Effects of ply drops on the fatigue resistance of composite materials and structures
by Mark Ethan Scott
A thesis submitted in partial fulfillment of the requirements for the degree of Master of Science in
Chemical Engineering
Montana State University
© Copyright by Mark Ethan Scott (1997)
Abstract:
Material thickness variations are required to optimize the design of laminated composite structures.
These thickness variations are accomplished by dropping layers of material (plies) along the structure
to match the load carrying requirements. Unfortunately, these ply drops produce internal stress
concentrations as a consequence of material and geometric discontinuities. This thesis provides a
parametric experimental investigation of ply drops in E-Glass stranded fabric reinforced polyester
composites and structures. These parameters include: ply drop location, laminate thickness, number of
plies dropped at one location, fabric type, loading condition, fiber content, and spacing between ply
drops. The damage which develops at ply drops is typically delamination cracks which propagate
between the layers of reinforcing fabric.
There were two parts to this study: (1) to examine delamination propagation rates at ply drops and
determine crack growth threshold levels, and (2) to determine the effect of ply drops on the lifetime of
various composite materials. Tests were conducted on both small coupons of material and beam
structural elements with ply drops in the flanges.
A strong sensitivity to ply drop position and manufacturing details is shown for fatigue damage
initiation and growth. The results indicate that it will be difficult to completely suppress damage and
delamination initiation in service. For 0° plies, single internal ply drops provide the greatest
delamination resistance. Multiple ply drops should be spaced at correct intervals so that the
delaminations from each do not overlap prior to arrest. It was found that, in most cases, there is a
threshold loading under which little growth after initiation is noted. Delamination retardation
techniques such as ply edge feathering, “Z-Spiking” and adhesive layers improve the delamination
resistance in many cases. After delamination has occurred, especially with exterior ply drops, it can be
repaired with adhesives. Ply drops adversely affect fatigue lifetime of low fiber content laminates more
severely than for high fiber content laminates. The choice of fabrics used in a laminate can have a
significant impact on delamination rates, but the lifetime of the laminate is insensitive to fabric type. 
EFFECTS OF PLY DROPS ON THE FATIGUE RESISTANCE
OF COMPOSITE MATERIALS AND STRUCTURES
by
Mark Ethan Scott
A thesis submitted in partial fulfillment 
of the requirements for the degree
of
Master of Science 
in
Chemical Engineering
MONTANA STATE UNIVERSITY 
Bozeman, Montana
August 1997
ii
APPROVAL 
of a thesis submitted by 
Mark Ethan Scott
This thesis has been read by each member of the thesis committee and has been 
found to be satisfactory regarding content, English usage, format, citations, bibliographic 
style, and consistency, and is ready for submission to the College of Graduate Studies.
Dr. John Mandell
jrperson, Graduate Committee
Approved for the Department of Chemical Engineering
Dr. John Sears
Department Head Date
Approved for the College of Graduate Studies
Dr. Robert Brown
DateGraduate Dean
iii
STATEMENT OF PERMISSION TO USE 
In presenting this thesis in partial fulfillment of the requirements for a master's 
degree at Montana State University, I agree that the Library shall make it available to 
borrowers under rules of the Library.
IfI have indicated my intention to copyright this thesis by including a copyright 
notice page, copying is allowable only for scholarly purposes, consistent with "fair use" as 
prescribed in the U.S. Copyright Law. Requests for permission for extended quotation 
from or reproduction of this thesis in whole or in parts may be granted only by the 
copyright holder.
Signature
I
Date
ACKNOWLEDGMENTS
I would especially like to thank Dr. Mandell and Dr. Cairns for their insight and 
direction. Dan Samborsky receives special thanks for all his time and suggestions with 
editing figures, efforts in helping with experiments and keeping the facilities running. 
Zane Maccagnano for his work using finite element analysis to back up the results 
obtained experimentally. To all of the students in the composites group for their input 
and good sense of humor during the frustrating moments of school and research. And 
finally, to my family who have always encouraged and supported me.
This work was supported by the U.S. Department of Energy and the State of 
Montana through the Montana DOE EPSCoR Program (Contract # DEFC02-91ER75681) 
and the National Renewable Energy Laboratory (Subcontract # XF-I-11009-5), and 
Sandia National Laboratories (Subcontract AN0412).
TABLE OF CONTENTS
LIST OF TABLES........................................................................................................  vii
LIST OF FIGURES ...........................................................................................................
ABSTRACT..................................................................................................................  xii
1. INTRODUCTION..........................................................................................................!
2. BACKGROUND ........................................................................................................ ...
Strain Energy Release Rate ................................................................................... 4
Tensile Testing of Dropped-Ply Laminates........................................................... 8
Compression Testing of Dropped-Ply Laminates .................  9
Damage Accumulation.................................................     10
Stress Analysis on Dropped-Ply Laminates......................     11
Delamination Prevention....................................................................................  13
Existence of a Resin Rich Region......................................................   14
Design Considerations......................................................................................... 15
Motivation for Thesis.............................. ; ......................... ............................... 18
3. EXPERIMENTAL METHODS AND MATERIALS................................................. 20
Test matrix and description of specimens........................................................... 20
Material Preparation............................................................................................. 29
Fabrics ...................................................................................... 31
Specimen Preparation ......................................................................................... 32
Test Facility and Development ................................................ 33
DCB and ENF test specimens............................................................................. 35
4. RESULTS AND DISCUSSION ................................................................................. 37
Delamination S tudy............................................................................................. 37
Effect of Ply Drop Location................ 37
Effects of laminate thickness and multiple ply drops at the
same position , , ........................................................................... 43
The ESE and ESF laminates ................................................................... 44
Effect of Ply Drop Spacing .........................   45
Effects of “Z-Spiking” .................................................. : ................... .... 48
Effect of Tough Adhesive at Ply D rop..................................................... 49
Effect of Butt-Joints............................  £0
Effects of ±45° Ply Drops ....................................................................... 51
Effect of A13 0 warp unidirectional fabric ............. -............................. 52
Other attempts to prevent delamination................................................... 53
Repairing Delaminated Samples ............................................................. 55
Comparing different laminates................................................................. 56
Effect of Ply Drops on Lifetime........................................................................... 57
Fiber Volume Content Effect................................................................... 57
A130 fabric vs. D155 fabric.................... ................................‘ ..............59
Resin Rich Region................................ ............................................. . . 62
Compression Testing........................................................................................... 64
Residual Strength of Coupons................................................................. 67
Ply Drops in Beam s........................................................................ ................. ; 70
Beam Geometry, Fabrication and Materials............................................. 70
Flange laminate E SA ............................................................................... 73
Beam 3 9 ....................................................................................... 73
Beam 4 4 ....................................................................................... 76
Beam 4 8 ....................................................................................... 78
Flange laminate E SB ............................................................................... 80
Beam 40 .................... ............................................................... 80
Flange laminate E SG .........................................................................   81
Beam 4 1 ....................................................................................... 81
Beam 4 2 ....................................................................................... 85
Flange laminate E SH ............................................................................... 86
Beam 4 9 ....................................................................................... 86
Beam 5 0 .......................................................... 89
Comparison of Delamination Rate in Beams and Coupons....................90
Strain Energy Release Rate and Modeling Results............................................. 94
Determination of Critical Strain Energy Release Rate ........................... 94
Static and Threshold Strain Energy Release Rates at Ply D rops............97
5. CONCLUSIONS AND RECOMMENDATIONS ..................................................100
Conclusions ........................................................................................................ 100
Delamination at Ply Drops......................................................................100
I-Beam Study..........................................................................................101
Effect of Ply Drops on Lifetime..............................................................101
Gc Tests ..................................................................................................102
Recommendations............................................................................................102
REFERENCES ............................................................................................................103
APPENDIX A
Coupons Tested.............................................................................................................108
APPENDIX B
Delamination Length vs. Cycles for Coupons................................................................118
LIST OF TABLES
Table Page
1. Lay-up of fiberglass materials with Ply Drops............................................................22
2. Test matrix................................................................................................................... 24
3. Comparison of delamination resistance of different ply drop configurations.. . . . . . .  42
4. Slenderness ratios calculated using Equation 4 .................................................. .... 65
5. Comparison of ESH laminate compressive strength with and without ply drops........66
6. Residual strength of ESH laminate after being fatigued (R=OT) ............................ .69
7. List of I-beams tested with ply drops..........................................................................72
8. Reference notation for Beams 39, 44 and 48 with ESA Laminate ............................74
9. Delamination length vs. fatigue cycles for Beam 3 9 .................. ............................... 75
10. Average delamination length vs. cycles for compression and tension
flanges for Beam 44...........................................    77
11. Delamination length vs. cycles for tension and compressive flanges on Beam 48. .. 79
12. Reference notation for Beam 40 with ESB laminate........................ ....................... 81
13. Reference notation for Beams 41,42 with ESG laminate ........................................ 83
14. Cycles vs. average tension and compression flange delamination
length for beam 4 1 .............................. 84
15. Cycles vs. average tension flange delamination for Beam 42.................................... 85
16. Cycles for tension and compression flange delamination on beam 49. . ; ................87
17. Reference notation for Beams 49 and 50 with ESH laminate.................................... 88
18. Cycles vs. tension flange delamination length for beam 50............. ......................... 91
vii
viii
19. Critical strain energy release rate for Mode I and Mode II cracks............................97
20. ESA and ESB critical static strength..........................................................................97
21. Static strain energy release rates ..............................................................................98
22. Threshold strain energy release ra tes........................................................................99
?
LIST OF FIGURES
Figure Page
1. Common structural elements with discontinuities from Ref. I ........ ........................... 2
2. The three modes of fracture from Broek [2] ................................................................5
3. Typical load vs actuator displacement for DCB specimen........................................... 7
4. Typical load vs actuator displacement for a ENF specimen.......................................... 8
5. Cross-sectional and edge view of ESB laminate showing ply drop........................... ; 28
6. Different delamination prevention techniques...............................................................30
7. Lay-up of laminate with ply drops in mold ...........................   31
8. DCB and ENF specimens ....................................................................................   35
9. Delamination of exterior zero degree ply drop.............................................................. 38
10. Typical static failure of laminate with ply drop......................................................... 38
11. Delamination length vs. cycles for ESA laminate, R = 0.1,
Exterior O0 Ply Dropped.....................     39
12. Delamination at an interior ply drop showing two delamination fronts.....................39
13. Delamination length vs cycles for ESB (Single interior 0°ply drop)
laminate, R=0.1.....................................................................................................40
14. Delamination length vs. cycles for ESC (Single center interior 0°ply drop)
laminate, R=0.1......................................................................................................41
15. Illustration of resin rich region in ESH laminate........................................................45
16. Delamination length vs. cycles for laminates ESB, ESH, ESF (all interior ply drops)
at a maximum running stress of 275 MPa, R=0.1 ..............................................46
ix
X17. Effect of different spacing between ply drops, R=O. I, ESI laminate (Two, 0° ply
drops) at 276 MPa, Vf=0.35 ............................................................ ................. 47
18. ESI coupons run at 276 MPa, R=0.1.  ................................................................47
19. Delamination length vs. cycles for ESJ “Z-Spiked” laminate compared to ESA
(Single, exterior 0° ply drop) laminate, R= 0.1.................................................... 48
20. Delamination length vs. cycles for laminates ESA, ESK, ESG, ESE (all exterior ply
drops) at a max. running stress of 138 MPa, R=0.1............................................. 49
21. ESL laminate in tension at the ply drop (Interior 0° ply drop with Hysol adhesive)
failed at 276 MPa................................................................................................. 50
22. Delamination length vs. cycles for ESA, JKA "Feathered" and JKA random
laminates at a maximum running stress of 207 MPa, R=0.1............................... 53
23. Delamination initiating along tows which were pulled out........................................54
24. Delamination length vs. cycles for ESB and JKB "Feathered" at a maximum
running stress of 275 MPa, R=0.1.......................................................................54
25. Delamination length vs. cycles at 207 MPa, R=0.1, initial ESA laminate
compared to repaired ESA laminate........................................... '........................ 55
26. Effect of fiber content on the normalized S-N Data, R=0.1, for DD materials
[0/±45/0]s compared to ESH laminate (Two interior O0 Ply drops).................... 58
27. S-N curve (R=0.1) comparing D155 Fabric (DD6) to A130 Fabric (DD11) control
materials with no ply drops, from Reference 50.................................................. 60
28. Fatigue life of ESH (Two D155 ply drops) vs. ESQ (Two A130 ply drops)
laminates (Maximum stress is on the thin side of the ply drop)...........................60
29. Tensile fatigue S-N data for ESQ (Two O0 internal ply drops) vs. control
DDl I (no ply drops) laminates, A l30 unidirectional fabric.................. ............. 61
30. Damage developing around thermoplastic beads as a result of fatigue
loading, R=OT...................................................................................................... 62
31. Photomicrograph of ESH laminate showing resin rich regions ahead
of ply drops................................................................... ........................................62
32. Side view of crack propagating through ESH laminate.......................... '.................63
33. Tensile fatigue (R=0.1) S-N curves for ESB (Single 0° internal ply drop)
and ESH (Two interior 0° ply drops).................................................................. 63
34. Compressive stress vs. percent strain for ESH 804 coupon........................................ 65
35. I-Beam testing apparatus ........................................................................................... 71
36. Beam coordinate system ...........  71
37. Beam 39 with ESA laminate for flange material...................................  76
38. Tension and compression flange on Beam 44. ....................................    78
39. Beam 48, ESA laminate, tension and compression flanges.....................   80
40. Beam 4 0 ..................................................................................................................... 82
41. Beam 4 1 ......................................................................................................................84
42. Beam 42, ESG laminate (Two exterior 0° ply drops) on flanges ................................86
43. Beam 49 with ESH (Two internal 0° ply drops) flange material.................................89
44. Beam 50 with ESH (Two interior 0° ply drops) laminate as flange material..............90
45. Beam (Tension Flange) vs. Coupon Data (R=OT) for ESA (Single exterior 0° ply
drop), R=OT....................................^ ........ ......................................................... 92
46. Beam (tension flange) vs. coupon data for ESH laminate (Two Interior ply drops),
R=OT..........................   92
47. Mode I coupon geometry.................................................  95
48. Mode II specimen geometry........................................................................................ 96
xi
xii
ABSTRACT
Material thickness variations are required to optimize the design of laminated 
composite structures. These thickness variations are accomplished by dropping layers of 
material (plies) along the structure to match the load carrying requirements. 
Unfortunately, these ply drops produce internal stress concentrations as a consequence of 
material and geometric discontinuities. This thesis provides a parametric experimental 
investigation of ply drops in E-Glass stranded fabric reinforced polyester composites, and 
structures. These parameters include: ply drop location, laminate thickness, number of 
plies dropped at one location, fabric type, loading condition, fiber content, and spacing 
between ply drops. The damage which develops at ply drops is typically delamination 
cracks which propagate between the layers of reinforcing fabric.
There were two parts to this study: (I) to examine delamination propagation rates 
at ply drops and determine crack growth threshold levels, and (2) to determine the effect 
of ply drops on the lifetime of various composite materials. Tests were conducted on both 
small coupons of material and beam structural elements with ply drops in the flanges.
A strong sensitivity to ply drop position and manufacturing details is shown for 
fatigue damage initiation and growth. The results indicate that it will be difficult to 
completely suppress damage and delamination initiation in service. For 0° plies, single 
internal ply drops provide the greatest delamination resistance. Multiple ply drops should 
be spaced at correct intervals so that the delaminations from each do not overlap prior to 
arrest. It was found that, in most cases, there is a threshold loading under which little 
growth after initiation is noted. Delamination retardation techniques such as ply edge 
feathering, “Z-Spiking” and adhesive layers improve the delamination resistance in many 
cases. After delamination has occurred, especially with exterior ply drops, it can be 
repaired with adhesives. Ply drops adversely affect fatigue lifetime of low fiber content 
laminates more severely than for high fiber content laminates. The choice of fabrics used 
in a laminate can have a significant impact on delamination rates, but the lifetime of the 
laminate is insensitive to fabric type.
ICHAPTER I
I
INTRODUCTION
Today’s need for stronger, lighter and cheaper structures has generated much 
interest in materials development, especially in composite materials. Fiber-reinforced 
composites have played a leading role in the technological advancement of structural 
material systems. Typically, fiber-reinforced composites are known for being light 
weight, high strength materials which are more durable than conventional materials. The 
use of composite materials in structural applications is rapidly increasing for commercial 
applications. With this increased use comes the need for a better understanding of the 
performance of the structures fabricated from composite materials, called composite 
structures. A large portion of composite structures are comprised of layered, laminated 
composite materials; thickness variations in such laminates are achieved by changing the 
number of plies in proportion to the thickness change. This requires the termination of 
layers, or plies, within the laminate, which then introduces a characteristic flaw into the 
material.
Laminated composites typically are fabricated from planar sheets of material, so 
that all fibers are oriented in a plane. Careful design and selection of the in-plane fiber
2orientation can create a laminate that is designed to carry the loads very efficiently in the 
plane of the fiber reinforcement. However, an inherent weakness of the laminate is the 
lack of fiber reinforcement in the direction normal to the fiber orientation . Consequently, 
the interlaminar direction, normal to the plane of reinforcement, is the weakest direction 
of the laminated material system. Therefore, any interlaminar loads that are applied to or 
induced within the structure are of particular concern in terms of structural integrity.
Figure I, from Ref. I, illustrates five structural elements used in laminated 
composite structures that produce interlaminar stresses. These common elements are the 
free edge, the open hole, the ply drop, and bonded or bolted joints. Free edges are 
unavoidable in many structures. Open holes are commonly employed to allow access to 
the internal parts of the structure. When the design calls for a laminate that is tapered in 
thickness, discontinuous layers or plies are utilized. It is also common to insert
Free Notch P ly  Bond Bolted
edge (hole) drop joint joint
Figure I. Common structural elements with discontinuities from Ref. I
3discontinuous plies to create a local build-up or thickening at high stress points. Finally, 
bonded or bolted joints are required to attach multiple sub-components of the structure. 
Each of the structural elements shown in Figure I develop significant out-of- plane 
normal and shear forces when the component is under load. These interlaminar loads are 
acting on the plane of minimum strength and toughness of the laminated structure. 
Therefore, each of these structural elements have the potential to cause a delamination of 
the individual layers. In addition, most analyses and failure models do not account for 
these interlaminar loads. The interlaminar performance of these critical structural 
elements provides a limit to the structural performance of the composite structure. It is 
important to note that the approach taken in this thesis is also directly applicable to some 
degree to all of the interlaminar stress risers illustrated in Figure I .
The approach in this work was to focus on the ply-drop configuration. This is an 
unavoidable flaw if the thickness is to be tapered, and has received limited attention in the 
literature with respect to low cost composites of this type under fatigue loading. A 
parametric experimental study of the influence of various geometric details of ply drops 
was carried out using laminate coupons, in terms of both the delamination resistance and 
the reduction in fatigue lifetime. The work is then extended to ply drops in the flanges of 
larger I-beam structures.
4CHAPTER 2 
BACKGROUND
This Chapter reviews the basic mechanics of delamination in terms of the strain 
energy release rate. Several key problem areas associated with thickness transitions in 
composite laminates are then identified and discussed. Issues in need of an increased 
research effort are identified.
Strain Energy Release Rate
Once a crack is initiated in a structure it can be further propagated in any of three 
different modes, or a combination of these. Figure 2 shows the three modes of crack 
growth. Mode I is an “opening mode” crack, which is caused by normal stresses. In­
plane shear causes Mode II or “sliding mode” cracks and Mode III cracks are caused by 
out-of-plane shear and are known as “tearing mode” cracks [2].
