(NASA) 33 p. CSCL 11D. Unclas. G3/24 0140764. FEBRUARY 1988. National Aeronautics and. Space Administration langley Research Center. Hampton.
NASA Technical Memorandum 100532 USAAVSCOM TECHNICAL MEMORANDUM 88-B-002
Jm A, HINKLEY,
Nm Jm JOHNSTON AND T I Km O'BRIEN
INTERLAMINAR FRACTURE TOUGHNESS OF THERNOPLASTIC COMPOSITES
(HASA-TF!- 100532) IUT EBLAMI WAR PR ACTUR E TOUGHNESS OF THERNOPLBSTIC COHPOSITES (NASA) 33 p CSCL 11D
N 88- 2 20 97
G3/24
Unclas 0140764
FEBRUARY 1988 Q
. National Aeronautics and Space Administration
US ARMY
langley Research Center
AVIATION SYSTEMS COMMAND
Hampton.Virginia 23665
AVIATION R I T ACTIVITY
1
INTRODUCTION
Thermoplastics are c u r r e n t l y being considered f o r use as m a t r i x r e s i n s i n h i g h performance composites.
Besides the potenti-al advantages o f
u n l i m i t e d s h e l f l i f e and rapid, inexpensive processing, a major reason f o r the i n t e r e s t i n these m a t e r i a l s i s t h e i r much g r e a t e r f r a c t u r e toughness re1 a t i ve t o t y p i c a l 350'F-cure
epoxy res1ns.
Improvements in r e s f n
toughness are desi r a b l e as one approach t o improvi ng del ami n a t i o n r e s i stance and damage to1 erance i n composite s t r u c t u r e s . Extensive t e s t i n g using t h e double c a n t i l e v e r beam (DCB) t e s t has shown a reasonable c o r r e l a t i o n between r e s i n toughness, as measured by compact t e n s i o n t e s t s on neat r e s i n castings, and composite i n t e r l a m i n a r f r a c t u r e toughness [ l ] . However, two c o m p l i c a t i o n s a r i s e i n t h e DCB t e s t i n g f o r i n t e r l a m i n a r f r a c t u r e toughness:
GIC
values as measured by t h e DCB
t e s t may be a f f e c t e d by f i b e r b r i d g i n g and t h e adhesion between carbon f i b e r s and the thermoplastic matrix may be poor. The present work examines the i n t e r l a m i n a r f a i l u r e process i n f o u r experimental composite m a t e r i a l s usi ng both the double c a n t i 1ever beam (DCB) and t h e edge delamination tension (EDT) t e s t s .
The EDT t e s t was chosen t o
g i v e an a l t e r n a t e measure o f i n t e r l a m i n a r f r a c t u r e toughness t h a t would n o t be i n f l u e n c e d by f i b e r b r i d g i n g , and would q u a n t i f y t h e i n f l u e n c e o f r e s i d u a l thermal stresses on delamination onset.
One s e t o f experiments was
conducted on composites w i t h a model thermoplastic r e s i n m a t r i x t o i l l u s t r a t e the e f f e c t o f v a r y i n g f i b e r surface p r o p e r t i e s on d e l a m i n a t i o n toughness.
.
The o t h e r s e t o f experiments was conducted on composites with
novel high-temperature thermoplastic polyimides.
The 1a t t e r t e s t s
i l l u s t r a t e the s i g n i f i c a n t e f f e c t o f thermal stresses on delamination onset s t r a i n i n composite laminates made a t h i g h c o n s o l i d a t i o n temperatures.
2
M a t e r i a1 s and Process4 ng The f i b e r s , AS4 (Hercules, I n c . ) and XAS (Hysol G r a f i l ) * were s u p p l i e d unsized, b u t w i t h t h e manufacturer's p r o p r i e t a r y s u r f a c e treatment. Previous work had shown t h a t these two f i b e r s behave very d i f f e r e n t l y toward polycarbonate i n s i ngl e - f i b e r adhesion t e s t s [ 2
1.
Polycarbonate r e s i n
(Lexan 101 from General E l e c t r i c Co.) was s e l e c t e d as a model tough t h e r m o p l a s t i c , l a r g e l y because i t s behavior i n n e a t r e s i n form i s well-characterized.
Prepreg was made by drum-winding where t h e r e s i n was
a p p l i e d as a 17% w/w s o l u t i o n i n a 50:50 m i x t u r e o f chloroform: chloride.
methylene
Before l a m i n a t i n g , prepreg was d r i e d f o r one hour a t 400°F i n a
forced-ai r oven.
