Gregory Ms
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METHODS FOR INTRODUCING CONTROLLED POROSITY
INTO IM6/3501-6 GRAPHITE FIBER
REINFORCED PLASTICS
A Thesis
Presented to
the Faculty of the Graduate School
Tennessee Technological University
by
Richard Eugene Gregory
In Partial Fulfillment
of the Requirements for the Degree
MASTER OF SCIENCE
Mechanical Engineering
December 2003
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CERTIFICATE OF APPROVAL OF THESIS
METHODS FOR INTRODUCING CONTROLLED POROSITY
INTO IM6/3501-6 GRAPHITE FIBER
REINFORCED PLASTICS
by
Richard Eugene Gregory
Graduate Advisory Committee:
Christopher D. Wilson, Chairperson date
Dale A. Wilson date
Jiahong Zhu date
Approved for the Faculty:
Francis OtuonyeAssociate Vice President forResearch and Graduate Studies
Date
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DEDICATION
This thesis is dedicated to my wife Kathryn who has been my support and
encouragement throughout this ordeal. I owe her big time!
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ACKNOWLEDGMENTS
Before I acknowledge anyone else, I must thank God who has blessed me with
the opportunity to pursue the things I enjoy. Now for the people God has surrounded
me with to make sure I succeeded.
First, I would like to thank Dr. Christopher D. Wilson who has been my
advisor and mentor for two years now. I have greatly appreciated his help and advice.
Also, I would like to thank the other members of my committee, Dr. Dale Wilson
and Dr. Jiahong Zhu for their help as well.
Next, I would like to thank the TTU Mechanical Engineering Department, the
Center for Manufacturing Research, and the National Center for Advanced Manufac-
turing whose funding made this research possible.
I would like to thank my fellow researchers Dr. Darrell Hoy, Dr. Sally Pardue,
Dr. Corinne Darvenne, Dr. Joe Richardson, Mike Renfro, Wayne Hawkins, Lance
Lowe, Brahmaji Vasantharao, Scott Smith, and Robert Matthews and especially Mark
Evans for their assistance with many tasks.
I would like to thank Harold Brewer and the late Bob Legg from the Aerostruc-
tures Corporation in Nashville, TN. Also, I owe thanks to James Walker at NASA
Marshall Space Flight Center in Huntsville, AL.
Finally I would like to thank Dr. Phillip A. Allen, without whose support and
encouragement I could not have succeeded.
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TABLE OF CONTENTS
Page
LIST OF TABLES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . x
LIST OF FIGURES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . xi
LIST OF SYMBOLS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . xvi
Chapter
1. Introduction & Terminology . . . . . . . . . . . . . . . . . . . . . . . . 1
1.1 Metals vs. Composites . . . . . . . . . . . . . . . . . . . . . 1
1.2 What is a Composite Material? . . . . . . . . . . . . . . . . 1
1.3 GFRP Fabrication Methods . . . . . . . . . . . . . . . . . . 2
1.4 GFRP Benefits & Applications . . . . . . . . . . . . . . . . . 3
1.5 GFRP Drawbacks & Weaknesses . . . . . . . . . . . . . . . . 5
1.6 A Look Ahead . . . . . . . . . . . . . . . . . . . . . . . . . . 6
2. Problem of Porosity . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7
2.1 Natural Porosity . . . . . . . . . . . . . . . . . . . . . . . . . 7
2.2 Definitions . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9
2.2.1 Pore vs. Void . . . . . . . . . . . . . . . . . . . . . . . . 9
2.2.2 Delamination . . . . . . . . . . . . . . . . . . . . . . . . 9
2.3 Effects of Porosity on Material Properties . . . . . . . . . . . 9
2.3.1 Tensile & Compressive Strength . . . . . . . . . . . . . . 9
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Appendix Page
2.3.2 Flexural Strength . . . . . . . . . . . . . . . . . . . . . . 10
2.3.3 Interlaminar Shear Strength (ILSS) . . . . . . . . . . . . 11
2.3.4 Fatigue Life . . . . . . . . . . . . . . . . . . . . . . . . . 13
2.4 Research Objective: Controlled Porosity Introduction . . . . 14
3. Material & Fabrication . . . . . . . . . . . . . . . . . . . . . . . . . . . 16
3.1 Raw Material . . . . . . . . . . . . . . . . . . . . . . . . . . 16
3.2 Double Vacuum Box Process . . . . . . . . . . . . . . . . . . 17
3.2.1 Process Description . . . . . . . . . . . . . . . . . . . . . 17
3.2.2 Equipment Description . . . . . . . . . . . . . . . . . . . 19
3.2.3 Lay-up Process . . . . . . . . . . . . . . . . . . . . . . . 22
3.3 Controlled Porosity Introduction . . . . . . . . . . . . . . . . 23
3.3.1 Standard Panels . . . . . . . . . . . . . . . . . . . . . . 25
3.3.2 Outer Box Pressure Variation . . . . . . . . . . . . . . . 25
3.3.3 Glass Microballoons . . . . . . . . . . . . . . . . . . . . 26
3.3.4 Bag Bridging . . . . . . . . . . . . . . . . . . . . . . . . 29
3.3.5 AIBN Foaming Agent . . . . . . . . . . . . . . . . . . . 30
4. Porosity Measurement Methods & Results . . . . . . . . . . . . . . . . 32
4.1 Porosity Measurement Methods . . . . . . . . . . . . . . . . 32
4.1.1 Density . . . . . . . . . . . . . . . . . . . . . . . . . . . 32
4.1.2 Matrix Digestion . . . . . . . . . . . . . . . . . . . . . . 33
4.1.3 Ultrasonic C-Scan . . . . . . . . . . . . . . . . . . . . . . 34
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Appendix Page
4.1.4 Micrography . . . . . . . . . . . . . . . . . . . . . . . . . 36
4.2 Porosity Measurement Results . . . . . . . . . . . . . . . . . 37
4.2.1 Specimen Preparation . . . . . . . . . . . . . . . . . . . 37
4.2.2 Digital Imaging . . . . . . . . . . . . . . . . . . . . . . . 38
4.2.3 Image Analysis . . . . . . . . . . . . . . . . . . . . . . . 39
4.2.4 Comparison with C-scan Data . . . . . . . . . . . . . . . 44
4.2.5 SEM Examination . . . . . . . . . . . . . . . . . . . . . 50
4.2.6 Pore Characterization and Measurement in SEM . . . . 57
5. Material Testing Methods & Results . . . . . . . . . . . . . . . . . . . 67
5.1 Equipment Description . . . . . . . . . . . . . . . . . . . . . 67
5.2 Material Characterization . . . . . . . . . . . . . . . . . . . 68
5.3 Tensile Testing . . . . . . . . . . . . . . . . . . . . . . . . . 69
5.4 Damage Threshold Test . . . . . . . . . . . . . . . . . . . . . 70
5.5 Fatigue Testing . . . . . . . . . . . . . . . . . . . . . . . . . 78
6. Conclusions & Recommendations . . . . . . . . . . . . . . . . . . . . . 84
6.1 Summary & Conclusions . . . . . . . . . . . . . . . . . . . . 84
6.2 Recommendations For Future Work . . . . . . . . . . . . . . 86
REFERENCES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 88
APPENDIX
A : TENSILE TEST DATA PLOTS . . . . . . . . . . . . . . . . . . . . . . . 93
B: LIST OF PHOTOGRAPHS INCLUDED ON CD . . . . . . . . . . . . . . 104
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Appendix Page
VITA . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 106
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LIST OF TABLES
Table Page
3.1 Vacuum Pressure Variation Porosity Comparison. . . . . . . . . . . . . 26
4.1 Porosity Measurement Results. . . . . . . . . . . . . . . . . . . . . . . 40
4.2 Vacuum Pressure Variation Porosity Comparison. . . . . . . . . . . . . 46
4.3 Specimen Thickness Variation of Middle Two Laminae. . . . . . . . . . 58
5.1 Summary of Material Characterization . . . . . . . . . . . . . . . . . . 69
5.2 Summary of Tensile Test Results. . . . . . . . . . . . . . . . . . . . . . 71
5.3 Damage Threshold Test Results . . . . . . . . . . . . . . . . . . . . . . 73
5.4 C-Scan Results from Damage Threshold Test . . . . . . . . . . . . . . . 74
5.5 Fatigue Test Results. . . . . . . . . . . . . . . . . . . . . . . . . . . . . 80
5.6 Changes in Stress and Strain from Fiber Misalignment . . . . . . . . . 81
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LIST OF FIGURES
Figure Page
2.1 Naturally Occurring Porosity in Cryogenic Fuel Tank (100) [3] . . . . 8
2.2 Naturally Occurring Porosity in Fuel Tank Nose-cone (100) [3] . . . . 8