The strain energy release rate, G, is based on the Griffith criterion [2], Griffith 
stated that crack propagation will occur if the energy released upon crack growth is 
sufficient to provide all the energy that is required for crack grpwth. The Griffith 
equation can be represented as
5dU dW
da da (I)
where U is the elastic strain energy and W the energy required for crack growth. G is 
equal to dU/da and is sometimes called the crack driving force. The energy consumed 
during crack propagation is denoted by R, which is equal to dW/da, and is called the 
crack resistance. Thus, R is equal to the critical strain energy release rate to cause crack 
extension.
There is a different critical strain energy release rate for each mode of crack 
growth. A subscript denotes the particular mode. The critical strain energy release rate is
Mode I
opening mode sliding mode
Mode 11 Mode 111 
tearing mode
Figure 2. The three modes of fracture from Broek [2],
denoted with a “c” in the subscript following the mode designation. Above this value of 
G, in simple linear elastic fracture mechanics [2], a crack will propagate unstably in the 
structure. If there is a mechanism which produces increased crack resistance as the crack
6extends, the crack will propagate according to some R-curve behavior, requiring higher G 
values as crack extension occurs [2].
To determine an opening mode, or Mode I, strain energy release rate for 
delamination, a double cantilever beam (DCB) specimen is used. The critical strain 
energy release rate can be obtained by determining the area enclosed by the loading and 
unloading curves on a load-displacement diagram, which is the incremental change in 
stored strain energy, U, with crack extension Aa. A typical loading-unloading diagram 
for a DCB specimen can be seen in Figure 3. Another method to determine G1 values
UP 2cQ2
EB2H3
(2)
uses an analytic formula (Eq. 2) proposed by Benbow and Roesler [3] and Gilman [4] 
which takes into account the strain energy generated due to the bending moment of the 
DCB test, where a is the crack length, E the modulus parallel to the crack direction, B the 
laminate width, h is the half height and Pc is the critical load. Many G values can be 
obtained from a single DCB specimen which allows a crack resistance (R) curve to be 
generated, indicating how ( and if) the resistance to crack growth changes with increasing 
crack length.
To determine Mode II crack growth resistance, it is necessary to use a different 
test method to determine the corresponding strain energy release rate. End notched 
flexure (ENF) tests apply a load to the center of the coupon; when the applied load
7Load vs. Actuator Displacement for a DCB Specimen
Area under load-unloading curve is equal to 
the change in strain energy during the test.
Actuator Displacement (inches)
Figure 3. Typical load vs. actuator displacement for DCB specimen.
reaches a critical value, the crack propagates suddenly toward the center where the load is 
applied. In this type of test, there is only one data point collected as compared with the 
many points for the DCB specimen due to the instability of crack growth in the ENF 
specimen. A typical load-actuator displacement graph is shown in Figure 4. Since the 
load-displacement diagram is unstable for Mode II tests, an analytic formula is necessary 
to determine a G value. The formula proposed by Russell and Street [5] to calculate Gllc 
is
^ c 2* 2 
\6Exw 2Zz3 (2)
8Load vs. Actuator Displacement for a ENF Specimen
40
Actuator Displacement (inches)
Figure 4. Typical load vs. actuator displacement for a ENF specimen.
Where Pc is the critical applied load, a is the initial crack length, Ex is the modulus in the 
long direction of the specimen, w is the width of the specimen and h is the half thickness 
of the specimen.
Tensile Testing of Dropped-Ply Laminates 
The DCB and ENF specimens are used to induce crack growth in pure Mode I 
and I I , respectively. Ply drops do not cause a pure mode of crack growth, but generally 
display a combination of at least Modes I and II, as discussed later. Lagace and Cannon 
[6] conducted an experimental program investigating the influence of discontinuous 
internal plies on the tensile response of graphite/epoxy laminates. The fiber orientation 
was chosen to minimize the occurrence of the edge delamination failure mode. For the
9configurations investigated, little change in the global stress-strain behavior and ultimate 
load carrying capability were observed relative to similar constant thickness laminates. 
The work illustrated an effect due to the placement of multiple ply drop-offs. When 
discontinuous plies were distributed along the specimen length, the laminate failed 
similar to a laminate without ply drops. However, when multiple plies were dropped at 
one position in the laminate, a delamination failure resulted. It was suggested that the 
apparent penalty in the in-plane response may be more severe for compression loading.
Wisnom [7] looked at the size effect of coupons on delamination rate. Larger size 
coupons showed a decreasing delamination rate. Also, when the number of plies dropped 
in one location was increased, there was a significant increase in the rate of delamination.
The effect of the fiber content on ply drop effect has not received much attention. 
The apparent reason for this lack of research is that the majority of work has been 
centered around the use of pre-preg laminates which generally have ply thicknesses of 
less then 0.2 mm, with fiber volume contents ranging narrowly in the fifty to sixty percent 
range. The fiberglass laminates being studied for use in this project ranged in fiber 
content from the low thirties to the mid-fifty percent range. Mandell, et al. [8] showed 
that the tensile fatigue resistance decreases with increasing fiber content. The cause of 
this is postulated as due to the decreased amount of resin between plies which acts as a 
buffer to reduce stress concentrations resulting from matrix cracks in adjacent layers.
Compression Testing of Plv Drop Laminates 
Grimes and Dusablon [9] investigated the static and fatigue compression behavior 
of graphite/epoxy composites with internal ply drop-offs subjected to severe operating
10
environments. A [±45/0/90]s family laminate with up to four dropped plies was tested. 
The static strength and stiffness were insignificantly affected by the discontinuities. 
However, the endurance limit was decreased by up to 30 percent depending on the 
orientation of the discontinuous layers.
Curry, et. al. [28] noted that dropped plies can cause a significant reduction in 
compressive strength. Also, the greater the change in axial stiffness between the thick 
and thin sections, the greater the reduction in strength. The reduction in compressive 
strength is much greater than the reduction in tensile strength.
Damage Accumulation
Ulman et al. [10] conducted an extensive experimental program with
graphite/epoxy laminates with single ply drops. Damage initiation and growth from both
open holes and ply terminations were evaluated during both static testing and constant
amplitude fatigue testing of a [0/±45/90] family laminate. The results indicated that
.
although the loading mode greatly influenced the damage development process, 
interfacial delamination was present in both static and fatigue loading before ultimate 
failure. Tapered specimens were shown to be dominated by the ply drop. Under tensile 
loads, the damage progressed from matrix cracking in the resin-rich zone, to transverse 
cracks at the ply termination. Continued loading induced delamination which increased 
in size until final failure. In fatigue loading, the damage development was similar, but, 
delamination growth dominated the last 75 percent of the structural life. A similar 
fracture response was illustrated for compression loading; however, the process occurred 
much more rapidly.
11
Reifsnider [11] showed that the residual (remaining) strength of typical 
composites does not deteriorate from the static value until a few hundred cycles before 
failure. This was because the accumulated damage in fatigue samples was similar in type 
and amount to that in quasi-static tests near failure.
Stress Analysis of Ply Drop Laminates
A numerical study of a 30-ply carbon fiber laminate containing two zero degree 
dropped plies at different locations through the thickness was performed by Adams, et al. 
[12]. A three-dimensional finite element analysis incorporating nonlinear orthotropic 
lamina properties was conducted to assess the effect of compressive load, moisture and 
temperature. Although the residual thermal stresses resulting from post cure cooling 
were found to be significant, the interlaminar stresses were negligible compared to the in­
plane stresses. It was concluded that the dropped plies have little effect on the in-plane 
stresses.
Kemp and Johnson [24] presented a failure analysis of an eight-ply quasi-isotropic 
graphite/epoxy laminate containing up to three dropped zero degree plies, in a single drop 
step. The finite element method was applied to determine the resulting three-dimensional 
state of stress. Interlaminar stresses were found to be significant at the location of the 
dropped plies. In-plane failure loads, were calculated based on both resin failure near the 
dropped plies and intralaminar failure in tension and compression. A maximum principal 
stress criterion was applied to predict resin failure and the Tsai-Wu criterion was applied 
to predict intralaminar failure. The analysis predicted initial failure to occur in the resin. 
No experimental data was generated to correlate with the numerical results.
12
Curry et al. [28] studied sixteen-ply graphite/epoxy laminates, containing four 
plies terminated at the midrplane. A stress based failure theory was applied to predict the 
axial compressive strength. The results were shown to under estimate experimentally 
determined failure loads by 33 percent. Experimental results indicated that both the 
tensile and compressive strength are reduced by the presence of a thickness discontinuity. 
However, the reduction was larger for compressive loading. In addition, the reduction in 
strength was shown to be inversely related to the increase in axial stiffness of the thick 
section relative to the thin section.
Fish and Lee investigated the tensile strength of tapered glass/epoxy laminates 
with multiple internal ply drops [13]. A three-dimensional finite element analysis was 
conducted and the average stress concept was applied to predict the initiation of failure. 
Strength predictions, based on the stress in the inter-ply resin layers and an averaging 
distance of one ply thickness, were found to correlate with the experimentally observed 
delamination failure. However, the predictions were valid only for delamination 
initiation and not delamination growth. Laminates both with and without significant edge 
effects were tested. The failure mode for those with negligible free edge effects was an 
unstable delamination growth across the width of the laminate, followed by stable 
delamination growth in the axial direction. One of the tapered laminates showed an 
increase in strength while the other showed a decrease. Alternatively, laminates 
containing 90° plies showed a significant free edge effect. The corresponding failure 
mode was a combination of free edge delamination and delamination growth in the axial 
direction [14].
13
A similar analysis for laminates with external ply drop-offs was conducted by Wu 
and Webber [15]. A two-dimensional finite element analysis was performed.
Interlaminar peel and shear stresses were found to peak in the comer regions of the 
external steps. The failure of the laminates in delamination was attributed to these peak 
stresses.
Unidirectional glass/epoxy and graphite/epoxy laminates with internal drop-offs 
were studied by Hoa et al. [16]. Under tensile loading, the laminates failed in 
delamination at load levels well below the in-plane strength. Stress and strain variations 
within the taper were obtained using a three-dimensional finite element model employing 
a quadratic displacement based element. An acoustic emission was found to correspond 
to the initiation of delamination while no change in the stress-strain response was 
observed. Reasonable correlation between experimental and numerical results was 
reported when the interlaminar stresses were compared to the interlaminar strength of the 
material.
Delamination Prevention
Another type of ply drop occurs in laminates with discontinuous interlaminar 
layers in the use of buffer strips [17,18]. A buffer strip is added or constructed by either 
discontinuing specific fiber layers of a laminate, usually the stiff zero degree plies, or 
inserting additional plies of lower longitudinal modulus. The objective is to create 
discrete region of increased local compliance within the component. The result is a 
superior construction in terms of damage tolerance. As damage in the main laminate 
grows, it can be arrested and controlled by the more compliant buffer strips. Such a
14
construction has been shown to reduce the sensitivity of composite structures to inherent 
damage and to increase the lifetime of the component [19].
Another method currently being used is called “Z- Spiking”. Aztex Co. 
(Waltham, Mass.) uses fibers positioned in foam which is then vacuum pressed into the 
laminate to reinforce laminates in the thickness direction. This reinforcement is 
especially useful for bonded joints, where the strength of the joint is matrix dominated. 
Aztex [20] reports a minimal decrease in the in-plane properties and a 30-fold increase in 
interlaminar fracture toughness. Tanzawa [21] showed that although the z reinforcing 
fiber content was approximately 0.6%, the normally brittle carbon fiber/epoxy plates had 
the same critical strain energy release rate as carbon fiber/PEEK plates. By increasing the 
critical strain energy release rate, the structure becomes more delamination resistant 
without having to increase the cost by using the more expensive PEEK thermoplastic 
matrix.
Chan [22] used narrow, tough thermoplastic interlayers to prevent coupon edge 
delamination. In another study. Masters [23] used an entire layer of adhesive to improve 
the impact toughness of composites. These interlayers provided a tough region where 
propagating cracks could be arrested before growing long enough to cause a catastrophic 
failure in the component.
Existence of a Resin Rich Region
It has been shown that during processing of a variable thickness laminate, a pure 
resin region develops at the end of the discontinued plies. Therefore it is the resin flow 
from surrounding layers during processing that creates this neat resin zone. After the
15
resin fills the mold, porosity can become entrapped in the resin in this area creating stress 
concentrations which can initiate delamination. The volume and dimensions of this 
region are clearly a function of the number of discontinuous layers at a given location and 
the orientation of the surrounding layers. The existence of this neat resin region has been 
observed in several works [24-28]. A common method employed to account for this 
micro structural detail is to inspect actual laminates with photo microscopy. Although 
this identifies the size and shape of the neat resin zone, there has been little effort to 
address the change in fiber volume content near a discontinuous layer. Chan and Ochoa 
[26] did attempt to account for the change in ply properties that occur local to the 
discontinuous plies.
Design Considerations
In metallic structures, damage tolerance technology has been used effectively to 
characterize crack growth under cyclic loading for a material, predict the rate of crack 
growth in the structure under service loads, and establish inspection intervals and 
nondestructive test procedures to ensure operational safety [29]. Because composite 
delamination represents the most commonly observed macroscopic damage mechanism in 
laminated composite structures, many efforts have been undertaken to develop similar 
procedures for composite materials by characterizing delamination growth using fracture 
mechanics [30-33]. Although this approach is promising, there are some fundamental 
differences in the way fracture mechanics characterization of delamination in composites 
may be used to demonstrate fail safe designs compared with the classical damage
tolerance treatment used for metals.
16
Many papers have been published recently where the rate of delamination growth 
rate with fatigue cycles has been expressed as a power law relationship in terms of the 
strain energy release rate, G, associated with delamination growth [1-4]. This fracture 
mechanics characterization of delamination growth in composites is analogous to that of 
fatigue crack growth in metallic structures, where the rate of crack growth with cycles is 
correlated with the stress intensity factor at the crack tip. However, delamination growth 
in composites, with a relatively high crack growth exponent, may change too rapidly over 
too small a range of load, and G, to be incorporated into a classical damage tolerance 
analysis for fail safe designs [2,34,35]. Where in metals the range of fatigue crack growth 
may be described over as much as two orders of magnitude in G, the growth rate for a 
delamination in a composite is often characterized over less than one order of magnitude 
in G. Hence, small uncertainties in applied loads may yield large (order of magnitude) 
changes in delamination growth. Different damage mechanisms may also interact with 
the delamination and increase the resistance to delamination growth. Delamination 
growth resistance curves may be generated to characterize the retardation in delamination 
growth from other mechanisms [36,37,38]. This delamination resistance curve is 
analogous to the R-curves generated for ductile metals that account for stable crack 
growth resulting form extensive plasticity at the crack tip. However, unlike crack tip 
plasticity, composite damage mechanisms such as fiber bridging and matrix cracking, 
may not always be present to the same degree.
One alternative to using the classical damage tolerance approach for composites 
would be to use a strain energy release rate threshold, below which no delamination
17
growth occurs, and design to stress levels below this threshold. Metals are 
macroscopically homogeneous, and the initial stress conditions that create cracks at 
particular locations in preferred directions cannot be easily identified beforehand. 
Composites, however, are macroscopically heterogeneous, with stiffness discontinuities 
that give rise to stress risers at known locations such as free edges, internal ply drops, and 
matrix cracks. Although these stress fields are not the classical variety observed at crack 
tips, and hence cannot be characterized with a single common stress intensity factor, they 
can be characterized in terms of the strain energy release rate, G, associated with eventual 
delamination growth [4].
The most common technique for characterizing delamination onset in fatigue for 
composite materials is to run cyclic fatigue tests on standard composite specimens, where 
G for delamination growth is known, at maximum load or strain levels below that 
required to propagate a delamination monotonically. A strain energy release rate 
threshold for delamination onset may be developed by running tests at several maximum 
cyclic load levels and plotting the cycles to delamination onset versus the maximum 
cyclic G, corresponding to the maximum cyclic load or strain applied[39-43]. This G 
curve may then be used to determine a threshold value of G for delamination and to 
predict delamination onset in other laminates of the same material, or at other points [44].
Uncertainty inherent in predicting service loads has generated concern for using a 
no-growth threshold design criterion for high cycle fatigue applications. If G values 
exceed no-growth threshold levels, a catastrophic delamination propagation may occur. 
O’Brien [45] outlines a damage threshold/fail-safety approach for composite fatigue
18
analysis that involves the following steps:
1. ) Predict the delamination onset thresholds using fracture mechanics.
2. ) Assume that surpassing the delamination threshold corresponds to complete
propagation.
3. ) Determine the remaining load carrying capability of the composite with
delamination present using composite mechanics (i.e., check for fail- 
safety).
4. ) Iterate on Steps I to 3 to account for multiple sources of delamination.
Step I may be used to demonstrate the delamination durability of any composite 
structure. Step 2 reflects a way to deal with relatively high exponent delamination growth 
observed for composites as compared to metals. An alternative would be to predict the 
delamination growth rate using growth laws that incorporate R-Curve characterization, 
thereby taking into account the resistance provided by other damage mechanisms. Finally, 
Step 3 acknowledges that the residual strength of the composite is a function of structural 
variables, and it is not uniquely a question of material characterization. This proposed 
damage-threshold fail-safety concept incorporates generic material properties of fracture 
mechanics and also takes into consideration the unique characteristics of laminated 
composites.
Motivation for Thesis
Previous research has focused on graphite/epoxy [6] or pre-impregnated E-glass 
and S-glass [7] laminates containing ply drops. Graphite/epoxy laminates have better 
fatigue resistance than do glass fiber composites [7]. The pre-impregnated composites 
usually have thinner plies than do typical laminates of the type used in this study, which
19
may minimize ply drop effects. To achieve efficient designs, wind turbine blades require 
severe thickness tapering to reduce the overall weight. Blades are also subjected to very 
high cycle fatigue loading, which can lead to delaminations over a period of many years. 
This, along with an absence of adequate data on the effects of ply drops in laminates 
using the E-glass stranded fabrics typical of blade and other low cost composite 
applications, gives motivation for this thesis.
20
CHAPTERS
EXPERIMENTAL METHODS AND MATERIALS
Test matrix and description of specimens 
This section provides a brief overview of the different cases'tested. Table I gives 
the lay-up of each case as well as other comments. Table 2 gives the test matrix used to 
determine the properties of the laminates listed in Table I, including the number of 
coupons tested at each stress level, and the loading conditions (tension or compression). 
In the following paragraphs, plies which are terminated (near mid-length of the coupon) 
are shown with asterisks. As described later, all laminates were resin transfer molded 
using stitched or woven fabric reinforcement.
The laminates were based on the DD set of laminates, which exhibited the best 
fatigue performance of the laminates previously tested [47]. Selected cases of ply drops, 
both interior and exterior (on the surface), were introduced into laminates. After initial 
delamination studies with single ply drops were completed, multiple ply drops at the 
same location, as well as spaced along the length, were also studied. Attempts to 
suppress delamination with the addition of adhesives, feathering and “Z-Spiking” were 
then studied. All of these attempts investigated dropping O0 plies. The effect of dropping 
single and multiple ±45° layers was then studied. In addition, A l30 woven fabric for the
21
zero degree layers was also substituted for the standard D 1,55 fabric used in the DD 
family of laminates.
The ESA laminate included a single exterior zero degree ply drop. The actual 
configuration of the laminate is [0*/0/±45/0/0/±45/0], where the angle given is relative to 
the applied load direction. Ply configurations follow standard laminate notation [46]. The 
ESB laminate has the configuration [0/0*/±45/0/0/±45/0], while the ESC laminate has the 
configuration [0/±45/0/0*/0/±45/0], with a central ply being terminated. In terms of 
percentages, the ESA, ESB and ESC laminates have a 14% drop in thickness, with 20% 
of the zero degree layers being dropped. The ESD laminate has two internal plies being 
dropped with the laminate configuration [0/±45/0*/0*/±45/0]. This is a 33% drop in the , 
total thickness, with 50% of the zero degree layers being dropped. Figure 5 shows an 
actual polished cross-section through an ESB coupon prior to testing. The edge of the ply 
drop is somewhat smeared during molding, with a resin rich area ahead of the drop. The 
thickness taper coincides with the ply drop as closely as possible, to give an 
approximately constant fiber content along the length.