Thermogravimetric a n a l y s i s o f t h e prepreg a f t e r d r y i n g
showed l e s s than 1%weight l o s s i n h e a t i n g t o 570°F. Laminating was performed i n a matched metal mold i n a heated press a t 200 p s i .
A 15-minute h o l d a t 50OOF was f o l l o w e d by a 2-hour h o l d a t 473OF.
Mu1t i d i r e c t i o n a l laminates showed small areas o f v i s i b l e f i b e r waviness i n t h e surface p l i e s a f t e r consol i d a t i o n , whereas u n i d i r e c t i o n a l panels had r e l a t i v e l y s t r a i g h t f i b e r s ( F i g u r e 1). T h i s f i b e r waviness has been commonly observed t o occur i n o t h e r t h e r m o p l a s t i c composi t e s
[ 3,4].
However, u l t r a s o n i c C-scan e v a l u a t i o n s i n d i c a t e d good consol i d a t i o n i n a l l I
panels, w i t h the exception of t h e c r o s s - p l i e d XAS l a m i n a t e s where some a t t e n u a t i o n was noted.
As w i l l be shown l a t e r , dye-penetrant-enhanced
radiography of the EDT specimens o f t h i s XAS/polycarbonate m a t e r i a l r e v e a l e d
*Use o f t r a d e names o r manufacturers does n o t c o n s t i t u t e an o f f i c i a l endorsement, e i t h e r expessed o r imp1 i e d , by t h e N a t i o n a l Aeronautics and Space Admi n i s t r a t i on.
,
3
extensive p l y cracking i n every ply direction.
These cracks were present
even i n an untested specimen. The polyimide blends consisted of 1:l mixtures of commercial LARC-TPI powder w i t h a polyamic acid (either LARC-TPI or polyimidesul fone)
were prepregged onto AS-4 fiber using a slurry process [ S I .
.
They
Each blend also
contained 2.5 weight percent of the diamic acid formed by the reaction of
' p-di ami nobenzene and
p h t h a l i c anhydri de.
Thi s addi t i ve had been shown t o
improve the flow of thermoplastic polyimides [ 6 ] .
For simplicity, the two
blends will be referred t o simply as TPI blend and polyimidesulfone blend. The prepreg was dried and imidized for one hour a t 500°F then 3" x 6" laminates were press-molded i n a 3 to 4-1/2 hour cycle w i t h a maximum temperature of 6 6 O O F and 1000 psi pressure.
Laminates prepared a t lower
pressures had a h i g h void content, as judged by C-scan.
Repressing these
panels a t 660°F and 1000 psi gave acceptable laminates. Test Procedures and Calculations Specimens for the polycarbonate double cantilever beam t e s t were cut from 24-ply unidirectional laminates containing a 0.005"-thick x 1.5" wide Kapton f i l m a t the midplane as a crack s t a r t e r .
Pin-loading was introduced
through 1/2" aluminum blocks bonded t o the beam ends. us1 ng the compl i ance cal i brati on method 17 3.
Results were analyzed
Crosshead speed was 0.05"/mi n ,
and specimen compliance was determined from the opening load vs. time record.
Crack propagation was steady.
In the case of the polycarbonate
matri x 1ami nates, extensive fiber bri dgi ng caused the apparent Gxc to
-
increase r a p i d l y as the crack advanced from the Kapton i n s e r t , typically reaching a steady-state value after a crack extension of only 0.5".
The
steady-state values were t y p i c a l l y 30-100% larger than the i n i t i a l values.
4
I n i t i a l values only are reported, s i n c e they a r e regarded as more c h a r a c t e r i s t i c o f t h e toughness which would be o b t a i n e d i n t h e absence o f f i b e r bridging.
For a l l o f the t e s t s , t h r e e t o f i v e specimens were t e s t e d , Procedures f o r t h e p o l y i m i d e m a t r i x DCB
and t h e r e s u l t s were averaged.
t e s t s were i d e n t i c a l , except t h e specimens were o n l y twelve p l i e s t h i c k . The data r e d u c t i o n took i n t o account geometric n o n l i n e a r i t y by t h e method o f D e v i t t e t a1 181. Edge delamination t e s t s were performed on (+452/-452/02/902)s o r 1" wide w i t h a 3" s e c t i o n between t h e g r i p s .
specimens t h a t were 0.5"
No
d i f f e r e n c e s i n delamination s t r a i n were seen between the two d i f f e r e n t widths.
