Composite Strength Testing

LAB REPORT 1:COMPOSITE STRENGTH TESTING

Date of Submission: 09 October 2006

by

________________________________Harsh Menon

[email protected]

Submitted to Dr. David Lanning, Jr. Department of Aerospace Engineering

College of EngineeringEmbry-Riddle Aeronautical University

In Partial FulfillmentOf the Requirements

OfES 321

Engineering Materials Science LabFall 2006

Embry-Riddle Aeronautical UniversityPrescott, Arizona

Composite Testing 1 4/8/2023

ABSTRACT

A carbon fiber composite consisting of unidirectional carbon tape and epoxy and a sandwich composite consisting of a polyether polyurethane core, E-glass fibers and epoxy was fabricated. The carbon fiber composite was subjected to a tensile test while the sandwich composite was subjected to a three-point bend test. Part of the fabricated carbon fiber composite was fractured and placed under a scanning electron microscope.

Data from the tests was used to determine material properties such as Young’s modulus, core shear stress and facing bending stress. The experimental values were then compared with theoretical values. The microscopy was used to obtain pictures of the composite which in turn were used to obtain volume fractions of the fibers and the matrix.

Differences between the theoretical and experimental values were attributed to either improper fabrication techniques or improper data analysis. Future recommendations involve spending more time on composite fabrication as well as experimentation with more than two specimen and with different configurations for the composites.


TABLE OF CONTENTS

ABSTRACT........................................................................................................................1

LIST OF FIGURES.............................................................................................................2

LIST OF TABLES...............................................................................................................4

LIST OF SYMBOLS...........................................................................................................5

LIST OF EQUATIONS.......................................................................................................6

1.0 INTRODUCTION...................................................................................................7

2.0 THEORY.................................................................................................................8

2.1 COMPOSITE FABRICATION..........................................................................................9

2.2 STRESS-STRAIN CURVE FOR FIBER-COMPOSITES....................................................10

2.3 THREE-POINT BEND TEST........................................................................................12

2.4 MICROSCOPY............................................................................................................13

3.0 APPARATUS AND PROCEDURES...................................................................15

3.1 APPARATUS....................................................................................................15

3.2 PROCEDURES.................................................................................................17

5.0 CONCLUSIONS AND RECOMMENDATIONS................................................27

6.0 REFERENCES......................................................................................................28

7.0 ATTRIBUTIONS..................................................................................................29

8.0 APPENDIX A: SAMPLE CALCULATIONS......................................................30

9.0 APPENDIX B: RAW DATA................................................................................31


LIST OF FIGURES

FIGURE 2.1.1: FLOWCHART FOR COMPOSITE FABRICATION...............................9

FIGURE 2.2.1: (A) STRESS-STRAIN CURVES FOR COMPOSITE 10

FIGURE 2.2.1: (B) STRESS-STRAIN CURVE FOR A METAL...................................10

FIGURE 2.3.1: THREE-POINT BEND TEST.................................................................12

FIGURE 2.4.1: FRACTURED CARBON COMPOSITE FROM A SCANNING

ELECTRON MICROSCOPE............................................................................................14

FIGURE 3.1: TENSILE TESTER.....................................................................................16

FIGURE 3.2: EXTENSOMETER.....................................................................................16

FIGURE 3.1: CARBON INSIDE EPOXY........................................................................19

FIGURE 4.1: STRESS-STRAIN CURVE FOR UNIDIRECTIONAL CARBON...........21

FIGURE 4.2: LOAD VERSUS DISPLACEMENT FROM THREE-POINT BEND

TEST..................................................................................................................................23

FIGURE 4.3: (A) CLOSER VIEW OF THE CRACK 24

FIGURE 4.3: (B) VIEW OF THE ENTIRE SPECIMEN.................................................24

FIGURE 4.4: CARBON COMPOSITE UNDER SCANNING ELECTRON

MICROSCOPE..................................................................................................................25

FIGURE 4.5: BLACK AND WHITE IMAGE..................................................................26


LIST OF TABLES


LIST OF SYMBOLS

A Area in2

E Modulus of Elasticity/Young’s lb/in2

Modulus

I Moment of Inertia in4

L Distance in

P Force lb

V Volume Fraction

b Sandwich Beam Width in

c Sandwich Core Height in

d Sandwich Beam Height in

t Sandwich Ply Height in

σ Normal Stress lb/in2

τ Shear Stress lb/in2

c,|| Composite (Parallel to fibers)

c Composite

f Fibers

m Matrix

max Maximum


LIST OF EQUATIONS

Equation 2.1

Equation 2.2

Equation 2.3

Equation 2.4

Equation 2.5

Equation 2.6

Equation 2.7


1.0 INTRODUCTION

In the simplest form, a composite is a combination of two or more phases. One of the phases is a generally a fiber phase and the other a matrix phase. Although composites appear to be a recent development, they have been around for centuries. In fact, Mongolian arcs were made of corn, wood and cow tendon composites (Gay, Hoa & Tsai, 2003). However, the recent development of composites pertains to the development of high performance composites which have the ability of being engineered for specific use.

