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THE EFFECTIVE PROPERTIES OF COMPOSITE MATEIUAL

BY

Hamid Ghaemi

Graduate Program [n

Faculty of Engineering Science

Submitted in partial fulfillment Of the requirement for the degree of

Master of Engineering Science

Faculty of Graduate Studies The University of Western Ontario

London, Ontario March 2000

O Hamid Ghaemi 2000

Abstract

Traditionally, the elastic and strength properties obtained fiom testing unidirectional

laminates were used in conjunction with Classical Lamination Theory to predict the

failure of multidirectional laminates. Many leading researchers have recognized that the

use of unidirectional laminate properties in such a lamina to laminate scheme results in

erroneous strength prediction. They have accordingly proposed methods to offset the

shortcomings of such a practice. The purpose of this research was to find the effective

elastic and strength properties of carbon and glass fiber epoxy laminae. It is argued that

this could result in more accurate predictions of the elastic and strength properties of

multidirectional laminates made fiom such larninae.

In this work a method for detemining effective larninae properties is outlined. which

uses the mle of mixture A cross-plied laminate was chosen for determining effective E2

. X. and Y First, eight samples of O" unidirectional were fabricated and tested to failure

to find El (125.5 GPa). To experimentally determine the onset of first ply failure. edge

replica technique was utilized. Three cross-plied sarnples were edge-replicated at various

intervals io allow for a visual determination of the load at which the 90" plies had failed.

Then. 12 samples were tested up to that failure point of the 90" layer where the Es

moduius of laminate was calculated (70.8 GPa). Knowing that each lamina orientation

contributes proportionally to the modulus of laminate, E2 was then determined to be 16.2

GPa Finally, 23 cross-plied samples were teaed to failure and the ultimate strength X

was determined to be 1 S IGPa. The CLT was used to determine the maximum transverse

strength Y. which was found to be 3 5.6 MPa.

It was found that the ultimate longitudifial and transverse strength of a cross-plied

laminate was smaller than the ultimate strength of unidirectional laminate with the same

number of O " plies. It is argued that the fracture toughness of composites, fracture mode

with respect to applied load, stress concentration at the crack tip at the interface of fiber

and matrix, orientation of the fiber with respect to the crack propagation direction, and

fiber hardening causes the discrepancy.

Keywords: Composite. Effective Properties. experimental Evaluation, Laminated

Materials. Unidirectional, Multidirectional, Fiber Reinforced, Composite Fabrication,

Carbon Fiber. Glass fiber.

ACKNOWLEDGEMENT

The Author would like to take this opponunity to express his appreciation for the support

received from the following faculty members and suppon staff

Dr. Z. Fawaz

Mr. Devin Ostram

Mr. J Karpynczyk

Mr Bob Pope

1 am especially indebted to my wife for her suppon and 1 would like to dedicate this to

rn!: wife

TABLE OF CONTENTS

CERTIFICATE OF EXAMlNATIO

ABSTRACT

AC KNO WLEDGEMENT

TABLE OF CONTENTS

LIST OF FIGURES

LIST OF TABLES

NOMENCLATURE

CHAPTER 1

CHAPTER 2 2.1 î 3 - , -

CHAPTER 4 4.1 4.1

CHAPTER 5

MECHANICS OF COMPOSITE MATENALS Introduction Mechanics of Composite Materials

LITRETURE REWEW Introduction Literature Review and Effective Propenies

EXPEEUMENTAL WORKS Introduction Manufacturing

4. I . I Compression rnould Fabrication 4.1 .S Autoclave Fabrication 4.13 Cutting and Orient ing Uncured Material 4 . 14 Step by Step Preparation of Laminate Lay-up 4.15 Vacuum Plate and Vacuum Assembly 41.6 Cure Cycle Testing 42.1 Testing Outlines 4.7.1 Sample Preparation 4.7.3 Tensile Testing Machine

NEW TESTING METHOD Introduction Transverse Modulus of Elasticity E90° Longitudinal Tensile Strength X Transverse Ultimate Strength Edge Replica

vii

CHPTER 6 EXPERIMENTAL DATA 6.1 Introduction 6.2 Young's Modulus

6.2.1 Young's Modulus Ex of Carbon Fiber Epoxy 6.2.2 Young's Modulus E, of Glass Fiber Epoxy

6.3 Longitudinal Ultimate Strength Longitudinal Ultimate Effective strength of Cross-Plied Carbon Fiber AS4-3 50 1 Longitudinal Ultimate strength of (O, +45, -45, O), Of Carbon Fiber Epoxy AS4-3501 Longitudinal strength of Quasi-isotropic Carbon Fiber EPOXY AS4-3 50 1 Longitudinal Ultimate Effective strength X of Cross -Plied Glass Fiber Epoxy Type 1003 Longitudinal Ultimate strength of (O, 145, -45, O), of Glass Fiber Epoxy Type 1003 Lonsitudinal Ultimate strength of Quasi-isotropie Glass Fiber Epoxy Type 1003

6 4 Transverse Ultirnate Strength (Effective) 6.4.1 Transverse Ultirnate Effective Strength Y of Carbon

Fiber Epoxy AS4-3 50 1 6 . 4 2 Transverse Ultimate Effective Strength Y of Glass

Fiber Epoxy Type 1003

6 5 Discussion of ResuIts

CHPTER 7 CONCLUSION

REFERENCES AND BIBLIOGR4PKY

LIST OF FIGURES

Figure Descript ion

Coordinate direction of composites

Symmetry about X 1- X2 and X2 - X1 planes

Normal S hear Coupling

Laminate Coordinate

Compression Mould

Photoelastic Oven and Box Jig

The Temperature Mechanism of Photoelastic Oven

The Autoclave fiont view

The Control Unit of Autoclave and Compression Mould

Actual Cure Cycle of carbon Fiber epoxy cross-plied laminate

Actual Cure Cycle of carbon Fiber epoxy angel-plied laminate

-4ctual Cure Cycle of Glass Fiber epoxy cross-plied laminate

Actual Cure Cycle of Glass Fiber epoxy angel-plied laminate

Test Sample Size and Configuration

The Front view of Test Specimen with Aluminum Tab

The side view of Test Specimen with Aluminurn Tab

Tensile testing Machine

The Calibration Curve of the Independent Load Ce11

The Calibration Curve of the tensile Testing Machine

The Gripper

Calibration Specimen with Strain Gauge for Machine Cross

Head Inspection

Calibration Specimen with Strain Gauge for Machine Cross

Head Inspection

The System Used for Machine's Cross Head inspection

Edge replica (x150) of a carbon fibedepoxy cross-plied

symmetric test specimen.

Stresdstrain Curve of Unidirectional O" AS4-350 1

Stress/strain Curve of Unidirectional O" AS4-350 1

Stresdstrain Curve of cross-plied AS4-350 1

Stresdstrain Curve of cross-plied AS4-3 50 1

Edge replica (x150) of a giass fiberlepoxy cross-piied

symmetric test specimen

Stress/strain Curve of cross-plied type 1003

Stressfstrain Curve of cross-plied type 1003

St.ress/strain Curve of cross-plied AS4-3 50 1

Stress/strain Curve of angle-plied AS4-350 1

Stresdstrain Curve of quasi-isotro pic AS4-3 50 1

Stresdstrain Curve of cross-plied Type 1003

Stresdstrain Curve of angel-plied Type 1003

Stresdstrain Curve of quasi-isotropie Type 1003

Work Hardening curve of carbon fiber epoxy

Work Hardening curve of carbon fiber epoxy

Failure of Multi-directiona and unidirectional carbon fiber

Epoxy laminate 88

Failure of Multi-directionai and unidirectional GIass fiber

Epoxy laminate 88

LIST OF TABLE

Table Description

Data Specification of Vacuum Assembly

The O" Unidirectional Modulus of Elasticity of AS4-350 1

The Modulus of Elasticity of Cross-Plied of AS4-350 1

The Young's Modulus Ex of Type 1003

The Ultimate Strength of cross-plied of AS4-350 1

The Ultimate Strength of unidirectional of AS43 50 1

The Ultimate Strength of angle-plied of AS4-3 50 1

The Ultirnate Strength of quasi-isotropic of AS4-350 1

The Ultimate Strength of cross-plied of Type 1003

The Ultimate Strength of angle-plied of Type l O03

The Ultimate Strength of quasi-isotropie of Type 1003

The Effective Properties of Carbon fiber and Glass fiber epoxy

Nomenclature

Basic Concepts:

O,, Components of stress tensor 0 Angular Orientation m Cos.8 n Sine E,, Components of strain [E] Strain Matrix [a] Stress Matnx

Constitutive Notation:

Stresses S trai ns Stiffness Coefficient (Contracted notation) transforrned Reduced Stifness Laminate Longitudinal Young's Modulus Unidirectional Young's Modulus Transverse Young's Modulus Major Poisson's Ratio Minor Poisson's Ratio S hear Modulus Ultimate longitudinal Strength Lrltimate Transverse S trength Laminate Mid Plane Strain Reduced Stiffness Matnx for K' layer Momentnength In-plane Forcdlength Laminate in-plane Stifiess Laminate Bending-stretching coupling Laminate Bending Stiffness

CHAPTER 1. INTRODUCTION

Classical lamina to laminate failure analysis is the most cornmon method for predicting

the failure of a multidirectional composite laminate. In this method the laminate failure is

dictated bp the failure of its individual laminae as they experience certain limit stresses.

Actual Lamina stresses used for this cornparison are computed using the classical

lamination theory (CLT) as will be described in the next chapter. The CLT also calculates

the laminate stifhess based on the lamina stifhess properties. Both the lamina elastic

propenies and experimental strength values are obtained from a set of standard test that

are most ly performed on unidirect ional. uniavial specimens.

W~ilt: ) ielding accuratr estimate of the laminate stifiess. the lamina to laminate analysis

procedure has quite oRen resulted in erroneous strength prediction. Different reasons

have k e n given to explain these discrepancies. Sorne of these reasons relate to the

inability of CLT to account for the out-of-plane stresses that m a y develop between

differently oriented laminae within rnuItidirectional laminates. These stresses can become

very significant around discont hui t ies and traction f?ee edges such as thox encountered

in straight edge test specimens. This phenornenon is oflen referred to as edge eeffect. The

incapability of the CLT to account for out-of-plane stresses is. by no means. the ody

reason for its Iack of accuracy in predict h g laminate strength. Another equally important

reason has to do with the difference between the elastic and failure properties of laminae

when presented in multiduectional laminate and when they are considered as

unidirectionai entities. The properties of individual lamina bonded together to form a

laminate. some times called in-situ properties. can be quite different Born the

unidirectional lamina properties. These properties are longitudinal and transverse

stifiess and strength of lamina. Reason for such differences have been adequately

discussed in the literature as will be outlined in the next chapter. Here is suficed to

mention that one major reason bas to do with the increase in crack growth resistance of a

given lamina when angle-plied or cross-plied with other laminae. This increase is mainly

apparent in matrix-dominated pro perties (transverse and shear propenies).

Since the accuracy of lamina-to-iaminate strength analysis depends greatly on its input

parameter. namely the lamina properties. the primordial issue becomes one of accurately

dctermining reprcsentative lamina propenies. To offset the disadvantages of

unidirectional test coupon some researchers have recomrnrnded the use of a number of

mu lt id irect ional specimen configurations to O btain the various e tTec t ive propenies

required to hilly characterize fiber-reinforced laminae. While the properties mrasured

are the same as those commonly used in the classical theories. those researchers suggest

the use of new specimens to obtain them. For example. the standard O" unidirectional

specimen usrd for obtaining trnsile properties of fiber-reidorced laminae according to

ASTM Test Method D3039-76 is deemed inappropriate to obtain an. properties but the

longitudinal Young's modulus E l and the associated Poisson's ratio VI? A number of

alternative specimrns are thus proposed to obtain the remaining properties of the lamina.

The comrnon feature among these specimens is that the target lamina (lamina for which

the properties are sought) is always constrained in one way or another thus. ensuring a

more prediction of the properties intended for use of the CLT within the context of a

lamina to laminate failure analysis.