In the next set of laminates thicker materials were used to investigate less severe 
thickness tapering. The ESE laminate had the configuration [0*/(0/±45/0)3], while the 
ESF laminate incorporated a single internal ply drop into the thicker laminate 
[0/0*/±45/0/0/±45/0/0/±45/0]. In each of these laminates the thickness was tapered 10%, 
while the percentage of zeros being dropped was 15%. The ESG and ESH laminates both
Table I. Lay-up of fiberglass materials with ply drops
Lay up of Fiberglass Materials with Ply Drops
Laminate Ply Configuration Description
ESA [0*/(0/±45/0)s] Single Exterior Ply Drop
ESB [0/0*/±45/0/0/±45/0] Single Interior Ply Drop
ESC [0/±45/0/0*/0/±45/0] Single Center Ply Drop
ESD [0/±45/0*/0*/±45/0] Double Central Ply Drop
ESE [0*/(0/±45/0)3] Single Exterior Ply Drop, with thicker laminate
ESF [0/0*/±45/0/(0/±45/0)2] Single Interior Ply Drop with thicker laminate
ESG [0*/0*/(0/±45/0)3] Two exterior ply drops with thicker laminate
ESH [0/0*/0*/±45/0/(0/±45/0)2] Two interior ply drops with thicker laminate
ESI [0/0*/0*/±45/0/0/±45/0] Two interior ply drops with different spacing between the ply drops.
ESJ [0*/(0/±45/0)s] “Z-Spiking” of a single exterior ply drop.
ESK [0*/(0/±45/0)s] Single exterior ply drop, Hysol EA9309,2NA adhesive applied to ply drop 
before polyester resin was put in the mold.
ESL [0/0*/±45/0/0/±45/0] Single interior ply drop, Hysol EA9309.2NA adhesive applied to ply drop 
before polyester resin was put in the mold.
ESM ' [0/0*/±45/0/0/±45/0] . Attempted “Z-Spiking” with an interior ply drop. The zero degree ply being 
dropped was slipped through a cut ±45 layer.
ESN [0* */±45/0/0/±45/0] Outside zero degree layer used as a butt-joint.
ESO [0/±45**/0/0/±45/0] Inside ±45° degree layer contains a butt-joint.
ESP [0/±45*/±45*/±45/0/(0/±45/0)2] Two interior ±45° layers being dropped.
ESQ [0/0*/0*/±45/0/(0/±45/0)2] Two interior ply drops with thicker laminate 
A13 O fabric instead of D15 5 fabric
ESR [0/±45/0/±45 */±45 */0/±45/0] Twq interior ±45 ply drops at the centerline
Lay up of Fiberglass Materials with Ply Drops
Laminate Ply Configuration Description
JKA
“Feathered”
[0*/(0/±45/0)s] Alternating tows were pulled one half inch 
past the adjacent tows in the ply drop layer.
JKA
Random
[0*/(0/±45/0)J Random mat laid down underneath the ply drop
JKB
“Feathered”
[0/0*/±45/0/0/±45/0] Alternating tows were pulled one half inch 
past the adjacent tows in the ply drop layer.
* Ply being terminated
** Ply contains a butt-joint oriented at 90° to the load direction.
The subscript s denotes a symmetrical lay-up about the location of s, while the notation ( )n indicates that the lay-up in ()  is 
repeated n times [46].
tou>
24
Table 2. Test matrix
Laminate
Configuration
Tension Coupons 
(R=OT)*
Compression Coupons 
(R=IO)*
Maximum. 
Stress (MPa)
# of 
Tests
Minimum 
Stress (MPa)
# of 
Tests
ESA Static 6 -207 2
207 3 -138 2
138 3
-
ESB Static 4
345 2
310 3
276 5
ESC Static 2
345 3
276 3
ESD Static I
138 I
ESE Static 2
207 2
138 2
121 I
ESF 345 I
276 4
207 I
ESG Static I
345 I
207 I
138 2
103 I
25
Laminate
Configuration
Tension Coupons (R=OT)* Compression Coupons (R=IO)*
Maximum. 
Stress (MPa)
# of Tests Minimum 
Stress (MPa)
# of Tests
ESH Static 8 -276 3
454 I -207 2 ■
414 5
345 3
276 7
207 6
ESIl 310 I
276 3
ESI2 310 2
276 3
246 I
ESB 276 3
241 2
207 I
ESI4 310 I
276 4
241 I
ESJ 276 I
207 3
ESK 276 I
207 2
138 I
ESL 276 2
241 2
26
Laminate
Configuration
Tension Coupons (R=OT)* Compression Coupons (R=IO)*
Maximum. 
Stress (MPa)
# of Tests Minimum # of Tests
Stress (MPa)
ESM 276 I
ESN 276 I
172 I
138 I
103 2
ESO 345 I
310 2
276 3
ESL Static 3
552 2
414 3
276 3
207 I
ESR 276 2
241 2
JKA
“Feathered”
276 . 3
JKA
Random
276 3
JKB
“Feathered”
276 3
* R = minimum load/maximum load
included more than one ply drop at the same position. ESG had two exterior zero degree 
layers dropped while the ESH laminate had two interior plies dropped. The laminate
27
configurations were [0*/0*/(0/±45/0)3] for the ESG laminate and 
[0/0*/0*/±45/0/0/±45/0/0/±45/0] for the ESH laminate. In both of these laminates the 
thickness was tapered 18% , while the percentage of zero degree layers being dropped 
was 25 percent.
The ESI laminate contained multiple single ply drops with each ply drop separated 
by multiples of 13 mm spans. For example, the ESI2 had a 25 mm spacing between ply 
drops, while the ESI3 laminate had a 38 mm spacing in between ply drops. The laminate, 
lay-up, [0/0*/0*/±45/0/0/±45/0], had a 25% thickness taper while the percent of zeros 
being dropped was 33%.
ESJ and ESK used special details shown in Figure 7 in an attempt to increase 
delamination resistance. The ESJ laminate used “Z-Spiking” [47 ], to provide a thickness 
reinforcement. This was accomplished by interlacing the exterior zero degree ply drop 
with a continuous zero degree layer. This helped to prevent the surface ply drop from 
peeling away when the coupon was loaded. The ESK laminate also was an exterior ply 
drop; to prevent delamination, a tough epoxy adhesive, Hysol EA 9309.2NA, was 
applied to the ply drop zero degree layer and the first continuous layer and then allowed 
to cure before the resin was introduced into the mold. The fabric lay-up for both the ESJ 
and ESK laminates was the same as for the ESA laminate, [0*/0/±45/0/0/±45/0]. The 
ESL and ESM laminates considered delamination prevention for internal ply drops. The 
ESL laminate used the same Hysol adhesive as the ESK laminate. Again the adhesive was 
applied both underneath and above the ply drop and allowed to cure before the resin was 
introduced into the fabric. The ESM laminate represents an attempt to “Z-Spike” an
28
Cross - sectional view
Dropped 0° ply
Edge view
Dropped 0°pl\
+/- 45° Layers
Figure 5. Cross-sectional and edge view of ESB laminate showing ply 
drop.
internal ply drop. This was accomplished by cutting the ±45° degree layer underneath the 
ply drop and pushing the layer being terminated through the ±45° layer. Both of these 
laminates used the same lay-up as the ESB laminate, [0/0*/±45/0/0/±45/0].
The ESN and ESO laminates incorporated butt-joints in plies. In order to use 
some glass fabrics which are only available as weft unidirectionals, butt-joints are 
necessary. In both the ESN and ESO laminate there are no dropped plies, but the butt 
joints cause a stress discontinuity similar to the effect caused by ply drops. The ESN 
laminate had an exterior zero degree layer as the butt joint, while the ESO laminate had 
an interior ±45° layer as the butt-joint. The actual lay-up in the ESN and ESO laminates 
was [0/±45/0]s
The ESP laminate had the lay-up, [0/±45*/±45*/±45/0/(0/±45/0)2], to evaluate the
29
effect of dropping ±45° layers. In this laminate the thickness was tapered 18%, with no 
zero degree layers being dropped. The ESQ laminate had the same lay-up as the ESH 
laminate, two interior zero degree ply drops, but the fabric being used for the zero layers 
was the warp unidirectional woven A l30. The last laminate is ESR, which had the lay-up 
[0/±45/0/±45*/±45*/0/±45/0]. JThis looked at the effect of dropping multiple ±45° layers 
in a thinner laminate. No zero degree layers were dropped, and the thickness was tapered 
25"%.
Additional coupons investigated the possibility of suppressing delamination by 
adding to or tailoring the laminate properties. The first configuration used the ESA 
laminate as a basis for comparison, and a second configuration used the ESB laminate as 
a basis. For the JKA coupons, random mat fabric was included between the ply drop and 
the first continuous zero layer. Random mat was not used in the internal ply 
configuration case. A second modification is called “feathering.” In this case, alternating 
tows were pulled one half inch past the adjacent tows to provide a less defined 
delamination site. This modification along with “Z-Spiking” can be seen in Figure 6.
Material Preparation
The various E-glass/polyester materials used in this thesis were all manufactured 
by RTM (Resin Transfer Molding), which consisted of a peristaltic pump forcing the 
resin into a closed, vented mold containing the reinforcing layers[48]. To incorporate ply 
drops into the manufacturing process, the following procedure was used to prepare the 
mold. First the mold was coated with a mold release (Frekote 700- NC), air dried, then 
the individual fabric layers were placed in the mold. The lay-up of a laminate with ply
30
drops is shown in Figure 7. Ply drop layers (4) were cut to shorter lengths and placed 
the desired position along with the continuous layers (3). Fluoro-Peel release film (2) 
added to the laminate to accomplish the thickness reduction in the desired area, with
"Z-spiking”
Top view of dropped layer
25.4mm
Side view lam inate  when "Z-epiked"
/ / / / / / / / / / / / / / / / / / / / / / / / / / / / / / / / }  xxxx\vx\\\\vv\\\vxx\\x<\\\x\<v:::':':::::::v\x<<\\x<\
xXXXV,X\XnXXXXX>X\XXXXX?XXXX\XxXXXX<XXX\\\x?XX<\\XXXV
0  Dropped ply 
Q  Contlnuoue piles 
88 "Z-epiked" region
Adhesive”
Plydrop
B  Glued a rea
Thick eectlon
Thin section
in
was
Figure 6. Different delamination prevention techniques.
31
2. Release Fabric
LExtraFabric 
(Removed Later)
/  / /  / /  / /  / /  / /  / /  77 / /  / /
/ /  / /  / /  / /  / /  / /  / /  / /  /y  /
5
5
3. Dropped Ply
4. Continuous Plies
Figure 7. Lay-up of laminate with ply drops in mold
fabric layers (I) added over the release film so that the fiber content was the same in the 
thin and thick sections. When the resin was injected into the mold, the porosity of the 
sample was low (<1%), except for the resin rich region in front of a ply drop which was 
often observed to have some localized higher porosity.
Fabrics
The E-glass fabrics used in these experiments were from Owens-Coming Fabrics 
(formerly Knytex). Three different types of fabric were used in these experiments. Wind 
turbine blades, because they are long and thin need to use a high percentage of continuous 
warp unidirectional fabric oriented with fibers in the length (0°) direction to give them 
stiffness. They also usually incorporate ±45° fabrics to control twisting. Unidirectional 
fabrics, from Owens-Coming, are available in either the warp or the weft direction. Warp 
fabric means that the individual tows of glass fibers are oriented parallel to the long
32
direction of the fabric when it is rolled up; conversely, weft fabrics have the 
unidirectional tows oriented so they are perpendicular to the length of the roll, which is 
generally only 1.3 meters wide. D155 fabric is a unidirectional fabric which weighs 
approximately 526 grams/ m2. The D155 fabric is available in the weft direction only, 
and while the properties are good, the current manufacturing lengths makes it unsuitable 
for structures longer than 1.3 meters. The second type of unidirectional fabric Used in this 
thesis is Al 30. This fabric is a woven warp unidirectional fabric which has good tensile 
properties, but because of the curves introduced from large thermoplastic beads 
connecting the tows, creating the weave, the compressive properties are significantly less 
than the D155 properties [49]. Finally the ±45° fabric layers involve the DB 120 fabric, 
the DB representing double bias fabric of 407 grams/m2.
The resin used in these experiments was CoRezyne 63-AX-051, an unsaturated 
orthophthalic polyester, manufactured by Interplastic Corporation. This resin was 
combined with methyl ethyl ketone peroxide (MEKP), 2% by volume, to catalyze the 
cross-linking reaction. Curing of the materials was at ambient conditions, followed by 
two hours at 60°C
Specimen Preparation
After the plates had been left to cure for at least five hours, they were removed 
from the mold. The edges were trimmed off to eliminate any edge anomalies. All 
specimens were cut using a water-cooled diamond saw. The coupon edges were then 
polished in sequential steps down to 400 grit emery paper from Buehler. This helps to 
suppress, but not eliminate, edge generated delamination. For coupons run to more than
33
IO5 cycles, fiberglass tabs were bonded onto the ends with Hysol EA 9309.2NA adhesive. 
The coupons were then post-cured a second time at 60 °C for two hours. Tensile coupons 
were 20 cm long by 2.5 cm wide with a gage section of 13 cm. Compressive coupons 
were 10 cm long by 2.5 cm wide with a gage section of 5 cm.
All coupons were tested for the fiber content using a matrix bum-off method 
described under ASTM D 2584. The only deviation from the ASTM standard was in the 
amount of material used in the bum off test, 15 to 20 grams rather than the prescribed 5 
grams.
Test Facility and Development
All of the coupons were tested using an Instron 8501 servo-hydraulic machine. 
This machine allowed maximum and minimum peak loadings to be accurately maintained 
while varying the loading wave forms and counting the number of cycles applied to the 
specimen. The specimens were clamped into the Instron using hydraulic grips. The 
amount of pressure used to clamp the specimens was varied according to the maximum 
stress experienced by the specimen, to prevent coupon crushing. If, during the course of 
the test, the grip hydraulic pressure gauges fluctuated, which could indicate sample 
slipping, the test would be stopped and additional pressure would be added to the grips. 
This method was used to avoid over-clamping the specimens and causing excessive 
gripping or tab failures.
An extensometer ( Instron 2620-525) was used to determine strains for 
calculation of the initial modulus. The extensometer was attached to the edge of the 
coupon via rubber bands. In all cases the initial longitudinal elastic modulus, Ex, was
34
measured on the thin section of the coupon by taking a least squares fit of at least five 
equally spaced load intervals. The modulus of the thin section then provided a basis for 
calculating the maximum running strain for a given stress.
A minimum of three static tests were performed to obtain an accurate ultimate 
tensile strength for each laminate. These static tests were performed under displacement 
control, with a displacement rate of 12.7 mm/sec. Fatigue tests used a sine-wave cyclic 
waveform with the testing machine in load control. The first coupons were set to run at a 
maximum stress of approximately 60% of the ultimate stress, with a minimum stress of 
10% of the maximum stress, giving ah R value of 0.1 ( R= minimum stress/ maximum 
stress). After these coupons failed, a best fit line through the two points was used to 
predict the approximate lifetime of coupons being run at different stress levels. All test 
coupons were run at frequencies where the coupon temperature did not rise more then 5 
°C above room temperature. A fan was placed approximately one meter away to provide 
additional cooling of the coupons. Fatigue tests were either performed until failure, 
which was defined as the inability of the coupons to carry the maximum load, or until 
delamination at the ply drop extended along the entire length of the coupon and into the
grips.
35
DCB and ENF test specimens
These experiments were run to characterize the pure Mode I and Mode II 
delamination resistance, Glc and Gllc. These lay-ups were [0]10, [(0)5/±45/(0)4], and 
[±45]l0. Figure 8 shows typical DCB and ENF specimens. The main requirement for 
double cantilever beam (DCB) specimens is to ensure that delamination propagates along 
the mid-plane. This was accomplished by inserting a 25 mm wide piece of Fluoro-Peel 
release fabric between the layers of interest to form the starter crack. Once the mold had
Double Cantilever Beam (DCB) Coupon
End Notch Flexture (ENF) Coupon 
Top View
^  Internal Teflon Film
Figure 8. DCB and ENF specimens
time to cure the plate was cut into test coupons with dimensions of 25 mm wide by 200 
mm long, and a nominal thickness of 5 mm. A starter crack, 3 mm long, was introduced
36
into the coupons, to bypass the resin-rich area ahead of the release fabric, with a flat head 
screwdriver. The area where the hinges were to be mounted for loading was then sanded 
to remove any release agent from the molding process. Typical piano stock hinges were 
then mounted to the coupon ends using Hysol EA 9309.2NA adhesive. The end notch 
flexure (ENF) specimens had the same dimensions as the DCB specimens, but with no
hinges.
37
CHAPTER 4
RESULTS AND DISCUSSION
This chapter is divided into four sections. The first section presents results for the 
delamination resistance of various laminates, where the delamination was allowed to 
progress over the entire specimen length. Comparison between various ply drop 
combinations are made in terms of delamination growth rates as well as initiation and 
arrest. The second section looks at the influence of ply drops on the fatigue lifetime of 
the laminates, including design knockdown factors. Appendix A contains the fiber 
content, elastic modulus, maximum running stress and the number of cycles to which 
each coupon was fatigued. Appendix B contains the delamination length versus cycles for 
all of the coupons tested in the first section. The third section presents results for 
composite I-beams containing ply drops, comparing results in these structural 
components with coupon results. The last section presents basic results for critical strain 
energy release rates in Modes I and II delamination using double cantilever beam and end 
notch flexure coupons, respectively.
Delamination Study
Effect of Ply Drop Location
Table I lists details of laminates ESA, ESB, and ESC, each of which contained a 
single O0 ply drop. The ESA laminate has a single exterior O0 ply drop. Static tensile tests
38
Delamination front
Figure 9. Delamination of exterior zero degree ply drop.
produced delamination of the dropped ply over the entire specimen free length, as seen in 
Figure 9, followed by a brooming type of failure of the remaining cross-section, shown in 
Figure 10. This type of dramatic failure is typical for all laminates tested statically, and 
Figure 10 is also typical of laminates without ply drops.
The ESA laminate was tested in tensile fatigue at various maximum stress levels 
at an R value of 0.1 to characterize the fatigue crack growth of the delamination. Figure 1
Figure 10. Typical static failure of laminate with ply drop.
11 shows results for delamination length versus cycles at three maximum stress levels. 
These are results for typical individual tests; results for all tests are given in Appendix B.
39
These coupons delaminated uniformly across the coupon width and along the gage length. 
Measuring the length of the delamination was straight-forward, because the initially
I  25 
£
c  20
CO
CQ.EE
s
Q
10
5
*
Single Exterior Ply
Dropped.
[0*/(0/±45/0)s]►►
* = Dropped Ply
i r  -
■ ► I
« >
_ _______ ▲—▲--------i k-------- A---------iK-------
20,000 40,000 60,000
Cycles
80,000 100,000
Max. Stress: ■207 MPa - * - 1 3 8  MPa - * - 1 2 0  MPa
Figure 11. Delamination length vs. cycles for ESA laminate, R = 0.1, Exterior 0° Ply 
Dropped.
Ply dropv Front Two /Front One\  I /
ESM107
Figure 12. Delamination at an interior ply drop showing two 
delamination fronts.
40
translucent coupons became opaque due to the delamination. In Figure 9, an ESA coupon 
is shown with the exterior 0° delaminated. This type of delamination is typical for all of 
the laminates with exterior ply drops. The coupons showed no other signs of fatigue 
damage. Unlike laminates with an exterior ply drop, interior ply drops developed two 
different delamination fronts. These two delamination fronts can be seen in Figure 12 for 
the ESB laminate [0/0*/±45/0/0/±45/0]. The crack initially started in between the +45° 
and -45° plies under the ply drop, labeled Front One in the Figure 12. The delamination 
then propagated along the length ahead of the second delamination front, labeled Front
[0/0*/±45/0/0/±45/0] 
* = Dropped Ply
50,00030,000 40,00020,00010,000
Cycles
310 MPa —A— 275 MPa344 MPaMax. Stress:
Figure 13. Delamination length vs. cycles for ESB (Single interior 0°ply drop) laminate, 
R=0.1.