Modern day composites have a certain advantages over metallic materials. Some of the advantages are higher strength and fatigue and corrosion resistance. In the field of aerospace, composites are playing an even major role with the latest commercial airplanes from Boeing using a lot of composites. The study of composites is still an actively researched field and lot can be learned from the fabrication of and from experimentation on composites.

The purpose of this lab was to investigate the behavior of two types of composites – a carbon fiber epoxy composite and a sandwich beam consisting of E-glass fibers, a polyurethane core and epoxy. The carbon composite was tensile tested while the sandwich beam was subjected to a three-point bend test.

This report documents the fabrication process and testing process of the different composites and lists important material properties such as Young’s modulus, shear stress and bending stress that were calculated from experiment. The experimental values are then compared to theoretical values. The results from viewing the fractured carbon composite under a scanning electron microscope are also documented in this report. Further recommendations are also discussed.


2.0 THEORY

A composite is “any multiphase material that exhibits a significant proportion of the properties of both constituent phases such that a better combination of properties is realized” (Callister, 2004). Some examples of composites are wood and bone (Gay et al., 2003). However, in the context of engineering materials science, composites take on a more specific definition. Composites refer to multiphase materials that are artificially created and whose constituent phases are chemically different and separated by a distinct interface (Callister, 2004).

Composites are used in a wide range of applications from electronics to space transports (Gay et al., 2003). The principal advantages of using composite materials are:

1. High strength and stiffness (Iatridis, 2006)2. Design Flexibility (i.e., fibers can be oriented to take maximum stress)3. Fatigue and Corrosion Resistance (Gay et al., 2003)

Askeland and Phulé (2006) classify composites into three major categories:

1. Particulate Composites: Particulate composites contain “large amounts of coarse particles that produce unusual combinations of properties rather than improve strength” (Askeland & Phulé, 2006).

2. Fiber Composites: Fiber composites consist of two phases – the fibers which constitute a reinforcing phase and the matrix phase. Fiber composites improve a lot of important parameters such as strength, Young’s modulus etc. (Askeland & Phulé, 2006). The fibers can be unidirectional, bidirectional or tridirectional (Gay et al., 2003).

3. Laminar-based Composites: Laminar composites are used primarily to “improve corrosion resistance while retaining low cost, high strength or light weight” (Askeland & Phulé, 2006).

The two composites used for this lab fall under the fiber composites and laminar based composites categories. The first composite – unidirectional carbon composite, is a fiber-reinforced composite. The second composite is a sandwich beam consisting of layers of E-glass and a polyether polyurethane core. Sandwich beams consist of layers of fibers attached to a lighter core. The main advantages of sandwich beams are their very light weight, high flexural rigidity and good thermal insulation characteristics (Gay et al., 2003).


2.1 Composite Fabrication

Figure 2.1.1: Flowchart for Composite Fabrication on the next page describes the steps involved in the formation of a composite.

Figure 2.1.1: Flowchart for Composite Fabrication.

(Gay et al., 2003)

In the figure above, mixing is followed by compaction which is can be thought of as the compression of the matrix and the fibers together. This is followed by polymerization, which is a process that converts a monomer into a polymer. A monomer is defined as the smallest repeating unit of a polymer (Australian Deparment of the Environment and Heritage, 2005). However, polymerization is required only in the case of polymer matrix compounds such as epoxies. After polymerization, the composite is taken from the mold. Finishing touches are then done to the composite.

This entire process is defined as the molding process. There are several ways to accomplish molding, some of which are listed below:

1. Contact Molding Contact molding involves placing the mixture of the phases on a mold and using rollers to compact the phases together (Gay et al., 2003).

2. Vacuum Molding


Vaccum molding involves placing the composites inside an enclosed region sealed with plastic and subjecting the area to a vacuum to compact the phases of the composite (Gay et al., 2003).

A manual method of composite fabrication involves hand lay up. This was done in the experiment and is described in 3.0 Procedures and Apparatus. The method used to fabricate the composites was a mixture of hand lay up and vacuum molding.