A layup c o n t a i n i n g 0" p l i e s was chosen t o minimize n o n - l i n e a r i t y i n
the s t r e s s - s t r a i n curve b e f o r e onset o f delamination, and s i x t e e n p l i e s were used t o i n s u r e t h a t t h e delamination s t r a i n was below t h e f a i l u r e s t r a i n o f t h e 0" f i b e r s [4,9,10].
Delamination was d e t e c t e d v i s u a l l y on t h e f r e e
edges, which had been coated w i t h water-based t y p e w r i t e r c o r r e c t i o n f l u i d . T o t a l c r i t i c a l s t r a i n-energy re1 ease r a t e s a r e c a l c u l a t e d v i a t h e formula 2
G,=-
- E*)
(1)
2
where c C i s the delamination s t r a i n , t i s t h e l a m i n a t e thickness, Elam
i s t h e o r i g i n a l l a m i n a t e modulus, and E* i s t h e modulus o f a completely delaminated specimen, which may be c a l c u l a t e d u s i n g t h e r u l e o f m i x t u r e s equation
*
1
E =FxEiti where E i and ti are the modulus and thickness, r e s p e c t i v e l y , o f t h e subl ami nates formed by t h e del ami n a t i on
[ 3,g ,113.
The subl ami n a t e modul i
5
may be calculated from p l y properties, or from the measured s t i f f n e s s of
symmetric sublaminates [ 3 ] .
For the materials used i n the present study,
these customary procedures had t o be modified s l i g h t l y .
For each material,
ply properties were obtained a s before from t e n s i l e measurements on ( O B ) ,
( g o l 2 ) , and (+45)2s laminates [ l l ] . The r e s u l t s are shown i n Table I .
These properties were used w i t h classical laminated plate theory t o predict the modulus o f the quasi-isotropic layup employed i n the EDT t e s t .
When
t h i s prediction was compared w i t h the measured Val ues E1 am, however, i t
was found that the actual specimen moduli f e l l below the predictions by 5 t o 30 percent.
Since i t was known t h a t the specimens contained wavy fibers,
the discrepancy between the measured moduli and the laminate theory prediction was attributed t o fiber waviness i n the quasi-isotropic Such waviness was also noted i n Ref [ 4 ] for AS-4/PEEK.
laminates.
The
rather large v a r i a b i l i t y i n Elam among specimens cut from the same plate was assumed t o be due t o variations i n the degree of f i b e r waviness across the plate - p r i m a r i l y i n the 0' plies, since they dominate the t e n s i l e
modulus.
I t f o l l o w s , then, that a proper value of E* for each specimen
should take account of t h i s v a r i a b i l i t y . was adopted.
Various values of E l i were used i n the laminate theory u n t i l
the measured value of Elam was obtained. El 1,
T h i s "adjusted" ply modulus,
was then used to calculate the corresponding sub1 aminate modulus E* for
t h a t specimen.
.
Therefore, the following procedure
Typically, the adjustment amounted to about 15% of Ell.
T h i s procedure gave the most accurate Values of (Elam
-
E*) f o r each t e s t
specimen, and yields the most r e a l i s t c values of Gc from E q . (1). The effects of thermal stresses ntroduced d u r i n g cool-down from the processing temperature were evaluated using a finite-element (FEM)
6
calculation [ 1 2 ] . a n a l y s i s [ 13
I,
Total G could have been calculated using plate theory
b u t FEM was used f o r convenience since i t was available.
The
mesh used i s shown i n Figure 2 i n Reference [12]. The rectangular mesh had 367 nodes w i t h 102 eight-noded parabolic elements.
This element size a t the
del ami nation t p has been shown t o yield accurate s t r a i n energy release r a t e components 114 The s t r e s -free temperature for each material was determined by fabricating a (02/906) laminate, which bowed upon cool-down. this laminate was heated i n an oven a t approximately 4.4"C/min
height of the arc was monitored w i t h a DCDT (Figure 2 a ) .
A strip o f and the
The temperature a t
which the strip flattened completely was taken as the stress-free temperature (Figure 2 b ) .
As expected, this temperature closely matched the
g l a s s transition temperature of the matrix resin.