The new tening philosuphy alluded to here and referenced in the next chapter was

formulated wit h advanced aerospace composites in mind. the kind composed of very

strong and stiff fiber in a very brittle matrix. As such new general guidelines are

proposcd that only applied to this class of maierials. The principal subject of this thesis is

there fore. to propose specific procedure for determining those lamina properties that

cannot be accuratel. determined fkom standard unidirectional tests. Ultimately. the aim of

this work is to show how these new effective properties can be used to accurately predict

the strength of an) mult idirect ional composite laminate under both static and fatigue

load. First. the new tens are described and the procedures for reducing the effective

lamina propenirs From these tests are outlined. h experimental investigation is then

carried out to compare the properties obtained using the new procedure with those

obtained using standard test methods and to quant. the improvement in laminate

strrngt h predictions whrn these effective properties are used. Finally. advantages

associated with the new test method are delineated and some recommendations are

tormuhed.

To obtain the necessary data. many test specirnens were fabricated using an autoclave

that the author of this thesis has contributed significant ly toward design and fabrication.

The fiber orientation test specimens chosen were (0. 90. 90. O), known as cross-plied. (0.

4 5 . -45. O), known as angle-plied. and (0. +45. -15. go), known as quasi-isotropie

laminate. These tests revealed significant information about the effective properties of

laminated composite in multi-axial arrangements.

CHAPTER 2. MECHANIC OF COMPOSITE MATERIALS

2.1 Introduction

This chapter provides a brief eqlanation of mechanics of composite materials. This

chapter is a review of some of the fundamental concept of mechanics of composite

materials. This chapter also establishes much of notation that will be used throughout this

thesis.

2.2 Mechanics of Composite Materials

Acceptancr and usage of composite structure requires knowledge of their load carrying

capacity and it is necessaq to predict the strength of a particular laminated composite

with certain amount of confidence for design purposes. Most Iarninate strength prediction

sc bernes available tel? on uniaxial strength data. However. a composite laminate is rareb

unidirec t ional. Indeed. consider the unidirect ional composite lamina in figure 2.1. the X

and the Y directions are the axial and lateral direction of the lamina where as 1 and 2

refers to t h r longitudinal and transverse.

Figure 2.1 coordinate direction of composites

As it is depicted in figure 2.1. the aiflkess and arength of fiber-reinforced

5

lamina are

directional. This means that laminated composite is stronger in the fiber direction (1) than

the transverse direction (2) thus. characteristics of composite materials are different than

conventional engineering rnaterials. Most conventional materials are homogeneous and

isotropic but composite materials are inhomogeneous and non-isotropic (onhotropic or

more generally anisotropic).

.-b inhomogeneous body has non-uniform properties over the body. i.e.. the pmperties

are a îünction of position in the body.

Figure 2.2 Symmetp about X I - X2 and X2 - X3 Planes.

.-ln r>rrhorropic bodv has material properties that are different in tree mutually

perpendicular direction at a point in the body and further. have three mutually

pirpend icu lar planes of material symmetry

An uonisotropic body has material properties that are different in al1 direction at a point in

the body. Thcre are no planes of material propeny symmetry.

A lamina is a flat arrangement of unidirectional fiber reinforced in a matrix and laminate

is a srack of Iaminae with various orientation of principal material direction in the

laminae. In study of composite rnaterials. the principal material direction rarely coincides

with coordinate direction. Thus. a relation is needed between the stresses and strains in

the principal direction and that of the body coordinate. This relationship is known as

transformation matrix. Using tensor notation:

- oij - aki ali D k l ski ali = 6,

. h d the transformation equation for fourth order tensor Cijki can be written as follows:

C9ijkl = a m i ani ark as1 Cm,,

In\ oking the following tensor notation and replacing aii by matrix representation it c m

be found the transformation matrix is a function of 8 fkom the X - a i s . Therefore. the

transformation is merely a rotation of stress and strain not material properties. Usine

direction cosine. the relationship between stresses and strains in the principal material

and global direction are given as {o) = [TI] { O ] , and ( & ) = [Tz] {E), which is the

generalizc Hooke's law. The transformation matrk of stress and strain differ by factors of

These transformation matrices are used to calculate the stifiess and cornpliance matrices

of lamina ftom material to global coordinate. It is ofien the case in the analysis of

composites that a condition of plane stress exists ( 2-D ) or is a very good approximation.

Thus the equation for the in-plane components of mess in tenns of transformed stiffiess

coefficient is as follows:

hhere the transformed reduced rnatrix coefficient Q ,j are defmed

(i, j = 1,2,3)

In tensor notation:

Mere Cijkl is called elastic rnodului or stifhess. Inverting the abovr cquation. gives the

strain in terms of stress in the following form:

Ekl = c-lijkl Ekl = Sijkl bij

Mere Sijkl is called cornpliance matriu.

The reduced stifhess matrix for orthotropic materials in principal matenal direction has

t e m C16 ( i * 6 ) equal zero. Hence. in principal material coordinate of an onhotropic

material. the plane stress constitutive equation has the following simplified form:

From the proceeding equation it can be seen that only four of the five material constant

Cor plane stress of an orthotropic material are independent. The constant that cm be

measured accuratelp in the laboratory are El. Ez. VI?. and elz. Accurate measurement of

Poisson's ratio v2l is often difficult because it is very small for many composites.

Tliere fore. the fo llowing equat ion is used to d e t e d e transverse Poisson's ratio from

other material constants.

The component of reduced stifbess matnu for orthotropic material in terms of

engineering constant is as follows:

The 2-D transformation equations for rotation about the 3' (z) avis are straightforward

simplification of the 3-D equation. As it was mentioned before. 2-D in-plane state of

stress is a very good approximation method for formulation of stress-strain relationship of

composite. We use the transformation matrix for stress [TI] and strain [Tz] ro transform

the plane stress constitutive equat ion in principal material direct ion to the global

coordinate.

The individual C,, tenns of the transformed reduced stifiess m a t h are:

O,, - = Cl, cosJ B + 7 ( ~ , ? + ZC6,)sin2 O cos2 0 + C7, -- sinJ 8

Q22 = Cl, sinJ 0 + I(c,, + 2c,)sin2 0 cos' 8 + C,, -- cosJ 0

Q,, = (cl, +c,, -- -4c,)sin2 B C O S ~ B + C , ~ ( S ~ ~ ~ e+c0sJe)

0, = (cl, +c,, -- -x,: - ~ ~ , ) s i n ~ ~ c o s ~ ~ + ~ , ( s i n ~ ~ + c o s ~ ~ )

Q : ~ = (c,, - c,: - sin ' B cos B + (c ,~ - Cl7 - + sin B cos' e

The reduced stifiess coefficients are fourth order n the sine and cosine function. The Q i 6

and QZ6 coefficients are very important in that they define the coupling between the in-

plane normal and shear responses. These two coefficients are zero for isotropic materials

and zero for orthotropic materials in the principal material coordinates. Therefore. there is

no coupling between shear and normal stresses in these cases. The presence or absence of

normal-shear coupling is demonstrated in figure 2.3. For isotropic and onhotropic

materials Ioaded in principal material direction there is no distortion of the original right

angle. However. for the unidirectional off-axis lamina coupling is clearly demonstrated

through the distortion of the original right angle.

Onhotropic

Figure 2.3 Normal-Shear Couplhg

O ff-asis lamina

Clussical Laminut ion Theon:

The stress-strain relations in an arbitrary coordinate are usefûl in the defmition of the

laminate stifiess. In fact. Classical Lamination nieory embodies a collection of

t rans formed lamina st i fhess and strength to calculate mult idirectional laminate nifiess

and strength. Classical Lamina-to-Laminate failure is the most cornmon method for

predicting the failure of multidirect ional composite lamhate. In such method. the failure

is dictated by the failure of its individual laminae as they experience theu failure stress.

The basic lamination theory is the stress-strain or constitutive relation and failure criteria

for the individual lamhae each consisting of a layer of unidirectional fiben. irnpregnated

and h l l y surrounded by matrk material. The strong stiff fibers provide the prirnary load

c q - i n g capability while the matriv protects and suppons the fibers and transfers the load

between them For the laminate in figure 2.4 we take the global x-y-z coordinate system

with z perpendicular to the plane of the laminate and positive downward. The origin of

the coordinate is located on the laminate mid plane. central between the top and bonom

surface. The larninate has N Iayers numbered fiom top to bottom. Each layer has a

distinct fiber orientation and as indicated in the diagram the z-coordinatte of the bonom

of the K ' ~ laver is ZA - Zk.1.

1

Figure 2.4 Laminate Coordinate

Lamination theory is the mathematical formulation for predicting the mcro-mechanical

behavior of a laminate based on arbitrary assembly of homogeneous orthotropic laminae.

Two-dimensional theory is most cornmon; while three-dimensional theory is very

complex. If the plane stress assumption is valid in lamination theory for a thin plate

subjected only to in-plane or membrane loads that do not cause any instability. the stress-

strain behavior becomes relat ively simple. Basic Assumpt ions made in C lassical

Lamination Theory are as follows:

i . The structure is restncted to be a thin plate or shell consisting of an arbitrary

combination of plies cured or consolidated into a single laminated plate

2 . Only plane stresses (0,) ) exist as a, is zero. Transverse shear stresses ( s- and r?,)

are neglrcted to meet the classic thin plate and shell theory

3. The e t k t of discontinuities such as holes. cut outs. stiffeners and nrighboring

edps are ignored to allow for a 'point" analysis of an effectively infinite large plate

and shell

4. Loading is assumed to be in-plane membrane stress and moment resultants

5 . The pcrfect bonding between the layes are assumed and forces of traction free

rdge (edge effect) are ignored.

The total strain (E), at any z-location in the laminate in ternis of mid plane straui {&O),.

and the curvature (r), is shown bellow. This is the fundamental equation of lamination

theory. The total strain is the sum of the mid plane strain and the strain associated with

curvature. It showed be noted that the equation derived up to this point is independent of

the type of material and the number of layer in the plate.

There fore. the stress at and z-location can now be determined simply by substitution of

the strain in to the plane stress constitutive equation.

In the above general equation the fust term correspond to the stress associated with mid

plane strain and the second term corresponds to the stress associated with bending strain

for K ' ~ layer. Please note that the subscript *x' is dropped to indicate that the equation is

valid for any arbitrary coordinate system The in-plane forces per unit length and

moments are defmed as the through thichess integrals of the in-planar stress in the

laminate.

Thus. combining the stresses and the in-plane force and moment gives:

A = I k : , k: .,

The aforemrntioned equation relates the in-plane forces per unit length to the mid-plane

strain and cm-ature. The [A] matrk represents the in-plane stifiess. the @3] matrir

defmes the bending-stretching coupling and ID] matrix defmes the bending a i f iess

matriu. From these equation it is evident that [A] rnatrix is a fùnction of iayer thickness

but is independent of stacking sequence. [BI and [DI matrices is dependent on the

stac king sequence of individual layers.

A,, is the extensional m embrane stif'fnes

D,, is the tlexural or bending stifiess

B , is responsible for the coupling between membrane and binding behavior

Escluding t hermomechanical propenies. the constitutive equat ion for a thin laminated

anisotropic plate can be written:

Or b) the following form

There is coupling between extensional de formation

the existence of the [BI matrix. Even within the

and bending deformation caused by

lirnits of a small deflection theory,

forced curvature within the laminate introduces in-plane loads M. through this type of

coupiing. In addition in-plane drains [E] would induce curvatures in the laminate as the

neutral axis and the rnid-plane of the laminate not being coincident. To be more

practical. fabricating the laminate symmetrically about the mid-plane so that an equal

number of identical plies in t e m of both thickness and fiber orientation located at the

samr distance abovr and below the rnid-plane so that the equations reduce to eliminates

the coupling matrix [BI:

.A laminate that is syrnmetrically laminated about the mid-plane is referred to as

"homogeneous" anisotropic. If the larninated is constructed with equîl number of pairs of

laminatr with symmetry about the coordinate (x.y) axis (angle ply laminate) the [A]

rnatri-x becomes onhotropic in nature ( A , , = A, = 0)

The [Dl rnatrix remains hlly populated and anisotropic in nature. If the laminate is

constructed with equal pain of laminates at angles of OQand 90' to both the x-y axes. the

[Dl matris wiil become onhotropic in nature. More practically. this equation may be

C'pon calculation of laminate stress and strain due to the applied external load per unit of

width Nx. the lamina stresses cari be calculated using CLT formulation. The mess is

considered as lamina transverse ultimate strength since the matrix dominated area faiis.

The failure of a lamina within the laminate occurs when its limit stress is reached.