2.2 Stress-Strain Curve for Fiber-Composites

The stress strain curves of fibrous polymeric composites have been experimentally shown to depend on the fibers and matrix of the composite, the interaction between the two, the direction of the load relative to the principal axes of the composite and environmental conditions (Tarnopol’skii, 1985). They also depend on the orientation of the fibers. The figure below shows an example of a stress-strain curve for a fiber composite, where the force was applied in the longitudinal direction (in the direction of the fibers):

(a) (b)

Figure 2.2.1: (a) Stress-Strain curves for Composite, (b) Stress-Strain curve for a Metal.

(Askeland, 2006, http://www.key-to-steel.com)

The stress-strain curve for a composite looks different from the stress-strain curve for an ordinary metal. The major difference being that the composites do not deform plastically or that composites deform only elastically. Composites generally also have a higher ultimate strength than metals.

Figure 2.2.1 (a) divides the stress-strain curve for a fiber composite into three different regions. The first region corresponds to a linear region where the stresses are low. The second region corresponds to an area of higher stresses leading finally to the third region of fracture. If we assume that the fibers in the matrix are continuous and unidirectional, then the longitudinal modulus of elasticity of the composite (along the direction of the


fibers) can be predicted using the rule of mixtures (Askeland & Phulé, 2006), shown in Equation 2.1 below:

Equation 2.1

where Ec,|| is the longitudinal modulus of elasticity, Em is the modulus of elasticity of the matrix, Vm is the volume fraction of the matrix, Vf is the volume fraction of the fibers and Ef is the modulus of elasticity of the fibers.

The volume fraction of a certain part of the composite is defined as the volume of that part divided by the total volume of the composite. Therefore, the volume fraction of the matrix is the volume of the matrix divided by the total volume of composite. However, if we assume that the composite, matrix and fiber phase lengths are all equal (Callister, 2004), then Equation 2.2 shown below holds.

Equation 2.2

where Em is the modulus of elasticity of the matrix, Am is the area of the matrix, Ac is the area of the composite, Ef is the modulus of elasticity of the fibers and Af is the area of the fibers.

This modulus of elasticity can also be determined experimentally from the first region of Figure 2.2.1 (a) and is equivalent to the modulus defined in Equation 2.1. Another important parameter that can be obtained from the stress-strain curve of composites is the ultimate tensile strength. The ultimate tensile strength is the highest point on the stress-strain curve and corresponds to the point of failure. Analytically, Askeland and Phulé (2006) describe the use of the rule of mixtures to approximate the tensile strength by using Equation 2.3 below:

Equation 2.3

where σc is the tensile strength of the composite, σf is the tensile strength of the fibers, Vf is the volume fraction of the fibers, σm is the tensile strength of the matrix and Vm is the volume fraction of the matrix.


2.3 Three-point Bend Test

The three-point bend test is generally performed on brittle materials and composites. These materials often break when placed in the grips of the tensile tester and therefore, the three-point bend test is used to determine their strength. The setup for the three-point bend test can be seen in Figure 2.3.1: Three-point bend test below:


Figure 2.3.1: Three-point bend test.

The figure above shows a specimen placed between two pins a distance L apart, subjected to a concentrated force P at the center of L. Using Appendix G of Gere’s Mechanics of Materials (2004), we can show that the maximum deflection at the center of the beam is given by Equation 2.4:

Equation 2.4

where P is the force acting at the center of the beam, L is the distance between the supports, E is the modulus of elasticity of the beam and I is the moment of inertia of the beam.

In the case of the three-point bend test, the maximum deflection is determined experimentally. Therefore, Equation 2.4 can be re-arranged to solve for the modulus of elasticity or the flexural modulus. However, Equation 2.4 was derived assuming a linear stress-strain relationship and therefore, Equation 2.4 holds only in the linear part of the composite material’s stress-strain curve.

Some other additional parameters from the ASTM C 393-00 that characterize the sandwich beam are the core shear stress and facing bending stress. The simplifying assumptions used in the calculations of these parameters are

1. The normal bending stresses are assumed to occur in the facings only and are assumed to be uniform in the facings (Gay et al., 2003)

2. The shear stresses are assumed to occur only in the core and are uniform in the core (Gay et al., 2003)

The core shear stress is given by Equation 2.5 on the next page.

Equation 2.5


where τ is the core shear stress, P is the applied force, d is the sandwich thickness, c is the core thickness and b is the sandwich width.

The facing bending stress is given by Equation 2.6 on the next page:


Equation 2.6

where σ is the facing bending stress, P is the applied force, L is the distance between the supports in the three-point bend test, t is the thickness of each ply, d is the sandwich thickness, c is the core thickness and b is the sandwich width.