Thermal expansion coefficients o f the laminae were assumed to be zero i n the fiber direction.
Transverse expansion coefficients were measured
near room temperature usi ng an interferometric technique, or calculated from the resin and fiber properties u s i n g the concentric cylinder model of Hashin 1151.
In the two cases where both were available, the measured and
calculated values showed f a i r l y good agreement, indicating the technique was reasonably accurate.
Results are shown i n Table 11.
RESULTS AND DISCUSSION
I n i t i a l G l c values for the AS-4/polycarbonate and XAS/polycarbonate obtained from the f i r s t increment o f crack growth beyond the i n s e r t i n the DCB t e s t , were 7.4 and 5.1 i n - l b / i n 2 ,
respectively.
Scanning electron
.
7
micrographs ( S E M I of t h e f r a c t u r e surfaces show, i n b o t h cases, e x t e n s i v e m a t r i x deformation ( F i g u r e 3 ) .
There does seem t o be a q u a l i t a t i v e
d i f f e r e n c e i n f i b e r / m a t r i x adhesion i n t h e two cases, however.
On t h e
AS-4/polycarbonate (PC) f r a c t u r e surface ( F i g u r e 3a) , many f i b e r s a r e judged by SEM examination t o be completely bare, s t r i p p e d of r e s i n , and threads o f drawn polymer l i t t e r t h e surface.
I n t h e XAS composite, by c o n t r a s t , most
f i b e r s seem t o r e t a i n shreds o f polymer ( F i g u r e 3b).
These observations a r e
c o n s i s t e n t w i t h the degrees of adhesion measured f o r t h e two systems u s i n g embedded-fiber fragmentation t e s t s [ 2 ] . from Reference 121.
Table 111 reproduces some r e s u l t s
The c r i t i c a l fragment l e n g t h lC, which p r o v i d e s a
q u a n t i t a t i v e measure of f i b e r / m a t r i x adhesion, seems t o be q u i t e l a r g e f o r AS-4/PCY i n d i c a t i n g very poor adhesion.
As a p o i n t o f reference, t h e
c r i t i c a l l e n g t h f o r AS-4 i n an epoxy r e s i n i s about 0.4 nnn.
The much
s m a l l e r c r i t i c a l l e n g t h f o r XAS/PC i s i n d i c a t i v e o f b e t t e r adhesion, a1 though one would need t o know the t e n s i l e s t r e n g t h s o f t h e two f i b e r s a t submi 11i m e t e r gauge 1engths t o c a l c u l a t e the r e 1 a t i ve s t r e n g t h s o f t h e i n t e r f a c i a l bonds.
A second, independent l i n e o f evidence, however, tends
t o c o n f i r m the d i f f e r e n c e i n adhesion t o t h e two f i b e r s .
This i s the
b i r e f r i n g e n c e p a t t e r n which a r i s e s around t h e broken f i b e r ends.
It i s
q u i t e d i s t i c t i v e , and shows c h a r a c t e r i s t i c features of poor adhesion w i t h AS-4 and good adhesion w i t h XAS.
Based on t h e data above, t h e r e f o r e , poor
f i b e r / m a t r i x bonding, such as t h a t between AS-4 and polycarbonate, appears t o y i e l d h i g h e r composite i n t e r l a m i n a r f r a c t u r e toughness as measured by t h e
DCB t e s t . caution.
T h i s i s a s u r p r i s i n g r e s u l t , and should be regarded w i t h I n Reference 1161, f i b e r / m a t r i x debonding i n DCB t e s t s was s a i d t o
c o n t r i b u t e t o increased GIC
by i n c r e a s i n g f r a c t u r e s u r f a c e area, by
8
promoting fiber b r i d g i n g , and by relieving s t r e s s t r i a x i a l i t y , thereby a l l o w i n g more resin deformation.
These e f f e c t s , and especially f i b e r
bri dgi ng, are normal l y observed a f t e r the del ami nation has grown some
distance from the i n s e r t [ 171.
However, i n the present study, increased
toughness was found for i n i t i a t i o n values.
A possible explanation would be
t h a t i n t h i s case fiber b r i d g i n g began to develop immediately a t the
insert.
T h i s conclusion must be regarded as tentative, though, because the
hypothesized b r i d g i n g was n o t detected directly. In the edge delamination t e s t s , delamination occurred, as expected, a t
the 0/90 interfaces w i t h the failure plane wandering back and f o r t h through the central 90" plies [3,9,11].