Two, which developed between the exterior zero ply and the ply drop layer. For the 
interior ply drop, in the ESB laminate, delamination required about twice as high a
41
maximum stress level as for the ESA laminate, apparently due to the two shear surfaces, 
as discussed later. The delamination rate for various stress levels for both the ESB and 
ESC laminates can be seen in Figures 13 and 14. The ESC laminate, with a central ply 
drop, showed a slightly higher delamination resistance than ESB.
Table 2 provides an approximate measure of delamination resistance based on the 
delamination test results. The Table rates delamination resistance for different laminates 
in terms of the maximum strain under tensile fatigue loading to produce a one-inch long
[0/±45/0/0*/0/±45/0] 
* = Dropped Ply
9,000 18,000 27,000 36,000 45,000
Cycles
275 MPaMax. Stress:
Figure 14. Delamination length vs. cycles for ESC (Single center interior 0°ply drop) 
laminate, R=0.1.
delamination in IO5 cycles. The ESA laminate with an exterior ply drop required only 
0.6% maximum strain in fatigue to produce a 25 mm delamination in IO5 cycles, while 
the ESB and ESC laminates required about 1.1% maximum strain. The stresses and 
strains in each case are the values measured and calculated for the thin cross-section side 
of the ply drop. The arrest strain values are obtained from the highest stress where the
Table 3. Comparison of delamination resistance of different ply drop configurations.
Laminate Lay-up % Strain for 25.4 mm 
delamination in IO5 cycles
Arrest
%
strain1
Threshold % 
strain2
ESA [0*/(0/±45/0)s] 0.6 0.5 0.4
ESB [0/0*/±45/0/0/±45/0] 1.1 LI 0.8
ESC [0/±45/0/0*/0/±45/0] LI LI 0.8
ESE . [0*/(0/±45/0)3] 0.6 0.4 0.4
ESF [0/0*/±45/0/(0/±45/0)2] 1.0 1.0 0.7
ESG [0 */0 */(0/±45/0)3] 0.4 —
ESH [0/0*/0*/±45/0/(0/±45/0)2] 0.7 0.6 0.5
ESI [0/0*/0*/±45/0/0/±45/0] I .Ia 1.0 0.7
ESJ [0*/(0/±45/0)s]
“Z-Spiked”
0.8 — —
ESK [0*/(0/±45/0)s] 
Hysol EA 9309.2NA
0.7 — —
ESM ' [0/0*/±45/0/0/±45/0] 
Hysol EA 9309.2NA
1 . 1 B — —
*- no further growth over most of the IO5 cycles 
2- no delamination after at least IO5 cycles.
Fabrics: 0°: D155; ±45°: DB 120 except as noted
Laminates ESO, ESR and ESP not shown, ±45° layers did not delaminate. 
A- Same as ESB, except multiple ply drops.
B- No delamination; however, failure of coupon occurred at ply drop.
43
delamination arrested after IO5 cycles. The threshold strain values are obtained from the
highest stress where no delamination had formed after IO5 cycles. These results are
limited to the IO5 cycle range , and could vary at high cycles.
Effects of laminate thickness and 
multiple ply drops at the same position
Two significant parameters in dropping plies are the percentage of the total 
thickness (and total O0 plies) which are dropped, and the number of plies which are 
dropped at the same position. These parameters were explored with laminates ESD 
through ESH. Laminate ESD had two interior 0°ply drops, while maintaining the same 
basic lay-up as the thin sections of the ESA, ESB and ESC laminates, without adding 
additional 0° plies to be dropped. This resulted in a 33 percent drop in thickness (a 50% 
drop in the percent O0 plies), whereas the ESA, ESB and ESC laminates were only 
tapered 14 percent, with 20% of 0°plies dropped. This interior ply drop did delaminate 
when statically loaded to failure, but no delamination occurred when the coupon was 
fatigued. The laminate simply broke through the thickness at the ply drop. The difference 
can also be seen in the moduli of the thick and thin sections, 22 GPa versus 15 GPa, 
respectively, resulting from a higher percent of 45° material in the thin section (Table I). 
This bounds the practical range of thickness taper that can be used successfully in a 
design. Having too much taper can prevent delamination, but only because the thin 
section can only support very low loads. Only two coupons were tested since no 
delamination occurred prior to tensile failure.
44
The ESE and ESF laminates
The ESE and ESF laminates incoiporated a single 0° ply drop into a base laminate 
that was thicker than the ESA laminate. The percent thickness change in these laminates 
was reduced to 10% as compared to the 14% for the ESA laminate. The ESE laminate 
had an exterior ply drop while the ESF laminate had an interior ply drop. The 
delamination rate Was slightly lower for the ESE laminate than for the ESA laminate 
(Figure 17), but the strain to produce a 25 mm delamination in IO5 cycles was about the 
same, 0.6% (Table 2). The ESF laminate also had a sightly slower delamination rate than 
the ESB laminate, but the strain value in Table 2 was slightly lower at 1.0% for ESF 
versus 1.1% strain for ESB. The delamination mode of both of these laminates was very 
similar to their thinner counterparts.
The ESG and ESH laminates were also thicker in cross-section, plus, in an 
attempt to simulate a manufacturing situation, two plies were dropped instead of one.
The ESG laminate contained two dropped plies on the outside while the ESH laminate 
contained two dropped plies in the interior (Table I). In both cases these configurations 
behaved poorly compared with their single ply drop counterparts. During testing of the 
ESH laminate, the resin rich area ahead of the ply drop was observed to crack extensively 
during the fatigue test. When the coupon was at the maximum stress of the fatigue cycle, 
fracture and fragmentation in the resin pocket occurred. No matrix material was left 
ahead of the ply drop after the initial cycles. This is illustrated in the Figure 15.
The ESG laminate, with a double external ply drop, delaminated more easily than
45
a single ply drop on the outside. The percent strain for a one-inch delamination in IO5 
cycles for ESG and ESH were 0.4% and 0.7%, respectively. These are significantly lower 
than the values for their counterparts single-ply-drop cases, ESE and ESF at 0.6% and 
1.0% strain, respectively (Table 2). A comparison of laminates having internal ply drops 
is presented in Figure 16. As noted, the ESH laminate with two interior ply drops had the 
highest delamination rate. ( The ESK laminate is discussed later).
"Window" through 
resin - rich region
Figure 15. Illustration of resin rich region in ESH laminate.
Effect of Ply Drop Spacing
The ESI laminate was constructed to look at the effect of multiple individual ply 
drops with spacing ranging from 13 to 48 mm, and are compared with the ESB laminate 
having the same ply configuration and a single ply drop. The results for the various cases 
in Figure 17 are scattered with no clear effect of spacing. The closest spacing, 12 mm, 
gives results which are essentially the same as for ESB laminate with a single, interior ply
46
ESB: Single Interior Ply Drop  
ESF: Thicker Lam inate, Single interior 0° ply drop 
- A —  ESH: Thicker Lam inate, Two  interior ply drops
10,000 20,000 30 ,000 40 ,000 50,000
Cycles
Figure 16. Delamination length vs. cycles for laminates ESB, ESH, ESF (all interior ply 
drops) at a maximum running stress of 275 MPa, R=0.1
drop. However, if the two delaminations merge together, then a single system is created 
which might propagate like the case with two dropped plies at the same location. 
Complete data to explore this question were not generated , but typical coupons 
containing multiple ply drops with delaminations can be seen in Figure 18. All the 
coupons were run at the same stress level, 275 MPa, with R = 0.1. A coupon with ply 
drops too close together can be seen at the top of Figure 18 right before both ply drops 
delaminated together. In the lower two coupons the delamination has arrested after some 
crack propagation, prior to the delamination combining into a single system.
47
40
_  35
E
E. 30 
|> 25
! 200
"-M
” 15
1  10a>
Q
5
0
0 10,000 20,000 30,000 40,000 50,000
Cycles
Figure 17. Effect of different spacing between ply drops, R=0.1, ESI laminate (Two 0° 
ply drops) at 276 MPa.
ESI laminate 
[0/0*/0*/±45/0/0/±45/0] 
* = dropped ply
-$—  E S H : 12 mm between ply drops 
* —  ESI2: 24 mm between ply drops 
-A—  ESI3: 36 mm between ply drops 
* —  ESI4: 48 mm between ply drops 
-  -  ESB: Single, interior ply drop
Ply drop spacing 13 mm
t ; .■«: I: ;iif
ESI 1103
Ply drop spacing 38 mm
ESI3104
Ply drop spacing 51 mm
ESI4103
Figure 18. ESI coupons run at 276 MPa, R=0.1.
48
35
30
I 25
I  20
I 15
.E
I 10 *
A
i i 
t i
— ESJ (Z-Spiked) at 344 MPa 
ESJ (Z-Spiked) at 207 MPa 
—A— ESA (Single Ext. 0° Dropped) at 207 MPa
i I  I I
__________ ____________ ____________ I
ii
i
i
_ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ — -------- ■PC
0 20,000 40,000 60,000 60,000 100,000
Cycles
Figure 19. Delamination length vs. cycles for ESJ “Z-Spiked” laminate compared to ESA 
(Single, exterior 0° ply drop) laminate, R= 0.1.
Effects of “Z-Spiking”
The ESJ laminate with “Z-Spiking” is not shown in Figure 19, which compared
laminates with exterior ply drops at a maximum stress of 138 MPa, because there was 
nodelamination until more than 200,000 cycles. Delamination at higher stresses in the
“Z-Spiked” laminate initiated locally around the tows and then proceeded into the ply 
drop. The delamination initiation took longer than for a normal exterior ply drop.
Delamination rates at various stress levels for the “Z-Spiked” laminate can be
seen in Figure 19 compared with data for the ESA laminate. At the lowest stress level 
shown using the “Z-Spiked” configuration (207 MPa), the ESA laminate completely 
delaminated rapidly, but the “Z-Spiked” laminate had reached a level where the
49
delamination arrested. The laminates with interior ply drops and “Z-Spiking” showed a 
definite improvement in delamination resistance over laminates that were constructed 
normally. The ESM laminate, with a “Z-Spiked” internal ply drop, was not successful in 
preventing delamination. A delamination immediately formed in the area where the ±45°
ESA  Single, exterior ply drop 
ESE  Thicker lam inate w /  single, ext. ply drop 
ESG  Thicker laminate with two, ext. ply drops 
ESK  Single, exterior ply drop w / Hysol AdhesiveE 30
c  15
50,000 100,000 150,000 200,000 250,000
Cycles
Figure 20. Delamination length vs. cycles for laminates ESA, ESK, ESG, ESE (all 
exterior ply drops) at a max. running stress of 138 MPa, R=0.1.
layer was completely severed to allow the ply drop to be inserted. The delamination then
propagated at the same rate as a typical ESB laminate.
Effect of Tough Adhesive at Ply Drop
The ESK and ESL laminates had single ply drops where the ply drop area was 
impregnated with Hysol EA9309.2NA epoxy adhesive prior to resin impregnation. In 
Figure 20, the ESK laminate with the Hysol adhesive is compared with three other 
laminates that incorporated an exterior ply drop. The ESK laminate showed a significant
50
Figure 21. ESL laminate in tension at the ply drop (Interior O0 ply drop 
with Hysol adhesive) failed at 276 MPa.
improvement over the other configurations tested. The delamination in the ESK laminate 
did not initiate until the other laminates had completely delaminated. The delamination 
did not initiate in the adhesive, but at the point right behind the adhesive, in the thicker 
section. Once the crack had propagated to the grip, the adhesive failed, totally separating 
the plies. The ESL laminate, with an interior ply drop with adhesive, showed no visible 
delamination prior to coupon failure. A typical failure can be seen in Figure 21. The 
region where the adhesive was allowed to cure did not delaminate prior to total laminate 
failure. Failure occurred near the ply drop location, apparently due to the stress 
concentration from the ply drop. This type of prevention technique provided significant 
delamination prevention, while not reducing the life of the coupon when compared with 
the ESB laminate at the same number of cycles, as discussed later.
Effect of Butt-Joints
The ESN and ESO laminates were used to investigate the effect of butt-joints on
51
the delamination properties of typical laminates used in this thesis. The ESN laminate 
contained an exterior 0° butt-joint. This had the same effect as dropping a ply, as the ply 
delaminated at the same rate as for the ESA laminate. However, the ESO laminate 
contained an interior ±45° layer butt-joint. The ±45° layer did not delaminate and caused 
no damage which influenced the lifetime of the coupon. This suggests that dropping 
±45° layers makes for a less delamination critical design. This is not suiprising since zero 
degree layers carry the majority of the load and have the greatest influence on stiffness. 
However, dropping ±45° layers can only be used to a limited extent, since they are not 
present in great numbers in this class of laminates.
Effects of ±45° Ply Drops
To confirm the effect of dropping ±45° degree layers on delamination rate, the 
ESR laminate was laid up similar to the ESC laminate. However, instead of incorporating 
two interior 0° ply drops, the ESR laminate incorporates two ±45° ply drops along the 
centerline. This laminate showed no signs of delamination into the thick section as did all 
the other laminates, as the delamination actually propagated into the thin section. This 
delamination may have been artificially influenced by the high degree of tapering, 25% 
of the thickness, which caused a severe shape change.
To get a better estimation of the influence of ±45° layer ply drops, the ESP 
laminate was manufactured with two ±45° ply drops (four total plies). The laminate 
showed no signs of delamination, even after 300,000 cycles at maximum stress of 276 
MPa, whereas the ESH laminate, with two interior 0° layers, had delaminated. However, 
after this many cycles the coupons had premature end failures. This laminate shows
52
results that are consistent with the laminate containing the ±45° butt-joint. The ±45° 
layers are much more compliant than 0° layers and are less prone to delamination at ply 
drops.
Effect of Al 30 warp unidirectional fabric
The ESQ laminate used the same laminate lay-up as the ESH laminate, but made 
use of the Al 30 unidirectional woven warp direction fabric in place of the D155 weft 
fabric used in other laminates. The delamination could not be measured with any 
consistency from coupon to coupon. The Al 30 fabric is constructed of unidirectional 
strands woven around a transverse thermoplastic bead every 25 to 30 mm. The 
delamination would propagate from one bead to the next and then arrest. After continued 
fatigue the delamination would jump to the next bead. However, the Al 30 fabric overall 
showed less delamination resistance than the Dl 55 fabric, with the delamination 
propagating unstably between beads. A typical example of this is that at 310 MPa level, 
a laminate using the Al 30 fabric would delaminate to the grips in under one thousand 
cycles, whereas a laminate using the Dl 55 fabric would not delaminate that far until after
30,000 cycles.
53
Other attempts to prevent delamination
The JKA “Feathered” and “Random” laminates used the ESA laminate with an 
exterior ply drop as a model to incorporate delamination suppression techniques, in 
addition to the “Z-Spiked” and adhesively bonded ply drops. The results in Figure 22 
show that the random mat had no effect on the delamination rate, while incorporating 
feathering in the laminate is very successful at retarding delamination. The delamination 
of the JKA “Feathered” laminate began at the tip of the tows that had been pulled out so 
they extended beyond the fabric edge, and propagated along those tows until it reached
E S A ,  Si ng l e  ex t e r i o r  ply drop
J KA  ( F e a t h e r e d ) ,  Al t e rna t i ng  t ows  pul l ed out  
12 . 5  m m  on d r o p p e d  ply.
J KA  R a n d o m ,  R a n d o m  m a t b e t w e e n  ply drop  
and con t i nuous  l ayer .
3 0 ,0 0 0 4 0 ,0 0 02 0 , 0 0 0  
C ycle s
I 0 , 0 0 0
Figure 22. Delamination length vs. cycles for ESA, JKA "Feathered" and JKA random 
laminates at a maximum running stress of 207 MPa, R=0.1.
the tows which were not pulled out. The delamination can be seen propagating along the 
pulled out tows in Figure 23. The JKB “Feathered” used the same scheme to prevent
54
tows that are pulled out
Figure 23. Delamination initiating along tows pulled out.
delamination in an internal ply drop laminate. The delamination progressed the same way 
as with the JKA “Feathered”. The JKB “Feathered” delamination rate is compared to a 
standard ESB laminate in Figure 24. The “feathering” has a significant effect on 
reducing the delamination length.
ES B , S ing le , interior O0 ply drop
JKB (Fea the red ) ,  Alternating tows on 
d ro p p e d ply a re pulled o ut 1 2 .5  m m .
4 0 ,0 0 020 , 000 60 ,0 00
Cycles
Figure 24. Delamination length vs. cycles for ESB and JKB "Feathered" at a maximum 
running stress of 275 MPa, R=0.1.
55
Repairing Delaminated Samples
If exterior ply drops are to be used in a design and an overload condition takes 
place, it would be desirable to repair the delamination. Two epoxy adhesives, Hysol EA 
9309.2NA and Hysol EA 9412, were used to repair the delamination. After the initial test 
was run on an ESA laminate, causing delamination, the adhesive was injected into the 
delamination and a C-clamp was applied to the specimen until the adhesive had cured. 
The specimen was then retested at the same stress level and the delamination rate was
ESA1 Single exterior 0° ply drop 
«— Repaired with Hysol EA 9309.2NA 
■*— Repaired with Hysol EA 9412
.2 15
10,000 20,000
Cycles
30,000 40,000
Figure 25. Delamination length vs. cycles at 207 MPa, R=O.I, initial ESA laminate 
compared to repaired ESA laminate.
compared to the original rate, which can be seen in Figure 25. From Figure 25, the two 
adhesives had a similar delamination rate and provided a significant reduction in 
delamination rate compared with the original laminate.
56
So far the only attempts to repair delamination have involved exterior ply drops, 
since interior ply drops present an access problem. One possible method of repair would 
be drill a hole into laminate that is smaller than the critical flaw size and draw a vacuum 
around the hole. This would draw the repair adhesive into the delamination and possibly 
extend the lifetime of the component. (This has not yet been demonstrated.)
Comparing different laminates
In order to make this work useful to a designer needing to incorporate ply drops 
into a design, Table 2 compares the various laminates, by listing the percent strain needed 
to cause a 25 mm delamination at IO5 cycles. By knowing the apparent penalty for using 
different ply drop configurations, more efficient designs can be developed. For example, 
by using an exterior ply drop instead of an interior ply drop, there would be an 83% 
reduction in the strain carrying capability of the laminate. Another example would be if 
two plies are dropped on top of one another instead of the plies being distributed along 
the length, there would be an 57 % reduction in the strain carrying capability of the 
laminate. Similar conclusions are evident for the strains to arrest a delamination and for
the threshold strain to initiate a delamination, also listed in Table 2.
57
Effect of Plv Drops on Lifetime
In the previous sections, different lay-ups were tested in relatively short fatigue 
tests to determine the rate of delamination. This section describes a study of the ESB and 
ESH configurations at extended fatigue cycling, in terms of the effect of ply drops on 
total coupon lifetime. The ESH laminate consisted of the lay-up 
, [0/0*/0*/±45/0/0/±45/0/0/±45/0], where the asterisk plies are being dropped. The ESB 
laminate, with a single ply drop, had a lower delamination rate than the ESH laminate in 
the delamination fatigue tests.