2.4 Microscopy

Looking at a fractured composite sample under a scanning electron microscope can give a lot of insight into the nature of composites. Figure 2.4.1: Fractured Carbon Composite from a Scanning Electron Microscope below shows a picture of the fractured unidirectional carbon composite as seen under a scanning electron microscope.

Figure 2.4.1: Fractured Carbon Composite from a Scanning Electron Microscope.

In the figure above, the lighter circular regions represent the fibers and the darker regions represent the matrix. A further analysis of this picture could yield the volume fractions of this matrix, using Equation 2.2 and the assumptions that come along with it. By determining the number of the circular regions and the average diameter of the lighter regions, one could estimate the total area of the fibers. The total area of the fibers divided by the total area of the image would then yield the volume fraction for the fibers. The volume fraction for the matrix could be determined using Equation 2.7:


Equation 2.7


3.0 APPARATUS AND PROCEDURES

3.1 APPARATUS

The following equipment was used in the experiment:

Two (2) layers of unidirectional carbon tape from Aircraft Spruce & Specialty Co.© with a width of 3 in obtained from a 12 ft roll,

Four (4) layers of medium weight standard E-glass industrial cloth from Aircraft Spruce & Specialty Co.© with a width of 50 in and a thickness of 0.009 in,

EZ-LAM 60 Epoxy from Aerospace Composite Products©, One (1) polyether polyurethane LAST-A-FOAM® from Aircraft Spruce &

Specialty Co.© with a density of 4.5 lb/ft3 and a thickness of 0.5 in,

Eight (8) rectangular strips of standard E-glass tape from Aircraft Spruce & Specialty Co.© with a width of 2 in and a thickness of 0.012 in, and

One (1) brush, One (1) tensile tester machine, as shown in Figure 3.1: Tensile Tester below:

Figure 3.1: Tensile Tester.

(Lanning, 2003)


The tensile tester is used to determine the resistance of a material to a static or slowly applied force (Askeland & Phulé, 2006).

One (1) extensometer, which can be seen in Figure 3.2: Extensometer below:

Figure 3.2: Extensometer.

(Lanning, 2003)

An extensometer is defined as an instrument that is used to measure the change in length of a tensile specimen (Askeland & Phulé, 2006).

One (1) three-point bend setup for the tensile testing machine, which can be seen in Figure 3.3: Three-point Bend Setup below:

Figure 3.3: Three-point Bend Setup.

(Lanning, 2003).


One (1) 320 Grinder Polisher Machine, as can be seen in Figure 3.4: Grinder Polisher Machine below:

Figure 3.4: Grinder Polisher Machine.

(Lanning, 2003)

One (1) 400 Grinder Polisher Machine,

One (1) 600 Grinder Polisher Machine,

One (1) 1200 Grinder Polisher Machine, and

One (1) Jeol Scanning Electron Microscope which can be seen in Figure 3.5: Scanning Electron Microscope on the next page.


Figure 3.5: Scanning Electron Microscope.

(Lanning, 2003).

3.2 PROCEDURES

The following procedures were used for the composite layout part of the experiment, which was completed on September 19, 2006 at Embry-Riddle Aeronautical University:

1.0 The class was divided into three (3) teams with two (2) students per team for the experiment.

2.0 The team obtained two (2) layers of unidirectional carbon tape (from Aircraft Spruce & Specialty Co.©) for the tensile test and four (4) layers of standard E-glass industrial cloth (from Aircraft Spruce & Specialty Co.©) for the three-point bend test.

3.0 The team obtained a polyether polyurethane foam core (from Aircraft Spruce & Specialty Co.©).

4.0 The team was given eight (8) rectangular strips of standard E-glass tape (from Aircraft Spruce & Specialty Co.©) to be used to protect the fibers of the tensile test coupon from the grips of the tensile test machine.

5.0 The team was given a cup half-filled with EZ-LAM 60 Epoxy from Aerospace Composite Products© from a ratio pump which mixed two parts of epoxy and one part of the hardener.

6.0 One (1) rectangular strip of the E-glass tape (to protect the fibers) was placed on the galvanized metal surface and epoxy was applied to the top side of the layer using a brush.

7.0 Another rectangular strip was placed on top of the first strip and epoxy was applied to the top side of the layer.


8.0 Steps 6.0 - 7.0 were repeated a horizontal distance approximately equal to the length of the unidirectional carbon tape layer away from the first two strips.

9.0 One (1) layer of unidirectional carbon tape was then placed on the four rectangular E-glass tape strips so that the rectangular strips were approximately at the ends of the unidirectional carbon tape.