Dye-penetrant x-ray photos (Figure 4a) show
t h a t i n the AS-4/PC composites, the 90' cracks do not extend much beyond the
delaminated region indicating t h a t they form a f t e r the delamination [ l l ] . In contrast dye penetrant-enhanced radiographs of the XAS/PC showed
extensive matrix o r interfacial cracks throughout the laminate width i n a l l plies.
As mentioned e a r l i e r , this cracking was present before any
mechanical l o a d was applied.
Micrographs of the 0/90 delamination fracture
surfaces of b o t h the AS-4/PC and XAS/PC EDT specimens are dominated by what appear t o be bare fibers and resin tracks l e f t by f i b e r s which peeled out cleanly (Figure 5 ) .
The resin i n the XAS/PC composite appears to have
fractured b r i t t l e l y (Figure 5 b ) .
T h i s apparent matrix b r i t t l e n e s s was
completely unexpected, and two possible explanations for i t have been considered.
I f polymer c r y s t a l l i n i t y was being nucleated by the f i b e r
surface 118 1, the apparent b r i t t l e behavior m i g h t have been understandable, b u t no evidence of matrix c r y s t a l l i n i t y was found by differential scanning
cal orimetry or by wide-angle x-ray scattering.
Another possi b i 1i ty
-
that
9
t r a c e i m p u r i t i e s on t h e f i b e r surface c o u l d c a t a l y z e r e s i n degradation
- was
a l s o r u l e d out, s i n c e r e s i n d i s s o l v e d from t h e XAS composite had n o t decreased i n i n h e r e n t v i s c o s i t y .
Thus t h e reason f o r t h e b r i t t l e appearance
o f the r e s i n i n F i g u r e 5b i s s t i l l n o t c l e a r .
The EDT r e s u l t s a r e compared
w i t h those from t h e DCB t e s t i n Figure 6, i n which t h e s u p e r s c r i p t s M and T stand f o r mechanical and thermal s t r e s s c o n t r i b u t i o n s t o Gc.
Both t e s t s
in d i c a t e t h a t t h e AS-4 composite possesses t h e h i g h e r in t e r l ami n a r
toughness.
Note, however, t h a t the t o t a l Gc as measured by t h e EDT i s
l o w e r than t h e DCB r e s u l t f o r
GIG.
T h i s t r e n d has been observed i n o t h e r
m a t e r i a l s [ 3 ] , b u t w i t h polycarbonate t h e d i f f e r e n c e i s s t r i k i n g .
Including
thermal stresses i n t h e EDT c a l c u l a t i o n s r a i s e s t h e EDT r e s u l t s , b u t o n l y by l e s s than 20% because t h e AT between t h e c o n s o l i d a t i o n temperature and room temperature i s n o t very l a r g e .
The l a r g e d i f f e r e n c e s between t h e DCB and
t h e EDT values i n t h i s case are t e n t a t i v e l y a t t r i b u t e d t o t h e e f f e c t s o f f i b e r b r i d g i n g i n t h e DC6 t e s t , although t h i s has n o t o f course been shown conclusively.
The processing [ l ] and o t h e r f a c t o r s which l e a d t o b r i d g i n g
i n u n i d i r e c t i o n a l specimens are not y e t f u l l y understood. AS-4/Polyimi de B1ends Double c a n t i 1ever beam specimens of these m a t e r i a1 s showed much 1ess obvious f i b e r b r i d g i n g than the polycarbonate, although an i n c r e a s e i n G I ~ w i t h crack l e n g t h was again noted.
Because t h e r e was no r e s i n squeeze-out
d u r i n g f a b r i c a t i o n o f these panels, t h e f i b e r s probably c o u l d n o t n e s t e f f e c t i v e l y , and the f r a c t u r e surfaces were q u i t e f l a t .
I n the edge delamination t e s t s , these two m a t e r i a l s behaved s i m i l a r l y , and t h e i r f r a c t u r e surfaces again show r e g i o n s o f bare f i b e r s and r e s i n t r o u g h s where f i b e r s have peeled out ( F i g u r e 7).
