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Fatigue damage propagation in graphite/epoxy composites Yavuz, Ömer 1984

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FATIGUE DAMAGE PROPAGATION IN GRAPHITE/EPOXY COMPOSITES by OMER YAVUZ B.Sc,  The U n i v e r s i t y Of Manchester Technology, Manchester,  I n s t i t u t e Of S c i e n c e And England, 1982  A THESIS SUBMITTED IN PARTIAL FULFILMENT OF THE REQUIREMENTS FOR THE DEGREE OF MASTER OF APPLIED SCIENCE in THE FACULTY OF GRADUATE  STUDIES  Department Of M e t a l l u r g y  We accept t h i s t h e s i s as conforming to the r e q u i r e d  standard  THE UNIVERSITY OF BRITISH COLUMBIA February 1984  ©  Omer Yavuz, 1984  In p r e s e n t i n g  t h i s t h e s i s i n p a r t i a l f u l f i l m e n t of  requirements f o r an advanced degree a t the  the  University  o f B r i t i s h Columbia, I agree t h a t the L i b r a r y s h a l l make it  f r e e l y a v a i l a b l e f o r reference  and  study.  I further  agree t h a t p e r m i s s i o n f o r e x t e n s i v e copying of t h i s t h e s i s f o r s c h o l a r l y purposes may  be granted by  department o r by h i s or her  the head o f  representatives.  my  It i s  understood t h a t copying or p u b l i c a t i o n o f t h i s t h e s i s f o r f i n a n c i a l gain  s h a l l not be  allowed without my  permission.  Department o f  nerALL_U»Q.GIC_AL  The U n i v e r s i t y o f B r i t i s h 1956 Main Mall Vancouver, Canada V6T 1Y3 Date  DE-6  (3/81)  " £ . M & i * J ££.A.I*J  Columbia  c  written  i i  Abstract  in  A  series  the  crack  composites  of t e s t s was performed t o determine the i n c r e a s e length  under  or  fatigue  damage  tension-tension  of  fatigue  graphite/epoxy using  microscope, tetrabromoethane enhanced radiography techniques. H e r c u l e s used  to  produce  traveling  and compliance  As/3501-6 g r a p h i t e f i b r e / e p o x y prepreg  laminates  which  were  compact t e n s i o n sample geometry. [ 9 0 / 0 ]  subsequently  was  cut i n t o  ply configuration  8i  was  used i n the p r e p a r a t i o n of the l a m i n a t e s . Fatigue  damage was observed  t o be i n the form of a 'damage  zone' r a t h e r than a s i n g l e c r a c k , which both  the  0°  and  90°  directions  increased  during  in  was  the  acceleration;  d e c e l e r a t i o n and the f i n a l damage. found  As  a  result  the  intermediate  The  first  stage was the  stage was the r e a c c e l e r a t i o n  of t h i s behaviour  in  f a t i g u e . The damage  f o l l o w e d three s t a g e s , d u r i n g the whole f a t i g u e l i f e . stage  size  of the  the P a r i s approach was  t o be n o n - v a l i d f o r t h i s m a t e r i a l . Change i n the compliance was observed  during  fatigue  and  t h i s change r e v e a l e d the same three stages as the development of the damage zone.  iii  Table of Contents T a b l e Of L i s t of L i s t of CHAPTER  Contents Tables Figures I. INTRODUCTION  1.1 G e n e r a l Review 1.2.1 T e n s i o n - T e n s i o n  i i i iv v  And  Tension-Compression  1 ..1 Fatigue 3 6 7  1.2.2 E f f e c t Of Frequency On F a t i g u e 1.2.3 E f f e c t Of Manufacturing On F a t i g u e 1.2.4 E f f e c t Of S t a c k i n g Sequence On Laminate Strength 1.2.5 Micromechanisms Of F a t i g u e 1.2.6 Damage Zone D e t e c t i o n Methods And Crack V e l o c i t y Studies 1.2.7 F o r m u l a t i n g The Decay In S t i f f n e s s During Fatigue 1 .3 Theory 1.3.1 F r a c t u r e Toughness Equations 1.3.2 F a t i g u e T e s t i n g Equations 1.3.3 Compliance T e s t s 1.4 Purpose . CHAPTER I I . PROCEDURE 2.1 2.2 2.3 2.4 2.5  Laminate And Sample P r e p a r a t i o n S t a t i c Tensile Fracture Testing Fatigue Testing Compliance T e s t i n g Determining The Weight And Volume F r a c t i o n Of Fibres CHAPTER I I I . RESULTS 3.1 T e n s i l e F r a c t u r e And Load Curve A n a l y s i s 3.2 F a t i g u e T e s t i n g R e s u l t s 3.2.1 F a t i g u e L i f e Of The M a t e r i a l 3.2.2 F a t i g u e Crack Length Measurements 3.2.3 Change In The Compliance During F a t i g u e 3.2.4 P a r i s P l o t Of The M a t e r i a l 3.2.5 Radiographs Of The F a t i g u e d Samples 3.2.6 F i n a l F r a c t u r e Morphology CHAPTER IV. DISCUSSION 4.1 4.2  S t a t i c Tests Fatigue Tests i ) Fatigue L i f e i i ) F a t i g u e Damage Propagation i i i ) Compliance Data i v ) F i n a l F r a c t u r e Morphology CHAPTER V. CONCLUSIONS REFERENCES  8 10 13 16 18 18 19 21 22 23 23 25 26 29 The 31 33 33 35 35 36 39 41 42 44 46 46 .47 47 48 51 52 54 57  iv  L i s t of Tables  I . F r a c t u r e Toughness v a l u e s f o r samples of L a m i n a t e 2 u s i n g ASTM compact t e n s i o n sample e q u a t i o n ( e q n . 3) .111 I I . F r a c t u r e Toughness v a l u e s f o r samples o f L a m i n a t e 4 u s i n g ASTM compact t e n s i o n sample e q u a t i o n ( e q n . 3) .112 I I I . F a t i g u e T e s t i n g Data f o r L a m i n a t e s 1,2,3,4, a n d 7 ...113 I V . Compliance Test R e s u l t s 116 V . Weight and Volume F r a c t i o n of t h e F i b r e s f o r E a c h Laminate 118  V  List  of F i g u r e s  1. I n c r e a s e i n the c r a c k l e n g t h d u r i n g compression f a t i g u e f o r u n i d i r e c t i o n a l and [0/90] c r o s s - p l y g r a p h i t e / e p o x y composite. 61 2 6  2. S t i f f n e s s decay d u r i n g f a t i g u e . T h i s i s the d a t a of Smith (22.5° t o the weave) used i n the development of Poursartip's theory. 61 2 7  3. Compact t e n s i o n specimen  used i n the e x p e r i m e n t s  62  of g r a p h i t e / e p o x y composite  63  4.  ( i ) C u r i n g Treatment  4.  ( i i ) The p o s i t i o n of the compact t e n s i o n samples t h e i r notches as they were c u t from a l a m i n a t e  and  64  5. T y p i c a l t e n s i l e l o a d i n g p a t t e r n of the t e s t samples l e a d i n g to f r a c t u r e using I n s t r o n t e s t i n g machine. ...65 6. The method of drawing new C-a l i n e s i f the c o m p l i a n c e s of the samples have the same notch l e n g t h from the same l a m i n a t e are d i f f e r e n t from each o t h e r 66 3  7. L o a d - d e f l e c t i o n data of sample F lengths  €  f o r various notch 67  8. Change i n the compliance w i t h the notch l e n g t h f o r Sample F  68  9. V a r i a t i o n of the compliance w i t h the cube of notch l e n g t h f o r sample F  69  e  6  10. F a t i g u e L i f e data  70  11. Crack Length vs c y c l e s f o r sample J min. load=222N,a = 24 . 5mm  max.  3  0  12. Crack l e n g t h v s . c y c l e s f o r sample K . min. load=240N, a = 24.8mm 3  load=4446N, .'....71  max.  0  13. a vs N f o r sample L . max. a o = 2 5mm  load=4688N, min.  14. a vs N f o r sample B . ao = 24. 5mm  load=4466N, min.  3  5  max.  load=4528N, ...72 load=142N, 73 load=l51N, 74  15. a vs N f o r sample C . max. Ioad-=4101N, min. a = 25mm. Low s t r e s s f a t i g u e t e s t  load=7lN,  16. a vs N f o r sample E . max. a 25mm  load=338N, 76  5  0  5  =  0  load=4822N, min.  75  vi  17. a vs N f o r sample F load=53N,a 25mm 0  max.  5  load=507lN,  min. 77  e  18. a vs N f o r sample H . a = 24 . 5mm 5  max.  load=4857N, m i n .  load=80N, 78  0  19. a vs N f o r sample K . max. load=3932N, m i n . load=80N, a =24.5mm. T y p i c a l low s t r e s s f a t i g u e . Sample d i d not fail 79 5  0  2 0 . a vs N f o r Sample A . High s t r e s s f a t i g u e 6  21.  max.  ( i ) a vs N for Sample C . load= 142N,a = 24 . 5mm 6  load=5204N, m i n . max.  load=80N.  80  load=3692N, m i n . 81  0  21.  ( i i ) a vs N for sample g . load=80N, a = 24.45mm 6  max. load=5249N, m i n .  81  0  22. a vs N f o r a = 24.8mm  Sample E .  max.  load=5266N, m i n .  load=l228N, 62  23. a vs N f o r Sample K . ao = 24.1 mm  max.  load=4902N, m i n .  load=142N,  6  0  6  83  24. a vs N f o r Sample L . max. load=5071N, m i n . a = 24.3mm. High s t r e s s f a t i g u e 6  load=142N,  84  0  25. a vs N f o r Sample C . max. load=4573N, m i n . load=15lN. a =24.9mm. Sample was not f a t i g u e d up to f i n a l f a i l u r e . A t y p i c a l medium s t r e s s f a t i g u e 85 6  0  26. a vs N f o r Sample F . max. load=499lN, min. l o a d = 7 l N , a =24.5mm. High s t r e s s f a t i g u e . E f f e c t i v e c r a c k l e n g t h showed a higher i n c r e a s e in v e l o c i t y than the o t h e r s . 7  0  27. a vs N f o r Sample I . a o = 2 4. 5mm 7  28. Change in the load samples L and H 6  max.  load=5l60N, min.  86  load=80N, 87  deflection  curves  during  6  fatigue  of 88  29.  Increase  in the compliance of Sample C  5  during  fatigue. 89  30.  Increase i n the compliance of Sample E  5  during  fatigue. 90  31.  Increase in the compliance of Sample F  5  during  fatigue. 91  vi i  32.  Increase  i n the compliance  of H  33.  Increase  i n the compliance  of J  34.  Increase  i n the compliance  of Sample K  35.  Increase  i n the compliance  of Sample E  36.  Increase  i n the compliance  of K  37. Increase  i n the compliance  of I  5  6  40. Sample C  6  41 . Sample E  6  5  and K  92  during fatigue  93  s  6  during  fatigue. 94  during  fatigue. 95  during fatigue  6  7  38. Load d e f l e c t i o n data of a t e n s i l e (failed accidentally) 39. P a r i s P l o t f o r Samples E  during fatigue  and F  7  during  fractured  96 fatigue. 97  sample  98 99  5  1 00 101  42 . Sample G  6  1 03  43. Sample K  6  1 04  44. Sample L  106  6  45. Sample C«  107  46. A sample f a i l e d by d e l a m i n a t i o n . N o t i c e the f a i l u r e i s caused from the damage i n the two p e r p e n d i c u l a r d i r e c t i o n s t h a t i s v i s i b l e i n the radiographs 109 47. F a i l e d sample geometry showing d e l a m i n a t i o n , debonding and p u l l - o u t 48. Sample f a i l e d  fibre  from the hole  49. Sample f r a c t u r e d under I n s t r o n incremental loading  -109 110  tensile  110  vi i i  Acknowledgement  The given  author  is  grateful  f o r the a d v i c e  and  encouragement  by Dr. J.S. Nadeau and Dr. E. T e g h t s o o n i a n . Thanks a r e a l s o extended t o my  f e l l o w graduate s t u d e n t s and  f a c u l t y members i n the Department of M e t a l l u r g i c a l The  assistance  of  p a r t i c u l a r Mr. R.C.  the  Engineering.  t e c h n i c a l s t a f f of t h i s department,  Bennett and Miss  N.  Talebian,  is  in  greatly  appreciated. Financial  assistance  a s s i s t a n t s h i p under gratefully  National  acknowledged.  received Research  in  the  Council  form of  of  an  Canada  is  1  Chapter I INTRODUCTION 1 .1 General Review It  was j u s t  plastics  i n the e a r l y 70's that carbon f i b r e r e i n f o r c e d  (CFRP) made t h e i r  debut  Given the v a r i e t y of m a t r i c e s , plastics  had  already  as  an  industrial  i n which carbon f i b r e c o u l d work,  emerged  as  the one s u p p o r t i n g  o f f e r i n g the most immediate promise. Since of  application  involving  in  material.  aerospace,  sports  material  then the wide v a r i e t y  goods  and  industry,  a l l the primary CFRP c h a r a c t e r i s t i c s have emphasised  the v e r s a t i l i t y of CFRP. These m a t e r i a l s have great promise i n r e d u c t i o n and  freedom  designer There  from  to t a i l o r  fatigue  weight  and c o r r o s i o n . They a l s o permit the  the m a t e r i a l to match  the  applied  loading.  a r e , on the other hand, s e r i o u s problems t o be overcome;  the cost of the m a t e r i a l , i t s b r i t t l e nature, to  of  erosion,  components  the and  variability  the  between  difficulty  of  i t s susceptibility  apparently  identical  making j o i n t s between sub-  assemblies. One sometimes wonders how many m a t e r i a l s would  never  have  been developed i f a l l the f a u l t s had been determined a t the time of  their  strength  discovery. as  researchers of  fracture  the have  Fr'om  main  a  guide  simple to  a  measurement material's  of  tensile  usefulness  g r a d u a l l y become accustomed to c o n s i d e r a t i o n s  toughness,  stress  corrosion  resistance  and an  2  everlengthening l i s t electrical The of  this  of other m e c h a n i c a l , p h y s i c a l , c h e m i c a l  and  properties.  evaluation trend  p r o p e r t i e s . As f a t i g u e property  of f a t i g u e r e s i s t a n c e  towards these  a  deeper  materials  should be c l e a r l y  is a  typical  understanding  have  vital  understood.  of  example material  applications  the  3  1.2.1  Tension-Tension And Tension-Compression  Ramani and fatigue  Williams  worked  1  on  'notched  Fatigue and  unnotched  behavior of a n g l e - p l y graphite/epoxy composites'.  i n v e s t i g a t i o n on the unnotched  [0/±30]  Their  graphite/epoxy r e v e a l e d  3  that f a i l u r e o c c u r r e d at s t r e s s l e v e l s t h a t were 40% lower the  average  static  to the e s s e n t i a l l y  s t r e n g t h . T h i s behavior i s i n sharp c o n t r a s t fatigue  insensitive  behavior  exhibited  u n i d i r e c t i o n a l graphite/epoxy at r e l a t i v e l y h i g h s t r e s s It  was  of the  observed that the f a t i g u e l i m i t static  percentage compression  of  According sufficient tensile  strength)  increases  0-degree  plies  the  cycling.  