10.0 Epoxy was applied to the top surface of the unidirectional carbon tape.

11.0 Another layer of unidirectional carbon tape was placed on top of the first unidirectional carbon tape and epoxy was applied to the top surface of this tape.

12.0 Steps 6.0 – 9.0 were then repeated but the strips were placed on top of the second unidirectional carbon tape layer as opposed to the galvanized metal surface.

13.0 Breather ply was then placed on top of the entire composite.

14.0 One (1) layer of medium weight standard E-glass industrial cloth was then placed on an unused part of the galvanized metal surface.

15.0 Epoxy was applied to the top surface of the E-glass layer.

16.0 Steps 14.0 – 15.0 were repeated for another layer of E-glass.

17.0 The polyether polyurethane foam was then placed on top of the second E-glass layer.

18.0 Steps 14.0 – 16.0 were repeated with the E-glass layers being placed on top of the foam as opposed to the galvanized metal surface.

19.0 Breather ply was then placed on top of the entire composite.

20.0 The composites from all the teams were covered with vacuum bagging ply.

21.0 The area with the composites inside the vacuum bagging ply was pressurized to 10 in of mercury for 24 hours.

The unidirectional carbon composite was then subjected to a tensile test, while the E-glass composite was subjected to a three-point bend test. The following procedures were used for composite testing which was completed on September 26, 2006 at Embry-Riddle Aeronautical University:

22.0 The carbon composite was placed in the tensile tester by ensuring that the grips only made contact with the part of the composite that had the extra rectangular strips.


23.0 An extensometer was placed on the unidirectional carbon composite.

24.0 The entire setup was covered by a wooden screen to prevent the fiber pieces from being scattered in all directions at failure.

25.0 The computer program was started and the tensile tester was run until the computer program beeped to signal the user to remove the extensometer.

26.0 The extensometer was removed and the test continued until the composite failed.

27.0 The data from the program was extracted into a Microsoft Excel© spreadsheet.

28.0 The tensile tester and the area around it were cleaned and the shrapnel from the composite were removed.

29.0 The tensile tester was then modified for the three-point bend test by the lab instructor, Dr. David Lanning Jr.

30.0 The E-glass composite was placed on the three-point bend test apparatus.

31.0 The distance between the bottom pins of the three-point bend apparatus was adjusted so that it corresponded to 6.5 in

32.0 Another computer program was started and the three-point bend test was run until the composite

Prior to the beginning of the experiment on September 26, 2006, about a quarter inch of the unidirectional carbon composite was cut off in the machine shop. This quarter inch of composite material was then bent till the material fractured, yet still remained one piece. This material was then placed in a cylindrical holder with epoxy and heat treated. The composite material was then used for the third part of the experiment involving microscopy. The following procedures were used for the microscopy part of the experiment, which was completed on October 3, 2006 at Embry-Riddle Aeronautical University:

1.0 The hardened epoxy was taken out of the cylindrical holder. The epoxy took on a cylindrical shape as can be seen in Figure 3.1: Carbon inside Epoxy below:


Figure 3.1: Carbon inside Epoxy.

2.0 The epoxy with the unidirectional carbon inside it was grinded using a 320 grinder until the edge of the composite was at the bottom surface of the epoxy.

3.0 The epoxy was then placed on the 400 grinder for two (2) minutes for further polishing.

4.0 The epoxy was then placed on the 600 grinder for two (2) minutes.

5.0 The epoxy was then placed on the 1200 grinder for two (2) minutes.

6.0 The epoxy sample with the carbon composite was then placed under the scanning electron microscope.

7.0 The fractured surface of the composite was analyzed and snapshots of the surface were taken at a scale of 10 microns.


4.0 RESULTS AND DISCUSSION

The data obtained from the tensile test experiment on the unidirectional carbon composite was converted into stress and strain and is shown in Figure 4.1: Stress-Strain curve for Unidirectional Carbon below:

Stress vs Strain(Tensile Test)

y = 7E+06x

-10000

0

10000

20000

30000

40000

50000

60000

70000

-0.001 0 0.001 0.002 0.003 0.004 0.005 0.006

Strain (inches/inches)

Str

ess

(psi

) Stress vs Strain

Linear Part

Corrected Linear Part

Linear (Corrected Linear Part)

Young's Modulus (E) = 7e6 psi

Ultimate Tensile Strength=60,000 psi

Figure 4.1: Stress-Strain curve for Unidirectional Carbon.

Figure 4.1: Stress-Strain curve for Unidirectional Carbon shows the data in blue. A linear curve was fitted to the data which can be seen as the pink line. However, the pink line did not have a y-intercept of 0 and was therefore corrected. The corrected linear part is the yellow line.