The SEM of the PISOZ/LaRC-
10
T P I surface i n F i g u r e 7b i n d i c a t e s some f i n e - s c a l e h e t e r o g e n e i t y , which may
r e s u l t from incomplete b l e n d i n g o f t h e LaRC powder with t h e PIS02 b i n d e r . Whether b e t t e r mechanical m i x i n g i s needed, o r whether these two polymers a r e thermodynamically i n c o m p a t i b l e i s n o t known. With t h e higher s o f t e n i n g temperatures of these p o l y i m i d e s r e l a t i v e t o
conventional epoxy matrices, the i s s u e o f thermal s t r e s s e s assumes much g r e a t e r importance, as w i l l be c l e a r from an examination o f t h e EDT r e s u l t s i n F i g u r e 8.
The s t r a i n - e n e r g y r e l e a s e i n c l u d i n g thermal s t r e s s e s GM+T i s
almost t w i c e as l a r g e as t h a t due t o t h e mechanical s t r a i n o n l y , GM.
As
mentioned e a r l i e r , one reason f o r u s i n g t h e EDT t e s t i s t o i l l u s t r a t e t h i s r o l e o f r e s i d u a l stresses i n delamination.
The EDT t e s t i s performed w i t h a
r e a l i s t i c laminate which i n h e r e n t l y c o n t a i n s r e s i d u a l thermal s t r e s s e s
[4,10,13]. I n the case o f these t h e r m o p l a s t i c polyimides, the c o n t r i b u t i o n o f r e s i d u a l stresses t o delamination i s s i g n i f i c a n t , amounting t o 28 and 32 p e r c e n t o f GcM+T f o r t h e T P I and p o l y i m i d e s u l f o n e blends, r e s p e c t i v e l y . F i g u r e 8 also shows t h e r e l a t i v e magnitudes o f G I ~ determined from t h e DCB t e s t and t o t a l GcM+T c a l c u l a t e d from t h e EDT t e s t .
For t h e
p o l y i m i d e m a t e r i a l s , u n l i k e t h e polycarbonate, one sees t h a t t o t a l Gc i s comparable i n magnitude t o G I ~ .
The l a y u p chosen f o r t h e EDT t e s t i n t h i s
work produces almost a pure mode I delamination, so t h i s would be expected. Furthermore, Johnson and M a l g a l g i r i [19] have shown t h a t f o r o t h e r tough r e s i n composi t e systems, in c l u d i ng t h e thermopl a s t i c PEEK APC2, GIC GII~=
Gc a t delamination onset.
=
11
SUMMARY AND CONCLUSIONS
The edge delamination t e s t and DCB t e s t gave similar toughness rankings for the fiber/resin combinations studied here.
Fiber b r i d g i n g seems t o
greatly affect the DCB results for the tough thermoplastic, even a t onset from the i n s e r t , although t h i s effect i s not fully understood. A procedure based on adjusting the modulus of 0' plies i n a laminate is
a reasonable way t o account for the e f f e c t s of fiber waviness on composite 1ami nate sti f fness i n thermoplastic materi a1 s.
T h i s procedure was needed t o
accurately measure Gc from the EDT t e s t . Thermal residual stresses must be considered when analyzing results of EDT t e s t s on higher-temperature matrix material s.
In one graphite/polyimide
system, the thermal contribution amounted t o 32% of the total Gc.
REFERENCES
1.
Hunston, D. L., Moulton, R. J . , Johnston, N. 3. and Bascom, W. D. "Matrix Resin Effects i n Composite Delamination:
Aspects
."
Toughened Composites.
Mode I Fracture
ASTM STP 937, Norman J
. Johnston,
ed. American Society for Testing and Materials, Philadelphia, 1987, pp. 74-94
2.
Bascom, W. D., Jensen, R. M. Cordner, L. W. Fibers t o Thermoplastic Polymers."
"The Adhesion of Carbon
International Conference on
Composite Materials (6th:1987:London) Matthews, F. L., London, 1987, p. 5.424
ed. Elsevier,
12
3.
Johnston, N.
J., O'Brien, T. K., M o r r i s , D . H., and Simonds, R. A.
"I n t e r l ami nar F r a c t u r e Toughness of Composites.
I1
-
Ref inement o f t h e
Edge Delamination Test and A p p l i c a t i o n t o Thermoplastics,"
28th N a t ' l
SAMPE Symp. and Exhib. 28, 502 (1983)
4.
O'Brien, T. K.,
"Fatigue Delamination Behavior o f PEEK Thermoplastic
Compos1 t e Laminates," Proceedings o f t h e American Society f o r Composites, F i r s t Technical Conference, Dayton, Ohio, October, 1986, Technomic P u b l i s h i n g Co.,
pp. 404-420 ( a l s o t o appear i n t h e Journal o f
Reinforced P l a s t i c s ) 5.
