore the thin-ply textile composites can easily be expected to improve mechanical performance with the superior suppression capability of the microcracking in the resin-rich matrix and delamination damage between the yarns. Finally, the present method also inspires a cost-effective manufacturing process of the composites tows. The automated gentle spreading process of the large tows, such as 100 K and 200 K tows, can be cost-competitive in making smaller tows such as 12 K tows rather than dealing with the smaller tows themselves. Acknowledgements This work was sponsored by the Itochu Corporation and Mitsuya Corporation through Think Composites. The authors thank Yasushi Kiyobayashi from Itochu Corporation, Kichiro Ishida from Mitsuya Corporation and Shigeru Tomoda from Industrial Technology Center of Fukui Prefecture in Japan for their valuable support. The authors also thank Ron Esterline of the University of Dayton Research Institute for sample preparation and testing. References [1] Pipes RB, Pagano NJ. Interlaminar stresses in composite laminates under uniform axial extension. J Compos Mater 1970;4:538–48. [2] Pagano NJ, Pipes RB. The influence of stacking sequence on laminate strength. J Compos Mater 1971;5:50–7. [3] Pagano NJ, Pipes RB. Some observations on the interlaminar strength of composite laminates. Int J Mech Sci 1973;15:679–88. [4] Pagano NJ. Stress field in composite laminates. Int J Solids Struct 1978;14:385–400. [5] Pagano NJ, Soni SR. Global-local laminate variational model. Int J Solids Struct 1983;19(3):207. [6] Pagano NJ, Soni SR. Models for studying free-edge effects. Interlaminar Response of Composite Materials 1989 [chapter 1]. [7] Kim RY, Soni SR. Experimental and analytical studies on the onset of delamination in laminated composites. J Compos Mater 1984;18. [8] Kim RY. Experimental observations of free-edge delamination. Interlaminar Response of Composite Materials 1989 [chapter 3]. [9] Rodini Jr BT, and Eisenmann JR. An analytical and experimental investigation of edge delamination in composite laminates. In: Proceedings of the 4th conference fibrious composites, San Diego, CA; November 1978. [10] Wang ASD, Crossman FW. Initiation and growth of transverse cracks and edge delamination in composite laminates: Part 1. An energy method. J Compos Mater 1980;14(Suppl.): 71–87.

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[11] Crossman FW, Warren WJ, Wang ASD, Law Jr GE. Initiation and growth of transverse cracks and edge delamination in composite laminates: Part 2. Experimental correlation. J Compos Mater 1980;14(Suppl.):88–108. [12] Kawabe K, Tomoda S, Matsuo T. A pneumatic process for spreading reinforcing fiber tow. In: The 42nd international SAMPE symposium & exhibition, Anaheim, CA; May 4–8, 1997.

[13] Sasayama H, Kawabe K, Tomoda S, Ohsawa I, Kageyama K, Ogata N. Effect of lamina thickness on first ply failure in multidirectionally laminated composites. In: The 8th Japan international SAMPE symposium & exhibition (JISSE-8), Tokyo, Japan; November 2003. p. 18–21. [14] BEPCC. BT250E-1 Epoxy Prepreg Cure Cycle; 1999. Available from: http://www.brytetech.com/webresources/ pdf/bt250e1cc.pdf. [15] BSS7260. Boeing specification; 1988.