temperature  growth  number of  0-degree are  were  observed  In graphite  immediately  contrast,  plies  to  by  made  and  static  these during  increase  Morris  f i b r e - r e i n f o r c e d composites  temperature  increasing and  different  its  tension-  from  the  and  fatigue  0-degree cycling  plies. and  asymptotically  4  a the  fracture. have  i n d i c a t e d that the  do not show any  r i s e even at the high t e s t  the to  c y c l e s and to remain  before the f i n a l  Owen  2  when a laminate c o n t a i n s a  3  constant v a l u e w i t h i n s e v e r a l thousand same u n t i l  with  laminate  was  controlled  measurements was  levels.  1  to Awerbuch and Hahn  strengths  Temperature  damage  in  by  ( i n terms of the percent  ultimate  cycling  tension-tension  than  significant  frequency of 7000  per minute at which s t r u c t u r a l s t e e l s become very hot.  cycles  4  Also, not  L i b e r and D a n i e l  exhibit  properties fatigue.  any  such  observed t h a t  5  significant  as  loss  of  deterioration  stiffness  and  strengths  and hence t h e i r  c o n t r o l l e d by the c o m p r e s s i v e Dharan  studied  6  the  t e n s i o n / c o m p r e s s i o n and limits.  f a i l u r e was  At  of  do  mechanical  strength  during  fatigue  in  of  reversed than  at  4x10  fatigue  flexural  fatigue  is  graphite/polyester  in  strength.  less  observed  tension/compression  75%  bending of  reversals.  s  tests  r e v e r s e d bending was observed  between  Results  indicated that  to  imposed  the u l t i m a t e s t r a i n  o c c u r r e d d u r i n g the compression p o r t i o n of of  materials  Compressive s t r e n g t h s have been found t o be much lower  than the t e n s i l e  stroke  these  be  from  failure  the c y c l e . less  the  always  The  severe  no  effect  than  in  fatigue  of  tension/compression. Bader  and  unidirectional and  Johnson  because  flexural  on the f r a c t u r e d s u r f a c e . For t h e i r  these two zones was  the  the  CFRP s p e c i m e n s . They observed two zones  compressive)  presence of  studied  7  compression  side  in  accord  of  the  the  expectations,  tension  during  in the  loading.  Beaumont tension  and  loading,  fractures static  study,  specimen was always  compression and the t e n s i o n s i d e always i n fatigue  with  (tensile  in  failure  Morris.*  Both  10  Harris  8  showed  carbon/polyester 7  cycles  stress. studies  for  This also  that  under  composites  stress levels was  also  exhibited  l e s s than 0 . 9 of  observed  showed that  repeated  by  Owen  pure no the and  the m a t e r i a l s behaved  5  w e l l under f l e x u r a l at Also  flexure  stresses  off-axis  significantly Ryder  loading; f a i l u r e occurred  loads  above  0.66  were  found  under c y c l i c and  Walker  9  of the to  i n 10  static  7  failure  reduce  the  studied  the and  fatigue  and  concentrations composites  and  reduce the  co-workers fatigue  a l t h o u g h the s t a t i c  f i b r e composite was  lower  stiffness  of  notched specimens  and  or compression-compression his  stress.  behavior  found that g e n e r a l l y t e n s i o n - c o m p r e s s i o n l o a d i n g was  Schutz  only  loading.  graphite/epoxy composites u s i n g p l a i n  than t e n s i o n - t e n s i o n  cycles  than  for  loading. * *  observed  11  strength  of  more severe 1  that  9  1 0  stress  graphite/epoxy  s t r e n g t h of the h i g h modulus the  high  strength  composite, the f a t i g u e s t r e n g t h of both appeared to be the  fibre same.  6  1.2.2  E f f e c t Of Frequency  Stinchcomb and frequency  during  On F a t i g u e  h i s co-workers fatigue  studied  12  c y c l i n g of composites.  parameter was found t o be a major f a c t o r and  governing  effect  tested.  composite the  High  frequency  d i d appear t o produce low  testing  of  of  The frequency the  s e v e r i t y of f a t i g u e damage i n the Boron/Aluminium  samples  in  the  extent  composite  Boron/Aluminium  s u b s t a n t i a l l y l e s s c r a c k i n g than  frequency c a s e . i . e . f a t i g u e damage appeared  t o be  more d i s p e r s e d at h i g h f r e q u e n c i e s and more l o c a l l y c o n c e n t r a t e d at low  frequencies  combined  in  different  modes a t d i f f e r e n t response  and  for  that ways  material. t o produce  The  the d i f f e r e n t  f r e q u e n c i e s a l s o produced  intermediate  failure  phenomena  different  which  fracture fatigue  e v e n t s . S t a t i c and dynamic  s t i f f n e s s of the sample was even reduced by 50% i n some c a s e s . T h e r e f o r e i t was concluded t h a t influence  or  even  control  f a i l u r e modes such as f i b r e and matrix c r a c k i n g .  frequency  of  cycling  can  the i n t e r a c t i o n and combination of breakage,  debonding,  delamination  7  1 . 2 . 3 E f f e c t Of M a n u f a c t u r i n g On F a t i g u e The  defects  production  was  of  early  of  degree p l i e s ;  voids  thicknesses  the in  fibres at  especially  of  a  .region  specimen  with  would  irregularities  be in  few  defects,  high;  revealed  if  at  The photos a l s o  revealed  non-uniform  a r e a s i n the 6 0 non-uniform plies.  of  The  local  static were  effect  located  strength  many  lower s t a t i c  ply  failure  the notches were  the  the notch r e g i o n  notch.  exposed  defects;  If  in  fatigue  sequence  there  a  that  the 0-degree outer  failure.  notched  the  automatically  interfaces;  these would be to randomize the  fatigue  adapt  of  many r e s i n - r i c h ply  the  to  inside  and  the  e v e n t s which l e d to g e n e r a l in  the  the t e s t .  with  in  the  developed f o r m e t a l s  tests,  irregularities  distribution  role  during  worked on doubly  1 3  previously  observing  during  developed  manufacturing  of  large  photomicrographs,  intervals  introduced  specimens  technique,  camera  Examination  cracks  a  graphite/epoxy  s  the  periodic  have  of the m a t e r i a l . P a p i r n o  photomicrographic which  irregularities  of composites  performance [0/-60/60]  and  of  defects  the and  strengths  were  expected. Therefore fatigue material  life as  applications.  there  would  of s i m i l a r fabricated  be  scatter  samples. P a p i r n o s ' is  a  poor  He a l s o suggests that  performed t o screen a g i v e n time-consuming and expensive  lot  of  in  failure  data  and  argument i s t h a t  this  candidate  static  for  notch t e s t s  composite  structural s h o u l d be  material  f a t i g u e programs are  before  initiated.  8  1.2.4  E f f e c t Of  Stacking  S e v e r a l authors have sequence tensile  on  fatigue  strength  laminates  groups  were  formulation,  the  rigorous  predicted  in  layer  regions,  behavior  The  regions  of  delamination  be  angle-ply  was  the  theory in  curve.  above phenomenon i s since  symmetric  triggered  portrayal  significant  by  identify  progressive  this  composites  arrangement.  of  interlaminar  the  possibility  the  delamination  i n boundary  that the  degradation  these i n t e r l a m i n a r  stress  stresses  are  unique  caused  by  s t r e s s e s . Severe  d e l a m i n a t i o n s have, i n f a c t , been witnessed by Foye who  in  i n [Reference 15,16] i n d i c a t e  realistic  to  and  pronounced  independent of s t a c k i n g  attributed  the  [±15°,±45°]  a  (LT),  stresses  It i s t h e r e f o r e a strong can  1  remote from a boundary, i t f a i l s  where  stacking  B a k e r " d e a l t with  effect  s o l u t i o n s presented  field  of  the p o s i t i o n s of the ±15°  lamination  that while LT g i v e s a very  developed.  combined  explanation  of  Strength  influence  throughout the e n t i r e S-N  under membrane l o a d i n g are More  the  which  reversed.  theoretical  beyond the scope  of  in  d i f f e r e n c e in s t r e n g t h The  studied  laminate s t r e n g t h . Faye and  boron-epoxy ±45°  Sequence On Laminate  and  Baker,  as the primary source of  strength degradation in f a t i g u e . Another dependence on adjacent or at an type  are  possible stacking  layers  on  mechanism  which  can  explain  sequence i s the c o n s t r a i n i n g  the propogation of a crack  i n t e r f a c e . A n a l y t i c a l r e s u l t s for p r e s e n t e d by Chen and  Sin  1 7  a  strength  i n f l u e n c e of  i n a given problem  of  layer this  f o r a laminate c o n s i s t i n g  9  of i s o t r o p i c l a y e r s . According  to  possibililites the  interlaminar should be  and  Pipes  1 8  there  f o r o p t i m i z a t i o n of laminate  stacking  interlaminar  Pagano  order. tension  The  stacking  are  several  s t r e n g t h by  arrangement  in the f r e e edge zone and  varying  should should  avoid  minimize  shear r e s u l t a n t s . I n t e r l a m i n a r compressive s t r e s s e s  i n t r o d u c e d to minimize the d e t r i m e n t a l e f f e c t  shear  s t r e s s e s . In the case of a [ 0 / 9 0 ] b i d i r e c t i o n a l  i t was  argued t h a t p u t t i n g 90°  give a s t r o n g e r  laminate.  l a y e r s at the o u t e r  of  the  laminate,  s u r f a c e would  10  1.2.5  Micromechanisms Of F a t i g u e  Prakash His  1 9  s t u d i e d the tension-compression f a t i g u e of  e x p l a n a t i o n f o r crack propogation i n CFRP i s as During  polymer  temperature  occurs.  Generation  of heat  the sample becomes g r e a t e r than the that p o r t i o n , because cracks,  dissipation  delaminations  dissipation  low  of  heat  on  a  rise  level.  This  which  the  i n temperature  permits  rate  fibres, of  heat  specified  ( s t a t e d simply: l o c a l h e a t i n g  With  occurs l e a d i n g to a  which leads to a f u r t h e r  local  causes  in  from  f i b r e s which are good thermal c o n d u c t o r s .  accumulation,  a  rise  i n c e r t a i n p o r t i o n s of  temperature. A f t e r some time the matrix shear  to  local  there are r e s i n - r i c h areas, broken  h i g h e r damping w i t h i n the specimen in  a  i s l e s s than i n the r e g i o n s c o n t a i n i n g the  amount of carbon heat  follows:  f a t i g u e , heat i s generated due to h y s t e r e s i s of the  m a t r i x . Heat accumulation l e a d i n g to  voids,  CFRP.  rise  modulus  falls  b u c k l i n g of the  fibres  local  softening  of  the  matrix and a l l o w s adjacent f i b r e s to buckle n o r m a l l y ) . As carbon fibres slot  are very b r i t t l e ,  (or c r a c k ) i n  itself  a  bad  the  the b u c k l e d f i b r e s break and produce a composite.  conductor  accumulation.  Furthermore,  elevation  stress  of  to  of  heat, broken  give  results  any more s t r e s s .  when  newly  causes  created  slot,  f u r t h e r l o c a l heat  fibres  produce  enough  r i s e to f a i l u r e of the adjacent  f i b r e s , which l e a d s to a deepening failure  This  of the c r a c k .  Finally  the remaining c r o s s - s e c t i o n cannot  total carry  11  The cumulative f i b r e accordance  with  the  f a i l u r e process d e s c r i b e d above i s i n  fatigue  failure  criterion  suggested by  Hashin and R o t e m . According t o t h i s c r i t e r i o n ,  i f 9 represents  the angle  between  and  direction  (for unidirectional  20  the  direction  of  loading  material)  for  the  fibre  0°^|(9<|2°,  the  specimen f a i l s by the process of cumulative f i b r e f a i l u r e ( i . e . fibre  f a i l u r e mode). For l a r g e r v a l u e s of 6,  that  of crack growth through the matrix, p a r a l l e l to the f i b r e s  the f a i l u r e mode i s  ( i . e . matrix f a i l u r e mode). Whitcomb and ply  21  [45/90/0]  i n v e s t i g a t e d f a t i g u e damage type  graphite/epoxy  c r a c k i n g were found to be  damage.  In  the  in  Off-axis graphite/epoxy  fibre  static was  dominant  types  and  fatigue  behavior  failure  angles  between  characteristics  0°  and  i n the form of  serrations  cracks.  Large  in  life  22  AS-3501-5A i n an e f f o r t  failure. 90°  Seven  were t e s t e d .  or v i s i b l e  damage.  v a r i e d with o f f - a x i s angle and  appeared  scatter  fatigue  i n t o adjacent  of  s t u d i e d by Awerbuch and H a h n  F a t i g u e f a i l u r e o c c u r r e d without any warning Matrix  of  orientation.  to c h a r a c t e r i z e the m a t r i x / i n t e r f a c e - c o n t r o l l e d off-axis  [0/45]  l a m i n a t e s . Delamination and  g e n e r a l , p l y c r a c k s d i d not propagate  p l i e s of d i f f e r i n g  different  notched  and was  axial  and  transverse  observed at a l l o f f - a x i s  a n g l e s . Micrographs  of f r a c t u r e s u r f a c e s showed a combination of  several  f a i l u r e modes such as f r a c t u r e of i n d i v i d u a l  fibres  independent  and f i b r e tows, matrix s e r r a t i o n  cleavage  and  (shear f a i l u r e ) ,  matrix/interface cracking p a r a l l e l  matrix  t o the f i b r e s .  12  M a t r i x s e r r a t i o n s i n c r e a s e d with l o n g i t u d i n a l shear the  absence  of  the  shear  s t r e s s , a cleavage  stress.  