However. the failure of the lamina is not indicative of laminate failure. Because the

matriv holds the fiber in tacked. the stress becomes redistributed evenly through out the

laminate some distance away fkom the first crack in the lamina. Therefore. a matrix

dominaicd lamina would crack in a progressive rnanner until its effect and contribution to

the laminate strength is zero. In this work al1 of these schemes are considered in

explaining the failure mechanism of laminated composite test specimen.

CHAPTER 3 LITRERURE REVIEW

3.1 Introduction

This chapter explains different approach taken by other researcher to study the failure

mechanism of composite matenals. First the major school of thought in this area is

discussed and then explains their differences. At last bnefly cites other works that are

experimental and are parallel to this study. The needs for effective properties are clearly

explained in this chapter.

3.2 Literature Review and Effective Properties

There are two schools of thought for the mechanics of composite matenals,

Micromechanics and Macrornec hanics. From Micrornechanics point of vi ew, fibrous

composites are made of two distinct constituents, fiber and matrix (epoxy) and the

mechanical properties of composite are traditionally believed to be dependent on the

volume fraction of the constituents. Although this assumption is tme, it is not the only

parameter that composite propenies are dependent upon. The overall strength and

stiffness are also dependent on the mechanical properties of interface since it occupies a

laye volume tiaction within the composite. This school of mechanics of composite

material includes self-consistent method, the differential self-consistent method, Mon-

Tanaka method. and generalized self-consistent method reference[l,2,3].

On the other hand, the Macromechanical work isdvided into two major categories,

numerical and experimental. The analytical (numerical) work usually involves computer

prograrnming. The program would then generate failure prediction based on a failure

model. The experimental work on the other hand relies on the data obtallied

experimentally. In experimental method occasionally theories would be improved and

slightly modified to correspond to experimental results. From engineering stand point.

there is not a universal defmition of failure criteria of composite rnaterials. In a more

~rneral fom the failure is governed by failure of any one of the constituents of composite b

in ful filling its structural hnction. Within the academic research comrnunity. a steady

strearn of work has k e n done for the last 50 yean. The assumption in macroscopic

failure theory is that individual larninae are homogeneous. orthotropic. and have known

strength for one-dimensional state of stress in material's principal direction. Application

of rnacroscopic failure theones to each layer of a laminate corresponds to frst ply failure

theory. That is to say that failure of laminate has happened when fust piy has failed. In

almost dl cases this is not true as multiple transverse cracking is expected to occur in any

layer at an angle to the primary ioad direction.

The Classical Lamina-to-larninate theory is used in computing actual lamina stresses. It

also calculates the laminate stifiess based on the lamina stifhess properties. Both the

lamina elastic properties and the experimental strength values of a lamina are obtained by

carrying out series of standard tests on unidirectional laminate. according to committee

D30 of ASTM. and committee D638 for carbon and glass fiber epoxy respectively.

These tests are carried out in such way to fmd the properties of unidirectional laminate. It

is well rstablished and agreed upon amongst researchers that the ASTM standard is not

sufficient to be used for fmding effective mechanical properties of fiber reinforced

materials reference[4]. This type of analysis quite often gives erroneous results in

predicting strength of composite materials in mult idirect ional configwat ion.

Nurnerous rnacroscopic failure theories have been proposed. To name a few of these

the0 ries. Hart-Smith introduced Maximum S train and Truncated Maximum Strain

theories reference[S]. McCartney used hcture-energy-based analysis on damage

mrchanism reference(61. Puck and Schurmann developed a critenon in which two

independent fracture criteria. fiber fracture and inter-fiber fracture were used including

the gradua1 loss of stifiess after initial crack formation in determining failure behavior

reference[ï]. Wolfe and Butlia employed a strain energy based theory in predicting the

failure of laminate reference[8]. Hashin used Rotem failure criteria to determine the

failure rnvelop based on the stress strain curve reference[9]. Tsai. afler over 30 years of

researck strongly believed that the quadratic failure criterion with linear terms called

Tsai-Wu failure criterion best described the failure behavior of composite materials

referencr[lO]. However. the most widely used failure theory today is maximum stress

theop and maximum strain theory which al1 of aforementioned theories are integral of

the two. The maximum stress criterion assumes that failure occun when any one

component of stress reaches the limiting stress value. Thus. the safe condition of the

maximum stress failure criterion is X, < oi < XT. Yc < o, < YT. and Z, < o, < Zr

reference[ll]. For off-êuis lamina the transformation on or after global to principal

material direction is used. The maximum nrain theory is identical to maximum stress

theop except that the limiting parameter is strain. Zinoviev reference[3] in Russia

developed maximum stress theory. This theoq has k e n fùrther developed and modified

to consider factors such as unloading behavior of cracked kminate. geometric non-

linearity caused by change in angle ply and rnany othen.

The prediction of laminate failure is believed to be based upon some failure criteria at

lamina level. Many researchers have followed the same method and have proposed

various theories to explain the behavior of fiber reinforced composite materiais. But the

success has k e n mainly dependent upon the mode of failure at lamina level best

predict ing the laminate failure refercnce(l21. However. in al1 cases the longitudinal and

transverse properties are cowidered as intrinsic properties of the composite material.

Thus. the main task to achieve successfùl and accurate strength prediction is strongly

depended on determinhg these intrinsic properties of composite rnaterials. Many have

artrmpted to experimentally measure these properties and relate them to the fiber

orientation. Such as the 10 % rule by Heart Smith et dl reference[l3]. in which he states

that the contribution of any orientation other than O" is considered to be IO-percent of its

unidirectional strength to the laminate strength. Or others run extensive tests to find

various factors for their analysis such as that of Tsai-Hill theory reference[I4]. Tsai

attempted to provide better correlation lxtween theory and experiment by including al1 of

stress components.

On the other hand Hashin gives methodology and issues involve in prediction of effective

propert ies. in which he uses exact variational method for chic properties re ference[ 1 51.

Naik and Murty suggea a simple criterion to identifj the dominant failure mechanism

reference[l6]. Tney propose that failure involve a combination crack syaem

representat ive of matrix cracking. fiber buckling, and rnatrix crushing. More recent

works by Bader and Curtis. the tensile strength of carbon-epoxy cross-plied

sqmmetrically at the initiation of transverse crack was found to be a fiinction of the

thickness of 90" layer reference[l7]. Bledzki and Kessler used vibration testing to f k d

the longitudinal and transverse properties of unidirectional composite reference[ 1 81. In

al1 these works. the lamina-to-laminate analysis is used to fmd the lamina stresses. Then

one can conc lude that although there are diflerences between aforementioned theories,

the? al1 have one common feature and that is the failure of the laminate is govemed by

the failure of the individual plies as they undergo a cenain limit stress or strain situation

according to a failure theory. Thereferenceore. the stresses in the individual laminae

uhich are used for a given failure theory are calculated using Classical Laminate to

Lamina Throry in conjunction with experimentally found intrinsic propenies laminate.

From the composite design point of view. a failure analysis based on fracture mechanics

criteria is an extremely involved method. To maintain the simplicity in strength analysis.

not compromising the accuracy. and accounting for eacture process would necessitate the

introduction of effective or in situ properties. This is a relatively new idea and more

recent work by McLaughlin et al reference[l9] showing that the lamina strength exceeds

those rneasured 6om unidirectional laminate test.

In other word. the properties of individual laminae bonded together to form a laminate

can be different than the unidirectional lamina properties. For such anisotropic or

O rthotro pic material analysis based on unidirect ional pro perties O fien result in erroneous

strength prediction Because of the directionality of the properties, namely the Poisson's

mismatch between difierently oriented layers, an inter-laminar shear stress is generated

when laminate is subjected to monotonie load. These stresses are very significant at the

discontinuity and free edge of the samples and quite ofien cause prernature fàilure of the

sarnple (called the edge effect). The incapability of classicd laminate theory to account

for out-of-plane stresses is not the only reason for the lack of accuracy in predicting

failure stresses. h o t h e r equally significant reason of inaccuracy is due to the fact tint

the behavior of lamina in a unidirectional laminate is different than its behavior when

used in multidirectional laminate. One reason for this is the crack growth resistance of a

given lamina when they are cross-plied or angle-plied with other laminar. Other reaçon

has to do with work hardening of reinforcing fibers and matrk soflening. thus increasing

the strengt h.

For the most highly constrained laminate (namely 0190). Flaggs and Kural found an in

situ properties 3.18 times that of unidirectional transverse test reference(201. This

increase happen mainly in matrix dominated properties. shear and transverse. From this

argument. one c m deduce that the accuracy of the lamina to laminate analysis is highly

dependent on the accuracy of lamina's properties used as input data. Therefore. the issue

becomes that of accuratel? determinhg the effective or in situ properties of lamina

CHAPTER 4 EXPERIMENTAL WORKS

4.1 Introduction

In this chapter the manufacturing methods are explained. A detailed explanation of

components that were designed for the purpose of fabricating test coupons is outlined. In

addit ion, the testing facility used for tensile testing is fully descnbed.

4.2 Manufacturing

The simplest and oldest rnethod of manufacturing composite stmcture is called wet-lay

up. In this process layer of fabric or mat are put together in a desired way and then the

mat is saturated with liquid resin. The layers are aacked together to achieve the desired

thickness and after the resin is appiied, the composite is transferred to the mould to be

shaped. A slightly better and superior method with less handling problems is one that

uses preimpregnated composite. In this method the sheet of reinforcing fiber is

preimpregnated in partially cured resin to form a so-called pre-preg lamina. The

advantages of the pre-preg method are that the resinhardener ratio is precise.

Manufacturer can also closely control the resin distribution per unit area. Finally the use

of pre-preg offers better consolidation and higher fiber volume fiaction than wet-lay up.

For this research, pre-preg materials are used to fabricate test coupons.

In general two methods of fabricating test Iaminate fiom pre-preg composite materials

can be adopted, namely compression moulding and autoclave consolidation. In

compression moulding, the specimen is under a direct pressure and the specimen is

sandwiched between the male and female part of the mould. In this process, the specimen

k i n g made would take the shape of the mould. The second type of fabricating composite

test samples is to use an autoclave in which the specimen is under hydrostatic pressure

and would take the shape of the vacuum plate. Both of these systems were designed and

implemented in houe such that common hardware was used for two processes in

fabricating test samples.

Advanced fiber reinforced composite materials have a unique cure cycle that has to be

iollowed closely if reliable test samples are to be obtained. The cure cycle varies

drpending on the type of material with the variables being the temperature. pressure and

processing t h e . The fabrication process is divided into three stages: pre-cure. cure. and

post-cure nage. During pre-curing the sample is heated to a k e d maximum temperature

while kept in a vacuum environment up to the recommended manufacturer temperature.

The vacuum pressure is released before heat liquefies the m a t h material when pre-preg

materials are used. In the curing stage. the sample is kept at the liquefyinp temperature

under a specified pressure for a specific period of t h e . Finally. in the pon-curing stage

the sprcimen is cooled down to room temperature at a specified rate under pressure. The

rate of temperature &op is such ihat no or minimum residual thermal stress would exist

once the room temperature is reached.

The above these three stages are essential parts of the manufacturing process and.

rcgardlrss of the method of fabrication these steps have to be followed. Consequently.

the main di fference between compression moulding and autoclave consolidation is the

fact that the means of applying pressure in compression moulding is by physical contact

whereas autoclave consolidation relies on hydrostatic pressure. To have versatility in

manufacturing means, both method of rnanufacturing were designed and built in our

laboratory. To ensure the reproducibility, both processes were automated. Each method

of manu facturing needed contro 1 board, analog/digital converter, cornputer, and data

acquisition system, Uiput/output modules and means of applying pressure. Furthermore,

each fabricating ce11 was designed to use the same control mechanism.

4.2.1 Compression Mould Fabrication

The mould can be made from almost any material that c m hold its shape under the

maximum 80 psi pressure and 350 "F temperature required for curing most composites.

Compression mould is made of two parts one is male (plug) and the other is fernale

(cavity). Normally the side of the part that should be smooth and glossy is put against the

rinished side rnould. since the mouid is fabricated with a very smooth surface see figure

4.1.