The figure above has the typical stress-strain shape of composites as discussed in 2.0 Theory. The reason why the graph shoots up at 0.005 in is because the extensometer (used to determine the strain) was removed at that point of time.

However, there appears to be a discontinuity at a strain of about 0.004 in extending to 0.005 in Some possible reasons for the discontinuity are:

1. Errors in Testing: Errors induced in the data due to the fact that the tensile tester slipped or lost its grip or due to improper placement of the composite specimen in the tensile tester.


2. Defects in the Composite: Errors induced in the data due to the presence of nonlinearities and deformities inherent in the composite.An analysis of Figure 4.1: Stress-Strain curve for Unidirectional Carbon yielded important results. The slope of the yellow line was determined to be 7.0 x106 lb/in2. The highest point on the blue line was taken as the ultimate tensile strength of the carbon composite and was determined to be 60,000 lb/in2. The results from this test are listed in Table 4.1: Experimental Data for Unidirectional Carbon below:

Table 4.1: Experimental Data for Unidirectional Carbon

Young’s Modulus 7.0 x106 lb/in2

Ultimate Tensile Strength 60,000 lb/in2

The experimental value of Young’s Modulus listed in Table 4.1: Experimental Data for Unidirectional Carbon was then compared to a theoretical value of Young’s modulus for a unidirectional carbon composite. Callister’s (2004) modulus of elasticity of high modulus carbon fibers with an epoxy matrix with a fiber volume fraction of 0.6 was chosen as the theoretical value (p. 743). The theoretical value of E was therefore determined to be 32 x 106 lb/in2. The experimental value was determined to be 7x106

lb/in2. These values are tabulated in Table 4.2: Theoretical and Experimental values for Carbon below:

Table 4.2: Theoretical and Experimental values for Carbon

Experimental Young’s Modulus 7.0 x106 lb/in2

Theoretical Young’s Modulus 32 x 106 lb/in2

Percent Difference 78%

This difference between the theoretical and experimental values listed in Table 4.2: Theoretical and Experimental values for Carbon can be attributed to

a) Improper application of epoxy resulting in a weaker composite (primary), or b) Propagation of nonlinearities in the data set into the equation of the corrected

trend line (secondary), orc) Both.

The results from the three-point bend test on the composite sandwich beam are plotted in Figure 4.2: Load versus Displacement from Three-Point Bend Test on the next page.


Load vs Displacement

y = 400.37x

0

10

20

30

40

50

60

70

-0.1 0 0.1 0.2 0.3 0.4 0.5 0.6

Displacement (in)

Lo

ad (

lb)

Load vs Displacement

Linear Part

Linear (Linear Part)

Figure 4.2: Load versus Displacement from Three-Point Bend Test.

The figure above shows a plot of the load versus the displacement for the sandwich beam during the three-point bend test. The data points are in blue and the pink line represents a fit to the linear portion of the plot.

During the testing process, a lot of pertinent observations were made that add depth to the figure shown above. Initially, in the linear portion of the graph, the composite beam was seen to deform slowly. This process continued over a period of time and finally leveled off when the displacement was 0.30 in But, at a displacement of around 0.38 in on the graph, cracking sounds were heard. At this point, the bottom layer of the composite started separating from the foam core. This corresponded to the drop in the load from 58 lb to 42 lb. This resulted in a crack on the beam which can be seen in Figure 4.3: (a) Closer view of the crack (b) View of the entire specimen on the next page.


(a) (b)

Figure 4.3: (a) Closer view of the crack (b) View of the entire specimen.

Figure 4.3 (a): Closer view of the crack shows the crack in greater detail. The fibers on the E-glass ply (layer) can be seen on a closer inspection. The crack represents a region where the fibers have been split. Figure 4.3 (b): View of the entire specimen shows the entire beam. The separation of the plies from the core was very minor and not as drastic as the other tests performed by the other groups. At the end of the test, the composite was still one piece. However, it had lost a lot of strength on one side.