In  type of f a i l u r e  p r e v a i l e d . F r a c t u r e s u r f a c e s which c o n s i s t e d of matrix-dominated and  interface-dominated  sudden  death  immediately serrations  r e g i o n s were  behavior before  and  the  was  attributed  final  microcracks  not  in  fracture. the  planar. to  the  r a p i d c r a c k growth  The  matrix  Finally  appearance  of  material possibly  c r e a t e d h i g h s t r e s s c o n c e n t r a t i o n s along the f i b r e s c a u s i n g them to f a i l  along some weak s p o t s .  13  1.2.6  Damage Zone D e t e c t i o n Methods And  Crack V e l o c i t y  Studies Mandell  and  McGarry  graphite/epoxy  specimens  crack v e l o c i t y  the  conductive  layer  studied  23  for  specimen and  static was  silver  with  recorded  when each s t r i p was  seemed  to  method can  an  be  so  with  a  non-  but  in  that a drop i n v o l t a g e  broken by  evaluation  p o s i t i o n of the crack  while  coated  the  crack.  The  a was  system  i t i s not c e r t a i n whether  this  tests  (NDE)  used  24  technique  in graphite/epoxy  . The  NDE  monitoring  a  to  samples  was  modified  x-ray  observe during  the  static  conducted i n r e a l  time  the f r a c t u r e specimens were under t e n s i l e ramp l o a d i n g  constant  the  be used f o r f a t i g u e t e s t s .  non-destructive  cyclic  in  To determine  s t r i p s were connected  Chang, Gordon and h i s c o - w o r k e r s  and  tests.  velocity  s t r i p s were then p a i n t e d ahead of  oscilloscope,  successful,  crack  first  the notch at s e v e r a l i n t e r v a l s . The circuit  the  amplitude  cyclic  l o a d i n g . Tetrabromoethane (TBE)  and was  a p p l i e d as an x-ray opaque a d d i t i v e at the t i p s of a s l i t  i n the  c e n t e r of the  Damage  initiaton, sequences similar  specimens  growth of  and  x-ray  studies  in  to  enhance  failure  pictures composites  the  mechanisms recorded  image.  were observed  during  testing.  from Some  have been l i m i t e d to r e a l - t i m e  s u r f a c e damage s t u d i e s or p o s t - f r a c t u r e TBE  flaw  evaluation,  but  using  gave a b e t t e r chance to observe the e n t i r e damaged r e g i o n .  14  Sendeckyj  and  Maddux  made  25  s i d e - b y - s i d e comparisons  damage i n d i c a t i o n s obtained by u s i n g TBE  enhanced x-ray,  t r a n s m i s s i o n u l t r a s o n i c C-scan, and h o l o g r a p h i c inspection  methods  of  various  photography  nature and  gave  planar  the  through  non-destructive  composite specimens c o n t a i n i n g  d i f f e r e n t amounts of damage. I t was x-ray  of  concluded  most  distribution  t h a t TBE  detailed  of  enhanced  i n f o r m a t i o n of  damage  in  the  graphite/epoxy  composites. Holography  using  thermal  l o a d i n g showed d e l a m i n a t i o n s  c r a c k s i n the s u r f a c e p l i e s and matrix  cracks  providing of  the  and  fibre  delaminations,  methods  for  u l t r a s o n i c c-scans without  giving  fractured  fractures  but  that  in  of  finding  Kunz  gave  the  be  and  be  done i n  selection  of  Through-transmission  either x-ray  picture used  distribution  to  extent  on of  best  method  anomalies  planar  results the  needs  and  in  of  delamination  matrix  c r a c k s or  radiography  using  of damage, so i t was studying  damage  in  composites.  and  graphite/epoxy compressive  The  work  both  the s u r f a c e p l i e s  resolution.  information  this  graphite/epoxy  pattern  showed  fibres.  more  damage  any  tetrabromoethane advised  capable  i n f o r m a t i o n on the t h r o u g h - t h e - t h i c k n e s s  i n t e r p r e t a t i o n of f r i n g e loading  was  and  Beaumont composites  loading.  investigated  26  They  in  notched  used  crack  beams  cross-ply  and  extension under  in  cyclic  unidirectional  specimens and observed  that crack propogation  of  a c c e l e r a t i o n . Fatigue crack extension i n  deceleration  and  underwent  periods  15  u n i d i r e c t i o n a l composites frequently  resulted  was  in  a result crack  composites  had lower compressive  fractures  resulted  cracking  from  of  axial  arrest.  Crossplied  s t r e n g t h s and  and  [0/90]  s p l i t t i n g of the 0° p l i e s , t r a n s v e r s e  i n the 90° p l i e s and delamination between the p l i e s . were  drawn  [0/90] samples u s i n g a t r a v e l l i n g microscope  1). The curves showed two corresponded Region  and  through-specimen  P l o t s of crack l e n g t h vs number of c y c l e s [0]  cracking  to  initial  II represented  stages  of  crack  growth.  for  (see f i g . Region  I  crack a c c e l e r a t i o n , then d e c e l e r a t i o n ;  crack  reacceleration.  Crack  growth  in  [0/90] composites  f o l l o w e d the same p a t t e r n as i n u n i d i r e c t i o n a l  composites,  it  but  was  more r a p i d i n the former. P a r i s  were then drawn showing l a r g e s c a t t e r of  the  Paris  samples had s t r a i g h t  Most  l i n e s and negative s l o p e s i n t h e i r  plots. It was  the  i n crack v e l o c i t i e s .  plots  noted that d u r i n g c y c l i n g corresponding to Region  crack that had  I  i n i t i a t e d at the machined notch grew r a p i d l y  under the a p p l i e d s t r e s s and then v i s i b l y slowed down. Continued c y c l i n g produced fracture  c r a c k s that o r i g i n a t e d  surface  and  grew p a r a l l e l  d i r e c t i o n . A f t e r some time c r a c k s fibres Region  with a v i s i b l e  at  the  newly  to the f i b r e s  proceeded  again  created  i n the a x i a l across  the  i n c r e a s e i n v e l o c i t y which corresponds to  I I . Cracks e i t h e r propogated  through the  specimen  f a i l u r e o c c u r r e d or d e c e l e r a t e d a g a i n and remained  until  arrested.  16  1 . 2 . 7 F o r m u l a t i n g The Decay In Poursartip,  Ashby  accumulation during was  monitored  tensor  fatigue  Beaumont  Fatigue  studied  the  damage  of composite m a t e r i a l s .  The  damage  27  by measuring the e l a s t i c m o d u l i , because b e i n g a  of the f o u r t h  distinguishing They  and  S t i f f n e s s During  and  rank,  the moduli o f f e r  monitoring  r e p r e s e n t e d the v e l o c i t y  different  the  possibility  components of damage.  of damage a s :  dD/dN = f ( A o , D ) where  D is a variable  of  (1)  which r e p r e s e n t s  Ao* i s the s t r e s s a m p l i t u d e and  f a t i g u e damage, N  is  the  number  of  cycles.  The above equation can be i n t e g r a t e d N  where  D  and D  ±  f  are the  f "  initial  t o g i v e the f a t i g u e  L_i5  life:  () 2  and f i n a l  amounts of damage  respectively. To  evaluate  experimental  dD/dN  measurement  The d a t a were S m i t h ' s decay d u r i n g found damage  to  fatigue  and  with in  curves a l l  an  of  stiffness  equation versus  which r e v e a l e d three  life.  The i n i t i a l  constant the  w i t h s t r a i n amplitude and stiffness  used  be both s t r e s s and s t r a i n  phase  dependent,  2 8  they  final had  number of  regions  of  damage growth  dependent. The  damage  based  growth  was  stress  f o l l o w e d an e x p o n e n t i a l  cycles.  stiffness phase  was  intermediate only  p h a s e , the damage r a t e little  on  stress  increased  dependence.  form. See F i g  The 2.  17  In Beaumont  agreement with the method of P o u r s a r t i p et a l , Kunz and 26  found that the i n c r a s e i n c r a c k l e n g t h  same p a t t e r n as the change in s t i f f n e s s d u r i n g  followed  fatigue.  the  18  1.3  Theory  While  studying  researchers  have  the  used  fracture Linear  of composite m a t e r i a l s some  Elastic  equations  directly,  as i n the  form of  materials,  and some o t h e r s e i t h e r  Fracture  Mechanics  b e i n g used f o r  converted  isotropic  these equations  for  composite m a t e r i a l s or c r e a t e d new e q u a t i o n s . 1.3.1  F r a c t u r e Toughness  Initially, for  it  was  the present p r o j e c t .  this  k i n d of geometry  Equations  d e c i d e d t o use Compact t e n s i o n samples The  fracture  toughness  from ASTM s t a n d a r d s  =  IC  where,  P=applied l o a d ; a=crack  (See f i g .  3) In other  ic  K  In  2 9  "  fracture  3  length;  w=width;  for  3 0  _i_^[29.6-185.5(a/w)+655.7(a/w) -1017.0(a/w) +638.9(a/w) wt 2  K  is:  equation  ]  A  (3)  t=thickness.  words,  - -!r < f  mechanics, s t r a i n  a / w  >  <*>  energy  release rate  is defined a s :  « where,  5 C / 6 a  For p l a n e  G  I  C  -  <«  i s the change i n c o m p l i a n c e w i t h c r a c k  length.  stress,  ^  ( 6 ) , hence  K*  =  m  £(§f)  E  (7)  where E i s a the homogenous n i s o t rYoung's o p i c m a tmodulus e r i a l the ofr e lthe a t i om n a tiesr i a l , b u t , 2 *  G  To use t h i s the  principal  straight  1-1  ic • ic K  in  1/2  te)  11/2 2 b +  12 2  + b  b  66  -  axes  of  m a t e r i a l symmetry such t h a t  nature  and  the  system  the  elastic  must  (8)  ll J  equation the c r a c k has t o c o i n c i d e with one  ahead. In t h i s c a s e ,  orthotropic  n  for a  it  extends  necessarily  coefficients  of  be in  19  e q u a t i o n 8 can be expressed elasticity  and the p o i s s o n ' s 1  1  E,  11  where  1  "  V  2 12  subscripts  respectively; E,=E  2  for  i n terms of  b  22  I  =  ( 1  2  ratio:  V  a [90/0]  which  12  ratio,  x,y  c a l l e d the  to: E' =  b  i  'Effective  fh  b  describing  directions (Note  that  +b  2 b  E'  for  composite  and i s  equal  n  1} V  I -n(fi) ' E  fatigue  Stress  <9  are  the  maximum  and  minimum  min  the  following  terms  are  important  in  cycle.  Amplitude,  o = (o a  Mean S t r e s s , Stress  »  Equations  and a  a  respectively, the  z  12  1 L/ /2 2 ^ - l  1 2 6 6T  b  max  stresses  b  Modulus'  ov ^ 2b  +  V ll/  1.3.2 Fatigue Testing if  23>' 66  modulus.  r e p l a c e d by  .2  cycling  V  and  /i=shear  1/2 T/v \ _22\ ll 22j  and  In  1+  composite.)  is  K  <  b  1,2 and 3 are p r i n c i p a l  v=poisson's  modulus of  3 5  " 23 >» 12  The term E i n e q u a t i o n 7 i s material  the p r i n c i p a l  Ratio,  -a max  a = (a m  +a max  R = a min  S t r e s s Range,  min  0  max  min  /o max  a min  )/2 )/2 -  -  -  -  (10)  20  To draw an S - l o g N graph e i t h e r Usually  crack length  drawn  to  determine  during  fatigue.  (a)  this  The slope of  relation  relationship  number  can be used as S .  max  of  cycles" plots  the r a t e and behavior of c r a c k  the a p p l i e d s t r e s s and the metals  vs  Aa or a  the curves crack  assumes  (da/dN)  length. the  form  propogation  is a function  Quite  3 6  are  of  a  often,  for  simple  power  wherein: da  m n  (ii)  _  • a & - o a - - - - - where m  2 - 7 , n =* 1 - 2  the  overall  ^ '  3 6  P a r i s p o s t u l a t e d that was  of  the  'stress  controlling  propagation process. This | | .  intensity  factor  relation A A K ° -  -  in  factor  the  range'  fatigue  crack  is -  where A i s a c o n s t a n t and AK=Stress  -  -  -  instensity  -  -  -  factor  (  1  2 )  range=K max  K  min  where K  max  calculated  i s c a l c u l a t e d u s i n g the max. l o a d  using  e q u a t i o n 9 can  the  be  min.  used  for  tensile these  load.  Either  calculations  and  K  min  equation 3 or while  using  composite m a t e r i a l s . A l s o the c o n s t a n t s A and m are f u n c t i o n s material stress  variables,  environment,  frequency,  is  temperature  of and  ratio.  The P a r i s p l o t u s u a l l y argued that to the i n d i v i d u a l  i s u s u a l l y drawn in l o g - l o g the crack growth r a t e  v a l u e s of K  or R max  min  form and  (da/dN)  is  it  is  sensitive  , and the s t r e s s  ratio.  21  1 . 3 . 3 Compliance T e s t s It  is  necessary  samples in order  to  measure  the  compliance of the  to be a b l e to use equations (7)  sample i s t e n s i l e  loaded to a c e r t a i n  and ( 9 ) .  f r a c t i o n of  its  test  A test fracture  l o a d and the d e f l e c t i o n of the sample along the l o a d i n g l i n e can be  measured  Loading material.  using  should The  various  be  methods  performed  inverse  the  using a c l i p - g a g e ) .  elastic  range  of  slope of the produced l i n e w i l l g i v e  compliance of the sample f o r  that notch  The t e s t can be c a r r i e d out on  in  (e.g.  the the  length.  f o r a number of  notch  lengths  the same sample and the compliance vs notch l e n g t h graph can  be drawn. The s l o p e of equations  (7)  and  (9).  t h i s curve g i v e s  (6C/6a)  which i s used i n  22  1.4  Purpose  As can be seen used d i f f e r e n t  from the l i t e r a t u r e  techniques t o determine  g r a p h i t e f i b r e r e i n f o r c e d composites. was  decided to determine  graphiteepoxy microscope,  composites  equipment i t was o n l y . The  In the present work, i t and damage i n  d u r i n g f a t i g u e , using  radiography and compliance  (See F i g . 3). Due  the f a t i g u e damage i n  the crack v e l o c i t y  t e n s i o n sample' geometry was purpose  survey, r e s e a r c h e r s have  found  travelling  techniques. The  'compact  t o be s u i t a b l e f o r t h i s  to the l i m i t e d  capability  of the  decided to work on t e n s i o n - t e n s i o n f a t i g u e  laminate having  the c o n s t r u c t i o n [ 90/3 j . was 8s  because t h i s  i s a simple  laminate to make and y e t , i t  approximates  the p r o p e r t i e s of more complex  laminates.  