Figure 4.1 Compression Mould

In the case of test coupons. both sides of the larninate have to be smooth therefore, both

plug and cavity were mac hined to a surtace finish of 20 to 50 p in. using surface grinding

machine (see figure 2.1). With pre-preg materials. the laminate is laid up and cut to size.

and then put in between the male and f e u l e parts. It was decided that the compression

moulds be made of high carbon steel for its high stiffhess and thermal properties. Once a

test panel was oriented in the desired direction and laid up, a porcupine roller was used to

make small holes in the laminate. The panel was then placed in a vacuum plate for a

period of 4 hours at room temperature under 20 inHg (68 Kpa) vacuum pressure. This

was performed to suck out the air pocket and some excess matrk fiom the larninate for

the right ratio of fiber and rnatrix. Since the matrix liquefies at high temperatures. the

ho les made by porcupine roller would disappear and the only function they perform is

that of providing charnels for air to escape under vacuum The pre-preg manufacturer

recommends the use of porcupine roller and it would not cause any damage to the fibers.

To apply the pressure. a box-jig was designed and built. The box-jig was made of

standard C-charnel and 3 6 in by 1/2 in rectmgular tension rods. Four precision holes

were machined in each male and female part of the rnould. Four guiding pins were used

to ensure male and female part of mould would press the laminate uniformly when the

pressure is applied. Then the mould was placed into the box-jig where a high temperature

resistant pneumatic cyiinder would be used to apply pressure onto the compression mould

see figure 4.2. Because air is unstable at elevated temperature. nirrcgen was used to

supply the necessary pressure to the cylinder. The mould and box jig assembly were

placed into a photo-elastic oven. The temperature gradient was controlled by the oven

controller, which utilizes a Cam. The Cam is cut fiom 1/4in Plexiglass. The temperature

controller has an arm that is connected to a pulley-like roller at the end. The controller

has a precision clock built in that has housing for c m to be mounted on. As the clock

works. the cam housing and consequently the cam rotates in such manner that one full

rotation is equal24 hours. The arm of the controller at the roller end follows the profile of

the c m as the cam rotates. Thus the tirneftemperature gradient is controlled precisely

providing the Cam is cut accurately, see figures 4.3.

Figure 4.2 Photoelastic Oven and The Box Jig.

Figure 4.3 The Temperature Mechanism of The Photoelastic Oven

To examine the temperature gradient more closely and to check the htegty of the c m

three thermocouples were used to measure the temperature within the oven and around

the rnould. The tirner of the cornputer was used along with the themocouple to

accurately monitor he ternperature gradient and to check if there was discrepancy

between the desired ternperature and actual temperature in the oven chamber. To

measure the applied pressure. a pressure transducer was utilized. The information fiom

thermocouples and pressure transducer was fed into the computer and data acquisition

system for instantaneous analysis and monitoring of fabrication process. In case of loss of

pressure. the pneurnatic control valve would apply more pressure otherwise it would

maintain the pressure by closing a solenoid. See appendix B for pneumatic circuitry.

Three safety pressure mechanisms were used in this system. A pressure relief valve. a

manual relief valve. and a pneumatic valve controlied by computer. The relief valve

would automatically release the pressure if the regulator of the nitrogen tank stans

malfunctioning. The toggle switch was used as a manual relief valve that could be

sxitchrd off if the pressure gauge on the pressure line would show a higher than desired

pressure. Finally, the computer receives data fkom a pressure transducer. In case of high

pressure. the pncumatic control board would change the control valve frorn pressurization

mode to rxhaust mode. Three safety measures were used to ensure that should any one or

two mschanisrn fails there is another safety feature Ieft to watch the fabrication process.

The main draw back of compression mould shaping was the size of the panel made. This

method of fabrication needed physical pressure and the capacity of high temperature

resistant pneurnatic cylinders was the limiting factor. Since the average ccylinders

available are capable of applying pressure not bigher than 200 psi ody small panels were

made allowing for two test coupons to be cut f?om each panel. Although this method

produced a very reliable specirnen but the number of run required for fabricating the

many specimens needed for this work. made this process very time consuming and

costly.

4.2.2 Au toclave Fabrication

Although the compression mould shaping process was very robust and repeatable. there

sriil were some minor variations in fabrication process kom run to m. Even though the

contribution of these minor variations was totaily negligible and did not cause variation

to mechanical properties of the test samples. it was a good practice to minimize them by

having a higher volume of samples being fabricated at once. Auto-calve was mainly w d

for manuîacturing most of the test specimens due to the fact that 8 sampies were

fabricated in one manu fact uring mn.

The project of designing and building an autoclave came about as a result of the need for

easy operation and higher productivity of test coupon Since a 55-liter thick walled s t e m

chest Kas available which aiready had a large port opening at one end and 4 srnaller ii2i.n

opening. the project of building an autoclave became possible. The large opening was

used as the led to access the inside chamber of the autoclave. A high temperature O-ring

was used to seal the opening and to keep the hydrostatic pressure for extended period of

time with minimum loss of pressure. Each of the four other srnaller openings were used

for safety and operating devices necessary for a pressure vesse1 to be safely converted

into an autoclave which is a fail-safe system.

One of the advantages of autoclave is that it allows the simultaneous imposition of heat.

pressure. and vacuum pressure. Since the pressure vessel was relatively srnall. the

uniformity of heat tlow was not a problem. The heat was supplied by 1 1 5 volts 500wat

coi1 element that was connected to a relay. The relay was controlled via cornputer to tuni

the power off or open the power line to the elements for the purpose of temperature

control. The heating elements were passed through ceramic tubes that acted as electrical

insulat ion. Furthemore. those tubes provided a shielding mec hanism against

concentrated heat fiom the elemeni and dispersed the heat evenly through out the vessel.

.-\ picture of the autoclave is presented in figure 4.4.

Figure 4.4 The Autoclave Front View

Three J type thermocouples were placed at three locations inside the autoclave, one at

each end of the vessel and one at the center. The data obtained from thermocouples

indicated temperature difference of 5 "F which was well within the acceptable range. The

pressure again was supplied by using nitrogen because air becomes unstable at high

temperature (+ 400 "F). Some samples were initially made using the Ryerson's air supply

line knowing t hat the maximum temperature reached during fabrication would not exceed

380 9. However. it was considered safer to use an inert gas such as nitrogen. One of the

pressure vessel's openings was used to pass the thermocouples inside the vessel. The

passing line was filled with high temperature silicone to prevent pressure escaping during

fabrication of the sarnple. A pressure transducer with 500 psi maximum capacity

monitored the pressure. Pressure transducer was comected to one of the vessel's opening

by a T-joint. On the same line a pressure gauge was installed to provide a visual reading

of pressure right inside the vessel. To control the pressure. a pneumatic control circuit

was designed which consisted of two four way control valves with built in flow control.

elrctrical ly controlled so leno id. spring loaded soleno id. toggle switch pressure gauge. and

pressure relief valve.

The Control board was assembled ont0 wooden board. The high pressure nitrogen tank

was comected to the toggle switc h on the control board. The nitrogen tank is located at

location away fiorn the autoclave (and photoelastic oven in the case compression

moulding fabrication). This board was placed inside of an enclosure that houses power

supplies. analog/digital converter module. input/output module board. a cornputer that

controls the whole fabrication process and data acquisition unit. This unit includes alI

components and is capable of running two system of manufacturing independent ly .

Please see figure 4.5.

Figure 4.5 The Control Unit for Autoclave and Compression Moulding.

1.2.3 Cutting and Orientating L'ncured Materials

Two common pre-preg materials were used in this study. namely AS4/3051 carbordepoq

bu Hexcel and TI003 glasdepoxy by 3M. Pre-preg materials were supplied as

unidirectional tape of l l in by 72 yard and 36in by 48 yard for AS4 and Type 1003

respectiveiy. These tape corne in roles with one side of the tape covered with a protective.

peelable wax paper. A clear glass of 5Oin by 36in by 1/4h was used as the laying up cell.

The unidirectional pre-preg tape was unrolled and placed flat on the clear glass with the

wax paper facing up. The desued orientation of the fiber was then marked on the wax

paper. A 30in alurninum niler was made for the purpose of providing a cutting edge.

Standard Knives with replaceable blades (utility knife or Stanley) were used to cut

uncured pre-preg laminate.

1.2.4 Step-by-Step Preparation of Laminate Lay-up

The stsps taken to Iay up laminates in both methods of rnanufacturing were the same. The

pre-preg was removed 6om fieezer and allowed to w m up io near room temperature

and then the following procedures were taken in laying up the plies:

I - Two layers of II in length and four layer of 1 l in length of the pre-preg were cut of the

supplied material. The width of the supplied pre-prep composite was Ilin. This

provided enough prc-preg material for cross-plied laminates. However. for other

angle-plied and quasi-isotropic laminates the size of the cut was different. In fact a

template was made for +JS" and 4 5 " layers.

2- The pre-preg Material was then mcasured and cut 0.15in larger than required panel

size for vacuum plaie. which was 20.5 in by 5 in. The slight over size cut was o d y

done because of the possible misalignment of layers as they were laid up and that

furthrr trimming would be needed. This step was proven to be a good practice as

misalignment trnded to occur due to the adhesive properties at room temperature.

3- The plies were stacked together one by one in sequence and desired orientation. To

reduce the chance of having air pockets between the plies. each top ply was nibbed

with a roller. supplied by De-Comp composite Inc.. very gently along the direction of

it's fibers.

When al1 the plies were stacked together. the uncured h i n a t e panel was cut to 20in

by 45in size. This size allowed for the fabrication of 8 samples of 10 in by 1 in with

an extra 0.5 in width to account for the waste caused by the thickness of diamond saw

cutter.

Both surfaces of the mould were sprayed with a mould releasing agent to prevent the

test sarnple fiom sticking to the mould. A temperature resisting Silicone was used as

a release agent.

To hrther enhance the surface fmish of the final product and to ensure an easy

separation of the part From the mould. two layers of my lar were used to sandwich the

composite hminate.

In case of compression moulding. the laminate was placed onto the female part of the

mould and the male part of the mould was placed on it. In the autoclave consolidation

procrss. laminate was sandwiched between layers of bleeder. breather. caul plate. and

vacuum bag. Then the vacuum plate assembly was placed into the autoclave for

curing.

XOTE: Excessive use of Silicone shouid be avoided as it cm cause a dent formation on

the surface of the test sample. This is due to the fact the thermal viscosity break down of

the Silicone is much higher than that of rnatrix. If excessive Silicone is used. number of

droplets may form because of the surface tension. These droplets would cause the dents

on the surface of test sample at high temperature by penetrating into the molten matrix

when pressure is applied. This mainly occurred with compression mould shaping.

However, less noticeable surface flows and dents were fonned in autoclave curing

process.

4.2.5 Vacuum Plate and Vacuum Assembly

The application of vacuum is to assist in cornpressing the plies. It is a necessary step in

pre-preg lay up The vacuum pressure provides dual advantage of pressing the laminate

together and at the same time sucking out al1 air pockets and volatile of the laminate.

Indeed, the application of vacuum allows the air and low density resin to flow out Beely

and provides debulking. The vacuum bag assembly consists of bleeder, breather, caul

palate. baggins film, and release film. Each of these components serves a purpose

essential to the tinal reliability of the composite coupon.

The release film covers the composite matenal and provides an easy release once the

composite is cured. Generally the release film is porous or perforated to allow the excess

matrix and air to escape. D 200 TFP. which is a porous, PTFE coated, plain weave fiber-

glass cloth was used as release agent. This cloth is resistant to ail solvent and resin

system normally used for composite rnanufacturing.

Bleeders / breathers, as their names imply, are mats that trap the excess matrix in vacuum

condition and prevent the excess matrix fkom flowing into vacuum pump. The breather

on the other hand, allows the air to travel fieely through the vacuum assembly. D3000-4

stretchable non-woven polyester was used. This materiai is a multi-directionai stretch,

low density. and lightweight construction which do not contain any binder that cm close

offauflow and stick to parts.

D316 bagging film was used to cover the vacuum plate and provide airtight seal by

means of double sticky tape. D3 16 is a green colored transparent heat resistant mylar that

is made from modified nylon 6 resin. It is recomrnendcd for composite manufacturing

and any other high-pressure application where so ftness, high strength and workability

are essential. The physical properties of each component are tabulated below:

Elongation i 1 I Over 300% i

Physical Properties 1 Release Film 1 1

I

Tcnsile Strength i 1 I / 12000 psi I

Bleeder / Breather

Tearing strength I i

j Mas using temperature ; 550 "F

Heat deflection I

Bagging File

9 0 G h l M i l 1

400 "F 1 400 "F i I 1 1 i !