Fiber Volume Content Effect
The results are represented in terms of conventional S-N fatigue data sets for 
control and ply drop cases. In Figure 26 the performance of the ESH laminate at two 
different fiber contents can be seen. The fiber volume content had little effect on the 
laminate with ply drops, contrary to the significant effect reported for control materials 
[50]. The low fiber content control DD5 laminate with no ply drop has an S-N curve 
slope of about 10% of the static strength per decade of cycles. This fatigue performance is 
optimal for glass fiber composites, and is used as a standard to compare other laminates 
[50]. The delamination rates and lifetimes for the laminates containing ply drops are 
about the same for both low and high fiber contents. This is contrary to fiberglass 
laminates without ply-drops previously tested (DD5 & DD7), Mandell, et.al. [50], which
58
show improved normalized fatigue performance at lower fiber contents. This trend can 
also be seen in Figure 26 for control materials. From this figure, the good performance at 
lower fiber content is lost due to the presence of ply drops. However, the performance of 
higher fiber content laminates containing ply drops is not affected. The effect of dropping 
plies can be seen when comparing the ESH laminate to the material performance of the 
DD5 laminate. While the DD5 laminate was capable of running at one percent strain at a 
million cycles, both ESH laminates were reduced to 0.63% strain at one million cycles,
S5
CO
</)
i f )0).b
CO
I
I
o
Z
xro
1.2
1
0.8
0.6
0.4
0.2
0
Laminate So, MPa Vf (%)
♦ DD7 832 54
■ DD5 724 36
a ESH High 798 52
•  ESH Low 747 35
r
AAm
A A
■  ■
♦
•^
A A A
<♦ A * ]
■■
•
1.E+00 1.E+01 1.E+02 1.E+03 1.E+04 1.E+05 1.E+06 1.E+07
Cycles to Failure
Figure 26. Effect of fiber content on the normalized S-N Data, R=0.1, for DD materials 
[0/±45/0]s compared to ESH laminate (Two interior 0° ply drops).
59
similiar to the high fiber content control material, DD7. By introducing ply drops, the 
effective strain carrying capacity of the laminates was reduced by forty percent.
The ESB and ESH laminates delaminated to the grips before the coupon failed. 
After coupons had been fatigued for IO6 cycles they were labeled as run-outs. In the 
majority of these coupons the delamination had not propagated to the grips. The first 
section of the thesis presented levels of stress where delamination had arrested, but the 
delamination might have been growing at a slow enough level where it just appeared to 
have arrested. Threshold levels of strain, where delamination did not initiate before IO6 
cycles, were not found. These strain levels would be low enough to prevent the laminate 
from being used.
A l30 fabric vs. D l55 fabric j
While the Dl 55 unidirectional fabric has the best fatigue properties at low fiber 
content of currently available fabrics, it is only available in the weft direction with a 
maximum width of 127 cm. While the D155 fabric is held together by stitching, which 
does not cause fiber waviness, the A130's have a thermoplastic bead which runs 
perpendicular across the tows to keep them together. Since, currently, this would be the 
fabric of choice for the zero degree layers in composites longer than 127 cm, the ESQ 
laminate was constructed using this fabric to see how the A130's would behave with ply 
drops in the laminate. When control laminates containing the two different O0 fabrics, 
DDl I and DD6 were compared, the fatigue properties were approximately the same, as 
can be seen in Figure 27.
60
700
_  600  
ra'
CL
5  500
M 
(Z)
2 400
55
I  300  
E
I  200  
100 
0
1.E+00 1.E+02 1.E+04 1.E+06 1.E+08
Cycles to Failure
Figure 27. S-N curve (R=0.1) comparing D155 Fabric (DD6) to A130 Fabric (DD11) 
control materials with no ply drops, from Reference 50.
Il 
11
_ L _
D D 6 & D D 1 1  Lay-up [ 0 / ± 4 5 / 0 l s
Il
Il
Fiber content of both 
laminates was 31 %
■
■■
■
106 percent strain for both 
laminates is 1 .0% ♦ DD6 ■ DD11
800
CL5
(Z )
55 400 
E3
E
x
CD
5  200
0
1.E + 00 1.E + 02 1.E + 04 1.E + 06
Cycles to failure
Figure 28. Fatigue life of ESH (Two D155 ply drops) vs. ESQ (Two A130 ply drops) 
laminates (Maximum stress is on the thin side of the ply drop).
♦ ESQ (Vf=31 %) 
■ ESH (Vf= 35%)
106 cycle strain: 0.7% for both ESH 
and ESQ laminates.
ESH and ESQ laminates 
[0/0*/0*/±45/0/0/±45/0/0/±45/0] 
*= Dropped Ply
ESQ used A 130 fabric for 0° layers. 
ESH used D155 fabric for 0° layers.
61
800
_  600  
ro
CL5
I
S  400 
E3
EI
5  200
!
▲ DD11: [0/±45/0]s
A ■  ESQ:[0/0*/0*/±45/0/(0/±45/0)2]
I l
▲ AA
M km a
■ ■  A k  A
■  ■ A
106 cycle strain; 
A 130 fabric 0° 's
ESH 0.8%, DD11 1.0%
DB 120 fabric ± 4 5 's , Vf= 31%
—
1.E+00 1.E + 02 1.E + 04
Cycles to Failure
1.E+06 1.E+08
Figure 29. Tensile fatigue S-N data for ESQ (Two, O0 internal ply drops) vs. control 
DDl I (no ply drops) laminates, A130 unidirectional fabric.
Introducing ply drops into the laminates appears to have a negligible effect. In Figure 28, 
the performance of the laminates with A130 and D155 O0 ply drops is very similar, with 
the greatest difference shown in the static (I cycle) strength, where the A130's were 
lower. Figure 29 compares the A l30 based laminate with and without ply drops; as in 
Figure 26, the ply drops at low fiber content lower the fatigue life, but to a lesser extent in 
Figure 29. The fatigue properties of the Al 30 fabric might be higher if the thickness of 
the thermoplastic bead could be reduced or the spacing increased. In Figure 30, the 
damage zone appearing around those beads can be seen.
62
thermoplastic bead.
Figure 30. Damage developing around thermoplastic beads 
as a result of fatigue loading, R=0.1.
Resin Rich Region
The size of the resin rich area at the ply drop has a large effect on delamination 
properties. In Figure 31, two resin rich areas can be seen, due to ply movement which 
caused the two ply drops to be misaligned. The delamination initiates at this resin rich
Figure 31. Photomicrograph of ESH laminate showing resin rich regions 
ahead of ply drops.
region and propagates into the ply drop layers. A typical crack in an ESH laminate can be 
seen in Figure 32. The delamination on the thicker side grows above and below the ply 
dropped layers, not between the two ply drops. Also, the crack propagates a short distance
63
into the thin cross-section and then arrests.
The more plies dropped, the worse the fatigue properties of the laminate, as can be
Figure 32. Side view of crack propagating through ESH 
laminate
seen in Figure 33. Here the fatigue properties of the ESB laminate are compared with the 
ESH laminate. The ESB laminate has only one interior ply drop, which forms a smaller
900
800  
„  700
CD
Q-
&  600
i 500
CD
• |  400
C
K 300  
x(D
2  200 
100
0
1.E + 00 1.E + 02 1.E + 04 1.E + 06 1.E + 08
Cycles to Failure
Figure 33. Tensile fatigue (R=0.1) S-N curves for ESB (Single 0° internal ply drop) and 
ESH (Two interior 0° ply drops).
resin pocket than the ESH laminate which drops two plies. The strain energy release rate
♦  E S B 1 [0 /0*/±45/0 / /0 /±45/0]
A ESH, [0/0*/0 */±45 /0 /0 /±45 /0 /0 /±45 /0 ], "=Dropped Ply
D 155 fabric 0°'s, DB120 fabric ±45's, Vf=44%
A
A
A #  ♦
------- A------
IO6 cycles strain; ESB @ 0.8% strain, ESH @ 0.6% strain 
---------------------- 1----------------------------- "----------------------------
64
will also increase by dropping more layers. The ESB laminate was able to operate at
0.8% strain out to a million cycles, while the ESH laminate was just above 0.6% strain.
Compression Testing
Due to the unsymmetric geometry caused by the ply drop, when a coupon is 
loaded in axial compression, it tends to buckle. The ESH laminate was studied since it 
had a thicker cross-section to resist buckling, while at the same time giving a worst case 
scenario, with two interior ply drops being terminated at the same location. Another 
problem also arises due to the short specimen used in compression testing. A 13 mm gage 
length specimen with a ply drop and delamination may be influenced by the grip pressure 
or stress concentrations associated with the grip. Considering these two problems, the 
challenge is to have a large enough gage length to create a uniform axial strain field, 
while avoiding buckling at the test load.
The first approach taken to determine the tendency of a compression coupon to 
buckle was by calculating its slenderness ratio. The slenderness ratio (SR) is calculated 
by [50]
SR=
M l ,
t (4)
Where Le is the effective length of the test coupon, which for fixed-fixed end conditions 
is equal to the length multiplied by 0.5, t is the composite thickness in the gage length. A 
study by Adams and Lewis [51] indicates that a slenderness ratio less than 30 was not
65
prone to buckling failure in glass epoxy unidirectional composites with uniform cross 
sections.
For the control coupons tested, Table 4 summarizes the results using Equation 
4. While the numbers calculated are less than the proposed numbers for buckling, strain 
gauges were used to confirm whether buckling was occurring.
Table 4. Slenderness ratios calculated using Equation 4
Slenderness Ratios calculated from Equation 4
Thickness (t) 13 mm Gage length 25 mm Gage length
Thick Section Control 4.06 mm 5.5 10.6
Thin Section Control 3.30 mm 6.8 13.1
Strain Gage  1 
Strain Gage  2
P e r c e n t  Strain
Figure 34. Compressive stress vs. percent strain for ESH 804 coupon.
66
BLH Electronics strain gages (FAE-25-12-513EL) were mounted on the front and 
back sides of the coupons to determine when the coupon starts to buckle. In Figure 34, 
the coupons were loaded to less than 70 MPa before buckling occurred. This had the 
result of reducing the ultimate compressive strength (UCS) of the coupons, from 
approximately 579 MPa for control material with no ply drop to 386 MPa with a ply drop. 
To verify the strength of the laminate without ply drops, control samples were made from 
the thick and thin sections of the ESH plate, using both 25 mm and 13 mm gage length 
coupons. The 13 mm gage length was tested to see if the 25 mm gage length was 
influencing the strength.. The results from these control samples showed that there was 
no decrease in strength in the thick control samples using either the 25 mm or 13 mm 
gage lengths. However, when the thin cross-section control coupons were tested using 
the 25 mm gage length the coupons decreased in UCS from 579 MPa down to 489 MPa. 
Therefore, when the coupon with the ply drop is tested, the section of the gage length 
containing the thin section is only 13 mm in height, which means it should not be 
affecting the strength. This suggests that the ply drop test coupon probably does not have 
a uniform strain field due to the presence of the ply drop.
Table 5. Comparison of ESH laminate compressive strength with and without ply drops
Comparison of ESH laminate compressive strength 
with and without ply drops
Gage length 
(mm)
Thin control 
section 
strength, MPa
Thick control 
section 
strength, MPa
Specimen with 
ply drop 
strength, MPa
Ply drop 
specimen 
bonded back- 
to-back, MPa
13 551 558 —1 —
25 489 558 386 550
67
Delamination of the coupons in compression occurred at a low number of cycles 
relative to expectations from tensile tests. In order to obtain more valid tests, two ESH 
coupons were bonded together back-to-back, using Hysol EA9309.2NA adhesive, to 
increase the stiffness of the coupon and avoid non-symmetry. These tests showed 
compressive behavior results that were similar to control cases in Table 5, and also 
similar to tests with ply drops located in the flange material of the beams described later. 
When compared to tensile results with ply drops, compressive behavior is less 
delamination resistant. At strain levels (0.6%) where delamination arrested in the tensile 
coupons, the compression coupons totally delaminated. The delamination of compression 
coupons did not show any stable delamination growth, similar to Mode II, ENF 
specimens discussed later. Typically, the fatigue coupons would not show any 
delamination until a few hundred cycles before complete delamination over the entire 
gage length. It was then decided to explore compression behavior in the context of beam 
flanges, which are more typical of blade structures.
Residual Strength of Coupons
Another interesting aspect of laminates containing ply drops is how well they 
retained their residual strength during fatigue, prior to failure at the fatigue condition. 
While the delamination in tension specimens was extensive, reaching into the tabbed area 
of the coupons, the overall strength after delamination was only reduced by 15% relative 
to the original coupon strength. In Table 6, the residual strength of ESH coupons is 
compared at various fractions of the cyclic lifetime (n/N0).
68
The high residual strength values give the opportunity to inspect for damage, such
as delamination from ply drops caused by overload conditions, while still maintaining the
strength to carry the load However, even though the strength might remain close to the 
• • *1initial strength tests in laboratory conditions, by introducing delamination, the laminate is 
now more susceptible to environmental effects, which can rapidly reduce the properties of 
fiberglass laminates.
I
Table 6. Residual strength of ESH laminate after being fatigued (R=OT)*
Residual Strength of ESH laminate after being fatigued
Coupon# Fiber
content
%
Max. Cyclic 
Stress, MPa
Cycles IVN0 Initial
Strength,
S05MPa
Residual 
Strength, Sr 
MPa
sys0
ESH 205 36 276 40,000 0.8 703 600 0.853
ESH 213 36 276 20,000 0.4 703 675 0.960
ESH 409 44 207 1.1E6 1.1 746 686 0.920
ESH 404 44 176 1.1E6 0.11A 746 717 0.961
A- Lifetime estimate used was IO7 cycles, however test was stopped at 106cycles after no delamination.
* Individual Specimen results
70
Plv Drops in Beams
Ply drops are a critical issue for the tapered sections of composite structures. 
Structures may differ from test coupons in several respects: (I) size, (2) stress field 
characteristics, and (3) manufacturing details. This part of the study was intended to 
validate the findings of the coupon studies in the context of a small I-beam sub-structural 
element developed in other parts of the research program [52,54], Ply drops were 
incorporated into the tensile and compressive flanges of the beam, which was then loaded 
in four-point-loading flexural fatigue. Thus, results were obtained from both the tensile 
and compressive areas of the beam simultaneously, for comparison with coupon data. The 
results for each beam are presented, followed by a comparison with coupon data 
with pictures showing typical delaminations, for specific ply drop data. Beams were 
loaded in one direction only (not reversed), so that the tensile flange experienced an 
R=0.1, and the compressive flange R=IO; these values correspond to the coupon test 
values.
Beam Geometry. Fabrication, and Materials.
Figure 35 shows the four-point beam testing apparatus, and Figure 36 indicates 
the beam coordinate system used in describing the various aspects of the beams. The 
tested beams are summarized in Table 7. Further details of the beams can be found in 
Reference [54]. The beam numbering system used here is relative to the overall beam 
study.
71
Material DD5 was modified to study the effect of dropping exterior and interior
End constraints 
Upper load platform
Figure 35. I-Beam testing apparatus
plies in the beam flanges. The ESA laminate configuration was used in beams 39,44 and
x = 0  mm
Tension flange^.
PfyJDrop x = 610 mm
I  T
Compression flange^ VPly Drop 
x = 114 mm x = 495 mm
= 6 5 -  75 mm 
=  0  m m
Tension flange Y
I ,
Compression flange
Figure 36. Beam coordinate system
Table 7. List of I-beams tested with ply drops
I - Beam Summary of Static and Fatigue Data
Beam Materials 
Flange / Web
Deflection 
Constant, K, kN/m
Max. / Min 
Load, kN
Max. / Min. 
Strain, %
Cycles Failure Notes
39 ESA / CH12 4,804 35.6/3.6 0.62 / -0.64 1,225,650 ESAply drop delamination
40 ESB/CH12 4,787 35.6/3.6 0.96/-0.85 264,137 Compression flange delamination
41 ESG/CH12 5,213 49.0/4.9 0.74 / -0.76 13,997 ESG ply drop delaminatioh
42 ESG / CHl 2 5,524 35.6/3.6 0.55/-0.54 436,508 ESG ply drop delamination
44 ESA/CH12 4,527 35.6/3.6 0.57/-0.58 600,000 ESA ply drop delamination
48 ESA/CH12 4,650 35.6/3.6 0.66 / -0.67 72,000 ESA ply drop delamination
49 ESH/CH12 6,625 57.8/5.8 0.76 / -0.73 25,500 ESH ply drop delamination
50 ESH / CH12 6,054 48.9/4.9 0.63 / -0.63 204,000 ESH ply drop delamination
88.2 (r) — I Compression flange
Web Material (3.0 mm) CH12- (±45/0/±45)s- Fabric 0's- D155 (39%),± 45's - DB240 (61%), Vf= 0.34, E n = 17.7 GPa
ESA -[O*/(0/±45/O)s I  Fabric 0's- D155 (71%),± 45's - DB 120(29%), Vf= 0.35, Ecoupon = 24.1 GPa
ESB - [0/0*/±45/02/±45/0], Fabric 0's- D155 (71%),± 45's - DB 120 (29%), Vf= 0.35, Ecoupon = 24.1 GPa
ESG - [0*/0*/(0/±45/0)3 ], Fabric 0's- D155 (73%),± 45's - DB 120 (27%), Vf= 0.44, Ecoupon = 33.1 GPa
ESH - [0/02*/±45/02 /±45/02/±45/0], Fabric 0's- D155 (73%),± 45's - DB 120 (27%), Vf= 0.44, Ecoupon = 33.1 GPa
(r) - residual ultimate strength
73
48, ESB in beam 40, ESG in beams 41 and 42, and ESH in beam 49 and 50. The beam 
flanges were fabricated with the smooth surface on the inside (web side), and the 
thickness steps on the outside. The respective material lay-ups are detailed in the 
following sections. During the fatigue tests, the length of delamination on the tension and 
compression flanges was determined by averaging the two flange edge delaminations, 
respectively, as each flange contained two ply drops.
Flange laminate ESA
Beam 39. Beam 39 was fatigued at a rate of 4 Hz with a maximum load of 35.6 
kN and a minimum load of 3.6 kN. The maximum load produced an initial maximum 
tension flange measured strain of 0.62% and a minimum compressive flange strain o f- 
0.64%. The initial beam stiffness was measured as 4,804 kN/m. The flange material was 
the ESA laminate, which had a single, exterior ply drop (ply I, Table 8). The ply drops on 
the tension flange were at x = 277 and 328 mm. The ply drops on the compression flange 
were at X = 292 and 314 mm. On the tension flange, the delamination started immediately 
between the surface 0° ply and first internal 0° ply (ply I and 2, Table 8). This 
delamination was uniform across the width of the tension flange. The delamination on 
the tension flange continued to grow at an average rate of 6.95E-5 mm/cycle (Table 9) 
until the beam was taken out of the apparatus after 1,225,650 cycles. The tension flange 
itself did not delaminate from the web. The compression flange did not start to show 
delamination until approximately 11,000 cycles. The compression flange delamination 
also occurred between the two top 0° plies (plies I and 2, Table 8) and grew 
approximately 30 mm in length over 600,000 cycles and then arrested for the duration of
74
the test. Table 9 lists the average delamination length versus cycles corresponding to the 
lines on the beam in Figure 37. No shear stiffener, torsional stiffener or other flange 
damage or delaminations were visible. Indicated materials such as DD5 and CH12 are 
described in Refs. 49 and 54.