3.1 Degas and Compaction Phases [15] . . . . . . . . . . . . . . . . . . . . 18
3.2 Double Vacuum Box Apparatus . . . . . . . . . . . . . . . . . . . . . . 20
3.3 Prestaging Cure Process Timeline . . . . . . . . . . . . . . . . . . . . . 21
3.4 Final Cure Process Timeline . . . . . . . . . . . . . . . . . . . . . . . . 21
3.5 Alternate Full Cure Cycle Timeline . . . . . . . . . . . . . . . . . . . . 22
3.6 Double Vacuum Lay-up Schematic . . . . . . . . . . . . . . . . . . . . 24
3.7 Peel-back View of Bagging Materials . . . . . . . . . . . . . . . . . . . 24
3.8 Graph of Void Fraction vs. Vacuum Level . . . . . . . . . . . . . . . . 27
3.9 Sample of microballoons at 800 . . . . . . . . . . . . . . . . . . . . . 28
3.10 Location of AIBN or Microballoons Insertion During Layup . . . . . . 28
3.11 Bag Bridging Technique . . . . . . . . . . . . . . . . . . . . . . . . . . 29
3.12 AIBN Chemical Chain [18] . . . . . . . . . . . . . . . . . . . . . . . . . 30
4.1 Inverted Light Microscope Photomicrograph of Panel 35 . . . . . . . . 41
4.2 Inverted Light Microscope Photomicrograph of Panel 51 . . . . . . . . 41
4.3 Inverted Light Microscope Photomicrograph of Panel 54 . . . . . . . . 42
4.4 Inverted Light Microscope Photomicrograph of Panel 57 . . . . . . . . 43
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Figure Page
4.5 Inverted Light Microscope Photomicrograph of Panel 61 . . . . . . . . 44
4.6 Mean Signal Strength Ultrasonic C-Scan of Panel 35 . . . . . . . . . . 45
4.7 Mean Signal Strength Ultrasonic C-Scan of Panel 51 . . . . . . . . . . 46
4.8 Mean Signal Strength Ultrasonic C-Scan of Panel 54 . . . . . . . . . . 47
4.9 Mean Signal Strength Ultrasonic C-Scan of Panel 57 . . . . . . . . . . 48
4.10 Mean Signal Strength Ultrasonic C-Scan of Panel 55 . . . . . . . . . . 49
4.11 Mean Signal Strength Ultrasonic C-Scan of Panel 61 . . . . . . . . . . 51
4.12 SEM Photomicrograph of Panel 35 Laminae Interfaces . . . . . . . . . 52
4.13 SEM Photomicrograph of Panel 35 Close-up of Laminae Interfaces . . . 52
4.14 Photomicrograph of Panel 51 Laminae Interfaces . . . . . . . . . . . . . 53
4.15 Photomicrograph of Panel 51 Close-up of Laminae Interfaces . . . . . . 54
4.16 Photomicrograph of Panel 54 Laminae Interfaces . . . . . . . . . . . . . 55
4.17 Photomicrograph of Panel 54 Close-up of Laminae Interfaces . . . . . . 55
4.18 Photomicrograph of Panel 57 Laminae Interfaces . . . . . . . . . . . . . 56
4.19 Photomicrograph of Panel 57 Close-up of Laminae Interfaces . . . . . . 56
4.20 Photomicrograph of Panel 61 Laminae Interfaces . . . . . . . . . . . . . 57
4.21 Photomicrograph of Panel 61 Close-up of Laminae Interfaces . . . . . . 58
4.22 Panel 35 Thickness Variation Measurement of Middle Two Laminae . . 59
4.23 Panel 51 Thickness Variation Measurement of Middle Two Laminae . . 59
4.24 Panel 54 Thickness Variation Measurement of Middle Two Laminae . . 60
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Figure Page
4.25 Panel 57 Thickness Variation Measurement of Middle Two Laminae . . 60
4.26 Panel 61 Thickness Variation Measurement of Middle Two Laminae . . 61
4.27 Representative Round Void in Panel 51 . . . . . . . . . . . . . . . . . . 62
4.28 Representative Elliptic Void in Panel 51 . . . . . . . . . . . . . . . . . 62
4.29 Representative Linear Void in Panel 51 . . . . . . . . . . . . . . . . . . 63
4.30 Representative Round Void in Panel 61 . . . . . . . . . . . . . . . . . . 64
4.31 Representative Elliptic Void in Panel 61 . . . . . . . . . . . . . . . . . 64
4.32 Representative Linear Void in Panel 61 . . . . . . . . . . . . . . . . . . 65
4.33 Representative Resin Rich Region in Panel 51 . . . . . . . . . . . . . . 65
4.34 Representative Resin Rich Region in Panel 57 . . . . . . . . . . . . . . 66
4.35 Representative Resin Rich Region in Panel 61 . . . . . . . . . . . . . . 66
5.1 MTS Test Frame with Specimen and Extensometer . . . . . . . . . . . 68
5.2 Composite Specimen with E-Glass Tabs . . . . . . . . . . . . . . . . . 68
5.3 Preliminary C-scan of Strip 35-07 . . . . . . . . . . . . . . . . . . . . . 74
5.4 Final C-scan of Strip 35-07 . . . . . . . . . . . . . . . . . . . . . . . . . 74
5.5 Preliminary C-scan of Strip 51-10 . . . . . . . . . . . . . . . . . . . . . 75
5.6 Final C-scan of Strip 51-10 . . . . . . . . . . . . . . . . . . . . . . . . . 75
5.7 Preliminary C-scan of Strip 54-09 . . . . . . . . . . . . . . . . . . . . . 75
5.8 Final C-scan of Strip 54-09 . . . . . . . . . . . . . . . . . . . . . . . . . 76
5.9 Preliminary C-scan of Strip 57-09 . . . . . . . . . . . . . . . . . . . . . 76
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Figure Page
5.10 Final C-scan of Strip 57-09 . . . . . . . . . . . . . . . . . . . . . . . . . 76
5.11 Preliminary C-scan of Strip 61-09 . . . . . . . . . . . . . . . . . . . . . 77
5.12 Final C-scan of Strip 61-09 . . . . . . . . . . . . . . . . . . . . . . . . . 77
5.13 Panels 36 and 37 Fatigue Plot, ult=10600 psi. . . . . . . . . . . . . . . 79
5.14 Saw Cut Edge of Composite Specimen . . . . . . . . . . . . . . . . . . 82
5.15 Failed Edge of Specimen 37-7 . . . . . . . . . . . . . . . . . . . . . . . 83
A.1 Tensile Plot for Specimen 35-1 . . . . . . . . . . . . . . . . . . . . . . . 94
A.2 Tensile Plot for Specimen 35-2 . . . . . . . . . . . . . . . . . . . . . . . 94
A.3 Tensile Plot for Specimen 35-3 . . . . . . . . . . . . . . . . . . . . . . . 95
A.4 Tensile Plot for Specimen 35-4 . . . . . . . . . . . . . . . . . . . . . . . 95
A.5 Tensile Plot for Specimen 51-1 . . . . . . . . . . . . . . . . . . . . . . . 96
A.6 Tensile Plot for Specimen 51-2 . . . . . . . . . . . . . . . . . . . . . . . 96
A.7 Tensile Plot for Specimen 51-3 . . . . . . . . . . . . . . . . . . . . . . . 97
A.8 Tensile Plot for Specimen 51-4 . . . . . . . . . . . . . . . . . . . . . . . 97
A.9 Tensile Plot for Specimen 54-1 . . . . . . . . . . . . . . . . . . . . . . . 98
A.10 Tensile Plot for Specimen 54-2 . . . . . . . . . . . . . . . . . . . . . . . 98
A.11 Tensile Plot for Specimen 54-3 . . . . . . . . . . . . . . . . . . . . . . . 99
A.12 Tensile Plot for Specimen 54-4 . . . . . . . . . . . . . . . . . . . . . . . 99
A.13 Tensile Plot for Specimen 57-1 . . . . . . . . . . . . . . . . . . . . . . . 100
A.14 Tensile Plot for Specimen 57-2 . . . . . . . . . . . . . . . . . . . . . . . 100
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Figure Page
A.15 Tensile Plot for Specimen 57-3 . . . . . . . . . . . . . . . . . . . . . . . 101
A.16 Tensile Plot for Specimen 57-4 . . . . . . . . . . . . . . . . . . . . . . . 101
A.17 Tensile Plot for Specimen 61-1 . . . . . . . . . . . . . . . . . . . . . . . 102
A.18 Tensile Plot for Specimen 61-2 . . . . . . . . . . . . . . . . . . . . . . . 102
A.19 Tensile Plot for Specimen 61-3 . . . . . . . . . . . . . . . . . . . . . . . 103
A.20 Tensile Plot for Specimen 61-4 . . . . . . . . . . . . . . . . . . . . . . . 103
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LIST OF SYMBOLS
Symbol Description
m micrometer microstrainCMC Ceramic Matrix Compositedg density of fiberdr density of resing fiber, weight %GFRP Graphite Fiber Reinforced Compositein Hg inches mercury, measure of vacuumMd measured density of specimenMMC Metal Matrix Composite
PMC Polymer Matrix Compositepsi pounds per square inchr resin, weight %SEM Scanning Electron Microscopet1/2 half life in minutesT temperature in KelvinV void content, volume %Vv void volume percent in the specimenVr volume percent of fiber in the specimenVm volume percent of resin in the specimen
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CHAPTER 1
INTRODUCTION & TERMINOLOGY
This chapter will begin with a short discussion contrasting metals and com-
posites. An overview of composite material types will follow along with a discussion
of current fabrication methods. Benefits and applications and drawbacks and weak-
nesses of composite materials will be discussed next. Finally, the outline for the
following chapters will be given.