Min using temperature 1 i

1 -80 "F I

1 1

Shrinkige Iess than l0/0 1

1 1

Table 4.1 Data Specification of Vacuum Assembly

I 350 "F 1

' Weight 1 1.9ozIYd , !

Thickness i inch

b

Supplier / De-Cornp 1

4.4 oz i Yd 1

i

j 1

De-Comp / Dr-Comp I

The caul plate provided s t E and fiat surface for the uncured laminate. An uncwd

laminate is very soft and conformable such that it can warp under pressure. The caul

plates sandwich the laminate imrnediately after the release film and breather to prevent

the test coupon 60m bending under pressure. In the preliminary stages of rnanufacturing.

a thin commercial perforated bras was used but the test coupon became warped under

pressure. Even wit h 12 hr of pon curing the results were not satisfactory. Because of this ,

problem a caul plate was made with holes of 1/32 in diameter with 1/16 spacing in

ktween.

S ince the sample is made of sevcral layers stacked together. the possibility of having ai.

pocket and void within the plies is very high. It is crucial to eliminate contamination and

air pockets by meûns of vacuurning the test sample. If volatile exists in laminate. they can

act as stress concentration areas and cause premature failure during testhp. The right

sequence and combinat ion of vacuum baç components are very essential to the separation

of sample fiom vacuum plate and very crucial to the right volume fkaction of fiber and

matri.; at cured state. Several combination of bleeder. breather. release film rnylar. etc

Kas used to determine the best combination for easy release. The fuial stacking sequence

for la'-up that gave the best result is as follows:

1. One breather laver 3 . One copper plate (caul plate)

3. Next breather layer

1. Release agent iayer

5. One release film with silicone

6. The actual test specirnen

7. Second release film with silicone

8. Release agent layer

9. Third breather layer

1 0. Nexi copper plate

1 1 . Last breather layer

12. Vacuum plate covenng mylar

For each sample, once al1 the components are stacked together, the plate was checked for

leakage and vacuumed. It is important to notice that vacuuming process is slightly

different in each fabrication process (autoclaving and compression moulding). Few

problems were encountered in lay-up stage of fabrication and in manufacturing stage.

During the iay up procedure, temperature in the lay-up Iab causes the matrix to become

soft. Because of the soflening of matrix. during the vacuum bagging some of earlier

samples ended up having one edge bent. In addition to the matrix soflening, the caul

plates caused some warping or out of Batness mainly due to their inability to provide flat

surfaces. and lack of stiffness to withstand the vacuum pressure.

In testing the initial samples, it was noted that the release film that came in contact with

the fiber-matrix specimen was not releasing adequately. Three types of releasing films

were tested simultaneously (keep a11 other components the sarne) to find the suitable

releasing film that can produce the ben results. M e r testing al1 possible mylars, it was

found that regardless of the type of releasing agent, a thin mat of silicone was useful and

beneficial to the final surface finish and separating action of the pan.

4.2.6 Cure Cycle

One of the most crucial steps in the fabrication of composite matenals is the post-curing

stage. It is important because there exists a difference in coefficient of thermal expansion

between the individual constituents of composite laminate. the fiber and the epoxy. In

carbon and glass/epoxy composite. the epoxy has higher coefficient of thermal expansion

in cornparison to the carbon or glas fiber. In the cooling stage. this difference causes a

larger shruikage in matrix than carbon fiber as the temperature decreases. This causes

compressive stress on fibers by matrix and tensile stress on rnatrix by fiben. which result

in a warped panel. These intemal stresses can contnbute to the pre-mature failure of test

samples during tensile testing. In fact if the guidelines of manufacturer of the pre-preg is

not followrd closely. the rnatrix can crack due to the residual thermal stress in the

laminate. Post-curing is a process of maintaining a somewhat lower temperature on the

larninatc for 12 ro 24 hours afier curing with the aim of relieving the above rnentioned

therrrial stresses. The temperature is t hen released slowly down IO room temperature.

To achirve the manufacturer recornmended pre-cure. cure. and post-cycle several heating

elemrnts with various heating capacities were teaed. A 1 10 V. 25 A heating elemen~

proved to produce the best-controlled temperature gradient for purpose of fabricating test

coupons based on the manufacturer recornrnendation. To control the curing cycle and to

moniior the progress of the fabrication two prograns were developed on QBASIC. The

functionality of both progrm is the same and both were developed to take the digital

input signal of the A . input output module. process these signals. and respond to a

probable interruption or problem. The firn program operates the compression mould

shaping. This program was developed to operate the necessaq components for

manufacturing a composite panel based on the elapsed tirne. This means that every action

(cg. energking valve for pressurization or shuttuig power to heater) takes place for a

predetermined period of tirne. For instance. in case of pressure &op, the solenoid for

pressurization is energized for 6 seconds. Because each progangs looping t h e was

approxirnately 1 .O minute and the pressure difference AP was set at 5 psi the signal

received by the computer is analyzed in a suficient t h e interval. If the pressure drops

below the set limit. it will energize the denoid for 6 seconds to increase the pressure.

The second program operates the autoclave for fabrication. This program, on the other

hand. controls the control board components based on the signal feed back to the system

This mcans that the pmgram checks the incorning signal and calculates the error. if the

rrror of the system is within the preset value. then the output signal would be zero

othenvise a corrective action would be taken. This system appears to be more efficient

than timr baxd program.

The solenoids operating control valves are 24 VDC and the relays operating the heating

elements are +5 VDC and kll5 VAC. The DC relay communicates with the computer

through Nil modules and the AC relay provides power to the heating element. It should

be mcntioned that each program foilows the curing and most importantly the post curing

part of fabrication very closely with regard to the manufacturer recornrnendation. Thus.

the quality of test samples was not compromised.

The cure cycle of carbodepoy composites is a function of t he . temperature. and

pressure. As per the manufacturer recommendation the cure cycle foi AS4 is as follows:

Recommended pre-cure cycle consists of rarnping the temperature by 3 to 10 "F

per minute. The sample is under vacuum pressure of 20 in Hg (68 Kpa) up to

1 50 "F. No vacuum pressure f?om 200 "F to 360 " F.

O Recornmended cure cycle is 2 to 3 hours at constant temperature of 340 to 360 F.

Recommended pst-cure cycle is I to a maximum of 2 F per minute f?om elevated

temperature to the room temperature.

The recommended cure cycle for pre-preg h i n a t e of glas fiber epoxy

Recommended pre-cure cycle is maximum 10' F per minute. The sample is under

vacuum pressure of 20 in Hg (68 Kpa) up to 200 OF. No vacuum pressure 60x11

200 "F to 360 " F.

Recornmended cure cycle is 1.5 to 3 hours at constant temperature of 340 to 360

Recornmended post cure cycle is the same as AS4 materials.

However. other method of post curing is to oven post cure the test specimen or structural

section for a period of 12 to 16 hours at 250 to 300 "F. In this case. the sample is kept at

elevated temperature for annealing. This process is more recommended for structural

section because of the difficulty involve in shaping a large section or structure (see

appcndix A for both program needed to run compression mould and autoclave.

(0,90,90,0) Cure cycle

Figure 4.6 Actual cure cycle of carbon fiber epoq cross-plied laminate

(0,+45,4,0) cure cycle

3%

à 2 - t ,, t 3 zm t a E pressure c lm

M

3

Figure 4.7 Actual cure cycle of carbon fiber epoxy angel-plied laminate

(O, 90,9û, O] Cure cycle

r CC

I V .--

!CC

:Y

i ix é .. C

Tcnp ' X 0 Pressure

?i

Figure 4.8 Actual cure cycle of Glass fiber epoy cross-plied laminate

(0, +45, -45, 0) Cure cycle 400

Figure 4.9 Actual cure cycle of Glass fiber epoxy Angle plied.

4.3. Testing

4.3.1. Testing Ou tlines

Selsction of test rnethod to determine the mechanical properties of composite materials is

one of the most important parts of teaing since the measurement of strength of these

materials is sensitive to configuration. The testing carried out was in accordance with

testing procedures outlined in ASTM committee D3039 for carbon and D638 for glass

tïbrr reinforced epoxy. This rnethod coven the steps for determinhg tensile properties of

high strength and high modulus reinforced composites. 1 n addition the ASTM calibration

standard for measuring instruments was used. The reference methods are the following:

*. ASTM committee D3039 and committee D 638 for detemiinhg the tensile properties

carbon and glass fiber reinforced materials respectiveiy.

* . ASTM E4 practice for load verificat ion of test ing machine

* . ASTM E83 Practice for verificat ion and classification of extemorneter.

Figure 4.1 0 Test Sample Size and Configuration.

Figure 4.1 1 The kont view of Ten Specimen with Aluminurn Tab

Figure 4.13 The side view of Test Specirnen with Alurninum Tab.

The tension specimen s h o w in figure 4.9 through 4.1 1. should be mounted in a gripper

capable of applying enough lateral force so that slippagr between the gripper and the tab

c m br prevented. The load and strain curve must be ploaed during the test if modulus

propenies are sought. The grippers should be aligned so that the longitudinal axis of the

test sprcimen coïncides with rhr line of action of the applied load. Thus preventing any

out of plane ( secondq) knding.

The ASTM testing standards recornrnends the use of a suitable load indicator capable of

showing the total load carried by the test specimen with an accuracy of + 1%. If the

tensile machine has constant speed. the longitudinal deformation of the system should not

be more than 25% of the deformation of the gauge length. The extension mesuring

system should be ftee of inertia lag with an accuracy of + 1% at the specified speed for

testing. The extension measuring system has to be calibrated as oflen as necessary. This

is to ensure that the reading is true and no error is resulted during the tende testhg due

to the destructive nature of the tests.

Load indicator should be calibrated and verified according to the ASTM E4 practice.

These practices cover the verification for static and quasi-static testing machine. Three

methods of verifications recommended by A S W are the use of standard weights. use of

equal arm balance and standard weights. use of elastic calibration devices.

The speed of application of the load has to be constantly monitored to ensure a constant

strain rate. To monitor the strain rate. a strain pacer should be used which gets attached to

the enrnsometer. The recommended strain rate by ASTM is between 0.01 and 0.02

idin.min. I f control of strain rate is not possible. the constant speed of the cross head of

the test ing machine is recommended.

The test specimrn has to be measured with an accuracy of k 1 %. The measurement for

width and thickness should be taken at several points on the specimen and the srnallen

should be used for calculating cross-sectional area of the test sarnple. The smallest cross-

sect ional area has to be used for calculating stresses. The sample is then to tx placed into

the gripper of the tension machine and secured in place. Care m u a be taken so that the

imaginary mid-line of test sample is aligned with the line of action of the force. The

Extensometer has to be mounted on the edge of the test coupon. Extensometer should not

cause notches on the edge of the samples as it composite materials are highly notch

sensitive.

In order to obtain information on load. stress. elongation, straul and percent strain a

universal tensile testing machine with 60000 Ib. capacity and a MTS impact resistance

extrnsometer with gauge length of ?- .Oh were utilized (the gauge length is variable).

4.3.2 Sample preparation

Both carbon fiber and g l a s fiber Epoxy test samples were made of the pre-irnpregnated

unidirect ional sheet made by Hercules and 3M respectively. The pre-preg materials were

laid up in 4.5 in by 1 Oin size sheet laminate in the desired orientation. A vacuum plate of

2Oin by 4.j in was designed and built to house 2 sheets at a t h e . Eight samples were

fabricated in each curing process. The cured materials were then cut to 1 in by 10 in

coupons. Since the matrix is hard. reguiar cuning machine cm burn the matrix. pull the

fiber out. or darnage the cutter. Special high speed diamond coated bench saw was used

to cut the coupons to the desired size.

Because of the brittle nature of the composites. in tensile testing they tend to break at the

gripping section. Composite materials also are very susceptible and sensitive to notches.

because these notches act as stress raiser. which can cause premature failure. To

rliminate the possibility of failure at the gripping section each test sample was tabbed

accordhg to ASTM standard. Tabs were cut fiom lin aluminum bar with thickness of

1!8in. Each tab was tapered at the leading edge. toward the gauge section to reduce the

spatial load transfer rate. The recornmended size for tab according to the ASTM

standards is 1 in by 2 . h by 1/8in. which was bonded to the gripping section of coupon

Initially a 5-minute epoxy made by DEVCON was used to glue the tab to test coupons.