Table 8. Reference notation for Beams 39, 44 and 48 with ESA laminate
Reference Notation for Beams 39, 44 and 48 with ESA Laminate
Ply
Number
Ply
Angle
Fabric Description
I 0° D155 Dropped Ply
2 0° D155
DD5 material
3 +45° DB 120
4 Lh
5 0° D155
6 0° D155
7 +45° DB 120
8 -45°
9 , o ° D155
Adhesive Layer Hysol EA 9309.2NA, 0.1 - 0.4 mm
10 +45° DB240
Web material CH12 
I - Beam shape, 3 mm thick, 
Vf = 0.35
11
OinI-
12 0 D155
13
OinI- DB240
14 +45°
75
Table 9. Delamination length vs. fatigue cycles for Beam 39
Cycles Average Tension 
Flange Delamination 
Length, mm
Average Compression 
Flange Delamination 
Length, mm
500 6 0
5,000 10 , 0
10,000 ' 14 0
41,000 20 6
50,000 22 7
93,000 28 7
■ 171,000 36 11
390,000 51 30
600,000 55 30
812,000 66 —
1,200,000 79 —
76
Beam 44. Beam 44 was fatigued at a rate of 4 Hz with a maximum load of 35.6 
kN and a minimum load of 3.6 kN. The maximum load produced an initial maximum 
tension flange strain of 0.57% an a minimum compressive flange strain of -0.58%. The 
initial beam stiffness was measured as 4,527 kN/m. This beam had the same exterior ESA 
ply drops as Beam 39, except the ply drops were positioned over the ends of the shear 
stiffeners. The tension flange ply drops were at x = 152 and 455 mm (Figure 36). The 
compression flange ply drops were at x = 222 and 383 mm. This ply drop position did not 
allow the flanges to reach a uniform stress condition away from the influence of the load
Cycles
X = IQSmm | x = 328 mm
x = 277 mm x = 406 mm
x = 314 mm
Figure 37. Beam 39 with ESA laminate for flange material.
77
introduction and geometry changes. The average delamination length versus cycles is 
listed in Table 10. The tension flange had a calculated average delamination rate of 
2.66E-5 mm/cycle, while the delammation on the compression flange did not start 
growing until 185,000 cycles, grew suddenly, and arrested at 5 mm. No shear stiffener, 
torsional stiffener or other flange damage or delaminations were visible. The beam is 
shown in Figure 38.
Table 10. Average delamination length vs. cycles for compression and tension flanges for 
Beam 44.
Cycles Average Tension 
Flange Delamination 
length, mm
Average Compression 
Flange Delamination 
length, mm
200 3 —
.2,000 4 - -
6,300 5 - -
40,000 8 —
79,000 11 - -
184,000 13 5
213,000 14 5
289,000 16 5
382,000 16 5
600,000 16 5
78
Beam 48. Beam 48, again with the ESA laminate configuration, was fatigued at a 
rate of 4 Hz with a maximum load of 35.6 kN and a minimum load of 3.6 kN. The 
maximum load produced an initial maximum tension flange strain of 0.66% an a 
minimum compressive flange strain of -0.67%. The initial beam stiffness was measured 
as 4,650 kN/m. Unlike Beam 44, the ply drops were situated over the constant moment 
section of the beam with the tension ply drops at x = 272 and 327 mm and the 
compression flange ply drops were at x = 295 and 320 mm. The delamination on the
Tension flange
x = 455 mm
I
Compression flange
I I
x = 222 mm x = 383 mm
50 mm
Figure 38. Tension and compression flange on Beam 44.
tension flange was uniform across the width and propagated at an average rate of 6.3E-4 
mm/cycle. The compression flange did not start to delaminate until approximately 2,500 
cycles and grew at an average rate of 5.513-5 mm/cycle. The web and shear stiffeners 
showed no damage after 72,000 cycles. The delamination length versus cycles is shown in
79
Table 11 and the beam is shown in Figure 39.
Table 11. Delamination length vs. cycles for tension and compressive flanges on Beam
Cycles Average Tension Average Compression
Flange Delamination Flange Delamination
length, mm length, mm
2,500 11 ——
4,000 13 —
6,000 16 —
7,500 16 —
10,000 19 —
16,000 22 —
37,500 34 4
72,000 45 6
80
Tension flange 
x = 272 mm x = 327 mm
x = 295 mm x = 320 mm
50 mm
Figure 39. Beam 48, ESA laminate, tension and 
compression flanges.
Flange laminate ESB
Beam 40. The flange laminate of beam 40 was ESB , which had a single, interior 
ply drop, that was located between plies I and 3, as shown in Table 12. It was fatigued at 
a rate of 3 Hz with a maximum load of 35.6 kN and a minimum load of 3.6 kN. The 
maximum load produced an initial maximum tension flange strain of 0.96% and a 
minimum compressive flange strain of -0.85%. The initial beam stiffness was measured 
as 4,787 kN/m. The tension flange ply drops were located at x = 145 and 452 mm, while 
the compression flange ply drops were located at x = 226 and 376 mm. The compression 
flange delaminated from the web after 264,137 cycles within the adhesive layer between 
x = 150 and 690 mm which led to subsequent web failure. No ply drop delaminations 
were visible after this failure. No other shear stiffener, torsional stiffener or other flange 
damage or delaminations were visible. Figure 40 shows the failed beam.
81
Table 12. Reference notation for Beam 40 with ESB laminate
Reference Notation for Beam 40 with ESB Laminate
Ply
Number
Ply
Angle
Fabric Description
I 0° D155
2 0° D155 Dropped ply
3 +45° DB 120
DD5 Material
4 -45°
5 0° D155
6 0° D155
7 +45° DB 120
8 OT
9 0° D155
Adhesive Layer Hysol EA 9309.2NA, 0.1 - 0.4 mm
13 +45° DB240
Web Material CH12 
I - Beam shape,
3 mm thick 
Vf = 0.35
14 -45°
15 0° D155
16 OT DB240
17 +45°
Flange laminate ESG
Beam 41. The flange of beam 41 was the ESG laminate, which had two exterior 
0° ply drops, (plies I and 2, Table 13). It was fatigued at a rate of 3 Hz with a maximum 
load of 49 kN and a minimum load of 4.9 kN. The maximum load produced an initial 
maximum tension flange strain of 0.74% and a minimum compressive flange strain o f- 
0.76%. The initial beam stiffness was measured as 5,213 kN/m. The tension flange ply 
drops were located at x = 282 and 327 mm and the compression flange ply drops were at
82
Tension flange
X= 145 mm
I I
x = 226 mm x = 376 mm
Compression flange
50 mm
Figure 40. Beam 40
x = 287 and 331 mm. On the tension side of the beam the delamination originally started
between the 0° ply (ply 2, Table 13) and the +45° (ply 3, Table 13), but continued 
between the 0° (ply 5, Table 13) and the -45° ply (ply 4, Table 13). The average 
delamination length versus cycles is listed in Table 14. The calculated average 
delamination rate on the compression side was 1.75E-3 mm/cycle, while the rate on the 
tension flange was calculated as 4.97E-3 mm/cycle. The test was stopped after the 
delamination length extended into the shear stiffener area of the beam. No shear stiffener, 
torsional stiffener or other flange damage or delaminations were visible. Figure 41 shows
the delaminated beam.
83
Table 13. Reference notation for Beams 41,42 with ESG laminate
Reference Notation for Beams 41,42 with ESG Laminate
Ply
Number
Ply
Angle
Fabric Description
I 0° D155 Dropped plies
2 0° D155
3 +45° DB 120
4 -45°
5 0° D155
6 0° D155
7 +45° DB 120
8 O
i"
9 0° D155
10 0° D155
11 +45° DB 120
12 Ln O
13 0° D155
Adhesive Layer Hysol EA 9309.2NA, 0.1 - 0.4 mm
14 +45° DB240
I - Beam shape, 
3 mm thick, 
Vf = 0.35
15 -K Ln O
16 0° D155
17 -45° DB240
18 +45°
84
Tension flange
x = 215 mm x = 327 mm x = 415 mm
x = 282 mm xStraingage 
Compression flange
1 x = 331 mm 
x = 287 mm
50 mm
Figure 41. Beam 41
Table 14. Cycles vs. average tension and compression flange delamination length for 
Beam 41.
Cycles Average Tension 
Flange Delamination 
length, mm
Average
Compression Flange 
Delamination length, 
mm
20 11 —
200 13 —
600 21 —
1,000 25 —
1,500 29 —
2,100 33 3
3,700 34 3
13,997 70 21
85
Beam 42. The only difference with Beam 42 as compared to Beam 41 is the strain 
level which the beam were run. Beam 42 was fatigued at a rate of 4 Hz with a maximum 
load of 35.6 kN and a minimum load of 3.6 kN. The maximum load produced an initial
maximum tension flange strain of 0.55% an a minimum compressive flange strain of - 
0.54%. The initial beam stiffness was measured as 5,524 kN/m. The tension flange ply 
drops were located at x -  280 and 325 mm and the compression flange ply drops were at 
x = 280 and 328 mm. On the tension side of the beam the delamination originally started 
between the 0° ply (ply 2, Table 13) and the +45° (ply 3, Table 13), but continued 
between the 0° (ply 5, Table 13) and the -45° ply (ply 4, Table 13). The average 
delamination length versus cycles is listed in Table 15. The average calculated 
delamination rate on the
Table 15. Cycles vs. average tension flange delamination for Beam 42.
Cycles Average Tension Flange 
Delamination length, 
mm
Cycles
(cont.)
Average Tension Flange 
Delamination length, 
mm
500 7 62,000 38
1,000 12 85,000 43
2,000 12 89,000 46
3,500 14 152,000 61
23,000 22 341,000 89
35,000 29 381,000 109
62,000 38 436,508 128
tension flange was 1.23E-4 mm/cycle. The compression flange did not start to delaminate 
until almost 100,000 cycles and grew 9 mm with an average calculated delamination rate
86
was 3.84E-5 mm/cycle for the remaining cycles. The delaminated beam is shown in
Figure 42.
Tension flange Cycles436.000
381.000
341.000
152.000
89.000
115 mm x -  280 mm x = 325 mm x = 452 mm
Compression flange
x = 280 mm x = 328 mm
50 mm
Figure 42. Beam 42, ESG laminate (two, exterior 0° ply 
drops) on flanges.
Flange laminate ESH
Beam 49. The flange material for beam 49 was the ESH laminate, which had two 
internal 0° ply drops (plies 2 and 3, Table 16). It was fatigued at a rate of 3 Hz with a 
maximum load of 57.8 kN and a minimum load of 5.8 kN. The maximum load produced 
an initial maximum tension flange strain of 0.76% an a minimum compressive flange 
strain of -0.73%. The initial beam stiffness was measured as 6,625 kN/m. The tension 
flange ply drops were located at x = 277 and 326 mm and the compression flange ply
87
drops were at x = 290 and 335 mm. The delamination on the tension side first propagated 
into the thin section of the flange, between the 0° and +45° plies (plies I - 4, Table 16), 
with a length of approximately 10 mm. After this initial delamination, the crack started 
to propagate into the thick section, starting two cracks between the 0° plies (ply I and 2, 
Table 16) and the 0° ply and the +45° ply (plies 3 and 4, Table 16). The delamination 
was uniform across the width of the flange and the average calculated tensile side 
delamination rate was 9.3E-5 mm/cycle. Delamination on the compression flange did not 
start until approximately 6,800 cycles. Once the delamination did start on the 
compression flange, it grew at an average calculated rate of 4.2E-2 mm/cycle until it 
reached the load pads at x = 150 and 455 mm. Even with the delamination on the 
compression side extending to the load pads, the stiffness of the beam remained ' 
approximately unchanged at 6,281 kN/m after 25,500 cycles. The average delamination 
length versus cycles is shown in Table 17. The delaminated beam is shown in Figure 43. 
Table 16. Cycles for tension and compression flange delamination on beam 49.
Cycles Average Tension 
Flange Delamination 
length, mm
Average Compression 
Flange Delamination 
length, mm
10 4 - -
1,000 4 —
2,000 4 —
6,800 8 12
12,868 20 19
25,500 20 19
88
Table 17. Reference notation for Beams 49 and 50 with ESH laminate.
Reference Notation for Beams 49,50 with ESH Laminate
Ply
Number
Ply
Angle
Fabric Description
I 0° D155
2 0° D155 Dropped plies
3 0° D155
4 +45° DB 120
5 OT
5 0° D155
6 0° D155
7 +45° DB 120
8 Oin
T
9 0° D155
10 0° D155
11 +45° DB 120
12 -45°
13 0° D155
Adhesive Layer Hysol EA 9309.2NA, 0.1 - 0.4 mm
14 +45° DB240
I - Beam shape, 
3 mm thick, 
Vf = 0.35
15 4^ Ul
16 0 D155
17
OmTj- DB240
18 +45°
89
Tension flange 
x = 277 mm
x = 326 mm 
x = 290 mm
Compression flange 50 mm
Figure 43. Beam 49 with ESH (two internal O0 ply drops) flange 
material.
Beam 50. Beam number 50 also contained ESH laminate flanges. It was fatigued 
at a rate of 3 Hz with a maximum load of 48.9 kN and a minimum load of 4.9 kN. The 
maximum load produced an initial maximum tension flange strain of 0.63% and a 
minimum compressive flange strain of -0.63%. The initial beam stiffness was measured 
as 6,054 kN/m. The tension flange ply drops were located at x = 265 and 325 mm and the 
compression flange ply drops were at x = 270 and 320 mm. Delamination on the tensile 
side started into the thin section and then arrested, just as the previous beams. The 
average delamination length versus cycles is shown in Table 18. The delamination on the 
tensile flange started into the thick cross-section and propagated uniformly at an average
90
calculated rate of 5.08E-4 mm/cycle. The compression side of the beam showed 
delamination being initiated at one comer of the ply drop and slowly working its way 
along the flange edge and then across the flange. Once the crack had propagated across 
the flange width, the delamination grew unstably to the load pads. The beam was then 
tested for residual strength after 204,100 cycles, with a residual load to failure of 88.2 kN
Tension flange 
x = 265 mm
I x = 325 mm
I
Compression flange
x = 270 mm x = 320 mm 50 mm
Figure 44. Beam 50 with ESH (two, interior 0° ply drops) 
laminate as flange material.
determined. The failed beam is shown in Figure 44.
Comparison of Delamination Rate in Beams and Coupons
Two comparisons were made between beam and coupon delamination rates, the 
first is the ESA laminate, a single exterior ply drop, and the second is the ESH laminate.
91
containing two interior ply drops.
Table 18. Cycles vs. tension flange delamination length for beam 50.
Cycles Average Tension Flange 
Delamination length, mm
500 4
2,500 19
7,000 24
12,000 30
31,800 ' 30
82,000 30
105,000 43
131,000 52
176,000 102
204,100 102
Compression flange did not have a uniform 
delamination front. Once the delamination 
propagated across flange width, the 
delamination rapidly proceeded to the grips.
92
Figure 45. Beam (Tension Flange) vs. Coupon Data (R=OT) for ESA (Single, exterior O0 
ply drop), R=0.1.
Coupon (ESH 213) run at 0.79% Strain 
Coupon (ESH 205) run at 0.79% Strain 
Beam run at 0.83% Strain
1 0 ------
10,000 15,000 20,000 25,000 30,000
Cycles
Figure 46. Beam (tension flange) vs. coupon data for ESH laminate (Two Interior ply 
drops), R=0.1.
93
These two comparisons are presented in Figures 45 and 46, respectively. Only one beam 
was run at each stress level. Therefore, no statistical tests were preformed, two beams are 
necessary to form a population, to correlate the beam and coupon delamination rates. The 
beam and coupon delamination rates, in tension, for both interior and exterior ply drops 
were similar. If the strain levels had been exactly the same, the delamination rates might 
have been even closer together.
In compression, no stable delamination took place with the interior ply drops, 
delamination typically showed up only a few hundred cycles before total delamination of 
the ply drops. However, one beam (Beam 50) did not show this behavior until the 
delamination had been initiated across the whole width of the beam flange. This may be a 
typical scenario, since Beam 49 was left unattended during the cycles when the plies 
delaminated to the load pads. This was also common with compressive coupons, the 
delamination would propagate during the last few cycles across the width and then the ply 
drops would delaminate to the grips. Exterior ply drops behaved differently in 
compression than did interior ply drops. The delamination for exterior ply drops would 
initiate but would arrest, whereas on the tension side of the beam the delamination would 
continue to grow. No figures showing the delamination rate of coupons versus beams is 
shown in compression due to the sudden nature of the delamination propagation. In 
general, delamination was observed on the tensile flange well before it was observed on 
the compressive flange.
The compressive ESH coupons that were bonded together back to back as
94
described earlier do correlate well with the compressive side of the ESH beam. Back-to- 
back coupons delaminated to the grips after 32,025 cycles at 0.62% strain, while an ESH 
beam compressive flange delaminated to the load pads after 25,500 cycles at 0.73% 
strain. At first it was thought that bonding the flange material onto the beam would be the 
only way to get accurate results. This was expected because the beam could restrain the 
coupon material from buckling. However, by increasing the coupon thickness, and 
reducing non-symmetry by back-to-back bonding, the coupon results were consistent with 
the beam data. The overall set-up time and the ability to run higher testing frequencies is 
greatly improved with back-to-back coupons versus beams.
Strain Energy Release Rate and Modeling Results 
Determination of Critical Strain Energy Release Rate.
Most of the modeling of delamination is done in a separate finite element study by 
Maccagnano. Modeling of delamination at ply drops requires basic Mode I and II critical 
strain energy release rate (Gc) data. These Gc values were determined for Mode I cracks 
using a double cantilever beam (DCB) coupon and for Mode II cracks using an end- 
notched flexure (ENF) coupon. Three different coupons configurations were tested using 
the D155 fabric for 0° layers and DB 120 fabric for the ±45° layers: the first laminate had 
the lay-up [0]IO with a fiber content of 36 percent, the second laminate consisted of 
[(0)5/±45/(0)4] with a fiber content of 38 percent, and finally a plate constructed of [±45]10 
with a fiber content of 26 percent. By using these three combinations it would be possible 
to determine Gc for cracks growing between a 0o/0°, 0o/±45° or a ±450ZMS0 interface.
A 2.22 kN load cell was placed in series with the standard load cell of the 8501
95
Instron testing machine to obtain the required load sensitivity. The hinges that had been 
mounted on the Mode I coupons were clamped in the machine and the coupons were 
adjusted so that they would remain perpendicular to the load cell during the test. The
Figure 47. Mode I coupon geometry.
Instron machine was placed in displacement control, with a ramp rate of 0.04 mm/sec. To 
avoid the crack propagating through the resin rich region ahead of the Fluro-Peel insert 
in both the Mode I and II coupons, which could artificially influence the value of G, a flat 
faced screwdriver was used to extend the crack 2.5 mm past the resin rich region. The 
initial crack lengths for the coupons were measured from the hinges to the crack front. A 
diagram of a typical Mode I specimen showing the dimensions used to calculate the strain 
energy release rate is shown in Figure 47. All three of the load-displacement diagrams 
were very similar. The Glc values were calculated using the methods described in Chapter 
2. During testing the crack stayed in the plane where it was initiated, except with coupons 
containing all ±45° layers. Therefore, values for these latter coupons are not reported.
96
crack
distance
Figure 48. Mode II specimen geometry.
Mode II coupons did not require hinges for testing. An ENF coupon with the
points of load application identified is shown in Figure 48. All the coupons were loaded
until the crack propagated from the initial length to the point of load application, except
the coupons containing all ±45° layers, which were to compliant and the crack did not
grow. One solution to this problem might be to add 0° layers and keep two ±45° layers
along the centerline, the stiffness of the coupon would be increased which would allow a
value of Gllc to be determined. The following table, Table 19, summarizes the results
found from the Mode I and II tests.
97
Table 19. Critical strain energy release rate for Mode I and Mode II cracks
Average* Critical Strain Energy Release Rate for Mode I and Mode II Cracks
Laminate Gt N-mm
n j Ic ----------------
mm2
Gnr N-mm
mm2
[OJio 0.49 ± 0.03 1.43 ±0.35
[(0)5/±45/(0)4] 0.78 ± 0.04 2.27 ± 0.53
*- Average of five specimens for each value.
Static and Threshold Strain 
Energy Release Rates at Plv Drops
The ESA and ESB laminates were used to determine the static G values. In Table 
20, the critical static stress causing delamination and the UTS of the ESA and ESB
Table 20. ESA and ESB critical static strength.