1.1 Metals vs. Composites
Metals can be forgiving materials to work with because they are well equipped
for dealing with discontinuities. In fact, metals are full of discontinuities. A metal
can slip around discontinuities and plastically deform near stress concentrations to
relieve stresses. This behavior in the presence of flaws is one of the characteristics that
makes metals an attractive engineering material for many fracture critical designs.
One of the fastest growing groups of materials used in the aerospace industry
today are composite materials. Composites materials are fundamentally different
materials and have a different benefits and weaknesses. Before that is discussed, an
introduction to composite materials is needed.
1.2 What is a Composite Material?
The name composite material is given to many engineering materials used
today. Strictly speaking, a composite is any material made up of two or more different
materials blended to make a new material system with two parts: a matrix material
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There are many different methods for fabricating GFRP components. One of
the most widely used is an autoclave which uses a combination of elevated temperature
and pressure to cure the material. Autoclaves can be used on large and small scales
and produces very high quality parts.
There are other methods used in industry that have specific applications. Pul-
trusion is very similar to extrusion processes for metals, except that the composite
is pulled through a die instead of pushed. Pultrusion is good for simple straight ge-
ometries such as L-shaped or T-shaped brackets. In filament winding, a continuous
fiber is wrapped around a mandrel to create a finished part. This is widely used for
round or spherical parts such as pressure vessels. Resin Transfer Molding (RTM) is
similar to injection molding for plastics or metals and uses a die or mold to form
the fibers and then injects the polymer matrix into the mold to cure. RTM can also
be used with a vacuum and is then called Vacuum Assisted Resin Transfer Molding
(VARTM). VARTM is used with specific composite systems that use thermosetting
resins.
A less widely used method in the aerospace industry is the simple vacuum bag
cure method. In this process, the composite is cured under a vacuum at elevated
temperatures. Vacuum bag curing is used much less because it is difficult to produce
a large-scale part with the same quality as an autoclave part.
1.4 GFRP Benefits & Applications
Usually, aerospace designers select a GFRP composite to replace a high strength
to weight ratio metal such as aluminum, titanium, or magnesium. The success of this
choice depends upon the basis of the design, strength, stiffness, weight, fracture re-
sistance, etc. Different composite systems have different strengths and weaknesses
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and may or may not be suited for a certain type of design. The benefits of composite
materials vary from one system to the next. In any composite system, the goal is
to combine the advantages of the two materials while balancing or reducing their
weaknesses. For any composite to be useful, it must have superior properties when
compared to the individual materials used separately.
Strength and stiffness are the first major advantages of GFRP composite ma-
terials. GFRPs possess a relatively high tensile modulus which makes them suitable
for applications where high stiffness and dimensional stability are required such as
airplane wing structures. GFRPs have very high tensile strengths which make them
suitable for high stress applications such as pressure vessels. Because of this high
stiffness and dimensional stability, they can be made to have a near zero coefficient
of thermal expansion which makes them attractive for cases that involve high tem-
perature gradients such as satellite platforms.
GFRPs also have very good fatigue properties. According to Strong [1],
GFRPs can retain as much as 60% of their strength in fatigue applications com-
pared to as little as 10% for some metals. This makes GFRPs a good choice for high
stress, long life applications such as helicopter rotor blades.
An obvious advantage of GFRPs is weight savings over comparable metallic
structures. This weight savings depends heavily on the part geometry and application.
Weight savings over a metal structure could range from 5% for supercritical structures
to as high as 40% for spacecraft applications according to Jones [2]. In designs where
weight is a controlling factor, such as spacecraft, even the smallest weight savings can
have tremendous savings later in reduced fuel costs and increased payload capacity.
An advantage of composites that is often overlooked is the potential for savings
on overall manufacturing and service costs by using a composite material [2]. Since
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a composite is formed during the curing process, a near-net shape can be produced
directly from the mold. Very little machining is required if the initial shape can
be molded close to the final shape. Reduced machining leads to increased material
utilization. Therefore, less material is required and less scrap is produced.
Composite systems tend to be more expensive when compared to noncomposite
alternatives due to higher per-pound material costs. This price increase is usually
offset by other savings or justified by some special characteristic that the composite
offers. In other cases, the raw material costs of the structure may be higher, but over
the life of the part, the total cost may be less due to reduced service requirements or
increased life span of the structure.
1.5 GFRP Drawbacks & Weaknesses
While the near-net shape of some composites is an advantage, production rates
can be a major disadvantage. Composite manufacturing methods are time intensive
and do not lend themselves well to mass production so they cannot produce parts in
the quantities required for large scale production. For this reason, composite materials
still do not have a strong presence in the automotive industry.
One notable exception to this is the use of composites in the sporting equip-
ment industry. Golf clubs, tennis rackets and fishing poles are examples of high
volume production parts that have been successfully produced with composites. This
is mainly due to the simplicity of the part being produced and advances in specialized
processing techniques that have reduced overall production times for these applica-
tions.
GFRPs can be very sensitive to flaws and may not behave well in the presence
of a flaw. This makes fracture mechanics and important consideration when designing
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with composites. Stress concentrations can severely degrade the performance of the
part because GFRP cannot deform plastically to relieve the stresses. Some common
flaws in GFRP include foreign object matter, kinked or broken fibers, and porosity.
The behavior of GFRP in the presence of these situations is called damage tolerance
and is the driving force behind this research.
1.6 A Look Ahead
In Chapter 2, the focus of this research will be introduced and discussed along
with previous work and publications in this field. Objectives for this research will bediscussed at the end of Chapter 2. The fabrication method and porosity introduction
methods used in this research will be discussed in Chapter 3. Chapter 4 will discuss
the porosity measurement techniques and results. The mechanical tests conducted
and their results will be discussed in Chapter 5. Finally, a summary of findings and
recommendations will be presented in Chapter 6.
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CHAPTER 2
PROBLEM OF POROSITY
First in this chapter, natural porosity samples will be examined to compare to
porosity created later in this research. Next, some terms used in this research will be
defined. Previous research will be examined to determine the effects of porosity on
the mechanical behavior of GFRPs. Finally, the objectives for this research will be
outlined.
2.1 Natural Porosity
Porosity is a naturally occurring phenomena in composite manufacturing pro-
cesses. Porosity has a variety of causes, such as incorrect curing pressure, and tooling
leaks. Figures 2.1 and 2.2 are examples of naturally occurring porosity from some
composite structures. These picture were taken by James Walker [3] at NASA Mar-
shall Space Flight Center in Huntsville, AL. These samples are of autoclave cured
fibrous PMC composites.
The location of the porosity should be noted here and will be of interest in
samples produced in this research. Figure 2.1 is part of a cryogenic fuel tank project
for NASA. The porosity in this sample is very large and is aligned along the laminae
interfaces. Figure 2.2 is part of the nose-cone for the Space Shuttles external fuel
tank. The porosity in this sample is smaller and more round and occurs within the
plys, not just at the interfaces. The void fraction in both samples was approximately
12%.
