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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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vii

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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viii

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




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xii

Figure Page


4.6 Mean Signal Strength Ultrasonic C-Scan of Panel 35 . . . . . . . . . . 45






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







4.22 Panel 35 Thickness Variation Measurement of Middle Two Laminae . . 59




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Figure Page



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



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







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Figure Page




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















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Figure Page








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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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8

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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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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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 4.17. Photomicrograph of Panel 54 Close-up of Laminae Interfaces


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Figure 4.19. Photomicrograph of Panel 57 Close-up of Laminae Interfaces


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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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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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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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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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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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80

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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