Laminate Critical Static 
Stress, MPa
Static Strength, 
MPa
ESA HO 427
ESB 496 700
coupons are given. Values for Gc calculated from ply drop static tests were obtained using 
Eqn 5 [53] for exterior ply drops and Eqn 6 [26] for interior ply drops.
j_p2 A 2E 2
IwA2E2A1E1
o2ht
4E(h-t) (6)
98
In Eqn 5, P is the critical static strength to cause the delamination, A is the cross-sectional 
area, E is the modulus in the longitudinal direction and w is the width of the coupon. The 
subscript 2 refers to thick cross-section properties, while the subscript I refers to the thin 
cross-section. In Eqn 6, o is the critical stress where delamination starts, h is the 
laminate thickness, t is the thickness of the dropped plies, and E is the modulus in the 
load direction, assumed to be the same for the thick and thin sides. The equations are 
based on simple mechanics of materials concepts. As the difference in modulus increases 
or the number of ply drops increase, the strain energy release rate will also increase. This 
shows the same trend as the experimental ESB and ESH coupons. Also, for a given ply 
drop case, increasing the thickness of the overall laminate reduces G. This is again 
consistent with the results for the ESB vs. ESH laminate. The static strain energy release 
rates for the ESA and ESB laminates are compared in Table 21 with values obtained from 
DCB and ENF coupons. The ESA laminate was compared to the critical strain energy 
release rate obtained from the DCB specimen with a O0ZO0 crack interface. However, the 
crack in the ESB laminate propagated via Mode II as opposed to Mode I in the ESA 
laminate. Therefore, an ENF specimen with a 07±45° interface was used as a 
comparison.
Table 21. Static strain energy release rates.________
Laminate Theoretical
(NZmm)
Experimental
(NZmm)
Percent Difference 
(%)
ESA 0.12 0.49 308
ESB 1.34 2.27 69
99
However, these equations do not correlate well the finite element predictions. Table 22 
compares the results obtained using the equations to the finite element results. From the 
table the percent difference between the formulas found in literature and the finite 
element methods vary greatly . By experimentally determining the threshold stress level 
for some of the cases tested, finite element methods were capable of predicting the 
threshold stress of other laminates.
Table 22. Threshold strain energy release rates_________
Laminate Theoretical Strain 
Energy Release 
Rate, N/mm
FEA Strain Energy 
Release Rate [55], 
N/mm
Percent Difference, 
%
ESA 0.08 0.12 51
ESB 0.34 0.73 114
ESH 0.15 0.74 380
For example, the ESB laminate had a threshold stress level of 276 MPa., which 
corresponded to a threshold strain energy release rate of 0.738 N/mm. The ESH 
geometry, with two interior ply drops, was then loaded at different stress levels until the 
threshold strain energy release rate of the ESH matched the ESB laminate. A threshold 
stress level for the ESH laminate of 200 MPa was obtained, which was only 3% lower 
than the experimental value of 207 MPa. The list of strains presented in Table 3 can be 
used in conjunction with finite element methods to predict threshold strain energy release 
rates for other laminates accurately.
100
CHAPTER 5
CONCLUSIONS AND RECOMMENDATIONS
Conclusions
Delamination at Plv Drops
The following detailed conclusions for delamination relate only to uniaxial tensile
fatigue, R = 0.1, at ambient temperature with no environmental effects considered. A
convenient comparison of results is given in Table 3 as the strain to produce a 25 mm
delamination in IO5 cycles, as well as arrest and threshold values.
o Internal ply drops are roughly twice as resistant to delamination as external ply 
drops.
o For the same number of ply drops, thicker laminates are better in resisting 
delamination.
o Dropping more than one ply at the same location increases the delamination 
growth rate.
o When the ply drops are separated, the delamination follows the same growth rate 
as a single interior ply drop until the delamination reaches the adjacent ply drop. 
The rate then increases.
o Dropping multiple ±45° layers is an effective way to taper the thickness without 
introducing delamination.
o The use of “Z-Spiking” and adhesives at ply drops in the manufacturing of parts 
substantially reduces the initial delamination rate.
o Incorporating adhesive into laminates containing interior ply drops prevents 
delamination, prior to tensile failure at the ply drop.
101
o Using “feathering” at the ply drops, on both interior and exterior ply drops, 
reduces the delamination rate.
o Random mat does not help and possibly has a detrimental effect on delamination 
resistance.
o Coupons can be successfully tested in compression, as long as care is taken to
ensure the stiffness and symmetry of the coupon is sufficient to prevent buckling; 
back-to-back bonding of coupons worked well and correlated well with beam 
results.
I-Beam Study
o The delamination results between coupons and beam flanges containing ply drops 
were at similar strain levels.
o Ply drops, both interior and exterior, on the tension side of the beam showed 
similar delamination rates to coupons.
o Exterior ply drops on the compression flange arrested after growing a short 
distance. Interior ply drops gave little warning in compression before total 
delamination. In all cases, delamination occurred earlier on the tensile flange than 
on the compressive flange.
Effect of Plv Drops on Lifetime
o Introduction of ply drops reduces the fatigue life of the lower fiber content
coupons more significantly than for the higher fiber content coupons.
o The fatigue resistance of lower fiber content coupons with ply drops was degraded 
to the level of the higher fiber content coupons without ply drops. The high fiber 
content coupons with ply drops did not show any significant loss in performance 
when compared to control coupons of the same fiber content.
o The A l30 fabric for 0° reinforcement with ply drops has a very erratic, strand by 
strand delamination response at strains well below the values for the D155 fabric.
o The lifetime of laminates with Al 30 fabric converged with that of the D155 fabric 
in cyclic fatigue after IO6 cycles.
102
Gc Tests
Determining the critical strain energy release rate was straight-forward, with 
Mode H values being 3 to 4 times higher than Mode I values. Testing of [±45]10 coupons 
did not provide any information due to the compliance of the coupons. Coupons that 
contained ±45° ply drops did not delaminate. Theoretical models for delamination growth 
are consistent with data trends, but disagree with some finite element results.
Recommendati on s
This set of data for ply drops establishes delamination rates and the performance 
of various lay-ups in cyclic fatigue loading at R=O. I and R=IO. The actual blade loading 
conditions need to be determined, so that coupon performance can be reevaluated under a 
worst loading case or spectrum loading conditions. Even though cracks form in 
composites which have no ply drops, the extensive delamination from ply drops causes 
even more severe cracks to form which makes the composite more susceptible to 
environmental attack. Therefore, an investigation of the environmental effects on the 
lifetime of laminates with ply drops should be investigated. Compression should also be 
explored in greater detail, along with understanding the influence of tapering coupons on 
buckling resistance.
Using the critical strain energy release rates from Mode I and II coupons, finite 
element modeling of laminates with ply drops should be used to predict static and 
threshold stress levels. Once the FEA results can be validated with the coupon data, ply 
drops can be incorporated into larger structures, such as wind turbine blades, with
confidence.
103
REFERENCES CITED
1. Wilkins, DJ., Eisenmann, J.R. Camin, RA., Margolis, W.S. and Benson, RA., “ 
Characterizing Delamination Growth in Graphite-Epoxy,” Damage in Composite 
Materials, ASTMSTP 775, 1982, pp. 168.
2. Broek, D. “Elementary Engineering Fracture Mechanics,” 4th edition, Kluwer 
Academic Publishers, 1986, pp 328-335.
3. Benbow, J.J. and Roesler, F.C., Process Physical Society, London, B 70, 1957, pp. 
201.
4. Gilman, J.J. Journal o f Applied Physics, Vol. 31,1960, pp. 2205.
5. Russell, A.J. and Street, K.N., “ Factors Affecting the Interlaminar Fracture Energy 
of Graphite/Epbxy Laminates,” Progress in Science and Engineering o f Composites, ed. 
Hayashi, T, Kawata, K., and Umekawa, S., ICCMIV, Tokyo, 1982, p. 279.
6. Lagace, PA., Cannon, R.K., “ Effects of Ply Dropoffs on the Tensile Behaviour of 
Graphite/Epoxy Laminates, 4th Japan-US. Conference on Composite
Materials, Washington, DC, June 27-29, 1988.
7. Wisnom, M.R., “Prediction of Delamination in Tapered Unidirectional Glass Fibre 
Epoxy with Dropped Plies Under Static Tension and Compression,” 74 th Meeting o f the 
AGARD Structures and Material Panel, Patras, Greece, May 24-29, 1992, pp. 25-1.
8. Mandell, J.F., Reed, R.M., Jr. and Samborsky, D.D., “ Fatigue of Fiberglass Wind 
Turbine Blade Materials, ”, Contractor Report SAND92-7005, Sandia National 
Laboratories, Albuquerque, NM, 1992.
9. Grimes, G.C., Dusablon, E.G., “ Study of Compressive Properties of Graphite/Epoxy 
Laminates with Discontinuities,” Composite Materials: TestingandDesign,, ASTMSTP 
787, 1982, pp. 513-538.
10. Ulman, D A., Bruner, R.D., Owens, S.D., and Miller, H.R., “Damage Accumulation 
in Composites (Volume II- Experimental Data.” Technical Report General Dynamics 
Technical Report AFWAL-TR-85-3070. General Dynamics Fort Worth Division. Fort 
Worth, TX. September 1985.
11. Reifsnider, K.L., “Damage and Damage Mechanics,” Fatigue o f Composite 
Materials, Elsevier Science Publishers, 1990, pp. IT.
12. Adams, D.F., Ramkumar, R.L., and Walrath, D.E., “Analysis of Porous Laminates in 
the presence of Ply Drop-Offs and Fastener Holes,” Technical ReportNOR 84-113.
104
Northrop Corporation, Hawthorne, CA. And University of Wyoming, Laramie WY 
May, 1984.
13. Fish, J.C., Lee, S.W., “Tensile Strength of Tapered Composite Structures,” AIAA 
Paper No. 88-2252, Proceedings o f the AIAA/ASME/ASCE/AHS/ASC 29th Structures, 
Structural Dynamics, and Materials Conference. Part I. Monterey, CA., April, 1988.
14. Fish, J.C., Lee, S.W., “Tensile Strength of Composite Structures with Internal Ply 
Drops,” PhD Thesis, University of Maryland, 1988.
15. Wu, C.M.L., Webber, J.P.H., “Analysis of Tapered (In Steps) Laminated Plates 
Under Uniform In-Plane Load,” Composite Structures, Vol. 5, Np. I, pp. 87,1988.
16. Hoa, S.V., Daoust, J., Du, B.L., and VuKhanh, T., “ Interlaminar Stresses in Tapered 
Laminates, ” Polymer Composites, Vol. 9, No. 5, pp. 337, October 1988.
17. Dharam, L .R., Seaton, C.P., “Analysis of a Buffer Strip Laminate with Fiber and 
Matrix Damage and Interlaminar Debonding.”, Composite Structures, Vol. 7, pp. 139- 
158, 1987.
18. Poe, C. C. and Kennedy, I. M., “Buffer Strips for Improving Damage Tolerance of 
Composite Laminates,” Journal o f Composite Materials Supplement, Vol. 14, p. 57,
1980.
19. Hess, T. E., Huang, S. L., and Rubin, H., “ Fracture Control in Composite Materials 
using Integral Crack Arresters,” Proceedings o f the AIAA/ASME/SAE17th Structures, 
Structural Dynamic and Materials Conference, pp. 52-69, King of Prussia, PA, 1976.
20. Fisher, K., “3D reinforcements: on the verge?”, High Performance Composites, 
March/April 1996, p. 32.
21. Tanzawa, Y., “Experimental Study of Interlaminar Delamination Toughness of 3-D 
Orthogonal Interlocked Fabric Composite,” Proceedings o f the 
AIAA/ASME/ASCE/AHS/ASC 38th Structures, Structural Dynamics, and Materials 
Conference and Exhibit and AIAA/ASME/AHS Adaptive Structures Forum, Vol. 4, pp. 
2611, Kissimmee, FE, 1997.
22. Chan, W.S., Rogers, C., and Aker, S., “Improvement of Edge Delamination Strength 
of Composite Laminates using Adhesive Layers,” Composite Materials: Testing and 
Design, ASTMSTP 893,1986, p. 266.
23. Masters, J. E., “Improved Impact and Delamination Resistance through Interleafing,” 
Interlaminar Fracture in Composites, E. A. Armanios, editor, Trans Tech Publishing, 
Switzerland
23., 1989.
105
24. Kemp, B. L., Johnson, E. R., “Response and Failure Analysis of a Graphite/Epoxy 
Laminate Containing Terminating Internal Plies, AIAA Paper No. 85-0608, Proceedings 
o f the AIAA/ASME/ASCE/AHS 26th Structures, Structural Dynamic and Materials 
Conference, Part I, April 15-17, 1985, pp. 13-24.
25. Chan, W. S., and Ochoa, 0 . O., “ Suppression of Edge Delamination in Composite 
Laminates by Termination of a Critical Ply Near the Edges,” Proceedings o f the 
AIAA/ASME/ASCE/AHS 26th Structures, Structural Dynamic and Materials Conference, 
Part I, April 18-20, 1988, pp. 359-364.
26. Ochoa, O. O and Chan, W. S.,“Tapered Laminates: A Study on Delamination 
Characterization,” Proceedings o f the American Society for Composites, Third Technical 
Conference, Sept. 25-29, 1988, pp. 633.
27. Hoa, S.V., Daoust, J., Du, B.L., and VuKhanh, T., “ Interlaminar Stresses in Tapered 
Laminates, ” Polymer Composites, Vol. 9, No. 5, pp. 337, October 1988.
28. Curry, J.M., Johnson, E.R., and Starnes, J. H., “Effect of Ply Drop-offs on the 
Strength of Graphite/Epoxy Laminates,” Technical Report CCMS 86-07, Center for 
Composite Materials and Structures, Virginia Polytechnic Institute and State University, 
Blacksburg, V A, December 1986.
29. O’Brien, T.K., “Towards a Damage Tolerance Philosophy for Composite Materials 
and Structures,” Composite Materials: Testing and Design (Ninth Volume). ASTM STP 
1059, S.P. Garbo, Ed., American Society for Testing and Materials, Philadelphia, 1990, 
pp. 7-33.
30. Wilkins, D.J., Eisenmann, J. R., Camin, R. A., Margolis, W.S., Benson, RA., 
“Characterizing Delamination Growth in Graphite-Epoxy,” Damage in Composite 
Materials, ASTM STP 775, American Society for Testing and Materials, Philadelphia, 
June 1982, p. 168.
31. Mall, S., Yun, K. T., and Kochnar, N. K., “ Characterization of matrix Toughness 
Effect on Cyclic Delamination Growth in Graphite Fiber Composites,” presented at the 
2nd ASTM Symposium on Composite Materials: Fatigue and Fracture, Cincinnati, OH, 
April 1987.
32. Russell, A J . and Street, K.N., “Predicting Interlaminar Fatigue Crack Growth Rates 
in Compressively Loaded Laminates,” presented at the 2nd ASTM Symposium on 
Composite Materials: Fatigue and Fracture, Cincinnati, OH, April 1987.
33. Gustafson, C. G., and Hojo, M., “Delamination Fatigue Crack Growth in 
Unidirectional Graphite/Epoxy Laminates,” Journal o f Reinforced Plastics, Vol. 6, No. I,
106
January 1987, pp. 36-52.
34. O’ Brien, T. K,, “ Generic Aspects of Delamination of Fatigue of Composite 
Materials,” Journal o f the American Helicopter Society, VoL 32, No. I, January 1987, pp. 
36-52.
35. Martin, R. H. And Murri, G.B., “Characterization of Mode I and Mode II 
Delamination Growth and Thresholds in Graphite/PEEK Composites,” NASA TM 
100577, National Aeronautics and Space Administration, Washington, DC, April 1988.
36. O’ Brien, T.K., “ Characterization of Delamination Onset and Growth in a 
Composite Laminate,” Damage in Composite Materials, ASTMSTP 755, American 
Society for Testing and Materials, Philadelphia, 1982, pp. 140-167.
37. Russell, A.J. and Street, K. N., “Moisture and Temperature Effects on the Mixed- 
Mode Delamination Fracture of Unidirectional Graphite/Epoxy,” Delamination and 
Debonding o f Materials, ASTM STP 876, American Society for Testing and Materials, 
Philadelphia, October 1985, pp. 349-370.
38. Poursartip, A., “ The Characterization of Edge Delamination Growth in Composite 
Laminates Under Fatigue Loading,” Toughened Composites, ASTM STP 937, American 
Society for Testing and Materials, Philadelphia, 1987, pp. 222-241.
39. O’Brien, T. K., “ Mixed-Mode Strain Energy Release Rate Effects on Edge 
Delamination of Composites,” Effects o f Defects in Composite Material, ASTM STP 836, 
American Society for Testing and Materials, Philadelphia, 1984, pp.125-142.
40. O’ Brien, T.K., “Tension Fatigue Behavior of Quasi-Isotropic Graphite/epoxy 
Laminates,” Fatigue and Creep o f Composite Materials, Proceedings, 3rd Riso 
International Symposium on Metallurgy and Materials Science, riso National Laboratory 
Roskilde, Denmark, 1982, pp. 259-264.
41. O’Brien, T.K., “ Fatigue Delamination Behaviour of PEEK Thermoplastic 
Composite Laminates,” Journal o f Reinforced Plastics, Vol. 7, No. 4, July 1988, pp. 341- 
359.
42. O’ Brien, T.K., Murri, G.B., Salepekar, S. A., “Interlaminar Shear Fracture 
Toughness and Fatigue Thresholds for Composite Materials,” Composite Materials: 
Fatigue and Fracture- Second Volume, ASTM STP 1012, American Society for Testing 
and Materials, Philadelphia, 1989, pp. 222-250.
43. Adams, D.F, Zimmerman, R. S., and Odem, E. M., “Frequency and Load Ratio 
Effects on Critical Strain Energy Release Rate Gc Thresholds of Graphite Epoxy 
Composites,” Toughened Composites, ASTMSTP 937, American Society for Testing and
107
Materials, Philadelphia, 1987, pp. 242.
44. O’ Brien, T. K., Rigamonti, M., and Zanotti, C., “Tension Fatigue Analysis and Life 
Prediction for Composite Laminates,” NASA TM 100549, National Aeronautics and 
Space Administration, Washington, DC, October 1988.
45. O’Brien, T. K., “Generic Aspects of Delamination in Fatigue of Composite 
Materials,” Journal o f the American Helicopter Society, Vol. 32, No. I, January 1987 pp 
13-18.
46. Agarwal, B.D. and Broutman, L.J., “Analysis and Performance of Fiber composites,” 
2nd edition, John Wiley & Sons, 1990, pp. 430-435.
47. Freitas, G. and Fusco, T., “Z-Fiber Technology and Products for enhancing 
Composites Design,” Aztec, Inc., Waltham, MA. pp. 1-7.
48. Hedley, C.W., “ Mold Filling Parameters in Resin Transfer Molding of Composites,” 
M.S. Thesis, April 1994, pp. 12.
49. Mandell, J.F., Samborsky, D.D., “ DOE/MSU Composite Material Fatigue Database: 
Test Methods, Materials, and Analysis,” Sandia Contractors Report, To be published.
50. Wung, E.C.J., Chatterjee, S.N., “On the Failure Mechanisms in Laminate 
Compression Specimens and the Measurement fo Strength,” Journal o f Composite 
Materials, Vol. 26, No. 13,1992, pp. 1885-1914.
51. Adams, D.F., Lewis, E.Q., “ Influence of Specimen Gage Length and Loading 
Method on the Axial Compressive Strength of a Unidirectional Composite Material,” 
Experimental Mechanics, March 1991, pp. 14-20.
52. Combs, D.W., “Design, Analysis and Testing of a Wind Turbine Blade 
Substructure,” MS Thesis, March 1995, pp. 24-36.
53. Ramkumar, R.L. and Whitcomb, J,D. “Characterization of Mode I and Mixed-Mode 
Delamination Growth in T300/5208 Graphite/Epoxy,” in Delamination and Debonding of 
Materials, ASTM STP 876, W.S. Johnson, ED., ASTM, Phi., 1985.