Using the maximum value of displacement and load in the linear part of the curve in Figure 4.2: Load versus Displacement from Three-Point Bend Test in Equation 2.4 and re-arranging that equation yields a value of 1.9x105 lb/in2 for the modulus of elasticity of the composite. The moment of inertia for the composite was calculated using only the dimensions of the two plies and the distance from the center of the plies to neutral axis. The other ASTM C 393.00 parameters were also calculated using Equation 2.5 and Equation 2.6. The core shear stress was determined to be 44 lb/in2

and the facing bending stress was determined to be 2700 lb/ in2. These values are listed in Table 4.3: Experimental Data for the Sandwich Beam below:

Table 4.3: Experimental Data for the Sandwich Beam

Young’s Modulus 1.9x105 lb/in2

Core Shear Stress 44 lb/in2

Facing Bending Stress 2700 lb/ in2


The data listed in the table above was then compared to theoretical values. Data obtained from CES EDUPACK (2006) yielded a modulus of elasticity of 3.8 x 106 lb/in2 for Epoxy/E-Glass Woven Fabric Composite with a Biaxial Lamina. Data about the polyurethane core was obtained CES EDUPACK (2006). The tensile strength of the polyurethane core was used as the core shear stress and was found to range from 36 lb/in2

to 44 lb/in2.The tensile strength of the Epoxy/E-Glass Woven Fabric Composite with a Biaxial Lamina was taken to be the facing bending stress and was determined to range from 54,000 lb/ in2 to 75,000 lb/ in2 (CES EDUPACK, 2006). Table 4.4: Theoretical versus Experimental Data for the Sandwich Beam lists these values below:

Table 4.4: Theoretical versus Experimental Data for the Sandwich Beam

Physical Quantity Theoretical Value Experimental Value Percent Difference

Modulus of Elasticity

3.8 x 106 lb/in2 1.9x105 lb/in2 95%

Core Shear Stress 36 lb/in2 - 44 lb/in2 44 lb/in2 Within RangeFacing Bending

Stress54,000 lb/in2 – 75,000 lb/in2

2700 lb/ in2 95% (using 54,000 lb/in2 as

theoretical)


The discrepancies between the theoretical and experimental values of the modulus of elasticity are most likely due to improper fabrication of the sandwich composite leading to a weaker composite material. The core shear stress falls within the range of the theoretical values. However, the facing bending stress is significantly lower than the theoretical range. Though improper fabrication could be the primary cause, another point of consideration is the fact that the theoretical values being compared against are the tensile strengths of the core (for core shear stress) and the composite layers (for facing bending stress). This could be another source of error.

The next part of the experiment required observing a fracture carbon composite sample under a scanning electron microscope. Figure 4.4: Carbon Composite under Scanning Electron Microscope below shows the fractured carbon composite:

Figure 4.4: Carbon Composite under Scanning Electron Microscope.

The figure above shows the fibers and the matrix in the carbon composite. The lighter white regions represent the fibers, while the darker regions represent the matrix. The image above shows that the majority of the area is taken up by the fibers. Also, the fibers seem to be closely spaced on top rather than the bottom. There appears to be a region at the bottom where the matrix seems to dominate.

Figure 4.4: Carbon Composite under Scanning Electron Microscope was converted into a black and white image using MATLAB to make be able to clearly distinguish the boundaries between the light and dark regions. Figure 4.5: Black and White Image on the next page shows the converted image.


Figure 4.5: Black and White Image.

The white regions in the figure above were then fitted with circles. A total of 70 circles were counted. The diameter of each of these circles was calculated and an average diameter of 0.59 in The total area of these circles was then determined to be 19 in Then the dimensions of the image were measured (excluding the legend) and the area of the image was determined to be 26.3 in The ratio of these two numbers yielded the volume fraction of the fibers, which was determined to be 0.73. The volume fraction of the matrix was then determined to be 0.27. The calculations are listed in Appendix A: Sample Calculations.


5.0 CONCLUSIONS AND RECOMMENDATIONS

This lab was aimed at determining the composite strengths for different types of composites. The first step of the lab involved fabricating composites. Unidirectional carbon was mixed with epoxy to create a carbon fiber, while E-glass fibers were placed on a polyether polyurethane fiber and held together with epoxy to create a composite sandwich beam.

Important parameters such as Young’s modulus were calculated for both composites and were compared to theoretical book values. All of the experimental values were lower than the theoretical values. The primary reason for this was improper fabrication of the composites. The lab was successful in the sense that the behavior of composites under a tensile test and a three-point bend test were documented and ASTM C parameters such as core shear stress and facing bending stress were calculated. The pictures from the scanning electron microscope helped determine the volume fractions of the carbon composite.

The major limitation of the lab was the time spent fabricating the composite. Not too much time was spent ensuring that the epoxy was spread evenly. The differences between the theoretical and experimental values could be have been reduced had more time been spent on composite lay-up and fabrication. Another potential recommendation for the future could be analyzing the data obtained from all the teams so as to be able to identify trends among material properties and be able to determine which materials are stronger than the others. Another option would be allowing the teams to experiment with more than just two samples and experiment with different configurations – such as a sandwich beam with Kevlar fibers on top, a polyurethane core in the center and carbon fibers at the bottom.