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Figure 2.1. Naturally Occurring Porosity in Cryogenic Fuel Tank (100) [3]
Figure 2.2. Naturally Occurring Porosity in Fuel Tank Nose-cone (100) [3]
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2.2 Definitions
2.2.1 Pore vs. Void
Much of the discussion in this research will center around the occurrence of
porosity and sites that will be called pores and voids. Porosity will be used as a
general term to describe the occurrence of any gas pockets inside a laminate. Pore
and void are used interchangeably in literature and will be so here. However, both
terms will be used to describe features that are at least as large as one fiber diameter,
approximately 4 m.
2.2.2 Delamination
Delamination is the separation of a cured laminate composite along the laminae
interface lines. This failure mode effectively splits the laminate into its separate layers.
With nothing holding them together, the layers are much weaker by themselves that
when they are joined together. Delamination is of interest when discussing porosity
because porosity tends to be located at lamina interfaces. Smaller voids can also
coalesce during the cure cycle and start behaving like a delamination.
2.3 Effects of Porosity on Material Properties
2.3.1 Tensile & Compressive Strength
Tensile strength is the property that determines or defines a materials ability
to withstand uniform tensile loading. Tensile strength is of interest because fibrous
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to introduce porosity. Porosity levels of 0.0% to 3.1% porosity were produced. Grels-
son noted that at higher porosity levels, the individual pores tended to cluster and
merge into larger voids that act like small delaminations. His test indicated a pat-
tern of reduced flexural strength with increased variation as porosity level increased.
Low porosity levels specimens had acceptable strengths. However, the trend was
that strength levels were decreasing with increasing porosity. This trend tended to
accelerate though as higher porosity levels were tested.
Work by Kan et al. [7] focused on molded composite flanges. In produc-
tion parts, it was noticed that porosity tended to occur where curing pressures were
nonuniform, such as in the region of a curved section. Specimens were taken from a
production part with naturally occurring porosity. The specimens were examined by
multiple methods to determine the porosity level and then tested under pure bending
conditions. The study suggested that a drop of 30% to 50% in interlaminar tensile
strength can occur with high levels of porosity.
2.3.3 Interlaminar Shear Strength (ILSS)
InterLaminar Shear Strength (ILSS) is the ability of a laminated material to
resist sliding between individual layers. ILSS mainly affects the matrix because the
matrix material is the only part of the composite that supports shear stresses. ILSS
can be easily illustrated by a stack of paper. The individual pages slide against each
other when shear stresses are applied to the upper and lower boundaries of the stack.
ILSS is greatly reduced when voids are present in a laminate because these voids
usually occur at the laminae interfaces where the laminate is most susceptible to
shear stresses.
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In some literature, ILSS has been called the Achilles heel of composite materials
since shear forces load the weakest part of the composite, the matrix. Pipes [8]
submitted short-beam specimens to fatigue loading to determine the reduction in
ILSS over the life of the composite. He found that while fiber tension possesses an
endurance of 80% of the ultimate strength, the ILSS fatigue limit diminished to less
than 55% after one million cycles.
Research done by Hancox [9] looked at the shear response of composite tubes.
Solid composite tubes with varying degrees of porosity were tested in torsion to de-
termine the change in shear modulus and shear strength. Previous studies in this
area predicted a drop of 10% to 50% in these properties with void fractions in the
5% range. However, Hancox found a steady reduction in these properties with a 70%
reduction at a 5% void fraction which was the limit of his research.
Judd and Wright [10] performed a survey of over 47 previous studies into the
effects of porosity on mechanical properties of composites. The majority of these
studies focused on porosity values between 0.0% and 5%. ILSS was the property
most often reviewed in the works and was quantified more completely. From this
survey, they determined that there exists a linear decrease in ILSS of 7% per 1%
increase in porosity with other properties being affected to a similar extent. This
relationship was found to be true regardless of the resin or fiber used. There were
indications that a more serious drop in ILSS may exist between 0.0% and 1%, but a
more accurate method of measuring porosity in needed to confirm this.
Yoshida et al. [11] looked at a statistical relationship between ILSS and poros-
ity. It was noted that mechanical reliability of composites decreases with increased
porosity. They used a chemical foaming agent to introduce porosity into their spec-
imens for study and verified the void fraction by density measurements. The ILSS
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amounts of porosity by varying the level of the bag vacuum and the autoclave pressure
during the cure cycle. Specimens with void fractions of 0.11% to 2.66% were produced
by this process and were submitted to static tensile tests and zero-max fatigue tests.
They noticed a 10% reduction in tensile strength. They also noticed a reduction in
fatigue life of 20% at low stress levels and up to 50% at high stress levels.
2.4 Research Objective: Controlled Porosity Introduction
More research is needed into the behavior of composite materials in the pres-
ence of porosity. Such research would benefit industry in the disposition of as-builthardware with naturally occurring porosity. Currently, such parts are discarded or
reworked when possible. A better understanding of the behavior of GFRP in the
presence of porosity would lead to clearer rules for using or disposing of production
components with porosity and more accurate design tools for engineers to work with.
The NonDestructive Evaluation (NDE) and manufacturing communities would
benefit from being able to produce composite specimens with controlled amounts of
porosity. Standards are needed to use with calibration and training. A well under-
stood and documented method of porosity introduction would assist in the production
of such standards.
To gain this better understanding or produce such standards, samples must
be produced for study. A method of creating porosity in a predictable and repeat-
able manner is needed to produce these samples. This research will concentrate on
evaluating proposed methods and selecting one method to be used in future research.
Four methods of porosity introduction will be studied in this research:
1. vacuum pressure variation
2. microballoon introduction
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3. bag bridging
4. chemical foaming agent introduction.
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CHAPTER 3
MATERIAL & FABRICATION
The previous chapter focused on the effects of porosity on the mechanical
behavior of GFRPs. This chapter will move on to the fabrication of GFRP samples
produced for this research. A description of the manufacturing method will be given
followed by an explanation of the four porosity introduction methods.
3.1 Raw Material
The panels fabricated in this study were made of IM6/3501-6, an intermediate
modulus (IM6) fiber, with a thermoset polymer resin (3501-6). The raw material was
a prepreg tape purchased from Cytec Engineered Materials on a 60.75 in. roll. The
prepreg had a 63% fiber content and had an areal weight of 143.3 g/m 2, commonly
referred to as grade 145. The particular pedigree was a formulation made for Bell
Helicopter, so much of the information about the resin composition and the fiber
treatment was not available.
The panels made for testing were eight layer laminates, [0/45/90/ 45]S, and
were approximately 0.050 in. thick when fully cured. Material properties were ob-
tained from the website of the University of California San Diego Composites and
Aerospace Structures Laboratory [14] which were using the same material system,
but a different processing technique. For this reason, independent material charac-
terization tests were conducted and will be discussed later in Chapter 5.
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3.2 Double Vacuum Box Process
3.2.1 Process Description
An autoclave process has historically been the preferred manufacturing process
in the aerospace industry for carbon and graphite fiber reinforced plastics. This pro-
cess uses elevated temperatures and pressures to produce a very high quality, nearly
void-free product. However, it requires some very large and expensive equipment,
making it impractical for producing very small parts such as those used for research
or damage repair applications. Glen Sherwin [15] at The Aerostructures Corporationin Nashville, TN, conducted a research project to investigate a non-autoclave process
for producing a partially cured, or staged, composite laminate for use as an in-field
repair patch.
A double vacuum out-of-autoclave process was investigated. It employs a
conventional vacuum bag to provide the first vacuum and a steel outer box to provide
the second vacuum. The cure cycle is divided into two parts: 1) staging and 2)
final cure. The staging process produces a partially cured laminate that can be fully
cured immediately or stored for future use. In either case, the staged laminate can
be reheated and formed to the contour of a damaged structure. The final cure cycle
produces a fully cured laminate that is ready for use.
The staging process consists of two phases: Degas and Compaction, illustrated
in Figure 3.1. In the Degas phase, a vacuum of approximately 27 in. Hg is applied to
both the inner vacuum bag and the outer vacuum box to remove any pressure gradient
between the two vacuum volumes. The lack of a pressure gradient effectively removes
the compacting force of the bag which allows adequate pathways for any gases to freely
escape the laminate. The vacuum also draws entrapped gases out of the laminate
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Figure 3.1. Degas and Compaction Phases [15]
instead of pushing it out as an autoclave would do. In the Compaction phase, the
outer box is vented to the outside atmosphere and the vacuum is maintained on the
inner vacuum bag producing a pressure gradient which applies a force from the bag.