used,  23  Chapter  II  PROCEDURE 2.1  Laminate And  Sample P r e p a r a t i o n  The H e r c u l e s AS3501-6 h i g h s t r e n g t h graphiteepoxy m a t e r i a l was corresponds  used  i n the p r e p a r a t i o n of the l a m i n a t e s .  to h i g h s t r e n g t h f i b r e s and  epoxy matrix was  r a t e d so that  prepreg 'AS'  '3501-6' means t h a t  the  i t m a i n t a i n s i t s mechanical  p r o p e r t i e s up to 350°F. The fibres to  prepreg m a t e r i a l was  i n the form of a r o l l  running along the l e n g t h of the r o l l .  prepare  intention  was  [90|0]B* type of laminates where 90° i s the d i r e c t i o n  p e r p e n d i c u l a r co the l o a d i n g l i n e and l i n e . The  The  with the  u n i d i r e c t i o n a l prepreg was  rectangular  90° and  0° i s along the l o a d i n g cut i n t o  sixteen  s i x t e e n 0° p l i e s which would be  together and then cured to produce A f t e r the lay-up procedure was  joined  the l a m i n a t e . f i n i s h e d by s t i c k i n g  the  p l i e s together f o l l o w i n g the above p l y sequence, a r e l e a s e f i l m was  a p p l i e d to both top and bottom of the laminate and a c a u l  plate  f o l l o w e d by bleeder p l i e s were stacked on top. The  p l a t e was  used  to produce  the b l e e d e r p l i e s was of  caul  a smooth s u r f a c e and the f u n c t i o n of  to c o l l e c t  the epoxy that was  the p l i e s d u r i n g the heat treatment. The  squeezed  laminate was  then  vacuum bagged onto an aluminum base p l a t e . Heating tapes were p r e s e n t under the p l a t e was  i n order to heat the laminate when i t  i n s i d e the a u t o c l a v e . (The d e t a i l e d o p e r a t i o n and  the  out  24  apparatus  i s described  assembly and  was  then  pressure  was  heated  i n the  recommended by  in detail  autoclave  Hercules  applied during  the  end  error  was  allowable during  the  tension  samples  was  No  this  cooled  removed.  (fig.3)  using  the  sample  a  After sharp at  crack  the  clip  cutting was  beginning  g a g e on  tensile  loading  About twelve 4(ii).  the  seven  a  of  the  sample  and  i n each  using notch  prepared  i n the  bag  cycle  100  of  80  p s i was  assembly  then  cut  psi applied  used  shown  a  (A  ±10°F  around  blade. t o be  the A  saw.  The During  the  i t s deflection  put holes.  diamond  groove  able  out  compact  m a t e r i a l was  sample w i t h  in order  to measure  into  taken  in f i g . 4(ii).  plastic  razor  was  diamond  introducing cracks  notch  produced  cure  1.25mm t h i c k  C.T.S., a  to avoid  vacuum  p o s t - c u r i n g was  I t was  of  under  The  A pressure  down t h e  drilling  the  37).  operation).  code numbers as  holes  the  f i g . 4(i)).  samples were g i v e n the  of  following  treatment.  autoclave  laminate  presence  240°F t e m p e r a t u r e  the  and  the  (see  until  When t h e  of  i n the  in reference  was  a  cut  to place during  saw  a  a  (fig.3). laminates  samples c o u l d  be  cut  were p r e p a r e d from  each  i n the  laminate  above manner as  shown  and  in f i g .  25  2.2 S t a t i c T e n s i l e F r a c t u r e In  accordance w i t h the a d v i c e of P a p i r n o i t 3  determine the f r a c t u r e  toughness of  of the l a m i n a t e . The d i r e c t i o n whether  Testing  the  crack  samples from  was d e c i d e d to various  parts  of the notches were v a r i e d to see  direction  a f f e c t e d the toughness  (see  fig.  For  the  4(ii)). Laminates 2 and 4 were former  laminate  the  used  notches  for  this  were cut by a diamond saw and no  sharp c r a c k s were i n t r o d u c e d by the razor c r a c k s were intoduced to the samples of The  static  tensile  tests  I n s t r o n t e s t i n g machine with a  purpose.  were load  blade  whereas  sharp  laminate 4. performed  cell  of  by  0-8896N  using  the  loading-  range and a c r o s s h e a d - d i s p l a c e m e n t r a t e of 0.25mm/sec. A t y p i c a l loading  curve  was r e c o r d e d f o r  is  shown in f i g .  each sample.  5. The maximum l o a d to  fracture  26  2.3 Fatigue Testing The MTS s e r v o - h y d r a u l i c tests. not  machine was used  designed was  to  apply  used f o r  compression  the s i x  changed t o l o a d c o n t r o l  for  fatigue  samples of  the i n c r e a s e in the crack l e n g t h s . gave very s h o r t  fatigue  f r o n t of  t e s t e d without  with  accuracy  were  of ±0.2mm.  there were u s u a l l y two c r a c k s running the l a r g e s t the  placing  crack was f o l l o w e d .  samples the other  it  was  watching  life.  The crack l e n g t h s an  Stroke  The s t r o k e c o n t r o l l e d samples  microscope was p l a c e d  the t e s t i n g assembly t o watch the c r a c k  the t e s t s .  intervals  loading.  laminate 1, but  S t a r t i n g with laminate 3 a t r a v e l l i n g  of  fatigue  the remaining samples.  Samples of laminate 1 were f a t i g u e  during  the  Only t e n s i o n - t e n s i o n l o a d i n g was used as the machine was  control  in  for  were  also  propagation  recorded It  from  at  certain  was d i s c o v e r e d the  notch.  Mostly,  S u r f a c e c r a c k s on the other  checked  by  that  side  s t o p p i n g the t e s t and  s i d e of the sample i n f r o n t on the  travelling  microscope. To  observe  the  damage zone i n s i d e the sample which c o u l d  not be seen by the naked eye, a r a d i o g r a p h i c the  crack  technique was u s e d .  Tetrabromoethane was put  into  tip  by  an  d u r i n g the f a t i g u e t e s t .  Due to the dynamic motion of  injector the sample  the l i q u i d p e n e t r a t e d i n t o the damage zone. The t e s t i n g was then stopped  and  the  sample was removed from the MTS and p l a c e d i n  f r o n t of a P h i l l i p s x - r a y machine equipped with a copper  target.  27  The x - r a y machine was then a d j u s t e d to A  circular  column of x - r a y s  image of  the damage zone  sample.  The  The  films  the  a  film  x - r a y s were a p p l i e d for  were  and 6 m i l l i a m p s .  the sample p r o d u c i n g placed  just  the  behind  the  15 seconds and d u r i n g  this  preferentially  then d e v e l o p e d ,  made. Crack l e n g t h s or  irradiated  on  s h o r t time tetrabromoethane  16 k i l o v o l t s  absorbed the  f i x e d and p o s i t i v e  were then measured e i t h e r  x-rays.  prints  were  by u s i n g the  films  prints.  After  this,  measuring  its  put back  into  the sample was s t a t i c a l l y  deflection the  MTS  to determine assembly  its  for  frequency  of  frequencies  gave r i s e  Around  mm  20  s e e i n g that changed  10Hz  was  tests  increase  in the  notch l e n g t h was used in the f i r s t  24.5±1mm  load  to enable the f a t i g u e  was This  each sample.  the  (under  while  fatigue.  times for in  to a temperature  no sample was f a i l e d  to  used  tested  compliance and i t  further  d e s c r i b e d procedure was repeated s e v e r a l A  tensile  as  higher  samples.  laminate,  control)  it  but was  c r a c k s or damage to  grow. Some samples were s u b j e c t e d fracture whereas  load other  (e.g.  loads  samples  (e.g.4500<Load<4700N) fatigue could  tests  low  fail  loads  5  for  to  The longest  long  period  of  s  to  their  medium  (Load<4500N).  which f a i l e d at 7 . 5 x 1 0 a  close  or equal to 4700 N)  subjected  run from two to nine d a y s .  not  than  ran for about one or two d a y s ,  taken by sample C did  greater  were and  to h i g h loads  High  whereas the p e r i o d of  cycles. time,  loads  Some  load  others time was samples  so they were not  28  fatigued  further.  Sample E  6  was t e s t e d with a r e l a t i v e l y  high minimum l o a d  and a high maximum l o a d (which was c l o s e to the f a i l u r e range)  in order  or s t r e s s range  to see whether graphiteepoxy sensitive.  load  i s maximum s t r e s s  29  2.4 Compliance T e s t i n g The  main  determine  aim  the  in  using  'effective  the  crack  compliance length'  i n c r e a s e i n compliance of the samples The  and  during  range  to  was  testing.  t h i s purpose.  with  to  measure the  fatigue  I n s t r o n or the MTS machines were used f o r  I n s t r o n was used in the 0-4448N load  method  a  The  crosshead-  d i s p l a c e m e n t r a t e of 0.25mm/minute whereas the MTS was used w i t h the  same l o a d range and displacement r a t e w h i l e  under s t r o k e  was o p e r a t e d  control.  The sample tensile  it  was  loading  placed  in  the  was a p p l i e d while  machine  and  d e f l e c t i o n was recorded by a  c l i p - g a g e which was p o s i t i o n e d on the sample ( f i g . 3 ) . gage  incremental  The  clip-  measured the d e f l e c t i o n at the b e g i n n i n g of the n o t c h .  'Thales t r i a n g l e '  r e l a t i o n was used to f i n d out  at  line.  the  loading  t a k i n g the elastic  inverse  The  slope of the  procedure  sample (see  fig.7).  their then  (equations  load-deflection  curve  in  the  7  compliance  to  9)  'effective  value  s e v e r a l notch l e n g t h s w i t h i n  lengths  used and  times to a sample from the  The c a l c u l a t e d c o m p l i a n c e s were then  notch be  determine the  easier,  was then c a l c u l a t e d by  was a p p l i e d s e v e r a l  each l a m i n a t e by c u t t i n g  could  deflection  range.  This  versus  compliance  the  The  were  (fig.  8).  determine  of  the  crack l e n g t h ' known.  To  plotted  The slope of t h i s  the  sample  same  fracture or  it  curve  toughness  c o u l d be used t o  of another sample i f  make the l a t t e r  compliance of the r e f e r e n c e sample  was  its  calculation  plotted  versus  30  the  cube  of  the  c e r t a i n notch l e n g t h all  (see f i g s  giving a straight  there was a s c a t t e r  same notch l e n g t h s .  to  the  notch l e n g t h and  C  with  (a)  length  sample  A  of  a  compliances C, and C  2  compliance  crack length' C,.  due  to  the  f o r other  c o u l d be c a l c u l a t e d by the new l i n e  The same i s true f o r  sample C .  belongs 0  at  the B  lengths.  increase its  for  samples  with the same notch  damage i n t r o d u c e d ,  6 was  specimen  w i t h a compliance v a l u e C  p a r a l l e l s were drawn to i t  it  samples w i t h  the middle curve  Hence, when sample B was f a t i g u e d and there was an its  tests  Then the method d e s c r i b e d i n f i g .  the compliance was known. If  reference  fatigue  i n the values f o r  used to compute the apparent c r a c k which only  l i n e up t o a  6 and 9 ) . When the c o m p l i a n c e s of  the samples had been measured b e f o r e the  was seen that the  notch l e n g t h ,  in  'effective  starting  from  31  2.5  D e t e r m i n i n g The Weight And Volume F r a c t i o n Of The  Fibres After  all  the  tests  were f i n i s h e d  s m a l l p i e c e s were  cut  from a sample of each l a m i n a t e . These p i e c e s were then b o i l e d concentrated N i t r i c resin  from the f i b r e s .  several  times  temperature. could  and  100°C f o r  an hour  The f i b r e s were then  dried  in  a  furnace  washed at  t o the r e a d i n g s due to presence of  of the  a relatively  ' f w  a n c  ^  w  a  r  e  t  m  and weight of  *  c  matrix  f  w  e  9  i  n  t  o  the matrix  P„  where p , p  i e  - P v f  c  and p  -  -  £  fibres  scale  Errors  some  low which  c o u l d be  resin  left  fibres:  w=w +w c r m c  water  1 3 , 1 4 , 1 5 and 16 were used to determine the weight  volume f r a c t i o n  w  with  fibres.  Equations  where  to remove the  The f i b r e s were weighed by a s e n s i t i v e  between the  f  t  + P V  f  respectively.  V  e  and V  -  -  -  -  c o m p o s i t e , wt.  -  -  of  the  -  -  (13)  respectively. -  -  mm  are the d e n s i t y  m  n  -  epoxy  i r  £  and  half  read the weight to an a c c u r a c y of ± 0 . 1 m g .  introduced  and  a c i d at  in  -  -  -  -  of the composite f i b r e s  are volume f r a c t i o n  of the  (14)  and fibres  m  matrix. Equation  14 can be r e w r i t t e n a s : p  To  determine  c  =  p  the  f f V  +  p  m -V  -  ( 1  volume  -  fraction  -  -  of  -  the  -  -  fibres  n e c e s s a r y to know the volume of the composite and d e n s i t y epoxy.  Then, v  m  m m  =— Pm m  ,  v f  m_ £  r  =Pft —  (15)  it of  is the  32  and  v where v i s  c  =v  c  f  + v  m  the a c t u a l volumes and V_ =v./v r  The d a t a are given i n t a b l e  V.  i c  33  Chapter  III  RESULTS 3.1 T e n s i l e F r a c t u r e And Load Curve The f r a c t u r e samples  from  Laminate  respectively. was used f o r No  l o a d and  cracks  (with  scatter  Sample  L  2  sharp  K  might  seemed  2  were  4  of  (eqn.  introduced  cracks)  be  of  the  samples  loads were h i g h e r  than  from  4893N  to  3)  of the  9118N.  c o n s i d e r e d e x c e p t i o n a l as no other sample l o a d . Samples A ,  B ,  2  have low f r a c t u r e  C ,  2  2  loads with respect to 2  which were a l s o on the edges were not so weak  2  II  Laminate 4. A l s o there was a  loads ranging  l e v e l of f r a c t u r e to  into  samples near the c e n t e r of the l a m i n a t e , but samples D , and I  the  are shown i n T a b l e s I and  the f r a c t u r e  in fracture  approached that and  and  values  the toughness c a l c u l a t i o n s .  