Johnston, N.
J. and
S t . C l a i r , T. L.
"Thermoplastic M a t r i x Composites:
LARC-TPI, Polyimidesul fone and T h e i r Blends."
1 8 t h SAMPE I n t ' l
Technical Conference Oct. 1986 6.
P r a t t , J . R.,
S t . C l a i r , T. L.,
"Polyimide Processing Additives."
Burks, H.,
D. and Stoakley, D. M.
32nd I n t ' l SAMPE Symposium.
April
1987 32, p . 1036.
7.
W i l k i n s , 0. J., Bensen , R. A.
Eisenmann, J . R.,
Camin, R. A.,
M a r g o l i s , W. S. and
"Charac t e r i z i ng Del ami n a t i o n Growth i n Graphi te-Epoxy
"
i n Damage i n Composite M a t e r i a l s , ASTM STP 775, K. L. R e i f s n i d e r , ed. American Society f o r T e s t i n g and M a t e r i a l s , 1982 pp. 168-183 8.
D e V i t t , D. F.,
Schapery, R. A. and Bradley, W. L. "A Method for
Determining t h e Mode I Delamination F r a c t u r e Toughness o f E l a s t i c and V i scoel a s t i c Composite M a t e r i a l st' J . Composite M a t e r i a l s v o l
. 14,
1980,
p. 270 9.
O'Brien, T. K.
"Mixed-Mode Strain-Energy Release Rate E f f e c t s on Edge
Delamination of Composites" i n E f f e c t s of Defects i n Composite
13
M a t e r i a l s , ASTM STP 836, American S o c i e t y f o r T e s t i n g and M a t e r i a l s , P h i l a d e l p h i a , 1984, pp. 125-142 Johnston, N. J . , Raju, I.S.,
10. O'Brien, T. K., Simonds, R. A.
M o r r i s , D. H.,
and
nComparisons o f Various C o n f i g u r a t i o n s o f t h e Edge
Delamination Test f o r I n t e r l a m i n a r F r a c t u r e Toughness" i n Toughened Composites, ASTM STP 937, N.
J. Johnston ed., American Society f o r
T e s t i n g and M a t e r i a l s , P h i l a d e l p h i a , 1987 11.
O'Brien, T. K.,
Johnston, N. J.,
Morris,
D. H. and Simonds, R. A.
"A
Simple Test f o r t h e I n t e r l a m i n a r F r a c t u r e Toughness o f Composites." SAMPE Journal, 18(4)8 (1982)
12.
Raju, I . S.
"Q3DG - A Computer Program f o r Strain-Energy Release Rates
f o r Del ami n a t i o n Growth i n Composite Lami nates,"
NASA CR 178205,
October 1986 (Computer program i s avai 1ab1e through COSMIC) 13.
O'Brien, T. K.,
Raju, 1. S.,
and Garber,
D.
P.
"Residual Thermal and
M o i s t u r e I n f l u e n c e s on t h e S t r a i n Energy Release Rate A n a l y s i s o f Edge Delamination" J
. Composites Technology
and Research v o l
.8
No. 2,
Summer 1986, pp. 37-47 14.
Raju, I. S., Crews, J . H. Jr., and Aminpov, M. A.
"Convergence of
Strain-Energy Release Rate Components f o r Edge-Delaminated Composite Laminates, NASA TM 8935, A p r i l 1987 15.
Hashin, Z.
" A n a l y s i s of P r o p e r t i e s o f F i b e r Composites with
Ani s o t r o p i c Consti tuentsll J . Appl 16.
Hibbs, M. F.,
. Mech.
Tse, M. K. and Bradley, W.
vol
L.
. 46,
pp. 543-550 (1979).
"Interlaminar Fracture
Toughness and Real-Time Fracture Mechanism o f Some Toughened Graphi te/Epoxy Composi t e s t ' Toughened Composites, ASTM STP 937, N.
J
.
14
Johnston, ed. American Society for Testing and Materials, Philadelphia, 1987 pp. 115-130 17.
Russell, A. J. and Street, K. N.
"Moisture and Temperature Effects on
the M i xed Mode Del ami nation Fracture of Unidirectional Graphi te-Epoxy" i n Delamination and Debonding of Materials, ASTM STP 876, W. S.