This force compacts the layers of the laminate together. Both of these phases are
carried out at elevated temperatures to ensure that the viscosity of the resin is low
enough so that it will bind the layers together. The temperatures in both phases are
kept below the full cure temperature of the matrix system to prevent complete curing
of the laminate.
Sherwin found that the critical parameters are the dwell temperatures at each
stages and the timing for switching from Degas to Compaction. After some experi-
mentation, Sherwin found that the optimal temperature for these phases was related
to the viscosity profile of the resin system. The optimal situation was to have the
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Figure 3.2. Double Vacuum Box Apparatus
and soak parameters so that the entire process can be programmed and run as one
continuous cycle or run as separate parts, whichever is needed. The stage profile is
shown in Figure 3.3, and the final cure cycle is shown in Figure 3.4. The alternate
full cycle is shown in Figure 3.5. All laminates in this study were made with the full
cycle option.
The vacuums were supplied by two Cenco Hyvac 14 vacuum pumps, which
were capable of producing a 27 - 29 in. Hg vacuum depending on the quality of the
seal. The pumps were controlled by the operator and had to be powered on and
off manually at the appropriate times during the cure cycle. This process could be
automated with the addition of a programmable logic circuit.
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0 50 100 150 2000
50
100
150
200
250
300
Time (minutes)
Temperature(F
)
Prestage Cure Timeline
Figure 3.3. Prestaging Cure Process Timeline
0 50 100 150 200 250 300 350 400 450 500 5500
50
100
150
200
250
300
350
400
Time (minutes)
Temperature(F
)
Final Cure Timeline
Figure 3.4. Final Cure Process Timeline
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0 50 100 150 200 250 300 350 4000
50
100
150
200
250
300
350
400
Time (minutes)
Temperature(F
)
Alternate Full Cycle Timeline
Figure 3.5. Alternate Full Cure Cycle Timeline
3.2.3 Lay-up Process
The laminate was bagged with several layers and kinds of materials to allow
for adequate air removal and resin flow. Figures 3.6 and 3.7 show the different layers
in the bagging process. The bag and laminate were laid up on a composite-backed
honeycomb base plate. This stiff base plate was needed so that the system would not
warp when the vacuum is applied. The bottom layer was a fiberglass N10 breather
cloth which provides pathways for degassing. Above the N10 cloth, was the heater
blanket that supplied the thermal energy for the process. On top of the heater blanket,
was a layer of porous teflon followed by a layer of non-porous teflon. These layersprotected the heater blanket from the resin since it runs when its viscosity decreased.
A thin aluminum caul sheet was placed next to support the laminate and to give
it a better surface finish. An adhesive backed layer of nonporous teflon was applied
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to the aluminum caul sheet to prevent the laminate or run-off resin from bonding
to it. Layers of nonporous and porous teflon film were layered above and below the
laminate to aide in controlling and facilitating resin run-off.
The nonporous layers had a tendency to wrinkle during curing which left im-
pressions in the cured panel. This wrinkling may be caused when the laminate shrunk
during curing. This layer did not shrink in the same manner but instead, stuck to the
laminate and was wrinkled in the process. No method of eliminating this phenomena
was found.
A second aluminum caul sheet with adhesive backed nonporous teflon was
placed on top of these layers between the compaction stage and the final cure to
further compact the laminate in the final cure and provide a better surface finish on
the top surface of the laminate. Another layer of N10 breather cloth was applied on
top to allow for degassing. Finally, the vacuum bagging material was applied over
the top and was sealed to the base plate with vacuum bag sealant tape. This sealant
tape was very sticky and had a putty-like consistency. It had a very high resistance
to shear so that it did not smear when the vacuum was applied. A two piece vent
base was used to pull the vacuum on the bag. The bottom piece was placed inside
the bag and a hose was screwed into it to connect to the vacuum pump.
3.3 Controlled Porosity Introduction
Before discussing panels and samples examined in this research, some explana-
tion is needed of the numbering system for panels, strips, samples, and photographs.
Panels in this research were numbered in the order they were made and labeled with a
two digit number, such as Panel 35. Strips were then cut from these panels for testing
or examination and numbered in the order cut. These strips were designated by the
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panel number followed by the strip number such as strip 35-5. Specimens cut from a
strip were numbered as cut and labeled with the panel, strip and specimen number
such as specimen 35-5-4. Finally, photographs taken of a specimen were numbered
in the order taken and labeled with the panel, strip, specimen, and picture number
such as picture 35-5-4-01.
3.3.1 Standard Panels
Before trying to introduce porosity, several panels were made to learn about
the manufacturing process. Through this exercise, the quality laminate that could beproduced using this process was better understood. As with Sherwins work, these
standard panels were found to be virtually free from porosity. The average thickness
of this panel was 0.045 in. with a standard deviation of 0.00058 in. measured in 24
locations.
All panels were examined using the ultrasonic C-scanning technique after fabri-
cation. The mean signal strength from these scans was used to determine the relative
porosity of the panel as well as to gauge the uniformity of the panel with higher
signal strength indicating lower porosity levels. This technique was used by fellow
researcher Lance Lowe [17] and is discussed further in Chapter 4 along with the scans
from each panel. All scans used in this thesis were conducted by researchers Lance
Lowe, Brahmaji Vasantharo, Scott Smith, and Robert Matthews.
3.3.2 Outer Box Pressure Variation
The first method of porosity introduction attempted was varying the vacuum
level on the outer box during the degas phase of the prestaging process. This technique
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was similar to a single vacuum processing method where the only vacuum applied
was to the bag. In this method, the pressure was varied from full vacuum up to no
vacuum. Nine panels were made in this way to determine what range of porosity
could be achieved with this technique. It was determined that with no outer vacuum
during the Degas phase, panels with approximately 3.0% maximum porosity could
be created. Panels with vacuums of 23, 20, 17, and 14 in. Hg were produced and
compared to panels produced with full vacuums. Table 3.1 Figure 3.8 show the
data for panels created with this method. The general trend was higher porosity with
lower vacuum level. The average thickness of this panel was 0.042 in. with a standard
deviation of 0.00044 in. measured in 24 locations.
Table 3.1. Vacuum Pressure Variation Porosity Comparison.
Panel # Vacuum Level (in Hg) Void Fraction21 0 2.8723 14 1.9224 14 0.2025 17 0.5226 17 0.9027 20 0.35
28 20 0.4029 23 0.0230 23 0.5817 27 0.0418 27 0.13
3.3.3 Glass Microballoons
The second method of porosity introduction tried was the inclusion of glass
microballoons. These microballoons are used in some marine and aerospace appli-
cations as a resin filler to reduce weight. To determine the effect the microballoons
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0 5 10 15 20 25 300
0.5
1
1.5
2
2.5
3
Vacuum Level (in. Hg)
VoidFraction(%)
Void Fraction vs. Vacuum Level
Figure 3.8. Graph of Void Fraction vs. Vacuum Level
on the ply interface, some information about their size distribution was needed. Five
small 300 m by 300 m slides of the microballoons were examined in a Scanning
Electron Microscope (SEM) by lab technician Wayne Hawkins and were found to
have a large size variation. The largest and smallest microballoons were found to be
76.88 m and 3.75 m, respectively with a mean and standard deviation of 27.13 m
and 13.34 m, respectively. Figure 3.9 shows a sample of the microballoons. The
large sphere in the center was 76.87 m in diameter.
In panel 54, a 3-inch wide strip of microballoons was introduced in the three
middle ply interfaces during layup (Figure 3.10). The microballoons were applied by
hand with a smearing action. The average thickness of this panel was 0.046 in. with
a standard deviation of 0.0016 in. measured in 21 locations.
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Figure 3.9. Sample of microballoons at 800
Figure 3.10. Location of AIBN or Microballoons Insertion During Layup
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Figure 3.12. AIBN Chemical Chain [18]
3.3.5 AIBN Foaming Agent
The final method attempted for introducing porosity was the application of a
foaming agent. AIBN is a expanding foam chemical agent often used as a catalyst in
many chemical reactions. It is a hydrocarbon chain with nitrogen attached at both
ends. As the temperature increases, the hydrocarbon chain begins to break down and
release nitrogen gas. Yoshida et al. [11] used a similar agent in their experiment to
create trapped air pockets called AZDN. According to Yoshida, 1 gram of AZDN will
give off 137 cm3 of nitrogen gas.