However. in the first 4 test. the interface bond between tabs and test samples failed which

caused slippage of test sample in the gripper. It was found that the "'two-ton" epoxy could

not provide sufficient bonding arength. FUially the high strength, 3.5 ksi shear strength,

glue made by DEVCON provided suficient strength for carbon fiber epoxy. However,

the *-two-ton" epoxy proved to provide necessary strength needed for g l a s fiber epoxy.

Al1 test samples were tabbed and were left under dead load of 5 Ib for over 24 hr to

ensure complete curing of the glue. To reduce the notches caused by cuning and

manufacturing. the edge of samples were sanded using 200. 400. and 600 grid wet-

sandpaper sequentially. Each sample was then carehlly examined to ensure that the

quality of the samples was not compromised.

4.3.3 Tensile Testing Machine

In this test. a tensile testing machine called "Data Matic" was utilized. The United

Calibration Corporation manufactured the Dat Matic and it contains a complete data

acquisition system. This machine was donated to Ryerson Polytechnic University in 1999

and %as originally designed for manual operation via the use of the knob (potentiometer)

to apply load by rnoving the cross head of the machine. The strain could be calculated by

ushg an eaemal extensometer and the speed of testing was measured by taking the

traveling distance of the cross head of the testing machine and the elapsed time of the

cornputer of the system. A digital indicator on the operating panel of the machine shows

both the position and speed. Because of the age of the computerized control system and

the failure of some key components. it was concluded that the universal testhg machine

was no t suitable for testing composite specimens. See figure 2.13.

DATA-MATIC The Complete Data Acquisition S y s t e m

United Calibratian

Corporation

Figure 4.1 3 Tensile Test ing Machine.

The problerns with this machine were the data collection fkom load ce11 and load indicator

failure. the optical encoder. extemorneter was missing. as well as the feed back system

for controlling the speed and applying load. Ir. order to obtain an accurate data f?om

instrumentation and to store the data into a file for further analysis. the testhg machine

had to be retrofitted. Therefore. the upgrading process included:

The complete replacement of the load control system with a new Pentium computer.

Installation of data collection card for the new computer system.

Construction of a load ce11 instrumentation amplifier for interfacing the new load cell

to the new data collection card.

Calibrat ion of new load cell.

Calibration of extemiorneter output to the data collection card.

Calibration of the optical encoder output to the data collection card.

Construction of a new control panel (the old control panel was discarded)

The initial testing revealed that the A/D convener of the testing machine was fully

îùnct ional. However. data acquisitions card. a signal amplification unit. and an electronic

filter were needed in order for the machine to comply with ASTM standard. A data

acquisition çard with 1 O charnels was used for receiving and storing data. The ASTM E4

standard calibration specifies the standard dead weight as the calibration method. Along

with the standard weight. another calibrated load cc11 was used in series with the tensile

machine's load ce11 to ensure the integrity of the process. Because a new load ce11 and

interface card was used. the load ce11 needed to be calibrated. This was accomplished by

applying a bown load to the load cell, the load ce1 then produces an m l o g output that

eets transferred to the AID convener. The digital output of the N D converter was read as C

a 12 bit word that was recorded. For each standard known load therefore. one digital read

out was recorded. The value of the applied load was known by recordhg the output fiom

the calibrated load ce11 that was attached to the sarne test cell. As the graphs show. the

linear regression was used to relate the digital output to the known load by mcasuring the

slope of the line and the X-intercept. This information then was used to write a sub-

routine that indicates the actual load in the unit of pound withh the operathg p r o h m

Load CeII Calibntion for Tensile Machine

Figure 4.14 the calibration curve of the Independent load ce11

- - -. -- - - -- -- Independant Lœd Ceil Tensile Machine Calibraüon

Figure 4.15 The calibration C w e of the Tensile Testing Machine.

A test was carried out with the same set up by hanging a precision dead load of the load

cross head of the tensile testing machine. The values of the independent load ce11 and

machine load ce11 were compared againa the dead load value. It was determined that the

machine load cell was in a very excellent agreement with the independent and precision

dead load. Once the load ce11 calibration was completed. a strain indicating unit

(enensometer) needed to be adopted. An MTS variable gauge length extensometer with

universal digital indicator was utilized. This extensometer has + 0.2500 in and - 0.2000 in

range at 1.000 in gauge length. The analog output of the indicating unit was used as an

input signal to the A/D convener of the testing machine. The signal was then used for

determination of strain and data were put into a file for further analysis. The fust step was

to calibrate the extensometer. The calibration was performed using a super micrometer

made be Praa & Whitney Canada capable of reading 0.0000 1 " (ten millionth of an inch).

The extensometer was mounted on the micrometer and the reading on micrometer was set

to zero. Then the extensometer was tested over its full operating range in 0.002 in

increments.

The optical encodrr of the machine was used to rneasure the displacement of cross head.

The distance traveled in conjunction with elapsed time gives the spced of testing. The

speed of testing according to ASTM should be such that the strain rate would be

approximately 0.0 1 to 0.01 idin min. The encoder was connected to the Lad screw of the

iesting machine directly via a pulley. It was measured that the encoder pulses 400 tirnes

per one full rotation of lead screw of the machine. Therefore. a subroutine was written to

take the pulses t o m encoder and divide it by t h e registered from cornputer tirner. The

output of this subroutine was compared with the recomrnended speed for testing. If the

calculated speed is high or low. the program would take the corrective action. Two safety

features were used in retrofitting tensile machine. One was a series of routine check

within the program for the maximum speed. load. and strain And the second was two

limit switches mounted on the two extreme ends the fiame. These switches were wired

inro the motor circuit so as to hnction as hardware safeties. Copy of the program is in

the appendiv A.

The available grip for this machine was a pair of self-adjusting wedge type grip. In early

stages of the testing. it was established that the bondinp between the tab and test coupon

uas failing bcfore enough lateral force was k i n g established between the specimen and

the grip. Thus slippage between grip and test specimen was inevitable. To overcome the

possibility of slippage. a set of mechanical grip was designed. see figure 4.16. This grip

was designed specifically to take a 0.053 in ihick specimens with 0.25 in total tab

thickness. Howevrr. the thickness of the samples could Vary and result in misalignment

of the crnterline of specimens and the line of action of force. which cause bending as

well as trnsile stress. It was concluded to use a bal1 and sockrt assembly to connect the

erip to the cross-head of the machine to compensate for the variation in thickness. Since C

the variation in thickness of the test samples were in the order of 0.000 1 to 0.000 15 in.

this arrangement tumed out to be the best solution.

To ten the grips according to ASTM standard. an aluminum sample with the same

configuration was mac hined. Three strain gauges were laid up on it. Two were laid upon

one side beside each other and parallel to the longitudinal axis of specimen. The third

strain gauge was laid up on the same location as two others but on the other side of the

specimen. Then the specimen was tested under uniaxial stress at a 180 N. increment up to

the maximum load of 2670 N. and the strallis of al1 three gaiiges were recorded. The

measurements of strains of al1 three gauges were almost the same within the accuracy of

strain gauge indicator see the figure 4.1 7% b and 4.1 8.

Figure 4.16 The gripper.

Figure 4.17(a) Calibration specimen with strain Gauge

For Machine's Cross head Inspection

Figure 4.17 (b) Calibration Specirnen with Strain Gauge

For cross Machine's head inspection

Figure 4.18 The System Used for inspection

Of Mac hirie's Cross Head.

Once the universal testing machine was retrofined in al1 aspects and the grip and other

auxiliary components were rnachined and tested. a new irnproved program was required

to operate the machine. Two programs were written for tensile testing purposes. These

programs were designed to rake the dimensions of the samples, the maximum expected

failure load. the maximum expected failure percent s t rah and the gauge length. Since the

extensometer was mounted on the edge of the specimen and the fact that the specimens

were thin the visual monitoring of strain rate became crucial. This would give the ability

to see on the monitor screen if there is a sudden jump in the strain or load. The expected

maximum load and percent nnin were needed to divide the screen accordingly for stress

strain curve.

The tirsr program was to test the specimen to failure in tensile. Basically the operator

would be asked to input the geometry of the test sample and the gauge llength then the

stress-strain cunre would be drawn on the screen sirnultaneously as the load king

applied. The load. stress. elongation strairi. and percent strah would be d o ~ n loaded to a

data file for tùnher analysis. ïhe second program was developed to be utilized for taking

rdge replica from the test specimens. This program is capable of loading a test sample to

a specified load and back to zero and pause. Then the operator has enough time to

analyzr the edge of the sample. .4nd with the stroke of a key. the load would be

incremented to the next specified load. At the end of the test. when the failure has

occurred. al1 the information would be stored into a data fde. See the appendix for a copy

of each program.

CHAPTER 5. NEW TESTING METHOD.

5.1 Introduction

This chapter explains the systematic method adopted here to find effective propenies.

The main emphasis of this chapter is to introduce the rule of mixture as an effective

method to find the transverse stiffness knowing unidirectional 0" and laminate stifiess.

The results of first ply failure obtained using edge-replica technique along with classical

lamination theory are used to find transverse strength.

5.2. Transverse Modulus of Elasticity E90°

It has been argued effectively in reference[3] that there is a need to replace the existing

procedure to obtain the transverse Young's modulus E2. Here, a simple procedure for

determinine this properiy will be outlined. It is proposed that a cross-plied Iaminate

longitudinal modulus (Ex) be expenmentally measured on a coupon specimen. It is then

assumed that Ex can also be approximately evaluated fiom a simple rule of mixture that

incorporates the volume ratios of the 0" and 90" larniae in the specimen:

Thus:

Es = V1 ( E 1) + V2 (E2 )

By rearranging the equation I

Where:

E2 = ( d - V i ( E i ) ) l V 2 -----.--------- ( 2 )

Ex(%> is the longitudinal Young's modulus of 0190 symmetric

El is the longitudinal Young's rnodulus of O"

Et is the longitudinal Young's modulus of 90"

V 1 is the volume fiaction of 0-degree fiber.

V? is the volume fiact ion of 90-degree fiber.

The Standard specimen proposed here is an eight layer (O0. 90'. 90'. OO)s coupon. Using

this configuration equation (2) yields the following expression for evaluation Ez.

E2=2E, - E l ----------------- ( 3 )

The important task became that of fmding the respective longitudinal Young's modulus

E, for cross-plied laminate.

5.3. Longitudinal Tensile Strength X

The cross-plied specimen will again be used here to obtain the longitudinal tensile

strrngth of fiber-reinforced laminae. This value is in fact obtained in a straightfoward

manner by loading the specimen to failure and s h p l y dividing the failure load by the

total cross section of the O" laminae. This derives ffom the assumption that the 90"

lamhae would have completely failed prior to the ultimate failure of the specirnen.

Hence. they do not contribute to the structural strength of the laminate. In fact. it is

argued here that their contribution is a functional one in that they act as amesters to any

initiai cracks that rxist in O* laminae. thus enhancing their strength. The assurnption that

the 90 layer would have filcd prior to ultimate fa i lw also signifies that the Stress state in

the tarpt O layer is pureiy uniaxial. This eliminates the need to use any correction

factors. as may be the case for the transverse failure strength Y discussed in the previous

section.

5.4. Transverse Ultimate Streagtb

There is a considerable difference between the transverse tensile arength of an

unconstrained unidirectional 90" lamina and tbat of constrained one. This difference is

easily explained by the notch sensitivity of the 90' lamina when tested alone. As has been

explained in many studies. reference [5 ] . srna11 initial microcrack in a 90" lamina will

propagate and cause rapid failure unless anested by fibers in adjacent layers. which is

what happens in a multidirectional laminate. Obviously, the strength of the 90" lamina

alone becornes of little usefulness and should be replaced by the effective strength of the

constrained lamina. As mentioned in reference(51. such practice has aiready k e n

adopted by at least two different organizations. These organizations opted to constrain

the 90' lamina with O" lamina Eom a more ductile material or with t 45' laminae from

the same material. The rational behind using these types of constraints is to allow the

target 90" layer to develop is strain limits. However. either procedure enhances the out of

plane stresses in the specimen. which may eventually lead to unwanted edge

delamination influenced failure. It is argued here that there is no need to artificially

enhance the strain in the target lamina since in real situations those strahs will be likely

limited by the presence of O' layers of the same material anyway. It is therefore proposed

that the same specimen described in the previous section be used for evaluating the

transverse tensile strength Y. A simple procedure to obtain these properties would be as

fo Il0 ws:

A) Subject the specimen to a monotonie load

B) Monitor the longitudinal modulus Ex given by cm / E

C) Rerecord the membrane load (N/rnm or l b h ) at the point where the lamina

longitudinal modulus ox / EX drops to the modulus of an equivalent laminate with

negligible modulus for 90 lamina i.e.