Laminate 2. As a r e s u l t  large  2  toughness  The ASTM compact t e n s i o n sample e q u a t i o n  sharp  samples  fracture  Analysis  F , 2  (See  J  2  the G , 2  fig.  4(ii ) ) . The d i r e c t i o n PS  and  load.  QR  of the notches f o r  of the laminate d i d not seem to e f f e c t  I n i t i a l damage at the t i p of  close  to  the samples near the edges  the  center  of  the notch  the sample whether  the  would  fracture  start  very  the notch was cut  towards the c e n t e r of the laminate or towards the edge. Fracture scatter  as  toughness v a l u e s f o l l o w e d the sample the f r a c t u r e  dependant to the f r a c t u r e  l o a d of  the samples as K  load ( i t  i s the h i g h e s t  pattern J C  is  and  strongly  parameter  in  34  eqn.  3). In  laminate 4 there was l e s s s c a t t e r  i n the f r a c t u r e  which gave some encouragement f o r  the f a t i g u e t e s t s .  was  results  ±670N.  reasonable  Fracture except  toughness  that  they  Laminate 2 perhaps due to the Also,  the  edge  samples  c e n t e r ones i r r e s p e c t i v e  have  were  plate,  with  the  lower  presence  seemed  scatter to  be  than the r e s u l t s  of  results  the  sharp  of  cracks.  c l o s e r t o those of  the  of the p o s i t i o n of the n o t c h .  In both laminates t h i c k n e s s e s the  also  The  loads  result  varied  that  slightly  throughout  the volume f r a c t i o n  of  the  f i b r e s v a r i e d from r e g i o n to r e g i o n . T h i s was b e l i e v e d to be the main source of s c a t t e r  i n the toughness v a l u e s .  As shown in f i g . 5 the l o a d i n g curves of linear.  Cracking  noise  was  peak shown. A f t e r  s e v e r a l peaks o c c u r r e d which meant that  d i d not l o s e i t s achieved a f t e r  samples  were  heard at about 85% of the maximum  l o a d which corresponds to a l i t t l e load,  the  load bearing c a p a c i t y .  the maximum  the c r a c k e d sample  Complete  fracture  1 . 5 - 2 minutes f o r most of the samples.  was  35  3.2 F a t i g u e T e s t i n g 3.2.1  Results  F a t i g u e L i f e Of The M a t e r i a l  There  seemed  to be a l a r g e s c a t t e r  i n the f a t i g u e l i f e  of  the samples. As most of the samples were f a i l e d by damage around the holes few data p o i n t s were l e f t graph  (see  figure  to draw  samples t e s t e d , under l o a d c o n t r o l . control  Samples of whereas  it  observed  depends  laminate 3 longer  mostly  Samples f a i l e d q u i c k l y  under  time f o r  below  the other  was t e s t e d  to  determine  sensitive stress  used. it  That  is  if  it  f a i l up to 5x10  suggests that Ao i s material,  but  c e r t a i n about  6  cycles  l a m i n a t e . Only one C ). 6  fatigue  life  range. Hence a h i g h and  a  low  stress  would s u r v i v e  the  material  was  f o r a very long time as the sample  did  c y c l e s and the o p e r a t i o n was s t o p p e d .  This  the important  further it.  s  the m a t e r i a l was maximum s t r e s s  a low s t r e s s range was a p p l i e d . It not  stress  would not s u r v i v e so long and i f  range s e n s i t i v e  3x10  whether  on the maximum s t r e s s or the s t r e s s  was  in  but they were not put on  failed  maximum s t r e s s c l o s e to the f r a c t u r e range  N  five  f a i l u r e was observed below 31.69 MPa (Sample  6  log  plot.  took  Sample E  vs  the  as shown i n Table I I I ,  the f a t i g u e l i f e  Ao  10). Most of the samples of Laminate 1 were  f a t i g u e d below 22MPa and no f a i l u r e was  stroke  the  was seen that  factor  i n the f a t i g u e of  this  experiments are necessary i n order  to be  36  3 . 2 . 2 F a t i g u e Crack Length Measurements Most of the crack length figs.  11  length' print  to  measurement  data  are  shown  27. Some readings are c a l l e d 'maximum damage zone  which corresponds to the l e n g t h measured from the or  damage  life,  that  damage  zone is  to f a i l u r e  it  rate.  initial  In  in  the  with respect stage of  tests. to i t s  fatigue  3  K ,  F  3  and L  This initial  cycling.  higher  be s a i d that  rate  damage  zone  speed then the  other  5  6  life,  but  in  the i n t e r m e d i a t e stage i s a c r a c k  stage.  samples B , 5  surface  H , 5  A ,  C ,  6  7  F  7  and I  7  the c r a c k running  of the sample was d e t e c t e d on the x - r a y  i s why the s u r f a c e readings are very c l o s e to zone  fatigue  i n c r e a s e d a g a i n , but at a slower samples J ,  but  measured.  stage of  intermediate stage of the f a t i g u e  can  deceleration  the  it  i n c r e a s e d with a r e l a t i v e l y  samples i n the  For  at the  i n t e r m e d i a t e and longest  than the i n i t i a l  general  visible  propagated very l i t t l e  close  lengths  was  f a t i g u e damage s t a r t e d e a r l y  zone  s i z e at the Then  x-ray  f i l m . As there was not a s i n g l e crack p r o p a g a t i n g  in g e n e r a l a number of c r a c k s , the maximum l e n g t h was The  in  length.  The  quality  of the x - r a y  the  film.  max.  on That  damage  f i l m a l s o a f f e c t e d the  readings. The s u r f a c e crack lengths are microscope three the  reading'  in  the  s t a g e s i n the f a t i g u e first  stage;  reacceleration  represented  a vs N g r a p h s .  life.  deceleration  An i n i t i a l in  as  'travelling  They a l s o high  followed  velocity  in  the i n t e r m e d i a t e stage and  i n the f i n a l s t a g e . The v a l u e s were u s u a l l y  lower  37  than the c o r r e s p o n d i n g maximum damage zone In sample B ,  one c r a c k s t a r t e d to propogate at the  5  stage of f a t i g u e , of  but a second crack s t a r t e d l a t e r  the notch and took the l e a d .  on the other  s i d e of  the  A l s o i n samples F  sample  started  became c l o s e to the maximum damage zone The  'effective  crack  length'  compliance method a l s o showed three the v a l u e s were c o n s i d e r a b l y two  types  of  lengths.  data  (In  to  stages during  method  the  by using a h i g h c r o s s h e a d displacement rate  it  surface.  the  fractured  sample  this  curve  sample  to  When I  other  crack  is  explained  by  tensile  loaded  the  was found compliance  was s t i f f e r  5  40.1mm (see  fig.  than a  38);  i.e.  was showing a b r i d g i n g e f f e c t which gave  the sample some l o a d b e a r i n g c a p a c i t y . in  but  i n s t e a d of a low one  was r e a l i z e d that  sample which had a notch l e n g t h of  and  fatigue,  c r a c k was measured and i t  to be 42mm long on the specimen performed  faster  fatigue  which was a c c i d e n t a l l y  was  cracks  which was c a l c u l a t e d by the  the  test  7  length.  T h i s c o u l d probably be  and f r a c t u r e d . The r e s u l t i n g  and C  5  lower than those given by the  this  s  from the edge  grow  c o n s i d e r e d to be a n o t c h ) . behavior of sample I  initial  a  depth of  When a saw cut  was  made  43.8mm the slope of compliance  f e l l c o n s i d e r a b l y behaving l i k e a normal n o t c h . T h i s might  be the case f o r  f a t i g u e d samples. The b r i d g i n g e f f e c t around the  damage zone does not l e t notch of e q u i v a l e n t effective lengths.  crack  the f a t i g u e c r a c k or damage behave as a  length does.  lengths  For samples E , 5  were F , 5  H , 5  That lower E  6  and  is  why  than F  7  the  values  of  the measured crack these  values  are  38  plotted  on  a  respectively).  larger  scale  (Figures  16,  17,  18, 22 and 26  39  3 . 2 . 3 Change In The Compliance During F a t i g u e . Most of compliance  the r e s u l t s are shown on F i g s . tests  were  showed the behavior applied i n the in  the  stresses initial  were,  for  When  the  on the f a t i g u e d samples they  that can be seen from f i g .  stage of  tests  performed  28 to 37.  28. Whatever  the  the compliance of the samples i n c r e a s e d fatigue l i f e .  This  i n c r e a s e slowed  the i n t e r m e d i a t e stage and c l o s e t o  down  failure,  the compliance i n c r e a s e d c o n s i d e r a b l y a g a i n . As the damage i n c r e a s e d the from the s t r a i g h t  line  load-deflection  form to a s l i g h t  curve.  line  changed  The i n v e r s e  slope  i n c r e a s e d as l o a d i n g proceeded. There was not so much d i f f e r e n c e in  the compliance f o r d i f f e r e n t  500N to  c y c l e s at low l o a d s e . g .  1000N. T h i s might be due to the b r i d g i n g  around  effect  caused  by the m a t e r i a l surrounding the damage zone or c r a c k s . Normalized were p l o t t e d  occurred  (C/C ) 0  values  f o r most of the samples.  t h e r e are t h r e e 1  compliance  at  stages of an  fatigue  The  life  average value of  vs number of data  cycles  revealed  that  of t h i s c o m p o s i t e . Region 0.2105 of the f a t i g u e  (N ) with an average i n c r e a s e of 0.168 in C / C .  Region  the l o n g e s t  time d u r i n g 0.49 of the samples l i f e  with an average  i n c r e a s e of  0.115 in C/C  0  0.35 i n C/C Sample continuous  0  F  5  took  whereas r e g i o n 3 showed an i n c r e a s e of  0  [ 0 . 2 5 in C/C  2  life  0  if  sample J  5  is  excluded].  which was f a t i g u e d under a h i g h s t r e s s showed a  i n c r e a s e in  i n t e r m e d i a t e stage i . e .  its  normalized  compliance  during  the  high s t r e s s gave q u i c k e r and more severe  40 o  damage.  Sample  The f i n a l  J  s  failure  clearly  region  sample was p r e f r a c t u r e d to determine i t s Sample  K  5  for  E  this  sample i s  very  6  which are t y p i c a l  of 0  f a t i g u e d up to the f a i l u r e , E  5  sample  their  final  I  7  the  As  they  stage i s not  intermediate  revealed  f o l l o w e d the p a t t e r n of  stage  a definite  d u r i n g the i n t e r m e d i a t e s t a g e . Sample F closely  low s t r e s s  (The rate  I. 7  7  fatigue  were  not  shown.  which was f a t i g u e d under medium s t r e s s  v e r y low i n c r e a s e in whereas  typical  load).  d i d not show an i n c r e a s e beyond 1.3 i n C / C .  Sample  fatigue.  i n the I n s t r o n at a v e r y low s t r a i n  fracture and  showed the t h r e e s t a g e s of  like  showed a  sample  J  5  i n c r e a s e i n compliance from the same  laminate  41  3 . 2 . 4 P a r i s P l o t Of The M a t e r i a l The fig.  Paris  39. This  P l o t s drawn  type of p l o t  u s u a l l y has a p o s i t i v e does  not  follow  Crack v e l o c i t y of  fatigue  sample final  K  velocity data.  If  reading  of a l l  is a visible  fatigue.  or the e f f e c t i v e lines.  and  ceramics,  system the m a t e r i a l figure.  and i n t e r m e d i a t e stage  stage i s not shown as i t  but there  to  failure.  incremental  decrease  Both p l o t s were drawn u s i n g the the  crack l e n g t h ,  For  was not c y c l e d  travelling  to in  x-ray  microscope  these methods would g i v e  There would a l s o be l a r g e s c a t t e r  from  sample  which can be understood by examining the a vs N data  the samples.  E q u a t i o n 9 was factor  metals  are shown i n  5  p a t t e r n as can be seen from the  the p l o t s were drawn using  sample  of  and K  then shows an i n c r e a s e c l o s e  during  different to  tests  5  s l o p e . In the p r e s e n t  that  the f i n a l  failure,  samples E  d e c r e a s e s d u r i n g the f i r s t  life,  5  in  for  ranges (AK)  used  to  calculate  the  stress  of the two samples shown i n f i g .  39.  intensity  42  3 . 2 . 5 Radiographs Of The F a t i g u e d Samples The p r i n t s 40  to  produced from the x - r a y  45 i n a l p h a b e t i c a l order of the samples. F i g u r e  the i n c r e a s e i n the damage zone of This  sample  had  beginning of direction that  f i l m s are shown i n  C  40 shows  during  6  fatigue.  a sudden i n c r e a s e i n c r a c k l e n g t h a t the very  the f a t i g u e  starts  test.  Notice  that  damage  in  the  0°  at some d i s t a n c e from the sharp c r a c k meaning  f a t i g u e damage s t a r t e d a f t e r  crack  sample  figs.  this  initial  increase  in  the  length. Fig.  41  shows  a 'low  The damage zone d i d  not  increase  cycles  revealing  that  the  damage  increased  more  in  perpendicular revealed  direction.  that  there  stress  is  range'  f a t i g u e of  much,  even  sample  was  far  sample E . 6  after  4.22x10  from f a i l u r e ,  but  the  0°  direction  than  Figs.  22  and  the same sample  damage  35 of  proceeding,  but  in  s  at  the  a  low  magnitude. Sample  G  6  showed a q u i c k e r damage 2.7X10  had 5  42) which was f a t i g u e d under a h i g h  i n c r e a s e i n the  propagated  cycles.  microscope  (fig.  damage  extensively  in  zone  unlike  6  The  both  directions  after  The c r a c k s that were d e t e c t e d by the  were  distinguishable  in  the  stress  damage  E .  