The AIBN was dissolved in acetone so that it could be sprayed onto the laminae
during the lay-up procedure. A model paint sprayer with an aerosol spray can was
used to apply the mixture. By doing this, the acetone evaporated quickly leaving
behind only the AIBN. The unfortunate problem with this method was the inability
to accurately determine the amount of AIBN applied to the laminae since some of
the mixture is lost to the surrounding atmosphere. The amount of AIBN applied
could not be measured because it was less than 1 gram. This difference was toosmall to be measured with any mass balance that was available for this research. The
average thickness of this panel was 0.043 in. with a standard deviation of 0.00091 in.
measured in 24 locations. Two panels were made with the addition of this mixture
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while keeping all other parameters constant. Two more panels were made with the
added change of removing the outer vacuum during the Degas phase as in the vacuum
variation method. It was believed that this combination might create a compounding
effect and boost the void content beyond what could be created with one method
alone.
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CHAPTER 4
POROSITY MEASUREMENT METHODS & RESULTS
The previous chapter focused on the manufacturing methods used to produce
the specimens for this research. This chapter will look at porosity measurement. Dif-
ferent methods of calculating porosity will be discussed first followed by the discussion
and results of the porosity measurements made in this research.
4.1 Porosity Measurement Methods
Over the years, many methods have been introduced to determine the relative
porosity, or void fraction, of a composite laminate. Each method has its strengths
and weaknesses. Methods used for this research will be discussed here along with
other methods that were considered but not used.
4.1.1 Density
The porosity measurement method in ASTM standard D2734 [20] uses the
density of the resin matrix, fiber reinforcement and total composite along with the
resin content of the total composite to determine the void fraction. Accurate mea-
surements must be made of the resin density and can be done using ASTM standard
D792 or D1505. Any error in this measurement will result in a large error later, per-
haps leading to the calculation of a negative void fraction in low porosity specimens.The density of the fibers should be measured using ASTM D792. Information from
the manufacturer on both resin and fiber densities may be used if the results are
certified for that batch of material. Resin content should be measured per ASTM
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which acts as a coupling agent between the two. A high frequency sound wave is sent
through the sample by a pulse receiver transducer and the echo of the signal is picked
by the same transducer. As the signal passes through the sample it is scattered or
absorbed by irregular features such as porosity and foreign matter inclusions. This
reduces the strength of the signal returned to transducer. As the number of these
irregular features increase, the overall mean signal returned decreases. The signal
amplitude is then plotted as a histogram. From this the mean signal strength is
determined which indicates the amount of porosity in the sample. In general, a
higher mean signal strength indicates lower porosity.
Steiner [19] studied the ability of C-scans to detect different types of defects
inside composite materials. Porosity was detected in samples with void fractions
ranging from 0.17% to 3.57%. The mean signal strength of the histogram of the scan
was found to depend greatly on the porosity content. Steiner noted the success of the
method at detecting the presence of porosity, but noted that in order to define the
level of the porosity, a number of known porosity level specimens would have to be
scanned to build a comparison database. This would also have to be done for each
material system examined. C-scanning was also used Grelsson [6] and Kan et al. [7]
in their work to study porosity in composites.
Another advantage to this method was that the entire specimen is examined.
The data gathered from this method produced a full color plot of the entire specimen
scanned so that areas of signal strength change can be easily identified. However,
as noted by Steiner, this method must be calibrated with standard samples. These
samples must be verified by other existing methods and therefore, cannot have an
accuracy better than that of the verification method as discussed by Judd [10].
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The verification of this method for this research was done with the help of Lowe
[17]. Several panels were fabricated, scanned, and then examined by light microscopy.
The porosity level was then calculated from digital image analysis software. Lowe then
compared these measured porosity numbers to the bulk and zoomed C-scan histogram
data that was collected to determine a mapping from signal strength to porosity level.
4.1.4 Micrography
Although not an ASTM standard, the microscopic examination method is often
used to determine porosity levels. In this method, a small specimen is cut, mounted,ground, polished and then examined with a microscope to quantify porosity levels.
In early studies, a point counting method was used where a very fine grid was placed
over the specimen and the void fraction calculated by counting the grid intersections
that coincided with a void and dividing by the total number of points in the grid.
Working with the point counting technique, Judd et al. [10] proposed that accuracy
for this method is about 0.50% and is as accurate as other current techniques such
as matrix digestion and density measurement.
In recent years, advances in computer hardware and software have produced
a number of graphics analysis programs that will count defined particles in a digital
image. This method can be adapted to measure porosity if there is sufficient contrast
between the voids and the surrounding material. Cilley et al. [22] found that this
method could be used in conjunction with other techniques and provide acceptable
results.
There are two major drawbacks to this method. First, like matrix digestion,
microscopic examination is a destructive measure. This is prohibitive in a production
environment where the area of interest cannot be cut and still be used. Second, it is
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eye. These sectioned pieces were then mounted in a two-part castable epoxy, or cold
mount, and allowed to cure overnight.
Grinding and polishing was done with a Buehler automated grinder-polisher.
ASM Handbook 9 [23] was consulted to determine the proper method for surface
preparation. The grinding was done with 320, 400, and 600 grit grinding discs while
flushing with water. Polishing was done with 6 m and 3 m diamond suspension
and 1 m cerium oxide on soft matte pads. The specimens were ground and polished
in 30 second intervals and examined for adequate material removal, then repeated as
necessary.
4.2.2 Digital Imaging
The specimens were photographed with a digital microscopy system and an-
alyzed with the computer software program Image-Pro Plus [24]. The photographs
are taken through an inverted light microscope at approximately 140 magnification.
Eight photographs were taken from each specimen, which were then examined indi-
vidually to determine their void fraction. The void fraction was determined by the
program using the built in particle counting function. The program differentiated
between the voids which appear black from the matrix and fibers which appear gray
and white. There was some level of subjectivity involved in determining the contrast
level to be considered a void. This could be overcome by the experience gained from
repeated measurements and familiarity with the process and the material. The aver-
age particle size, number of particles and void fraction for each image was calculated
by the program. The void fraction was then calculated as the particle area divided by
the total specimen area. The average void fraction of the eight images was taken to
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get the void content for the total specimen. Four specimens from different locations
out of each panel were examined to increase the validity of the measurement.
This method was verified by the manual point counting method described by
Judd [10]. The comparison was done on one sample in work done by Hoy et al.
[26] using a computer program called Global Lab [25]. A verification study was done
to determine if sufficiently accurate results could be obtained using the Image Pro
Plus software. One sample was examined using both programs by the same user.
Measurements of 2.57% and 2.86% were obtained for the chosen test piece with the
two different programs. This difference is well within the accepted 0.5% accuracy
of the method.
4.2.3 Image Analysis
The porosity measurements from Image-Pro Plus are listed in Table 4.1. The
measured void fraction of the chosen standard panel was actually the largest of any
standard panel created in this research. It is likely that some features counted as
porosity were really not porosity. The vacuum pressure variation panel had a lower
measured void fraction than expected and did not appear to be very different from
the control panel. However, examination in the ESEM shows that these panels are
actually quite different in their porosity content. Panels 54 and 61 have high mea-
sured porosity contents as expected. Panel 57 was expected to have a high measured
porosity content but apparently does not. The nature of the porosity, or lack thereof,
is of great interest and is examined later in this chapter.
In Figures 4.1 to 4.5 the third and fifth layers are oriented parallel to the cut.
This made measuring porosity in these areas difficult if not impossible. These layers
have many dark regions where fibers are dropping in and out of the plane that appear
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Table 4.1. Porosity Measurement Results.
Panel # Method Void Fraction (%) Mean Signal Strength35 Standard 0.75 49.39
51 Pressure Variation 1.05 17.0054 Microballoons 2.01 38.8757 Bag Bridging 0.36 40.8961 AIBN 5.93 8.10
Note: Higher mean signal strength indicates lower porosity.
to be porosity to the software. Other features like scratches from metallographic
preparation can also get counted as porosity. This all contributes to the 0.5%
accuracy that is cited by Judd [10].
The standard panel specimens from panel 35 had a void fraction of 0.75%.
However, further examination of the photomicrographs shows that the actual value
is much less than that which was evident from the C-scan data. In Figure 4.1, the
cross section of a standard panel is shown. Notice that there were no voids visible
anywhere on the cross-section. Further, the individual laminae were touching each
other and are difficult to distinguish in some places. There were also no noticeable
resin rich regions. All of these factors define a good quality standard panel with low
porosity.