EX = V I (El)

The recorded load would then be used in conjunction with classical lamination theory to

determine the stress in the 90" laminae. This stress is eventually considered as effective

transverse failure stress Y of a unidirectional lamina. It is important to note at this stage

that the state of stress determined in the 90" lamina based on the above described

procedure is not purely uniaxial. In fact for materials showing very high level of

aniso trop? signi ficant compressive stress component will develop in the 90" lamina

parallel to the fiber. In such cases a correction factor may need to be applied to the value

of Y obtained with the aforementioned procedure. Altemtiveiy, it may prove to be more

appropriate in these cases to obtain this property using a specimen capable of taking a

biaxial load. thus offsetting the negative stress in the 90" layer. This however is beyond

the scope of the present work.

5.5. Edge Replica

The edge-replica technique. as its name implies, replicates the pattern on the edge of the

sample on a film where any crack formation on the edge can later be examined. The

rdge-replica procedure was camed out in the foilowing manner. A cellulose acetate film

was attached to a conformable rubber pad (in this case a cut-off section of standard

eraser). At each load interval. with the aid of eyedropper, two or three drops of acetone

were applied to the edge of the test specimen. By pressing the cellulose acetate film to

the edge of the test sample for approximately 15 to 20 seconds. acetone would imprint the

detail of e d p of the sample ont0 the thin film. Tensile test was carried out in an

incremental manwr with an initial load of 500 Ib. Then the load was incremented by

3001b and back to zero. At each step, the imprint of the edge of the sample was

replicated and recorded. This procedure was repeated until the test specirnen failed under

the load. Any micro-crack fonning in a layer other than O" (mostly in matrix dorninated

layrr) would be detected by making an imprint of the edge. Edge-replica procedure was

carried out on two samples to establish the failure load and consequently the failure stress

and strain of 90' layer of carbon fiber and glass fiber epoxy. An additionai Edge-replica

was taken to check and establish the failure stress and strain of 90" and 45" layers when

they are laminated with 0' at the onset of crack. Edge replicating was used for test

specimens laminated to form a quasi-isotropic laminate and to conform the validity of

cxperimental data.

CHAPTER 6. EXPERIMENTAL DATA AND DISCUSSION

6.1. Introduction

Chapter 6 introduces the results of experimental test perfomed on different test

specimens. The average values of the experimental results are calculated and each result

has been tested for their validity using statistical means. Also the explanation for

discrepan cy between experimental and anal ytical is provided in detail.

6.2 Young's Modulus E

6.2.1 Young's Modulus Ex of Carbon fiber Epoxy (AS4f350)

Edge-replica technique to detemine the furt ply failure load was the first step in finding

longitudinal Young's modulus Es. It was determined that the fmt crack occurred at stress

beween 15 1 -7 and 158.6 GPa in 90" layer resulting in a noticeable change in the modulus

E, Therefore. for al1 intents and purposes, 151.7 GPa was taken as laminate failure

stress and the Modulus Es aas calculated using data corresponding to this stress.

Figure 6.1 Edge replica (x 150) of a Carbon fiber epoxy cross-ply symmeûic test specimen. Failure of rnatrix occurs in between differently orienteci

layer due to monotonic Load Fx

Figure 6.1 depicts the results of replicated edge taken fiorn a test cmied out on (0, 90,

90. O), cross-plied carbon fiber epoxy test specimens. Once the failure stress level of

15 1.7 GPa was determined for the 90" layer of carbon fiber epoxy cross-plied laminate.

ten specirnens were tested to 15 1.7 GPa iimit stress to fmd the longitudinal modulus of

elasticity. Table 6.2 is the surnmary of the experimental data. The longitudinal Young's

modulus was determined using the stress strain relationship (Hooke's Law) assurning

rlastic behavior up to the fust ply failure point. The method of Least Square was utilized

to calculate Ex using straight iine cuve fitting. In order to use the mie of mixture the

modulus of elasticity of 0" was needed. Eight (8) identical specimens of unidirectional O"

were manufacturrd using the very same manufacturing set-up that was used to fabricate

other angle-plied test coupons. This was done to ensure that the effects of curing on

mechanical propenies of al1 samples are taken into account since cure cycle has an

important trffect on mechanical properties of composite materials. These specimens were

made up of 6 layer of O". which complies with ASTM committee D-3039. Table 6.1

suinmarizes the result of the tensile test of unidirectional test specimens.

Longitudinal Young's Modulus EOO (GPa) 121.20 122.31

Table 6.1 The 0' unidirectional modulus of elasticity of AS4-350 1

Starisricallrl

Young's Modulus

Standard Deviat ion

E,= 125.4 GPa

E,,= 18187146 psi

S = 3.51 GPa (409706 psi)

Ilongitudinal Young's Modulus E, (GPa) 1

Table 6.2 The modulus of elasticity of cross-plied of AS4350 l

Smfi.~r icd!\-

Young's Modulus

Standard Deviation

Es ,,,,= 70.79 G Pa

Ex,,,,= 10266627.45 psi

S= 3.2 G Pa (4680 14 psi)

With 9g0/o confidence. the population means are between 125.388 GPa - l25.40? GPa

and 70.809 GPa - 70.740 GPa for 0" and cross-plied laminate respectively. In other word.

the average value of Modulus of elasticity of 0" and cross-plied laminate were found to

be between the above range that is the range of population's mean. To rely on the

experimental result. the Goodness of Test result was checked by means of natistical

method. To examine the validity of the results, the ratio of the deviation to the standard

deviation of tabulated data for aforementioried samples was given as 3.25. Thus. if the

ratio of deviation and standard deviation is srnaller than 3.25 for each sample, the

experimentally found value is an accepted value. This process proved that al1 the

cxperimental values are within the statistically acceptable envelope. Therefore, the

average value for modulus of elasticity for 0" was found to be 125.4 GPa ( 1 8.19 Msi) and

average modulus of elasticity of (0. 90. 90. 0)s was found to be as 66.74 GPa (10.27

Msi). Thus. by detennining the unidirect ional 0' and cross-plied elast ic properties. the

e f ec t ive transverse Young's Modulus c m be found as following:

E90 = 2 (70.79) GPa - t 25.4GPa

Ego = 16.2 GPa (2.35 Msi)

Published value for unconstrained Young's modulus is:

Eso = 9.5 GPa (1.38 Msi)

O degree AS4-3501

Figure 6.2 Stress-grain curve of unidirectional O" of AS4-350 1

Figure 6.3 Stress-~train curve of unidirectional O" of AS4-350 1

Figure 6.4 Stress-arain curve of cross-plied of AS4-350 1

Figure 6.5 Stress-~train curve of cross-plied of AS43 50 1

Figures 6.1 and 6.3 show the typical stress vs. strain curve for 0" laminate rested

according to the guide lines as outlined in ASTM handbook cornmittee D30-39. Figures

6.4 and 6.5 show the stress vs. drain curve for the cross-plied laminate. The linear curve

fittinp using method of Least Square was used to €id the slope of the line.

6.2.2 Young's Modulus of E, Glass Fiber Epoxy (Type 1003)

The fust step in fmding longitudinal Young's modulus was to examine edge-replica

result and to determine the fust ply failure load and strain. It was experimentally obtained

that the fun crack was initiated at approximately 37.92 MPa stress in the 90' layer and a

subsequnit change in the longitudinal Young's Modulus EL Therefore, 37.92 GPa was

taken as the laminate failure stress and the Modulus, Ex, was calculated using data

correspondhg to this stress. Figures 6.6 and 6.7 depict the stress strain curve of two

moss-plied test specimens. At the stress value when fyst 90" layer crack occurs, the dope

of the line changes slightly up to approximately 0.5% strain whae the stiffiess changes

more drastically. However die slope at each section is linear, hence of piece wise linear

andogy is suficient for calculation. As figures 6.7 show, the &op of stifiess in Glass

fiber epoxy is more noticeable and drastic than Carbon fiber epoxy.

Figure 6.6 (x 150) The Edge replica of a Glass fiber Epoxy cross-ply syrnmeîric test specimen. Failure of matrix occurs in benveen differently oriented

layer due to monotonic Load F,

Figure 6.6 shows the results of the edge-replica of a (O, 90, 90, O), test specimens. As the

fi_rmre 6.6 depicts, the fmt crack had been fomed in 90". After determinhg the failure

stress level of 90' at the on set of crack, eleven specimens were tested to the 37.92 GPa

limit stress in orda to fmd the longitudinal modulus of elasticity. Table 6.3 is the

summary of the experimental data of cross-plied test coupon. Although the mechanicd

properties of glass fiber is dflerent than carbon fiber. to find the Young's Modulus the

same process (rule of mixture) as for carbon fiber epoxy was used.

4ooooo00 Cnss-Pîied Glass Type 1003 I

35000000 .

Figure 6.7 Stress-strain Curve of cross-plied of Type 1003

Cross-Plied Class Type 1003

Figure 6.8 Stress-strain Curve of cross-plied of Type 1003

ILong itudinal Young's Modulus Es (GPa) 1

Table

Stut isticull is

Young's Modulus

Standard Deviation

6.3 The longitudinal Young's Modulus E, of Type 1003

Ex,,,,= 25.61 GPa

E, ,,,,= 37 1430% 1 psi

S= 871 MPa (126316 psi)

With 99% confidence. the true means of population lies between 26.60 and 26.61 GPa.

.As the csperirnental values of cross-plied laminate shows. the average value is within the

population mean value. Therefore. the average vaale for modulus of elasticity Ex was

found to be 25.6 GPa (3.71 Msi). Similar to carbon fiber epoxy. data for glass fiber epoxy

was tested for Validity of result. The ratio of deviation to standard deviation of this set of

test specimens for 1 1 samples was f o n d to be 25.6 GPa The ratio of deviation of each

value to the standard deviation was found to be less than 25.6 GPa and therefore. the

samples were nat ically acceptable.

Using equat ion #3 resulted in the fo llowing:

Et = 2 (25.6) GPa - 39GPa

E2 = 12.22 GPa (1.77 Msi) Experirnental

Published value for unconnrained Young's modulus is:

E2 = 8.5 GPa (1.23 Msi)

6.3 Longitudinal Ultimate Strength

6.3.1 Longitudinal Ultimate Effective Streogth of Carbon Fiber Epory AS43501

To determine the ultimate çtrength of the AS4-350 1 laminate. twenty test specimens were

teaed to failure while the maximum loads king recorded. At the ultimate load, the 90"

layers in the laminaie would have failed and no longer would contribute to the overall

strength of the laminate. The ultimate failure loads coupled with cross sectional area of 0°

were used to calculate the ulrimate strength X. The data obtained in the experiment is

tabulated below. Figures 6.9 represent the stress vs. strain curve of cross-plied test

specimens. In addition to the cross plied test samples. 8 samples of 0 unidirectional

specimens were tested to find the ultimate strength of unidirectional laminate. The result

of this test oniy provided value so that the result of cross-plied laminate cm be compared

against

O 0002 0004 a m om O 01 O 032 O 014 Stnin

Figure 6.9 Stress-strain curve of cross-plied of AS4-350 1

Table 6.4 The ultimate strength of cross plied of AS4-3501.

.Ultimate cross-plied Laminate Stress (MPa) 752.5 1

siut i.siicalli.

Longitudinal ult imate strength

UltUnate Stress in O" layer (ma) 1505.0

Standard Deviation

X,,,,=1510 MPa

X,,,, = 2 18704.6 psi

S= 114 MPa (16605.7 psi

- -- -- -

mimate unidirectional laminate O" stress ( M D ~ ) 1

1 1950 1 Table 6.5 The ultimate Strength of unidirectional of As4-3501.

Sraristically

-Longitudinal Ultirnate Strength

Standard Deviat ion

a= 1950 MPa

a= 282757.14 psi

S= 3667.74 psi (25.3 MPa)

As table 6.5 indicates. the average uitimate strength of 0' was found to be 1950 GPa.