travelling zone  of  the  radiographs. Fig.  43 shows length  another of  the  high  increase  in  interval,  which was a l s o v i s i b l e  stress  fatigue.  Notice  two c r a c k s i n the 2 . 1 7 x l 0 on the s u r f a c e of  the  6  the  cycles sample.  43  Normalized increase  compliance  (see f i g .  Sample L stress  again  direction 28 of  of  shows  this  sample a l s o r e v e a l e d  44) which an  was  also  increase  in  fatigued  45  6  (load-deflection shows  data)  the p e c u l i a r  f a t i g u e d sample, C . Not only 7  by  a  this  high  crack l e n g t h a l s o i n  the notch as w e l l as a damage zone  sample L Fig.  of  36).  (fig.  6  data  increase.  reveals this  behavior  the Fig.  change.  of the medium s t r e s s  the two c r a c k s s t a r t i n g  from  the  c o r n e r s of the notch but other c r a c k s p a r a l l e l but away from the notch i n c r e a s e d i n l e n g t h d u r i n g In  all  initial  fatigue.  of the samples the damage zone grew q u i c k l y  stage of f a t i g u e l i f e  which i s c o n s i s t e n t with  and C / C - N d a t a . A l s o the i n d i v i d u a l 0  in  the  fatigue,  high  stress  fatigue  tests;  i n the low  the damage was in the form of a zone of  than a s i n g l e  crack.  the  c r a c k s were c l e a r l y whereas,  cracks  in  the S-N  visible stress rather  44  3 . 2 . 6 F i n a l F r a c t u r e Morphology The  final  geometry  of  samples i s shown in F i g s .  the  tensile  fatigue  failed  46 to 49.  The s e p a r a t i o n of the f a t i g u e d from  and  samples  into  two  started  the c o r n e r s and not from the center of the n o t c h . T h i s was  i n good agreement with the r a d i o g r a p h s . As mentioned b e f o r e , most of the samples hole.  The  failed  This  was  probably  by  some  MTS  machine.  the hole was l a r g e r  i.e.  layers  of  p o r t i o n was l e f t .  than the p i n .  46 i s a t y p i c a l d e l a m i n a t i o n  playing a big role  failure.  the sample seemed to be d e l a m i n a t e d . The of  the  T h i s geometry, i n d i c a t e s that the damage  which was seen in the 0° d i r e c t i o n  Fig.  impact  of the h o l e .  sample was not d i v i d e d i n t o equal h a l v e s , but some p a r t lower  the  48).  The geometry in f i g . Only  the  o c c u r r i n g on the upper h a l f  impact would be severe i f  (See f i g .  from  region was near the edge of the sample where  the p i n was pushed upwards loading  failed  in the  radiographs  in the f a i l u r e of the p l y  is  also  configuration.  47 r e v e a l s d e l a m i n a t i o n , f i b r e debonding and p u l l - o u t  in the f a t i g u e d region of the f r a c t u r e d region  is  clearly  region.  The l a t t e r  distinguishable  surface.  from  the  region i s smoother than the  The  fatigued  final  fracture  former.  Fatigue  f r a c t u r e again s t a r t e d from the corner of the notch not from the sharp c r a c k .  45  The t e n s i l e  f r a c t u r e d samples showed the smoothest  fatigued  samples  (see  fig.  failure.  Unlike  the  49) f r a c t u r e  exactly  i n the c e n t e r of the notch a l o n g the sharp c r a c k .  started  46  Chapter  IV  DISCUSSION 4.1  Static  Tests  The s t a t i c  tensile  and the f r a c t u r e same  laminate  tests  toughness of were  not  r e v e a l e d that  the samples that  because  no  introduced.  not  bad  average f r a c t u r e were  40.18  and  scatter  so  toughness v a l u e s f o r  case).  The  (Sample L  2  load  from the  was worst on the  sharp  cracks  were  on Laminate 4. The  laminate 2 and l a m i n a t e  29.9 MPa/m r e s p e c t i v e l y  t e n s i o n sample equation exceptional  was  fracture  were cut  u n i f o r m . The s c a t t e r  samples of Laminate 2 p o s s i b l y The  the  is  u s i n g the ASTM compact  excluded  difference  4  as  it  was  an  in the toughness v a l u e s  p r o b a b l y due to the presence of a sharp c r a c k i n the samples  is of  laminate 4. Also,  when the t h i c k n e s s e s of the seven prepared l a m i n a t e s  were measured, i t from  was seen t h a t  there was s c a t t e r  in the  values  laminate to laminate (±0.6mm) and w i t h i n the sample p l a t e .  Hence the r e s i n  content  and  fibre  volume  fraction  was  not  uniform even w i t h i n the same l a m i n a t e . T h e r e f o r e as d i s c u s s e d by Papirno the s c a t t e r tests  would  material.  cause  a  scatter  in  the f a t i g u e  i n the  static  t e s t i n g of  this  47  4.2 F a t i g u e T e s t s i)  Fatigue  Life  S c a t t e r was observed i n the f a t i g u e that for  were  tested. Fatigue  life  of  the  samples  f a i l u r e s o c c u r r e d below 3 x 1 0  laminate 3 which was t h i n n e r  cycles  s  than the laminates 6 and 7  (the  t h i c k n e s s e s of  laminates 5, 6, and 7 were 4 . 1 , 4 . 4 5 , and 4 . 9  respectively)  whereas  laminates. longer  was above 3x10  Specimens w i t h h i g h e r  fatigue  If  it  5560N  load.  not  fail  the s t a t i c  r e s i n content appeared t o  have  i s taken as the normal f r a c t u r e  This  Beaumont and H a r r i s . does  other  life.  f a i l u r e s o c c u r r e d when c y c l i c fracture  the  not  in  agreement  A c c o r d i n g to those  8  i n 10  failure  is  7  points,  points  c y c l e s for  levels  the the  the work of the  material  l e s s than 0 . 9 of  curve was  using  the  in the data and too few  test  from  the  drawn  holes),  but  it  was  to be a shallow l i n e as d e s c r i b e d by D h a r a n ; Awerbuch 6  2 2  maximum  with  workers  stress  as there was s c a t t e r  and H a h n ; and O w e n . In  of  stress.  (as most samples f a i l e d  observed  load, a l l  l o a d range was above 71.2% of  No s t r e s s vs c y c l e s to f a i l u r e test  s  cycles for  mm  fig.  38  39 the s t r e s s  s t r e s s because of  range was  rather  the behavior of sample E  very high maximum l o a d ( c l o s e a high minimum l o a d (hence,  plotted  6  than  which had a  to the average f r a c t u r e  low s t r e s s  range)  the  during  load)  with  fatigue.  48  The  a-N  and  compliance  data f o r t h i s specimen showed that the  maximum l o a d d i d not a f f e c t the sample at a l l ,  because  low  i n 5x10  stress  range  and the sample d i d not f a i l  of the cycles.  6  Therefore the m a t e r i a l can be s a i d t o be s t r e s s range s e n s i t i v e , but more work i s necessary ii)  F a t i g u e Damage Propagation  Although life  to be c e r t a i n .  t h e r e was s c a t t e r  i n the s t a t i c  t e s t s and  data the measurement of crack propogation  clearly  fatigue revealed  the c h a r a c t e r of the f a t i g u e p r o c e s s . Most r e s e a r c h e r s have used one method t o measure f a t i g u e damage. On the other three  methods  compliance  (i.e. travelling  microscope,  methods) i n c r e a s e d the scope  of  hand,  using  radiography  the  data  and  and the  methods complimented one another. It  was  propagation  observed  that  there  are  d u r i n g f a t i g u e . An i n i t i a l  three  a c c e l e r a t i o n i n the f i r s t  stage; d e c e l e r a t i o n i n the intermediate stage reacceleration  towards  the  final  stages of crack  (small  r e a c c e l e r a t i o n of the crack l e n g t h , c l o s e to  is  a  composites, during  different  behavior  failure.  This  i n metals and i n glassepoxy  i n which c r a c k s u s u a l l y a c c e l e r a t e  or  stay  steady  fatigue.  The  crack  length  produced an a vs N curve other  than  of  stage may or may not o c c u r ) ;  and  very  amount  methods  measurement using x-ray that was higher than  because the x-ray technique  f i l m s or p r i n t s  obtained  by the  not only r e v e a l s the  c r a c k s on the s u r f a c e , but a l s o the c r a c k s produced i n the inner  49  l a y e r s . The two at  cracks  showed a damage zone  seen on the s u r f a c e ) which was  (rather  than  the  not  only  propagating  r i g h t angles to the l o a d i n g d i r e c t i o n , but along the  direction the  radiographs  as  well.  loading  When high s t r e s s e s were used in the  tests,  i n d i v i d u a l c r a c k s that were v i s i b l e on the s u r f a c e c o u l d  distinguished  on  the x-ray  f i l m s and  these c r a c k s seemed to be  growing f a s t e r than  the damage zone. In the low  tests  was  the  growth  be  stress  fatigue  i n the form of a damage zone r a t h e r  than  well defined cracks. The  s i z e of the damage zone was  initial early  c o n s i d e r a b l y l a r g e r at  stages of f a t i g u e , r e v e a l i n g that f a t i g u e damage s t a r t s  i n the t e s t which was F r a c t u r e noise was  cycles  of the t e s t . The i n the 90° p l i e s  resists  the  load  a l s o observed  a l s o heard  cracking  at  in  the  by  the  Papirno. first  few  thousand  ( t r a n s v e r s e c r a c k s ) as only the  i n t h i s d i r e c t i o n . R e i f s n i d e r and  first  1 3  i n i t i a l c r a c k i n g i s known to c o n s i s t of  showed that the number of these c r a c k s load,  the  c y c l e and  increases  that with  resin  Jamison  rapidly  3 9  with  i n c r e a s i n g c y c l e s the  s p a c i n g of t r a n s v e r s e c r a c k s approaches a value equal  to the p l y  t h i c k n e s s . Thus the f i r s t  the  length  stage of r a p i d  and compliance must correspond  change  in  crack  to t h i s r a p i d i n c r e a s e i n  the number of t r a n s v e r s e c r a c k s . The  presence of the t r a n s v e r s e  cracks  in  causes two  t h i n g s to happen i n the zero degree  i.  ( r e s i n c r a c k s ) occur p a r a l l e l  ii.  Splits  Delaminations  begin  the  90°  plies.  to the  to grow between the two  fibres, plies.  plies  50  The  intermediate  gradual  growth  projection line.  The  of  sections  Delaminations were  revealing x-ray  delaminations  f a i l u r e occurs when  large  fig.43(iii).  the  as  well  as  a  TBE  itself  steady  clear  of  the  the  delaminations specimen  in  some  radiographs  penetrated  more  into  to  link-up separate.  especially  the delaminated  in  areas  as a continuous dark zone i n the p r i n t s of  the  films. The  'effective  compliance  crack  fatigue  lengths'  cracks  c a l i b r a t i o n was  (see  using a-N  the data,  based on notches, whereas  are r e g i o n s t h a t are not completely severed  are t h e r e f o r e c o n s i d e r a b l y tougher length  calculated  method f o l l o w e d the lowest p a t t e r n i n the  because the compliance  f i g . 38).  than  Nevertheless  a  notch  effective  of  the  crack  t h r e e stages of  length  the  pattern  of  p l o t t e d u s i n g the x-ray continuous  deceleration  the data and  tested  metals and a  samples  did  the final  composite by Kunz and Beaumont fatigue).  not  ( f i g . 3 9 ) . The curves were crack  velocity  showed  a c c e l e r a t i o n . Hence the  m a t e r i a l does not obey P a r i s law. S i m i l a r P a r i s p l o t s drawn  compression  the  fatigue.  The P a r i s p l o t drawn f o r the two follow  and same  f o l l o w e d the same p a t t e r n as the other two methods r e v e a l i n g  this  to  the t r a n s v e r s e c r a c k s p a r a l l e l to the main crack  final  causing  of  stage of f a t i g u e probably corresponds  25  for  had a negative s l o p e ( f o r  51  There crack  i s not much previous  extension  during  B e a u m o n t ; and K i m . * 26  for compression behavior  as  fatigue  Fig.1  0  fatigue.  observed  after  the  reveals  the  the  present  case.  in  the i n i t i a l  the  same  growth of  cubic  relation  to  relation  shows in f i g .  load  s i m i l a r to metals which i s 1  9. T h i s  normalized  of  When a notch  the  last  compliance v a l u e s  (C/C ) 0  the  magnitude of  contrast  with  the  work  there  2 2  and  Liber  i s no l o s s  fatigue. there  ;  In  and D a n i e l .  stiffness  5  stresses  (i.e.  the  in  stress  They supported the idea  that  3 Awerbuch and  Hahn  only  of  fatigue  and  i n s t r e n g t h or s t i f f n e s s  reference  during  of the m a t e r i a l even  i s an i n c r e a s e i n t e n s i l e modulus a f t e r  low  point  Awerbuch  l i t t l e o b s e r v a b l e change in s t i f f n e s s at  cubic  p l o t t e d vs number  upon  in  this  i s a l s o in agreement with the LEFM t h e o r y .  depending  is  length  test  various  amounts,  obeyed  i n agreement w i t h  in  3  history,  tests.  of c y c l e s r e v e a l e d the decay  Hahn '  the  was d e s c r i b e d  the width (W) of the specimen was t e s t e d ,  range. T h i s  and  the compliance on the c r a c k l e n g t h  was not v a l i d as the behavior  The  of  Compliance Data  the study of Mostovoy and h i s c o - w o r k e r s . * closer  work  fatigue  and not dependent upon the  The dependance of a  pattern  K i m ' s work a l s o  present  although change was not observed in the s t a t i c iii)  work of Kunz and  This  in  as a m a t e r i a l p r o p e r t y  except  determined  shows the data of Kunz and Beaumont  r e v e a l e d the p a t t e r n d e s c r i b e d deceleration  work a v a i l a b l e which  asserted  fatigue.  