Johnson, ed., American Society f o r Testing and Materials, Philadelphia, 1985 18.
Kardos, J . L., Cheng, F. S., and Tolbert, T. L. Interface in Graphi te-Rei nforced Polycarbonate
.I'
"Tailoring the Polymer Engi neeri ng
and Science, vol. 13, p. 455 (1973) 19.
Johnson, W. S. and Mangalgiri, P. D.
"Influence of the Resin on
Interlaminar Mixed-Mode Fracture," i n Toughened Composites, ASTM STP 937, N. J . Johnston, ed. American Society f o r Testing and Materials, Philadelphia, 1987
15 TABLE I PLY PROPERTIES
Ms i
--
AS4/Polycarbonate
17.7
E 612 1.10 0.37
XAS/Polycarbonate
19.0
1.37
0.44
20.8
1.64
18.4
1.29
MATERIAL
AS4/LARC-TPI
+ LARC-TPI Powder
AS4/Polyimidesulfone + L A R C - T P I Powder
E11
22
"12
FIBER .VOLUME AVERAGE PLY,THICKNESS % h, x10 i n 55.8
6.08
0.37
55.1
5.63
0.91
0.34
56.5
5.21
0.78
0.34
52.0
6.55
0.71
16 TABLE I 1 THERMAL EXPANSION PROPERTIES L I N E A R COEFF IC IENT OF THERMAL EXPANSION, m/m/"C LAM I NATE
MEASURED STRESS-FREE TEMPERATURE, O C
RESIN
CALCULATED COMPOSITE a2
MEASURED COMPOSITE a2
AS4/Polycarbonate
130
67.5
41.2
36.7
XAS/Polycarbonate
--
67.5
51.5
--
229
45
29.3
--
243
45*
31.9
27.9
AS4/TPI
blend
A S 4 / P o l y i m i d e s u l f o n e blend *Assumed
17 1
.
TABLE I 1 1 EMBEDDED FIBER FRAGMENTATION RESULTS
F IBER/MATR IX
C R I T I C A L FRAGMENT LENGTH*, mm
BIREFRINGENCE PATTERN
AS-4/PC
0.74
Poor Adhesion
XAS/PC
0.36
Good Adhesion
*The c r i t i c a l fragment length i s i n v e r s e l y proportional t o t h e f i b e r / m a t r i x i n t e r f a c i a l shear strength if f i b e r breaking strength i s constant
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Report Documentation Page 3. Recipient's Catalog No.
2. Government Accession No.
1. Report No.
NASA TM- 100532 USAAVSCOM TM 88-B-002 5. Report Date
4. Title and Subtitle
February 1988
I n t e r l a m i n a r Fracture Toughness of Thermoplastic Composites
6. Performing Organization Code
8. Performing Organization Report No.
7. Authork)
J. A. Hinkley, N. J. Johnston, and T. K. O'Brien 10. Work Unit No. 9. Performing Organization Name and Address
NASA Langley Research Center Hampton, VA 23665-5225 and A e r o s t r u c t u r e s D i rectorate, USAARTA-AVSCOM NASA Langley Research Center, Hampton, VA 23665-5225
505-63-01-01
-
11. Contract or Grant No.
13. Type of Report and Period Covered
12. Sponsoring Agency Name and Address
N a t i o n a l Aeronautics and Space A d m i n i s t r a t i o n Washington, DC 20546-0001 and U.S. Army A v i a t i o n Systems Comnand S t . Louis, MO 63120-1798
Technical Memorandum 14. Sponsoring 4gency Code
I n t h e model thermoplastic system (polycarbonate m a t r i x ) , surface p r o p e r t i e s o f t h e g r a p h i t e f i b e r were shown to be s i g n i f i c a n t . C r i t i c a l s t r a i n energy r e l e a s e r a t e s f o r two d i f f e r e n t f i b e r s having s i m i l a r nominal t e n s i l e p r o p e r t i e s d i f f e r e d by 30-60%. The reason f o r t h e d i f f e r e n c e i s n o t c l e a r .
18. Distribution Statement
17. Key Words (Suggested by Authorls))
ThermoDlastic-matrix composites I I n t e r 1ami n a r f r a c t u r e F i b e r / m a t r i x i n t e r f a c e Polyimide composi t Carbon f i b e r Edge d e l ami n a t i o n Double c a n t i l e v e r beam
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