54. Mandell, J.F., Samborsky, D.D., Combs, D.W., Scott, M.E., and Cairns, D.S.
“Fatigue of Composite Material Beam Elements Representative of Wind Turbine 
Substructure,” NREL Contractors Report, To be published.
55. Maccagnano, Z., Personal Interview, 8 June 1997.
APPENDIX A 
Coupons Tested
109
The coupon tests were given a letter which identified the material and individual 
coupons. This appendix contains results from cyclic fatigue tests using a constant stress 
amplitude sine waveform with R-values of 0.1 and 10, where the R-value is defined by:
_ Minimum cyclic stress 
Maximum cyclic stress
and the compressive stresses are negative.
The individual test results are listed and summarized using eight columns with the 
following data structure:
Col. I & 2 Col. 3 Col.4 Col.5 Col. 6 Col.7 Col. 8
Sample Max. Running E(GPa) %e Cycles Notes Rvalue
ID # Stress (MPa)
Col. I: List the laminate being tested.
Col. 2: Lists the MSU test reference number.
Col. 3: This column indicates the maximum stress in megapascals (MPa) which was applied to 
the coupon. A positive number indicates tension while a negative number indicates compression. 
For a compression test the stress listed as maximum is actually the minimum stress.
Col. 4: Lists the initial measured elastic modulus (E) of the coupon in gigapascals (GPa).
Col. 5: Indicates the initial absolute maximum fatigue running strain (e) in percent or the percent 
strain to failure for a static test.
Col. 6: Number of cycles the coupon was fatigued
Col. 7: Lists any other notation for comments.
The notations used in column 6 are summarized below:
F- Failure of coupon
RO - Run out, coupon has significant fatigue cycles but has not yet failed, test stopped.
Z - Double coupon thickness, two coupons bonded together to increase the thickness.
—  Indicates that a value was unavailable
DA - Delamination arrested 
EF - End failure of coupon
no
Other notations used in the test material summary tables include: 
Vf - Fiber volume content of the material in percent
Laminate Number Max. Stress (MPa) E (GPa) % Strain Cycles Comments R value
ESA 113 452 23.1 1.96 I — —
ESA 108 207 — 0.90 2,001 - - 0.1
ESA 109 138 — 0.60 110,000 - - 0.1
ESA 111 121 — 0.52 175,000 RO 0.1
ESA 112 103 — 0.45 400,000 DA 0.1
ESA 102 207 23.4 0.90 1,301 — 0.1
ESA 104 207 — 0.90 1,001 0.1
ESA HO 138 23.4 0.60 85,000 — 0.1
ESA 103 138 0.60 40,000 — 0.1
ESA 107 -276 — -1.19 318 — 10
ESA 101 -207 — -0.90 20,000 -- - 10
ESA 105 -138 -0.60 252,222 — 10
ESA 106 -172 — -0.75 102,000 -- - 10
ESA 202s 410 - - - 1.77 I — —
ESA 203s 416 — 1.80 I — —
ESA 204s 436 — 1.89 I -—
ESA 205s 423 — 1.83 I — —
ESA 207s 477 — 2.07 I — —
(Vf=35%)
ESB 112 524 22.1 2.27 I — —
ESB 106 207 24.1 0.90 90,001 — 0.1
ESB 109 345 — 1.49 7,326 EF 0.1
ESB 114 276 24.1 1.19 50,000 -- - 0.1
ESB 115 276 . 1:19 45,000 — 0.1
ESB 101 345 -- - 1.49 10,000 — 0.1
ESB 102 310 — 1.34 20,100 — 0.1
ESB 113 310 23.4 1.34 15,000 — 0.1
ESB 105 310 22.8 1.34 15,000 0.1
ESB HO 276 — 1.19 44,342 — 0.1
ESB 103 -276 — -1.19 409 F 10
ESB 111 -241 — -1.04 30,996 F 10
ESB 107 -241 — -1.04 187,000 — 10
Laminate Number Max. Stress (MPa) E (GPa) % Strain Cycles Comments R value
(Vf=35%)
ESB 409 753 25.5 2.95 I F ~ ~
ESB 411 788 26.2 3.09 I F
ESB 405 758 — 2.97 I F
ESB 407 414 - -- 1.62 2,401 F 0.1
ESB 403 207 — 0.81 1,060,000 RO 0.1
ESB 404 414 25.5 1.62 3,567 F 0.1
ESB 414 207 — 0.81 1,277,000 F 0.1
ESB 401 276 „25.5 1.08 58,501 F 0.1
ESB 402 276 — 1.08 164,000 F 0.1
(Vf=44%)
ESC 101 555 22.8 2.37 I
ESC HO 345 24.1 1.47 8,700 — 0.1
ESC 109 345 23.4 1.47 4,000 — '— 0.1
ESC 108 276 — 1.18 41,000 ------- 0.1
- ESC 104 276 --- 1.18 20,000 — 0.1
ESC 107 345 — 1.47 10,000 0.1
ESC 105 276 ~ ~ 1.18 25,001 —Trf 0.1
‘(Vf=35%).
ESD HO 396 14.5 2.13 I i—VW
ESD 109 138 18.6 0.74 300,000 RO 0.1
(VL=37%)
ESE 116 207 — 0.68 5,000 ~ ~ 0.1
ESE 112 655 30.3 2.16 I
ESE 114 207 — 0.68 5,810 0.1
ESE HO 138 — 0.45 100,000 0.1
ESE 111 138 32.4 0.45 10,000 0.1
ESE 106 121 — 0.40 101,000 ------- 0.1
(Vf=40%)
Laminate Number Max. Stress (MPa) E (GPa) % Strain Cycles Comments R value
ESF HO 207 — 0.67 34,033 EF 0.1
ESF 105 276 31.0 0.89 74,428 EF 0.1.
ESF 113 276 — 0.89 45,000 — 0.1
ESF 106 276 — 0.89 45,000 — 0.1
ESF 109 345 31.7 1.11 8,646 EF 0.1
ESF 108 276 — 0.89 9,785 EF 0.1
(Vf=40%)
ESG 104 345 29.6 1.16 I — . 0.1
ESG 102 207 — 0.70 150 - - - 0.1
ESG 107 138 — 0.47 30,000 RO 0.1
ESG 105 138 29.6 0.47 50,000 — 0.1
ESG 106 103 — 0.35 225,000 — 0.1
ESG 101 121 0.41 42 - -_ 0.1
(Vf=46%)
ESH 101 276 — 0.93 24,000 — 0.1
ESH 102 454 29.6 1.53 I - - —
ESH 105 207 — 0.70 50,000 — 0.1
ESH 106 276 29.6 0.93 25,000 •— 0.1
(Vf=46%)
ESH 202 634 33.09 1.88 I F 0.1
ESH 204 630 33.78 1.87 I F 0.1
ESH 208 345 - - 1.02 1,300 F 0.1
ESH 209 345 — 1.02 1,350 F 0.1
ESH 213 276 • — 0.82 20,000 RS 0.1
ESH 205 276 0.82 40,000 RS o.i
ESH 211 379 — 1.12 57 F 0.1
ESH 212 . 379 — 1.12 63 F 0.1
ESH 207 276 — ■ 0.82 15,000 F 0.1
ESH 210 207 — 0.61 92,479 F _ 0.1
Laminate Number Max. Stress (MPa) E (GPa) % Strain Cycles Comments R value
(Vf=45%)
ESH 301 733 32.40 2.42 I F
ESH 312 770 31.02 2.54 I F
ESH 313 798 — 2.63 I F
ESH 304 414 - ----- 1.36 1,935 F 0.1
ESH 305 414 1.36 2,021 F 0.1
ESH 303 414 — 1.36 2,693 F 0.1
ESH 307 207 — 0.68 953,762 F 0.1
ESH 311 345 30.33 1.14 3,929 F 0.1
ESH 310 207 ----- 0.68 526,064 F 0.1
ESH 306 276 31.71 0.91 39,000 F 0.1
ESH 314 276 — 0.91 82,921 F 0.1
ESH 309 276 ~~ 0.91 56,938 F 0.1
(Vf=52%)
ESH 413 747 2.64 I F
ESH 407 207 27.58 0.73 731,521 F 0.1
ESH 412 728 28.27 2.58 I F
ESH 411 414 ------- 1.46 813 F 0.1
ESH 406 414 — 1.46 748 F 0.1
ESH 410 276 0.98 28,019 F 0.1
ESH 405 276 27.58 0.98 30,689 F 0.1
ESH 404 207 28.27 0.73 1,093,000 RO 0.1
ESH 409 172 — 0.61 1,100,000 RO 0.1
ESH 408 207 27.58 0.73 968,675 F 0.1
.ESH 403 207 — 0.73 685,000 F 0.1
(V£=35%)
ESIl 101 276 31.0 0.89 42,500 EF 0.1
104 310 — 1.00 9,500 0.1
105 276 — 0.89 0.1
I 103 276 — 0.89 96,000 EF 0.1
Laminate Number Max. Stress (MPa) E (GPa) % Strain Cycles Comments R value
ESI2 107 310 31.0 1.00 1,730 EF 0.1
103 310 — 1.00 3,829 EF 0.1
102 276 — 0.89 19,000 EF 0.1
104 241 — 0.78 90,000 0.1
106 276 — 0.89 19,100 EF 0.1
ESI3 107 276 31.0 0.89 25,000 0.1
104 276 — 0.89 52,000 0.1
103 241 0.78 136,000 — >— 0.1
108 207 — 0.67 1,121,809 ■------- 0.1
106 241 — 0.78 130,326 .-w rv 0.1
102 276 — 0.89 87,090 0.1
ESI4 108 276 31.0 0.89 9,500 EF 0.1
106 276 — 0.89 51,000 . 0.1
107 276 0.89 10,475 EF 0.1
104 310 — 1.00 5,447 EF 0.1
103 276 — 0.89 36,000 0.1
101 241 0.78 162,000 -------- - 0.1
(ESI I -ESI4, Vf=36%) ,
ESJ 101 207 — 0.79 90,000 0.1
ESJ . 108 276 — 1.05 31,213 EF 0.1
ESJ 104 207 — 0.79 80,000 0.1
ESJ 106 207 26.2 0.79 110,000 ~ — 0.1
(Vf=37%)
ESK 104 207 — 0.71 30,000 —' 0.1
ESK 105 276 — 0.95 809 0.1 .
ESK 107 207 — 0.71 10,000 — 0.1
ESK 102 138 29.0 0.48 210,000 ------- _ 0.1
(Vf=37%)
Laminate Number Max. Stress (MPa) E (GPa) % Strain Cycles Comments R value
ESL 103 241 29.0 0.83 10 EF 0.1
ESL 107 241 — 0.83 365,000 0.1
ESL 102 276 — 0.95 83,400 0.1
ESL 101 276 — 0.95 80,000 0.1
(Vf=44%)
ESM 107 276 — 142,000 0.1
(Vf=44%)
ESN 101 276 24.1 1.14 I 0.1
ESN 102 172 24.8 0.71 1,500 0.1
ESN 103 138 --- - 0.57 17,000 0.1
ESN 104 103 — 0.43 102,000 0.1
ESN 105 103 — 0.43 187,000 0.1
(Vf=36%)
ESO 101 276 24.8 1.14 120,000 EF 0:1
ESO 102 345 25.5 1.43 2,443 F 0.1
ESO 103 310 ---- 1.29 2,297 EF 0.1
ESO 104 310 _ — 1.29 . 8,379 EF 0.1
ESO 105 276 - - 1.14 41,808 EF 0.1
ESO 106 276 — 1.14 19,500 EF 0.1
(Vf=36%)
ESL 102 955 33.1 I F
ESL 104 907 32.4 I F
ESL 105 887 — 2.86 I F
ESL 114 414 31.7 1.33 4,600 F 0.1
ESL 108 414 — 1.33 4,005 F 0.1
ESL 109 414 31.0 1.33 4,263 F 0.1
ESL 116 276 32.4 0.89 64,613 F 0.1
ESL 112 . 172 --- 0.56 998,920 RO 0.1
Laminate Number Max. Stress (MPa) E (GPa) % Strain Cycles Comments R value
ESL 115 207 0.67 922,651 F 0.1
ESL HO 276 31.7 0.89 116,740 F 0.1
ESL 113 276 — 0.89 108,767 F 0.1
ESL 111 552 — 1.78 311 F 0.1
ESL 106 552 — 1.78 381 F 0.1
(Vf=44%)
ESQ 612 172 20.0 0.68 1,217,000 RO 0.1
ESQ 607 544 22.1 2.13 • i F 0.1
ESQ 603 552 — 2.17 i F 0.1
ESQ 606 548 •— 2.15 i F 0.1
ESQ 608 276 21.4 1.08 12,405 F 0.1
ESQ 602 276 — 1.08 10,207 F 0.1
ESQ 601 310 23.4 1.22 . 2,704 F 0.1
ESQ 613 310 ----- - 1.22 5,632 F 0.1
ESQ 614 276 — 1.08 19,000 F 0.1
ESQ 601 172 — 0.68 1,106,000 F 0.1
ESQ 611 241 22.1 0.95 210,000 F 0.1
ESQ 605 241 0.95 32,000 F 0.1
(Vf=44%)
APPENDIX B
Delamination Length vs. Cycles for Coupons
De lam ina tion  Length  vs. C yc les  fo r ESA  lam ina te , R=O. 1
■121 MPa 
103 MPa 
207 MPa 
138 MPa 
207 MPa 
207 MPa 
138 MPa 
138 MPa 
121 MPa
90,000 180,000 270,000
Cycles
360,000 450,000
De
la
m
i
Delam ination Length vs. Cycles for ESB laminate, R=O. 1
50.0
40.0
?
E
xz
g80.0
C
™20.0 
10.0 
0.0
/► 344 MPa -* -3 1 0  MPa
J
275 MPa 
- x -  344 MPa 
—*— 275 MPa 
-# -276  MPa 
—t—310 MPa/  ,
J J  / /
a ------------- ir ____  ________ _I___________________, :
&
0  IV---------  * —  - ■ J
0 10,000 20,000 30,000 40,000 50,000 60,000
Cycles
D
el
am
in
at
io
n 
L
en
gt
h 
(m
m
)
Delamination length vs. Cycles for ESC and ESE laminates, R=O.1
4 0 .0
30.0
20.0
10.0
0.0
20,000 40,000 60,000 80,000 100,000 120,000
Cycles
Delam ination Length vs. Cycles fo r ESF laminate, R=O. 1
3 0 .0
^ 25.0
E
£ 20.0
CD
C
0
^  15.0 
o
CD
I  10.0
CD
0
O 5.0
0.0
0 5,000 10,000 15,000 20,000 25,000 30,000 35,000 40,000
Cycles
— ESF 109 (275 MRA)
- e -E S F 1 1 0 (2 0 7M P a )
ESF 105 (275 MPa)
ESF 106 (276 MPa)
ESF 109 (345 MPa)
- # - E S F  108 (275 MPa)
-------
—t f-------------—
io
n 
Le
ng
th
 (
m
m
).
Delamination Length vs. Cycles for ESG laminate, R=O. 1
60.0
20,000 40,000 60,000 80,000
Cycles
100,000 120,000 140,000
Delam ination length vs. Cycles for ESH laminate, R=O. 1
60.0
40,000 80,000 120,000
Cycles
ESI 1
ESM 101 (276 MPa) ESI 1 104 (310 MPa)
LOW CR.CYCLES UP CK. LOW CK. CYCLES UP CK.
0 0 0 0 0 0
2,000 0 5 250 0 0
3,700 0 7 500 4 4
5,000 0 7 1,500 5 4
8,000 0 8 2,600 5 5
10,000 0 8 3,600 5 15
16,000 0 8 5,000 10 15
21,000 0 8 8,000 10 15
25,000 0 8 9,000 20 15
31,000 0 8
ESI2 107 (310MPa) 
CYCLES UP CK LOW CR.
ESI2103 (310 MPa) 
CYCLES UP CK LOW CK.
0 0 0 0 0 0
500 5 5 500 2 10
1,730 10 10 1,000 4 17
1,750 EF 1,600 8 17
2,500 10
3,829 EF
17
ESI1 105 (275 MPa) 
CYCLES UP CK1
0 0
3,400 5
5,400 5
8,000 8
ESI2102 (275 MPa) 
CYCLES UP CK.
0 0
1,200 0
3,000 0
5,000 9
10,000 9
19,000 EF
to
V i
LOW CK.
0
8
8
8
LOW CK.
0
5
9
15
15
ESI2 104 (241 MPa) ESI2 106 (275 MPa)
CYCLES UP CK. LOW CK. CYCLES UP CK. LOW CK.
0 0 0 0 0 0
10,000 2 5 300 1 0
41,000 5 5 1,000 5 10
90,000 5 25 3,500 10 10
8,200 15 17
10,000 15 17
15,000 15 17
19,100 15 17
ESI3 107 (275 MPa) ESI3 104 (275 MPa)
CYCLES UP CK LOW CK. CYCLES UP CK. LOW CK.
0 0 0 0 0 0
400 0 5 1,500 4 4
800 0 7 3,000 5 5
1,700 0 8 6,000 5 5
5,000 5 10 IOiOOO 5 7
11,000 10 10 20,000 15 10
15,000 10 10 41,000 25 20
20,000 15 12
25,000 20 15
ESI3 108 (207 MPa) 
CYCLES UP CR.
0 0
74,650 1
535,000 3
647,000 3
931,000 3
ESI3 103 (241 MPa) 
CYCLES UP CK.
0 0
9,000 5
15,000 5
50,000 10
75,000 15
100,000 15
136,000 25
K>Q\
LOW CR.
0
6
14
14
15
LOW CK.
0
5
- 5
10
20
20
25
ESI4107 (275 MPa) ESI4 106 (275 MPa) ■ ESI4 104 (310 MPa)
CYCLES UP CR. LOW CR. CYCLES UP CR. LOW CR. CYCLES UP CR.
0 0 0 0 0 0 0 0
3,600 0 3 4,000 0 3 ' 50 0
5,600 5 5 6,500 0 5 200 0
7,600 5 5 8,500 0 5 400 0
11,000 0 5 800 0
20,000 0 5 1,300 0
30,000 0 5 2,300 0
51,000 0 5 3,000 0
60,000 0 5 5,000 0
LOW CR. 
0
10
11
12
12
15
to
- J
ESI4 103 (275 MPa)
CYCLES UP CR. LOW CR.
0 0 0
900 0 7
2,500 3 10
3,600 3 12
5,000 3 12
7,600 4 15
10,000 5 17
15,000 5 20
21,000 7 20
26,000 8 21
36,000 9 23
ESI4 101 (241 MPa)
CYCLES UP CR. LOW CR.
0 0 0
15,000 3 5
20,000 4 9
50,000 8 20
75,000 8 20
89,000 8 20
100,000 8 20
162,000 8 20
^
 
CD 
CO
Delamination length vs. cycles fo r ESJ and ESK laminates, R=O. 1
E
E
40
35
_c 30 +■»O)
m 25
I  20
roc
'E
CO 
0)
Q
45
15
10
50,000 100,000 150,000 200,000 250,000
Cycles
Delamination Length vs. Cycles fo r ESQ and ESM laminates, R=O. 1
ESQ 608 (275 MPa) 
ESQ 602 (276 MPa) 
ESM 107(276 MPa)
.5 30
40,000 80,000
Cycles
120,000 160,000
D
el
am
in
at
io
n 
Le
ng
th
 (
m
m
)
Delamination Length vs. Cycles fo r ESA laminate with feathering, random mat, and
repaired, R=O 1
Repaired ESA (207 MPa)
Repaired ESA (207 MPa)
JKA "FEATHERED" (207 MPa) 
JKA "RANDOM" (207 MPa)
JKA "FEATHERED" (207 MPa) 
JKA "RANDOM" (207 MPa)
5,000 10,000 15,000 20,000 25,000 30,000 35,000