during that  There was  i n the samples t e s t e d  compliance has the o p p o s i t e e f f e c t )  like  52  sample K ,  but  5  after  4.9xl0 The  clearly  still  the normalized compliance  cycles  6  (see  compliance that  magnitude  data  of the f a t i g u e  deceleration  intermediate  stage,  but  deceleration  and  slow  a  fatigue  varied  iv) Most tension  2 7  this  high  and S m i t h .  stresses  the  seemed  reacceleration  in  finish  this  with  the  to p i n  especially  in  loads  Therefore  the  holes.  loading.  aircraft could  than t h i s  agreement  Jamison. splitting  3 9  with  Therefore  This  is  applications  effect  a  pin  type of f a i l u r e ,  the  a  results  of  the  the m a t e r i a l i s  very  serious  where hole  disadvantage  various  of  this  the  d e l a m i n a t i o n morphology  0°  direction  (fig.  46).  types  of  composite.  behavior.  d e l a m i n a t i o n was found to  work of W h i t c o m b ,  D e l a m i n a t i o n was not found along  The  improve  be the predominant f a i l u r e mechanism of the m a t e r i a l , in  stage.  Morphology  more work i s necessary to improve t h i s  Other  cause  samples f a i l e d from the h o l e s of the compact  of  susceptible  to  the  2 8  F i n a l Fracture of  material.  the damage. The  specimen even though attempts were made to  surface  fatigue  of  from sample t o sample i n  c o m p l i a n c e data seems to be in agreement Poursartip,  1.3  t e s t e d samples showed  d e c e l e r a t i o n and r e a c c e l e r a t i o n of  of  to  fig.34).  there are 3 stages of  Acceleration,  changed  in  21  all  clearly  which  is  and R e i f s n i d e r  and  the  plies,  a f f e c t e d the  but final  53  Or.  some  mechanisms  failed  were  also  samples visible.  debonding The  and  final  fibre  pull-out  fracture  r e g i o n was  found to be smoother than the f a t i g u e d zone. F a t i g u e c r a c k s i n i t i a t e d at the c o r n e r s of the notch and/or at  the c o r n e r s of the sharp c r a c k i n t r o d u c e d whereas  cracking  started  smoother f r a c t u r e samples.  transverse  at the c e n t e r of the sharp crack l e a d i n g to a s u r f a c e i n the  statically  tensile  fractured  54  Chapter V CONCLUSIONS The  following  conclusions  can  be  made  i n l i g h t of  the  found i n the v a l u e s  of  present e x p e r i m e n t a l d a t a : 1. There was a c o n s i d e r a b l e s c a t t e r the  fracture  load  and  fracture  toughness  from  laminate to  laminate and on the samples of the same l a m i n a t e ,  although  same  the m a t e r i a l .  manufacturing schedule was a p p l i e d to a l l  There was fraction  also in  necessary  a  variation  all  to  the  of  tested  produce  thickness  samples.  materials  of  of  and  fibre  Larger  uniform  the  volume  a u t o c l a v e s are thickness  and  properties. 2.  The  presence  r i s e to d i f f e r e n t sharpness  is  of  a sharp crack and a b l u n t  magnitude of the f r a c t u r e  an  important  factor  loads  notch gave  meaning  in the f r a c t u r e  toughness  the m a t e r i a l . Sharpness i n c r e a s e s ' s t r e s s c o n c e n t r a t i o n 3 . As a r e s u l t life  showed  sensitive  a  of  scatter  rather  and  it  tests,  appeared  be s t r e s s  revealed  microscope,  that  initial  range  and  there are three stages of life  of  the  compliance  fatigue crack material.  stage the c r a c k or the damage zone a c c e l e r a t e d ;  the i n t e r m e d i a t e stage i t reaccelerated.  fatigue  sensitive. radiography  propagation d u r i n g the whole f a t i g u e the  to  The  d e c e l e r a t e d and in the f i n a l  damage  was  in  of  factor'.  i n the s t a t i c  than maximum s t r e s s  4. The t r a v e l l i n g methods  the s c a t t e r  that  the  form  of  a  In in  stage  it  zone  of  55  t r a n s v e r s e c r a c k s i n the 90° p l i e s and s p l i t s Effective  crack l e n g t h  data  method  gave  tougher  than a notch of e q u i v a l e n t  because  measured  by  the  length. »  . The m a t e r i a l does not obey P a r i s of  the  acceleration  and  law, _  d e c e l e r a t i o n of  damage. The P a r i s p l o t c u r v e s had a n e g a t i v e and in  the  first  crack v e l o c i t y  two stages of  fatigue  was only v i s i b l e  the  plotted  a l s o r e v e a l e d the three stages of  increase  in  compliance  or  decrease  in low s t r e s s  7. The compliance of showed  that C v a r i e s  as i s the case f o r  isotropic  versus  CTS  number  f a t i g u e by showing in  stiffness.  s e v e r a l notch  with the cube of the notch  The  fatigue  lengths length  materials.  8. Most of the f a t i g u e d samples f a i l e d from the  the  fatigue.  the m a t e r i a l f o r  linearly  slope  failure.  compliance changed c o n s i d e r a b l y d u r i n g h i g h s t r e s s range and change very l i t t l e  fatigue  varying  c l o s e to the f i n a l  0  cycles  the  l i f e and an i n c r e a s e i n  6 . The normalized compliance (C/C ) of  compliance  the lowest a vs N curve as the f a t i g u e c r a c k s were  •  5.  i n the 0° p l i e s .  specimen meaning that  the  holes  the m a t e r i a l i s s e n s i t i v e  to  of pin  loading. Delamination dominant  and  mechanism of  resin  cracking  fatigue  failure.  was  found  to  be  the  D e l a m i n a t i o n f a i l u r e was  r e v e a l e d by the morphology of the f a i l e d samples.  56  Some samples f a i l e d with d e l a m i n a t i o n , debonding and pull-out.  The  final  fracture  region  was  smoother  fibre  than the  f a t i g u e d zone. All  the f a t i g u e c r a c k s s t a r t e d at the c o r n e r s of the  or even at the c o r n e r s of the sharp c r a c k , whereas in tensile c e n t e r of  fractured  samples  transverse  statically  c r a c k i n g s t a r t e d at  the  the  the sharp crack l e a d i n g to a smooth f a i l u r e .  9 . More work seems to be n e c e s s a r y to e x p l a i n the of  notch  graphite/epoxy  composite under  The  production  equipment s h o u l d be improved i n order  t o g i v e uniform  properties  to the m a t e r i a l . S t a t i c t e s t s should  be  time consuming and expensive Although, of  there  the g r a p h i t e / e p o x y ,  fatigue  is scatter it  performed  its  before  the  tests.  i n the s t a t i c  i s found t o  m a t e r i a l and seems promising f o r  fatigue.  behavior  be future  a  and dynamic fatigue  tests  resistant  applications.  57  REFERENCES 1.  Ramani, S . V . and W i l l i a m s , D . 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C . , ' F r a c t u r e of Composite M a t e r i a l s ' , USA-USSR Symposium, R i g a , USSR, 1978.  32.  Beaumont, P.W.R. and T e t e l m a n , S . , 'The F r a c t u r e S t r e n g t h and Toughness of F i b r o u s C o m p o s i t e s ' , F a i l u r e Modes i n Composites, The M e t a l l u r g i c a l S o c i e t y , AIME, pp. 4 9 - 8 0 .  33.  K e l l y , A . , ' I n t e r f a c e E f f e c t s and The Work of F r a c t u r e of a F i b r o u s C o m p o s i t e ' , P r o c . Royal S o c , London, A. 319, 1970, pp. 95-116.  34.  Beaumont, P.W.R., 'A F r a c t u r e Mechanics Approach to F a i l u r e i n Fibrous C o m p o s i t e s ' , J . Aveston, Gordon and Breache Science P u b l i s h e r s L t d . , 1974, V o l . 6 , pp. 107137.  35.  S i h , G . C . , H i l t o n , P . D . , B a d a l i a n c e , R., Shenberger, P . S . and V i l l a r r e a l , G . , ' F r a c t u r e Mechanics of F i b r o u s C o m p o s i t e s ' , ASTM S p e c i a l P u b l i c a t i o n s , No 521, 1973, p. 98.  36.  H e r t z b e r g , R.W., Deformation and F r a c t u r e Mechanics of E n g i n e e r i n g M a t e r i a l s ' , John Wiley and Sons, 1976, pp. 465-472.  37.  R a d f o r d , D.W., ' F r a c t u r e Toughness of a Carbon Fibre/epoxy Composite M a t e r i a l ' , M . A . S c . T h e s i s , The U n i v e r s i t y of B r i t i s h Columbia, Vancouver, B . C . , 1982.  38.  Owen, M . J . , ' F a t i g u e of Carbon F i b r e R e i n f o r c e d C o m p o s i t e s ' , F r a c t u r e and F a t i g u e , Academic P r e s s , New Y o r k , 1974.  Cambridge U n i v e r s i t y ,  Engineering  Proc,  of  First  60  39.  R e i f s n i d e r , K . L . and Jamison, R., ' F r a c t u r e of F a t i g u e loaded Composite L a m i n a t e s ' , I n t . J . F a t i g u e , V o l . 4, Butterworth and Co ( P u b l i s h e r s ) L t d . , O c t . 1982, pp. 187197. -  40.  Kim, R . Y . , ' E x p e r i m e n t a l Assesment of S t a t i c and F a t i g u e Damage of Graphite/epoxy L a m i n a t e s ' , U n i v e r s i t y of Dayton Research I n s t i t u t e , A i r Force M a t e r i a l s L a b o r a t o r y , 1978.  41.  Mostovoy, S . , C r o s l e y , P . B . and R i p l y , E . J . , 'Use of C r a c k l i n e Loaded Specimens for Measuring Plane S t r a i n F r a c t u r e T o u g h n e s s ' , J o u r n a l of M a t e r i a l s , V o l . 2 . , Sept. 1967, p p . 6 6 1 - 6 8 1 .  61  i 1000  i * (cyclM)  i  MOO  Figure 1 - Increase i n the c r a c k l e n g t h d u r i n g c o m p r e s s i o n f a t i g u e f o r u n i d i r e c t i o n a l and [0/90] c r o s s - p l y graphite/epoxy composite. 2 6  Figure 2 - S t i f f n e s s decay d u r i n g f a t i g u e . T h e s e a r e t h e d a t a of Smith ( 2 2 . 5 ° to the weave) used i n the development of Poursartip's theory. 2 7  62  P W=56mm,  r — * — •  6 =72mm, H=62mm, P = LOAD,  0 = 19 or24.5mm,  Figure  1  d=8mm,  3 - Compact t e n s i o n specimen used in experiments.  g=32mm,  the  63  AS/3501-6  / TEMPERATURE  /:  i • ! , • , g  1 hour  1  J  1 f  |  I  3 - 5*F/nif  i i 1 i i ii 1 1 • 1  1 1 1 1 1 1 1 3|- 5*F/«jin • i  j  200'F — ^  2 hours  J 1 - 2'F/nln  |  •  1  TIME  Figure  4 -  (i)  C u r i n g Treatment of graphite/epoxy  composite  A.  1 H.i  'i c z - : - : : : 3 = LAM. 2  »  F  ,  R  i LAM. 4 and others a  F i g u r e 4 - ( i i ) The p o s i t i o n of the compact t e n s i o n samples and t h e i r notches as they were cut from a l a m i n a t e .  65  CROSSHEAD  DISPLACEMENT  Figure 5 - T y p i c a l t e n s i l e l o a d i n g p a t t e r n of the t e s t samples l e a d i n g to f r a c t u r e u s i n g I n s t r o n t e s t i n g machine.  66  SAMPLE A*REFERENCE SAMPLE,  Figure 6 - The method of drawing new C - a l i n e s i f the c o m p l i a n c e s of the samples have the same notch l e n g t h from the same laminate are d i f f e r e n t from each o t h e r . 3  67  DEFLECTION (8) mm  Figure  7 - L o a d - d e f l e c t i o n data of sample F lengths.  6  for various  notch  i  18  22  26  3CT NOTCH  Figure  r —  34 LENGTH  1  "I  38 (a) mm  8 - Change in Compliance with notch l e n g t h f o r  ~T5  42  Sample F  er» s  00  Figure  9 - V a r i a t i o n of compliance with the cube of n o t c h for sample F 6  length  STRESS ro  o  o  T"  o o C fD  > 13 I  **1  -  m  O I  m co  2.  0» rt  2°*  C ro  m  fl>  RANGE  a. a! 3 O  T  (AO")  MPa O 1  O 1—~1  r  • 2  r m o -  co > -D  rn co  2  • • • •  & 0) rr 0)  Q  i  OL  i  i  i  i° i  i  i  i  1  I 45 h 340  1  SAMPLE 1  J  O--MAX. DAMAGE ZONE LENGTH Q=T. MICROSCOPE  READING  it 3 5 h o  .j  UJ 3 0 * o  <  25  o  • o =24.5mm 0  cc 2 0 o 10 15 CYCLES (N) Figure  20 (XlO )  11 - Crack Length vs c y c l e s for sample J min. load=222N,a =24.5mm 0  25  N  4  3  max.  load=4446N,  f  SAMPLE £ E  O-MAX. DAMAGE ZONE ••T. M I C R O S C O P E  0.5  3  LENGTH  READING  LO CYCLES  K  L5 (N)  2.0  %  (xlO ) 3  F i g u r e 12 - Crack l e n g t h v s . c y c l e s for sample K . max. load=4528N, m i n . load=240N, a =24.8mm. There i s c o n s i d e r a b l e d i f f e r e n c e i n the two types of r e a d i n g s . (N represents hole f a i l u r e ) 3  0  to  SAMPLE 1  i  1  1  O'MAX. DAMAGE ZONE  45  •=T.  MICROSCOPE  L  LENGTH  READING  ^40 X35  H O  o  z30  •  •  UJ  -j  O  o  N = 93 320 cycles, f  u CYCLES  (N)  (XlO )  8  4  F i g u r e 13 - a vs N for sample L . max. load=4688N, m i n . load=142N, a =25mm. There i s not a r a p i d i n c r e a s e i n c r a c k speed as f a i l u r e n e a r s . 3  0  to  SAMPLE B E E *o 40 X 35 hO z 30 lxl1 °o < 20 o 15  ,  1  0«MAX. OT.  1  DAMAGE ZONE  MICROSCOPE  o*:@ -I  s t  —  r  LENGTH  READING  • 2ncrock A  crock  A«EFFECTIVE  5  CRACK  LENGTH i  JL  2  J  3 CYCLES  (N)  4  -5  L 5  Figure 14 - a vs N for sample B . max. load=4466N, min. load=151N, a 2 4 . 5 m m . A second crack formed a f t e r some time which had a v e l o c i t y g r e a t e r than the f i r s t c r a c k . 5  0  e  SAMPLE 1 1 1 0=MAX. DAMAGE ZONE LENGTH Q T. S  -40  MICROSCOPE  A* E F F E C T I V E  (L 9  r  READING  C.  LENGTH  -O  • A  I -  2  Figure  15 -  3 CYCLES  4 5 (N) (xlO ) 6  a vs N f o r sample C . max. load=41u1N, min. load=71N, a =25mm. Low s t r e s s f a t i g u e t e s t . 5  0  Ni  E E  SAMPLE E,  13  X h- 35 O z UJ 30 -J < 25 CC o 20  O'MAX. DAMAGE ZONEi LENGTH  Q- T. MICROSCOPE READING 4-  E 28 E• -J 27  y  A O  •  o u: 26 u. UJ  eff.  JL  CYCLES  (N)  9 10 (XlO )  I  Ji II  12  5  F i g u r e 16 - a vs N for sample E . max. load=4822N, m i n . load=338N, a =25mm. X - r a y v a l u e s were n e a r l y c o n s t a n t throughout the f a t i g u e l i f e . 5  0  a.  EFF. C. LENGTH mm  LL  CRACK L. (a) mm  SAMPLE H.  £ E S  35 0  30  O  o w 25  MAX. DAMAGE ZONE  •=T. MICROSCOPE  LENGTH  READING  20  d  E E 28 5 27 ? U.' Li. Ui  26  N,  25 5 6 7 8 CYCLES (N) (XlO )  10  II  5  F i g u r e 18 - a v s N f o r sample H . max. load=4857N, m i n . load=80N, a =24.5mm. S u r f a c e and x-ray r e a d i n g s a r e t h e same. 5  0  CO  SAMPLE  K. 5  Figure 19 - a vs N for sample K . max. load=3932N, min. load=80N, a =24.5mm. T y p i c a l low s t r e s s f a t i g u e . Sample d i d not f a i l . 5  0  V£>  SAMPLE i  E 45 E -5 4 0 x  I  0=MAX. DAMAGE ZONE  6  r  LENGTH  •= T. MICROSCOPE READING  35  e>  g 30 * o  i  A  2  D  8  N «4.59xl0c.  5  < 20 cc  o  0.5  1.0  1.5 CYCLES  2.0 2.5 (N) (XI0 )  3.0  3.5  F i g u r e 20 - a vs N f o r Sample A . max. load=5204N, m i n . load*80N. High s t r e s s f a t i g u e . F a i l e d n o r m a l l y . Two types of data are i n good agreement. g  oo o  SAMPLE 1  o=  E E  40 35  N6TH 1  o  30 25  LU  tACK  -J  IX.  -  l • "  T  O  1  1  1  I -T—  1  '"T  »  •= T. MICROSCOPE READING  r  -  o d  L  9  —  —  "  D  n  LJ  a = 24.5 mm 0  1  1  1  1  2  3  i  i 4 N  35 30  !  MAX. DAMAGE ZONE LENGTH  C  —  i  i  6  7  8  9  i 1  i1  0.5  1.0  1.5  t  —  B  CYCLES  i  10  5  SAMPLE  i  - J  i  0  Q  25  5 (XlO )  i  i1 2.5  1 I  2.0 (N)  24.45 mm  V I  3.0  iL _ 3.5  (XI0)  Figure 21 - i i ) a vs N for Sample C . max. load=3692N, min. load=142N. At the very b e g i n n i n g , the sample had an immediate f r a c t u r e and hence a s h o r t e r l i f e , i i ) a vs N for sample G . max. load=5249N, min. load=80N. High s t r e s s f a t i g u e . 6  6  SAMPLE i 1 1 1 OMAX. DAMAGE ZONE LENGTH  % 40 e>  •=T. MICROSCOPE READING  35  O O  30  S  2 5  o  •  o •  •o •  •o •  -1  5 20  J  CYCLES  (N)  eff.  -6  (x|0)  Figure 22 - a vs N for Sample E . max. load=5266N, min. load=1228N, a =24.8mm. Sample d i d not f a i l . High max. s t r e s s and a low s t r e s s range was used. Sample showed the behavior of a low s t r e s s f a t i g u e i . e . not max. s t r e s s s e n s i t i v e . f  0  SAMPLE ,  E 45 h E 40  1  0 » M A X . DAMAGE •«T.  1  ZONE  MICROSCOPE  x 35 —J  r  LENGTH  READING  •  A  25 20  6  O  8  30  K  _  A'EFFECTIVE  15  C. L  0.5  Figure  23 -  a vs N for  1 0  CYCLES  Sample K . 6  (N)  max.  2.0  « 1-5 (XlO ) 6  load=4902N, m i n .  load=142N.  CD CO  SAMPLE 1  £  £  X H  40 35  1  L  s  r -  0 MAX. DAMAGE ZONE LENGTH s  • » T . MICROSCOPE  READING  e> 30 z UJ -J  25  tt  < 20  A = EFFECTIVE  C. L.  JL  a 15  10 15 CYCLES (N)  Figure  24 -  20 (XlO )  25  4  a vs N f o r Sample L . max. load=5071N, m i n . load=142N, a =24.3mm. High s t r e s s f a t i g u e . 6  0  CO  SAMPLE  C  7  §40 335 -  P:  Q  •  •  • A  A 0=MAX. DAMAGE ZONE LENGTH • « T . MICROSCOPE READING  - - O b o c k s i d e T.M.R.  A « E F F E C T I V E C. L  CYCLES (N)  o =24.9 mm o  r5  (x|0 )  Figure 25 - a vs N for Sample C , . max. load=4573N, m i n . l o a d « 1 5 1 N , a, >=24.9mm. Sample was not f a t i g u e d up t o f i n a l f a i l u r e . A t y p i c a l medi urn stress fatigue. 00  SAMPLE F (  35 30  ,  j  = MAX. DAMAGE ZONE LENGTH »T. MICROSCOPE READING  ^n  7  1  p 0^24.5™™  •  -• •  25 20 26 25 24 6 8 CYCLES (N)  10 (X|(5 )  12  5  F i g u r e 26 - a vs N f o r Sample F . max. load=4991N, min. load=71N, a =24.5mm. High s t r e s s f a t i g u e . E f f e c t i v e crack l e n g t h showed a h i g h e r i n c r e a s e in v e l o c i t y than the o t h e r s . 7  0  00  cr>  SAMPLE l UJ(D  E  O  T  35 30 "P  z 25  UJ -J  7  20  < 15 or  0 •  •  •  — - A  •A  Q  0  A  A-  0*MAX. DAMAGE ZONE LENGTH  -  •=T MICROSCOPE o =24.5 mm  READING  A=EFFECTIVE C. L.  0  o  1  1  ;  I  8 10 12 14 CYCLES (N) (XlO )  L _  16  18  5  Figure  27 -  a vs N for  Sample I . max. a =24.5mm. 7  load=5160N, min.  load=80N,  0  00  1  1  1  1  I  1  1  1  I  I  1  I  I  1  I  1  1  1  I  I  DEFLECTION (8) mm F i g u r e 28 - Change i n the load - d e f l e c t i o n c u r v e s d u r i n g of samples L and H . 6  6  fatigue  2.087x10 rnn/H N.* 7,506,660 eyes.  2 Figure  29 -  Increase  3 CYCLES  (N)  4  -« (XlO )  5  6  in the compliance of Sample C  5  during  fatigue, CD VO  F i g u r e 30 - Increase in the compliance of Sample E  5  during  fatigue.  vo o  NORMALIZED  COMPLIANCE  (C/C ) 0  £6  T  1  r  o  T  —i 1 r SAMPLE K.  1  1  r  u 1.3 o QL  1.2  2 O  u  C = I.996xl0" mm/N 4  0  .. 1.1  o  J  LO  I  CYCLES  Figure  34 -  Increase  I  I  (N)  I  I  I  6  (XlO )  in the compliance of Sample K  5  during  fatigue.  L  Figure  35 -  Increase  in the compliance of Sample E  s  during  fatigue.  2  4  6  8 10 12 14 CYCLES (N) (XlO )  16  18  20  9  F i g u r e 36 -  Increase  in compliance of K  6  during  fatigue  vo  CJ  I  i  F i g u r e 37 -  i  Increase  r  1  1  in the compliance of  I  I  7  i  and F  7  during  i  i  fatigue. vo  i  1  1  1  1  1  1  1  r  DEFLECTION IS) mm F i g u r e 38 - Load d e f l e c t i o n data of a t e n s i l e (failed accidentally).  f r a c t u r e d sampl  CRACK  66  VELOCITY (log da/dN) m/sec.  100  (ii)  C  6  N=9xl0 U3.6)  5  cycles  101  * : -1  (iii) E  6  , N=4.22X10  6  cycles  (x3.7)  103  Figure 4 2 Sample G ( i ) N=5.3X10 (x3.3) 6  (ii)  G  cycles  11  6  N=2.7X10  (x3.6)  5  cycles  F i g u r e 43 Sample K ( i ) N=1.66x10 (x3.5) 6  (ii)  cycles  s  K N=3.28X10 cycles (x3.5) 6  5  105  (iii)  K, 6  N=2.17X10  6  cycles  (x3.5)  106  F i g u r e 44 Sample L ( i ) N=1.36x10 (x3.3) 6  (ii)  cycles  s  L N=2.07x10 cycles (x3.6) 6  5  107  F i g u r e 45 Sample C ( i ) N=6.65x10* (x3.6) 7  (ii)  C N=1.93x10* cycles (x3.75) 7  cycles  108  (iii)  C  7 f  N=4.2X10  5  cycles  U3.75)  109  \  F i g u r e 46 - A sample f a i l e d by d e l a m i n a t i o n . N o t i c e the f a i l u r e i s caused from the damage in the two p e r p e n d i c u l a r d i r e c t i o n s that i s v i s i b l e in the r a d i o g r a p h s .  F i g u r e 47 -  F a i l e d sample geometry showing d e l a m i n a t i o n , f i b r e debonding and p u l l - o u t .  Figure  49 -  Sample f r a c t u r e d under I n s t r o n t e n s i l e loading.  incremental  111  Table  I  F r a c t u r e Toughness v a l u e s f o r s a m p l e s o f L a m i n a t e 2 u s i n g ASTM compact t e n s i o n sample e q u a t i o n ( e q n . 3)  Sample Number  * B  cD  2  2  2 2  E F  2 2  G 2  H I? J  2  2  K L  2 2  Notch Length a (mm) 18 19 18 19  19.8 19.5 21.5 20.5 21.0 20.0 19.8 19.3  Notch t o Thickness Width r a t i o a/w t (mm) 0.321 0.339 0.321 0.333 0.350 0.345 0.375 0.353 0.363 0.357 0.350 0.342  3.985 4.130 4.01 3.970 4.250 3.990 4.050 4.235 3.950 4.090 4.010 3.970  Fracture load P (N)  Fracture Toughness ^(MPai'in)  5693 5071 5560 6574 7384 7006 6427 6503 5960 5271 4893 9118  36.565 32.867 36.893 43.212 46.590 47.406 45.733 41.566 41.850 35.915 33.396 61.116  f  112  Table  II  F r a c t u r e Toughness v a l u e s f o r s a m p l e s o f L a m i n a t e 4 u s i n g ASTM compact t e n s i o n sample e q u a t i o n ( e q n . 3)  Sample Number  A. B,  c.  D, E» Fft G. H, I« J« Kft Lft  Notch Length a (mm) 18.3 19.2 20.0 20.0 19.4 19.4 19.7 19.8 19.8 19.5 20.0 20.0  Notch t o Thickness Width r a t i o a/w t (mm) 0.3327 0.349 0.357 0.364 0.353 0.353 0.358 0.354 0.354 0.355 0.357 0.357  4.970 5.250 4.880 4.910 4.895 4.895 4.960 4.970 4.860 4.990 5.010 4.935  Fracture load P (N)  Fracture Toughness K^MPav/m)  5591 5898 5427 5471 4653 4559 5471 5769 5493 5017 5489 5293  30.029 31.093 30.359 31 .249 26.233 25.785 31.060 32.686 32.697 28.176 30.523 29.882  f  113  Table I I I T e s t i n g Data f o r  Fatigue  L a m i n a t e s 1,2,3,4, a n d 7  i ) LAMINATE 1 Sample Number  A, C, D, E, F, G, H, Ii Ji Ri L,  (SO (SC) (SC) (SC) (SC) (SC)  Notch Length a (mm) 18.5 22.0 21.0 22.0 21.0 21.0 21.0 20.0 20.0 18.5 18.5  Max. Load £>ax  Min. Load Range Load < > Pmin <»> AP(N)  4435 4946 4626 4385 4644 4404 4515 4003 4359 4092 4510  N  911 943 1059 400 495 854 890 979 756 778 133  4324 4003 3567 3985 4199 3550 3625 3024 3603 3314 4377  Stress Range Ao Mpa  Cycles to Failure  19.1 18.3 16.03 18.83 18.87 15.95 16.29 14.2 16.92 15.22 21.01  6.64xl0« 4990 690 4320 1250 45570 4360 4.24x10' 4.8x10' 5.1x10 5x10' s  (NF)  (NF) (NF) (NF) (NF)  'SC = S t r o k e C o n t r o l 'NF' = No f a i l u r e N o t e : Samples C,, D,, E,, F,, G,, H, were f a t i g u e d under ' S t r o k e C o n t r o l ' . Sample B, was t e n s i l e f r a c t u r e d . Load t o f r a c t u r e was 5756N. i ) LAMINATE 3 G Ij J K L 3  3  3  3  24.5 29.0 24.5 24.8 25.0  4484 4048 4466 4528 4688  356 303 222 240 142  4128 3745 4224 4288 4546  33.68 33.67 34.28 34.53 37.22  1 12240 95640 269210 212960 (HF) 93320  'SC = S t r o k e C o n t r o l 'NF' = No f a i l u r e 'HF' = H o l e f a i l u r e N o t e : Sample A was c o m p l i a n c e t e s t e d . Sample C f r a c t u r e d a c c i d e n t a l l y . Below samples of Laminate 3 was o n l y c y c l e d and c r a c k l e n g t h was not measured. Most of them f a i l e d by h o l e failure. 3  B D E F  3 3 3 3  19.8 20 20 24.5  3  4484 6263 6192 4995  294 543 712 907  4190 5720 5480 4088  28.16 38.66 39.02 31.58  1.15x10' 38360 4860 38400  (HF) (HF) (HF) (HF)  114  iii)  LAMINATE 5  Sample Number  Notch Length a (mm)  Max. Load P (N) max  B  c  s  5  D E F G H  s 5 5 5 5  Js  K  5  24.5 25.0 25.5 25.0 25.0 25.0 24.5 25.5 24.5  Min. Load min  4466 4101 4973 4822 5071 4617 4857 3994 3932  151 71 80 338 53 80 80 80 80  Load Range AP(N)  Stress Cycles to Range F a i l u r e Ao Mpa f  4315 4030 4893 4484 5018 4537 4777 3914 3852  32.17 31 .69 38.16 34.28 38.51 34.79 36.02 31 .28 31 .04  N  459560 7506660 425260 1077550 431630 705630 1102290 1148910 4.9x10  (HF) (HF) (HF) (HF) (HF) (NF)  s  'HF' = Hole f a i l u r e 'NF' = No f a i l u r e N o t e : Samples A and L were compliance t e s t e d . Sample J was i n i t i a l l y loaded to i t s f r a c t u r e s t r e s s l e v e l to observe i t s Load to f r a c t u r e (P £ 5 5 1 6 N ) . That i s why i t had a s h o r t l i f e with a low s t r e s s . Sample I was a c c i d e n t a l l y f r a c t u r e d d u r i n g s t a t i c tensile loading. 5  s  5  5  iv)  c  E G H K L  LAMINATE 6  6 e 6 6 6  6  24.2 24.5 24.5 24.8 24.45 24.5 24. 1 24.3  5204 4448 3692 5266 5249 5089 4902 5071  80 80 1 42 1 228 80 80 142 142  5124 4368 3550 4038 5169 5009 4760 4929  'HF' = Hole f a i l u r e ' NF' = No f a i l u r e Note: Sample F was compliance t e s t e d . b e g i n n i n g of the f a t i g u e t e s t i n which was Sample I . Sample J was not u s e d . an immediate c r a c k i n g at the b e g i n n i n g 6  6  6  38. 18 32.86 26.81 29.0 37.51 37. 18 33.93 36.53  458880 4325830 981380 5X10 6  -  (NF)  Sample D f a i l e d at the Max. Load was 5305N. So Sample C seemed t o have of the t e s t . 6  6  1 1 5  v)  LAMINATE 7  Sample Number  Notch Length a (mm)  Max. Load P (N) max  c F  7 7  I T  Min. Load P (N) min  Load Range AP(N)  S t r e s s C y c l e s to Range Failure Ao Mpa N  2 4 . 9  4 5 7 3  151  4 4 2 2  2 9 . 4 8  2 4 . 5  4 9 9 1  71  4 9 2 0  3 0 . 9 2  2 4 . 5  5 1 6 0  8 0  5 0 8 0  3 1  _  -  . 4 6  Note: Sample G was compliance t e s t e d . Sample A the very b e g i n n i n g of the f a t i g u e t e s t . 7  f  7  f r a c t u r e d at  116  Table  IV  Compliance T e s t  i ) V a l u e s of Compliance (C) sample A .  Results  vs Notch Length (a)  and C vs a  3  for  5  Notch Length (a) mm  Compliance (C) mm/N (xlO"*) 1.994 3.066 5.418 8.783  24.5 28.0 34.2 39.5 ii)  of C vs a  3  6  (for  L ) 5  3  from Laminate 6 1.062 1 .940 2.543 3.073 4.078 5.341 9.772 18.080  17.5 24.6 27.05 29.00 31 .8 34.95 40.0 44.0  3  16003.00 23076.10 29076.05 34805.63 41709.72 51895.12 64481.20 82029.36  u s i n g the f i r s t f i v e notch l e n g t h s C=1.3193881x10' a -1.226322x10"*  For sample F  C vs a  14706.13 21952.00 40001.69 61629.88  2.135 2.884 3.740 4.496 5.578 7.566 14.577 17.273  7  iii)  3  a g a i n from Laminate 5  For sample 25.2 28.47 30.75 32.65 34.68 37.3 40.1 43.45  Eqn.  Cube of Notch Length mm  5359.38 14886.94 19792.55 24389 32157.43 42691.51 64000.00 85184  equation f o r F using the f i r s t 6 notch l e n g t h s C=1.143664x10' a +2.9448x10'" 6  7  3  is:  is:  117  iv)  For sample G  Notch Length mm  7  (a)  Compliance (C) mm/N ( x l O - " )  24.5 26.4 29.15 31 .77 33.82 35.9 40.5 44.45 C vs a  3  2.078 2.399 3.156 4.044 4.984 5.933 9.837 18.196  Cube of Notch Length mm 3  14706.13 18399.74 24769.41 32066.51 38683.06 46268.28 66430.13 87824.42  equation f o r G u s i n g the f i r s t 6 data p o i n t s C=1.2166865x10* a +1.4785x10-" 7  7  is:  3  v) For sample A from Laminate 3 . (The r e s u l t s were not used f o r ' e f f e c t i v e crack l e n g t h ' measurements. 3  18.8 27.8 31.2 35.0 39.0  1.18 3.0230 4.456 6.519 11.18  6644.67 21484.95 30371.33 42875 59319  118  Table Weight and Volume F r a c t i o n of  Fiber Lam. Comp. Num. Mass(gm) Mass(gm) (m ) (m ) c  f  V the F i b r e s  Mass Fraction of f i b r e s (mf/m ) c  for  Volume of f i b r e s (v ) (cm ) f  3  Each L a m i n a t e .  Volume Volume of comp F r a c t i o n <v ) of f i b r e <v ) c  f  (cm ) 3  1 2 3 4 5 6 7 Note -  6.7666 3.4900 5.5993 3.8129 3.5117 6.0794 7.9994  5.0797 2.6622 4.1124 2.6728 2.5369 4.316 5.6516  0.751 0.763 0.736 0.701 0.722 0.710 0.702  p = 1 . 9 g m / c c , ^ = 1 . 1 9 gm/cc f  2.6735 1.4012 2.1644 1 .407 1 .335 2.2716 2.9556  4.0911 2.0968 3.4139 2.365 2.1542 3.7534 4.9588  0.653 0.668 0.634 0.595 0.620 0.605 0.596  

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