The vacuum pressure variation specimen from panel 51 had a void fraction of
1.05%. In Figure 4.2, notice the dark spots that were visible in the middle region
of the specimen. These were voids in the laminate. They were primarily located at
the laminae interfaces and were generally elliptical in shape and sometimes group
together to form larger voids. There were very few resin rich regions visible in this
specimen.
Panel 54 contained microballoons and, as expected, had a higher void fraction
of 2.01%. However, the porosity created in this panel was unlike the porosity created
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Figure 4.3. Inverted Light Microscope Photomicrograph of Panel 54
in panel 51 with the vacuum pressure variation. In Figure 4.3, the cross-section of a
specimen from panel 54 is shown. The dark spherical objects were the microballoons
that were included in the layup between the plys. Also, the microballoons were
still located at the laminae interfaces and did not migrate through the layers. Their
presence also created large resin layers at each interface where they were present. The
presence of this large deposit of resin was of great concern, because it changes the
way the laminate behaves, likely reducing the interlaminar shear strength.
Panel 57 was the bridged sample. This specimen was expected to have a very
high void fraction, but when measured, was found to have a very small void fraction
at 0.36%. Surprisingly, this measurement was less that the standard panel. However,
notice in Figure 4.4 the angled layers in the the top of the laminate. In this figure,
the specimen was getting thicker on the right side by approximately 0.015 in. This
phenomena was curious since the specimen was believed to have been cut directly
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Figure 4.4. Inverted Light Microscope Photomicrograph of Panel 57
through the bridged region. Very little porosity was visible, but other problems exist
that were not yet evident and will be seen in the Section [?].
Panel 61 was sprayed with the AIBN foaming agent and was also expected
to have a high void fraction. The measured porosity was 5.93%, the highest of any
panel manufactured in this research. A typical cross section is shown in Figure 4.5.
The poor quality of this specimen was immediately evident. Large voids along with
significant pockets of resin were visible throughout the specimen. In earlier specimens
with the other methods, voids were only found at the laminae interfaces. Here, they
were found in all locations through the panel and not just along the three middle
interfaces where the AIBN was sprayed. This spread indicates that the nitrogen
gases from the AIBN did migrate during the cure cycle.
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Figure 4.5. Inverted Light Microscope Photomicrograph of Panel 61
4.2.4 Comparison with C-scan Data
Figure 4.6 shows the C-scan of panel 35. In this figure, small blue lines were
noticed in the midst of the green surrounding areas that indicated a drop in signal
strength. The green indicated a higher mean signal strength while the blue repre-
sented a lower mean signal strength. The blue lines were caused by the impressions
left by the wrinkled non-porous plys during curing and were visible on the surface of
the panel. The red spots at the top were most likely caused by an air bubble trapped
underneath the panel in the water tank. This scan also indicated that this panel
was fairly uniform. From Table 4.1, the mean signal strength was 49.39, the highestof any of the five panels used in this study which would indicate that the porosity
possibly lower than the 0.75% measured by micrography.
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Figure 4.8. Mean Signal Strength Ultrasonic C-Scan of Panel 54
The C-scan of panel 57 is shown in Figure 4.9. As with the microballoons, the
area where the bridge was applied is easily seen from the C-scan. A first look at the C-
scan would imply that there was a large region of porosity under the bridge. However,
the unevenness indicated that any created porosity was not as evenly distributed as
with the microballoons or the vacuum pressure variation. The panel also had thickness
variations of up to 40% in this middle region. Since the time of flight of the sound
wave was used to calculate the mean signal strength, this thickness change may be
the reason behind the mean signal strength change. This observation is supported by
the void fraction of 0.35% measured by micrography.
However, it was later discovered that the panel was manufactured incorrectly,
with the bridged region running parallel to the outer layer fibers instead of perpen-
dicular. The first clue was that the low signal region in Figure 4.9 was perpendicular
to those in the microballoon panel and the AIBN panel. At first, this was thought
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where:
t1/2 = half life in minutes,
T = temperature in Kelvin.
At the 225F prestaging temperature, the half life of the AIBN is only 2.75
minutes. Since it took the laminate 52 minutes to heat up to this point from room
temperature and then dwelled there for 25 minutes, it seems very likely that the
nitrogen gas was escaping from the laminate during the prestage phase.
Therefore, panel 61 was made the same way as panel 55 except the outer
vacuum was left off during the degas phase to attempt to capture the nitrogen within
the layers. Also, only the middle three interfaces were sprayed with AIBN in panel 61.
The C-scan of panel 61 can be seen in Figure 4.11. In this figure the area where the
AIBN was applied can be clearly distinguished from the surrounding area implying
that the nitrogen gas was entrapped in this region and producing significant porosity.
This conclusion was supported by the 5.93% void content measured by micrography.
4.2.5 SEM Examination
One specimen from each panel was further prepared for examination in the
Scanning Electron Microscope (SEM). The specimens were coated with a thin layer
of gold to discharge electrons in the chamber since the specimens are nonconductive.
The purpose of examining in the SEM was to determine the shape and size of the
porosity produced and to determine the effect the different methods might have on
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Figure 4.12. SEM Photomicrograph of Panel 35 Laminae Interfaces
Figure 4.13. SEM Photomicrograph of Panel 35 Close-up of Laminae Interfaces
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Figure 4.14. Photomicrograph of Panel 51 Laminae Interfaces
Notice that the porosity in this panel was located predominately at the laminae
interfaces and furthermore, mainly at the laminate centerline. It is not certain if
the location of these sites was a property of the method used or a function of the
amount of porosity present. All panels manufactured by this method exhibited a
similar pattern.
Specimen 54-5-2 was examined to further determine how the microballoons
affected the laminate. From Figure 4.16, the microballoons can be seen to be widely
distributed along the middle three interfaces. In the SEM, they glow white because
the electrons build up on the sharp edges, causing them to charge. A closer look in
Figure 4.17 shows that there are many more microballoons present than were visible
in lower magnification pictures. Some broken pieces that have been filled with resin
are clearly visible and may likely create a structural weakness in the panel. Also,
the resin rich laminae interface region is the same thickness as the microballoons,
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Figure 4.16. Photomicrograph of Panel 54 Laminae Interfaces
Figure 4.17. Photomicrograph of Panel 54 Close-up of Laminae Interfaces
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Figure 4.18. Photomicrograph of Panel 57 Laminae Interfaces
Figure 4.19. Photomicrograph of Panel 57 Close-up of Laminae Interfaces
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Figure 4.20. Photomicrograph of Panel 61 Laminae Interfaces
much of what was seen earlier from the photomicrographs from the light microscope
(Figure 4.5. Figure 4.21 shows a close-up of the cross-section. Here the extent of the
porosity can really be seen. Laminae are almost destroyed in some places by resin
layers and voids. Laminae thickness also vary greatly from point to point. Laminae
interfaces were also not consistently lined up along the length as seen in panel 35.
The voids have permeated every layer of the material. This was definitely a high
porosity, low quality specimen.
4.2.6 Pore Characterization and Measurement in SEM
The pictures taken in the SEM were then measured using built-in functions of
the software. Pictures from each specimen were examined to determine the changes in
specimen thickness and ply thickness as well as to measure the size of resin rich regions
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Figure 4.22. Panel 35 Thickness Variation Measurement of Middle Two Laminae
Figure 4.23. Panel 51 Thickness Variation Measurement of Middle Two Laminae
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Figure 4.30. Representative Round Void in Panel 61
Figure 4.31. Representative Elliptic Void in Panel 61
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CHAPTER 5
MATERIAL TESTING METHODS & RESULTS
The previous chapter looked at the porosity measurement techniques and
results. This chapter will discuss the mechanical testing performed on the test speci-
mens. Tensile testing, damage threshold testing and fatigue testing are all discussed
in this chapter.
5.1 Equipment Description
A MTS 810 servohydraulic test frame with a 20000 pounds maximum load
cell was used to perform tensile and fatigue testing in this research. The frame was
equipped with MTS 647 hydraulic wedge grips with interchangeable grip faces to grip
flat or round samples. The system was run by a MTS TestStar IIs controller. The
extensometer used for tensile testing was a MTS model 632.25B-20 which has a gage
length of 2 in. The load frame and setup used are shown in Figure 5.1.
Tensile and fatigue specimens were prepared per ASTM Standard D5687 [27].
The panels were cut with an abrasive cutoff wheel with water coolant. This method
was compared to a ban saw and was found to produce a cleaner edge. The specimens
were tabbed with E-glass fiberglass tabs to improve gripping and reduce damage done
to the specimen by the grips. The tabs were bonded with 3M AF167 Structural Weld
which cures at 250F for 90 minutes. A tabbed specimen is shown in Figure 5.2. No
tab bevel angle was used and was the only noted exception to the ASTM standard.