6.0 REFERENCES

Askeland, D.R., & Phulé, P.P. (2006). The Science and Engineering of Materials: Fifth Edition. Toronto: Thomson.

Callister Jr., W.D. (2004). Materials Science and Engineering: An Introduction. New York: John Wiley & Sons, Inc.

CES EDUPACK 2 [Computer Software].(2006). Cambridge, United Kingdom: Granta Design Limited.

Gay, D., Hoa, S.V., & Tsai, S.W. (2003). Composite Materials: Design and Applications. New York: CRC Press.

Gere, J.M. (2004). Mechanics of Materials. Toronto: Thomson & Brooks/Cole.

Iatridis, J.C. (2006). Introduction to Composites Retrieved October 5, 2006 from the University of Vermont, ME257 Composite Materials web site: http://www.cems.uvm.edu/~iatridis/me257/

Lanning Jr. D. (2003). Lab Photos. Photos posted on T-drive under ES321 Folder and inside Lab Photos Folder archived at T:\ES321\Lab Photos.

Tarnopol’skii, Y.M., & Kincis,T. (1985). Static Test Methods for Composites (G. Lubin, Trans.). New York: Van Nostrand Reinhold Company Inc.

(2005). Degradable Plastics: Glossary of terms. Retrieved October 5, 2006 from the Department of the Environment and Heritage of the Australian Government web site: http://www.deh.gov.au/settlements/waste/degradables/glossary.html

(2005). Engineering Stress-Strain Curve. Retrieved October 5, 2006 from web site: http://www.keytosteel.com/fr/Articles/Art43.htm


7.0 ATTRIBUTIONS


8.0 APPENDIX A: SAMPLE CALCULATIONS


9.0 APPENDIX B: RAW DATA

Table 9.1: Raw Data.

TESTNUM

POINTNUM TIME POSIT FORCE EXT

1655 1 0 -00.06041

8 -0.00354

1655 21.31821

20.00153

70.42292

8 -0.00222

1655 31.86746

7 0.002950.78543

8 -0.00086

1655 41.97731

80.00326

30.90627

4 -0.00056

1655 52.52657

30.00466

31.32920

20.00076

4

1655 63.24060

40.00646

21.81254

80.00243

1

1655 73.95463

60.00822

52.17505

80.00400

6

1655 84.55881

60.00973

72.59798

60.00529

6

1655 95.16299

60.01121

23.02091

40.00655

1

1655 105.65732

60.01247

53.44384

10.00757

7

1655 116.37135

70.01423

73.92718

80.00903

8

1655 126.81076

10.01532

54.41053

30.00994

2

1655 137.57971

8 0.017254.83346

10.01155

6

1655 148.12897

30.01858

75.25638

90.01269

6

1655 158.56837

70.01971

35.67931

60.01364

6

1655 168.95285

50.02067

56.04182

60.01447

4

1655 179.55703

60.02218

76.46475

40.01572

5

1655 189.88658

90.02298

76.82726

30.01640

4

1655 1910.4358

40.02433

87.18977

20.01749

1

1655 2010.6555

50.02492

57.55228

10.01797

2

1655 2111.2597

30.02643

8 7.91479 0.01925

1655 2211.5892

80.02726

2 8.27730.01995

91655 23 11.9737 0.02822 8.63980 0.02077


6 5 9 2

1655 2412.3033

1 0.029059.00231

80.02150

4

1655 2512.8525

70.03038

89.36482

70.02264

1

1655 2613.0722

70.03091

29.72733

60.02309

1

1655 2713.6215

20.03228

810.1502

60.02429

7

1655 2813.9510

80.03311

210.5127

70.02502

1

1655 2914.2806

30.03393

7 10.93570.02579

2

1655 3014.9397

30.03558

811.3586

3 0.02728

1655 3115.2692

9 0.036411.7211

40.02803

1

1655 3215.7636

20.03763

712.2044

80.02915

3

1655 3316.2579

50.03893

812.8086

60.03032

8

1655 3416.7522

8 0.0401513.1711

70.03143

4

1655 3517.1916

80.04123

713.5336

80.03241

8

1655 3617.6310

80.04228

714.0774

40.03339

5

1655 3718.0704

90.04342

514.5003

70.03445

1

1655 3818.5098

9 0.044514.8628

80.03542

4

1655 39 18.94930.04557

515.3462

20.03640

4

1655 4019.2788

5 0.046415.7087

30.03717

9

Documents

Composite Strength Testing