However, according to Adams et al. [28] there is widening acceptance that a non-
beveled edge, or bevel angle of 90F, is a usable geometry and has even been adopted
67
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68
Figure 5.1. MTS Test Frame with Specimen and Extensometer
Figure 5.2. Composite Specimen with E-Glass Tabs
by the International Organization for Standardization (ISO) as their standard test
configuration.
5.2 Material Characterization
The material system chosen for this study is widely used in research and in-
dustry. As a result, the material properties of the system are fairly easy to find in
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69
the public domain. One resource used for this research was the website of The Uni-
versity of California, San Diegos Composites and Aerospace Structures Laboratory
[14]. This site contained a complete set of material properties for IM-6/3501-6 with a
fiber volume of 63.5%. However, it was desired to verify these properties before using
them in any further research since different processing techniques were used.
Two unidirectional 8-ply laminates and one [45]8S laminate were made to
verify some of the material properties. The two unidirectional panels were cut so
that one had transverse outer layer fibers and the other longitudinal outer layer
fibers. The other panel was cut so that the top layer was oriented at +45. A three
element 45
rosette strain gage was applied to one specimen of each lot for tensile
testing.
The results of this verification study are shown in Table 5.1. The values for
E11,E22, 12, and 23 were in close agreement with those obtained from the UCSD
website. A value for G12 could not be obtained because the equipment needed to
make the measurement was not available.
Table 5.1. Summary of Material CharacterizationProperty Measured UCSD [14]
E11 22.15 106 psi 23.30 106 psi
E22 1.187 106 psi 1.395 106 psi
12 0.285 0.296523 0.285 0.2965G12 not measured 0.9161 10
6 psi
5.3 Tensile Testing
Samples from each panel were prepared for testing according to ASTM D3039
[29]. Each specimen was measured for thickness and width variation before testing.
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72
by the porosity introduction would be evident in this test. In each test, the specimen
was aligned in the machine the same way so that gage 1 was on the back side and gage
2 was on the front. This was done so that any bending effects would not be missed
since because of normal differences between the two gage readings. A final strain
load of 12000 was to be applied after the previous scan. However, the first two
specimens tested broke before reaching this strain level. It was decided to proceed
with each specimen to failure and record the final strain reading.
The results of this test are shown in Table 5.3. No apparent pattern could be
seen in the data. It would perhaps have been more beneficial to perform this test
in specimens with a notch or central hole so that the porosity was in the vicinity of
a stress concentration. This scenario might give a better indication as to the flaw
sensitivity of the material.
Table 5.4 shows the C-scan results for these specimens after the first three
strain levels. In most cases, the initial mean signal strength was lower than that
measured after the first strain level. It is unknown why this occurred. However, after
the first level, the trend was as expected, decreasing signal strength with increasing
damage. Figures 5.3 through 5.12 show the C-scans before any testing and after
7500 for each specimen. In each sample, the strain gauge and lead wires were
visible in the second picture.
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73
Table5.3
.
DamageThresholdTestResults
Sample
2500
5000
750
0
10000
Gag
e1
Gage2
Load
Gage1
Gage2
Load
Gage1
G
age2
Load
Gage1
Gage
2
Load
35-7
2488
2612
779
4860
5036
1525
7366
7538
2304
10011
10160
3041
51-10
2545
2485
832
5002
5070
1672
7575
7650
2516
10068
10160
3264
54-9
2515
2472
882
5050
5044
1796
7540
7077
2688
10119
10060
3541*
57-9
2478
2510
895
5100
5072
1840
7545
7563
2747
10068
10021
3587*
61-9
2500
2450
855
5040
4945
1724
7555
7460
2583
10166
10077
3459
Note:Allloadsmeasuredinpounds.
Note:*Deno
tesaudibledamageduringloa
ding.
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74
Table 5.4. C-Scan Results from Damage Threshold Test
Sample
Mean Signal Strength at Strain Increment
0 2500 5000 7500 35-7 28.34 34.41 34.82 33.60
51-10 18.22 15.79 17.81 14.9854-9 29.55 36.44 39.27 33.6057-9 37.25 36.89 38.46 32.3961-9 6.48 6.48 5.67 4.86
Figure 5.3. Preliminary C-scan of Strip 35-07
Figure 5.4. Final C-scan of Strip 35-07
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Figure 5.5. Preliminary C-scan of Strip 51-10
Figure 5.6. Final C-scan of Strip 51-10
Figure 5.7. Preliminary C-scan of Strip 54-09
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Figure 5.11. Preliminary C-scan of Strip 61-09
Figure 5.12. Final C-scan of Strip 61-09
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Tab
le5.5.
FatigueTestResults.
Sample#
%
FailureLoad
StressLe
vel(psi)
#
Cycles
Notes
36-6
0.9
957
27
404
nodelam.
36-7
0.8
850
90
2308
signific
antdelam.,somelong.splitting
36-8
0.7
744
54
58906
signific
antdelam.,somelong.splitting
36-9
0.6
638
18
1000000*
extrem
edelam.,somelong.splitting
36-10
0.8
850
90
1288
somed
elam.
37-1
0.75
850
90
1438
littled
elam.
37-2
0.75
850
90
97
badtestrun,
failedbybuckling
37-3
0.7
744
54
7224
somed
elam.
37-4
0.7
744
54
8344
somed
elam.
37-5
0.7
744
54
7631
somed
elam.
37-6
0.75
638
18
2559
somed
elam.
37-7
0.65
638
18
16795
somed
elam.
37-8
0.65
638
18
20527
signific
antdelam.
37-9
0.6
531
82
78773
extrem
edelam.,significantlong.splitting
37-10
0.6
531
82
65370
extrem
edelam.,significantlong.splitting
Note:Av
erageFailureStress106000psi(From
specimens35-1,2,3,4,
5)
Note:Lo
adRatioR=
0.1
Note:*Denotestestrunout.
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89
[1] Strong, A. B. Fundamentals of Composites Manufacturing: Materials,Methods,and Applications. Dearborn: Society of Manufacturing Engineers, 1989.
[2] Jones, R. M. Mechanics of Composite Materials, Second Edition. Philadelphia:
Taylor & Francis, Inc., 1999.
[3] Walker, J., Telephone conversation and emails, October 15, 2003 to November5, 2003.
[4] Gay, D., Hoa, S. V., and Tsai, S. W., Composite Materials, Design and Appli-cations. Boca Raton: CRC Press, 2003.
[5] Garrett, R. A., Effects of Manufacturing Defects and Service-Induced Damageon the Strength of Aircraft composite Structures, Composite Materials: Testing
and Design, ASTM STP 893, pp. 5-33, 1986.
[6] Grelsson, B., Correlations Between Porosity Content, Strength and UltrasonicAttenuation in Carbon Fibre Laminates, Materials & Design, Vol. 13, no. 5,pp .275-278, 1992.
[7] Kan, Han-Pin, Bhatia, Narain M., and Mahler, Mary A., Effect of Porosityon Flange-Web Corner Strength, Composite Materials: Fatigue and Fracture,ASTM STP 1110, pp. 126-139, 1991.
[8] Pipes, R. B., Interlaminar Shear Fatigue Characteristics of Fiber-ReinforcedComposite Materials, Composite Materials: Testing and Design, ASTM STP546, pp. 419-432, 1974.
[9] Hancox, N. L., The Effects of Flaws and Voids on the Shear Properties of CFRP,Journal of Materials Science, Vol. 12, pp. 884-892, 1977.
[10] Judd, N. C. W., Wright, W. W., Voids and Their Effects on the MechanicalProperties of Composites - An Appraisal, SAMPE Journal, January/February,pp. 10-14, 1978.
[11] Yoshida, H., Ogasa, T., and Hayashi, R. Statistical Approach to the RelationshipBetween ILSS and Void Content of CFRP, Composites Science and TechnologyVol. 25, pp. 3-18, 1986.
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0 0.002 0.004 0.006 0.008 0.01 0.012 0.0140
2
4
6
8
10
12x 10
4
Strain (in/in)
Stress(psi)
Stress vs Strain for Specimen 513
Figure A.7. Tensile Plot for Specimen 51-3
0 0.002 0.004 0.006 0.008 0.01 0.012 0.0140
2
4
6
8
10
12x 10
4
Strain (in/in)
Stress(psi)
Stress vs Strain for Specimen 514
Figure A.8. Tensile Plot for Specimen 51-4
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