With 99% confidence. the population means is between 1918 GPa - 1980 GPa for the

ultimate strength 0". It can be said that we are 99% confident that the true population

means strength is being 191 8 GPa - 1980 GPa. Furthemore. the average longitudinal

strengh X of a cross-plied Iaminate was found to be 15 10 MPa (2 18.7 ksi). Using 99%

confidence limit. the me population ultimate strength of cross-plied laminate was found

to be between 1579 and 1434 MPa.

6.3.2 Longitudinal Strength of (0, +45, -45, O ) , X4S of Carbon Fiber E p o q AS43501

As it was mentioned. the purpose of this work was to predict the effective or in situ

propenies of laminate using simple. accurate. and ushg least rigorous method that does

not invo lve testing unrealistic number of test coupon It was assumed that in the case of

cross-plied laminate the 90" layer would not contriiute to the strength of laminate once

multiple cracks is fidly formed in these layers. This assumption should hold tnie for uiy

angle-plied laminate. To examine this assumption 24 angle-plied (0. +45. -45. O), of

Carbon fiber epoxy laminates were fabricated using the same set-up and curing process

as that of cross-plied. which was outlined by manufacturer. These t e s coupons were

teaed under quasi-aatic and monotonie load condition to failure. From 24 sarnples, 3

samples failed at the gripping section and 21 fàiled at the gauge section. Ody the data

frorn samples that failed at gauge section were used in the analysis. Table 6.6 shows the

results.

U lt imate ang le-plied laminate stress (MPa) Ultimate Stress in O" layer ( M ' a )

Table 6.6 The ultirnate strength of angle-plied of AS4350 1.

Sror isr icallv

.-\veragr longitudinal ultimate stress

Standard deviat ion

a = 1550 MPa

o = 225 Ksi

S = 23.3 MPa (3374.8 Psi)

As data indicate. the ultirnate strength of ( 0. +45, -45. O )z laminate is 1550 MPa (225

Ksi). It is seen that the difference between the cross-plied ultirnate strength and angle

plied is wry small about 2.2%. Figure 6.10 shows a stress-grain c w e of (0, +45, -45,O).

-

800000000 Angle-Pied As43501 j l

@ 0.002 0.004 0.006 0.008 0.01 0.012 Strain

- - - - - - - - - - - - - -

Figure 6.1 0 The stress-strain curve of angel-plied of AS4-3 50 1 .

6.3.3. Longitudinal Strength of Quasi-lsotropic Carbon Fiber EpoxyASJ-3501

To fùrther examine the validit y of effective elastic propenies found experimentally. eight

8 samples were fabricated with the same curing cycle. These coupons were then tested to

failurr and results were compared against data obtained by testhg cross-plied coupons.

As the table 6.7 shows. the tende ultirnate strength of Quasi-isotropic is substantially

hig her then other sample tested.

Ultimate quasi-isotro~ic laminate stress (MPa) I~ltûnate Stress in O" laver ( MPa) 1

I 1830 1 Table 6.7 The ultimate strength of quasi-isotropie of AS4-3501.

Siah! icailv

Average longitudinal ult imate stress

Standard deviation

o = 1830 MPa

o = 265.7 Ksi

S =22.7 MPa (3290.3 psi)

Figure 6.1 1 The stress-strain curve of quasi-isotropie of AS4-3501.

6.3.1. Longitudinal Ultimate Effective Streogth X of Ghss Fiber E p o q Type 1003

To determine the ult imate effective strength X. fourteen cross-plied specimens were

teaed to failure and the maximum loads were recorded. As it was explained. then the

90" layers would not contnbute to the ovenll strength of the laminate once the complete

failure of 90 has occurred. Foiiowing the same procedure as for AS4, the u h h t e

longitudinal effective strength of Glass fiber epoxy was calculated to be 73 1 MPa. The

experimental data obtained in the experiment is tabulated below. As figure 6.12 shows.

the change in stifhess or rnodulus of elasticity (slope of the line) occun at 37.9 MPa

stress with a correspondhg strain of .O0225 mmhrn. It can be seen that the change in

siope of the stress-strain curve is more noticeable than that of AS4.

Table 6.8 The ultimate strength of cross-plied of Type 1 003

Ultimate cross-plied laminate stress (MPa) 354.3 7 366.35 367.57 369.65 363.70 367.50 3 54.29 365.55 363.90 366.92

Average Longitudinal ult imate strenph

Ultimate Stress in O" layer (MPa) 709.1 732.7 a

73 5.2 739.3 727.3 73 5 .O 708.6 731.1 727.8

l

733.8

Standard Deviat ion

X,,,, =73 1 M Pa

x,,,, = 105994.16 psi

S= 10.7 MPa (1555.1 psi)

Thus the average longitudinal strength X of a cross-plied laminate was f o n d to be 731

MPa (-1 O6 ksi). The published average strength of O" is recorded as 965 MPa. The

longitudinal ultimate strength X wwas found to be lower than the published data as it was

for AS4-350 1.

Figure 6.12 The stress-~train curve of cross-plied of Type 1003

6.3.5. Longitudinal Strength of (0, +45,45, O), Xdg of Class Fiber epoxy Type 1003

It \sas found earlier that the interaction between glass fibers and epoxy is slightly

different corn the interaction between carbon fiber epoxy in that there is no fiber

hardrning in case of g las fiber. To examine the fact that the longitudinal ultimate

strengt h is the same regardless of the orientation of other layen than O", 14 angle-plied of

tjpe 1003 samples were fabricated with regard to manufacturer curing process. These test

coupons were tested under quasi-static and monotonie loading condition In the case of

Glass fiber epoxy. 12 samples failed at the gauge section thus the data obtained kom

these were used for analysis. The results of 12 test are summarized in table 6.7.

-

Ultimate angle-plied laminate stress (MPa) Ultimate stress% O" layer (MPa) 393.28 786.6

Table 6.9 The ultimate strength of angle-plied of Type1 003

Srur isr icallï

Average longitudinal ultimate stress

Standard deviat ion

o = 756 MPa

o = 109.6 Ksi

S = 17.1 MPa (2478.6 psi)

The ultimate strengh of ~ l a s s Fiber epoxy angle-plied was determined to be

approximately 755.7 MPa (109.6 Ksi). The experimental result is in an excellent

agreement with the result of cross-plied laminate. Figure 6.13 shows the stress strain

curve of angle-plied laminate. As the c w e depicts. the shape of the c w e can be defmed

as piece wise linear.

35000000 Angle-Plied Clas s Type 1003 1

O 0.005 o.or Strain o.ors 0.02 0.025 -- -- - ---

Figure 6.13 The stress-strain curve of angle-plied of Type 1003

7.3.6. Longitudinal Strength o f Quasi-lsotropic laminate of Glass Fiber e p o q

As for carbon fiber epoxy. eipht (8) glass fiber epoxy samples were fabricated under the

same condition as other glass fiber epoxy test coupons. These test coupons were then

tested to failure to determine ultimate strenpth. The results are show in table 9 and

figure 6.14 shows the behavior of stress vs. strain of quasi-isotropie laminate.

IlTltirnate auasi-isotro~ic laminate stress (MPa) Ult imate Stress in O" layer ( MPa)

Table 6.10 The uftirnate strength of quasi-isotropie of Type 1003.

Average longitudinal ultïxnate stress a = 955.6 Mf a

o = 138.6 Ksi

Standard deviation S = 20.09 MPa (2914 psi)

mooomo.m Quasi-Isotropie Glass Type 1003 I

Figure 6.14 The stress-~train curve of quasi-isotropic of Type 1003.

6.4 Transverse Ultimate Strength Y

6.4.1. Transverse Ultimate Strength Y of Carboa Fiber Epoxy (AS4 350)

.4ccording to the method outlined in section 5. the cross-plied specimens were subjected

to a rnonotonic load. The average membrane loads were recorded for the use in

conjunction with Larninate Theory to calculate laminate grain. To calculate the larninate

strain the larninate stiffness was used in conjunction with applied load.

Knowing the laminate nrain and lamina stifhess matrk. the lamina maximum strength

was calculated.

Therefore. The transverse strength Y was calculated to be 35.6 MPa (5168.2 psi). As the

experimental data indicates. the ultimate strength of unidirectional laminate is published

as 55 MPa ref1231. which is greater than the effective value obtained fiom cross-plied

spec imens.

6.4.2. Transverse Ultimate Strength Y of Glass Fiber Epoxy Type IO03

AS for Carbon Fiber epoxy. the mid plane strain of the laminate was calculated ushg

Classicai Lamination Theory.

And the ultimate transverse strength is

As ihe calculated lamina strength shows. the strength is approximatriy 19 MPa (1.2 Ksi ).

But the published data for ultimate transverse strength is given as 35 MPa req231.

Clearly it can be seen that transverse strength is reduced as a result of being plied by O"

6.5 Discussion of Results

As it was shown in the previous section the result of rule of mixture indicated that the

unconstrained transverse Young's Modulus is srnaller than the effective (constrained)

one. One reason is the fact that O" oriented layers are acting as crack arrester and stops the

propagation of crack. Another reason for the increase in Modulus of Elasticity is the

strained hardening of O" layer on laminate. The interaction of O" wit h the laminate lirnits

the strain of 90" layer and therefore. 90" layer would undergo lower strain level than

unidùec t ional specimens for a given load. There fore. this phenornenon causes an

increase in modulus of elasticity. The reinforcing carbon fibers are strain hardened as

the- undergo stresses. The work hardening of fibers plays an important role in increasing

the stifiess Ez. To find a quantitative measure of hardening of 0". the data obtained Born

O" test and cross-plied were compared. This analysis showed that d e r relatively small

stress. the cross- plied laminate started to strain higher than the unidirectional test

specimen. The difference in strain between the cross-plied and unidirect ional increased

by about 0.09% to the point of fust ply failure. However. since progressive crack in

rnatrk dominated layer or stiffhess reduction to multiple cracking is imminent. the arain

diffierence continues to increase up to the point of complete failure of off-axis layer. From

this point on. the nrain of unidirectional and cross-plied larninated was the same taking

into account the fact that actual stress in cross-plied is twice as the data obtained f?om

test. In other word the stress was calculated based on the appiied load and the entire

cross-sectional area of the sarnple. When the 90" fail and no longer contribute to the

strength. only O0 layers provide strength for the test sarnple. Figure 6.1 5 and 6.16 show

the work hardening effect of carbon fibers on transverse s t f i e s s of cross-plied laminate.

Work harùeniog effeît of unidirecfiod SmX3

m

24[31300

nam

Figure 6.15 Work hardening curve of carbon fiber epoxy

Wrkhanlenig effect o f m e

Figure 6.16 Work hardening cuve of carbon fiber epox?

After the fust crack uiitiated in 90" layer. the strain difference increases at a slower rate

until the point of complete failure of off-auis layer. Nevertheless. when the mess is

Uicreased by factor of 2. the stress-main relationship of cross-plied laminate becornes the

same as that of unidirectional laminate within the accuracy of eqerimental results.

Therefore. it can be deduced that work hardening of fiber causes an Uicrease in transverse

Young's Modulus by hardening the laminate.

In the case of G l a s fiber epoxy. the Young's Modulus E2 in constrained statr (effective)

was also calculated to be higher than the unconstrained modulus. This is again the resulr

of the interacrion betwern O" oriented Iayers and the 90" oriented layers. This

phenornenon can be explained by looking at the 0" layer that hnctions as crack arrester

when it is plied with 90". G l a s fiber epoxy is not known to have work hardening effect in

fact some researchers have clairned that g las fibers loose their hardness as a result of

annealing when the laminate goes throuçh curing process reference[21]. Therefore. the

reason for the increase in transverse Young's modulus is the interaction of O" with 90"

when laminated. Thus. the longitudinal decline. as expected. in a symmetric laminate was

exptximentally determined to appear at strain of 0.22%. A sudden increase in strain and a

drop in stifthess did not appear up to about 0.45% strain. As figures 6.7 and 6.8 show.

afier 0.45% arain the change in strain and stifbess was noticeable.

As the data Born expriment in tables 6.4 and 6.8 shows. the longitudinal strength of

cross-plied laminate is less than the unidirectional one in both carbon and g l a s fiber

cpoxy cases. Xab. = 15 10 MPa and pub!ished value is 1950 MPa also Xdw = 73 1 MPa

and published value is 965 MPa reference[23]. The reduction of longitudinal arength

between unidirectional and cross-plied can be explained by looking at the contribution of

90" layers. AIthough it was argued earlier that 90' layen are not contributing after the

complete failure stress lirnit has reached. the 90" layers have a fùnctional contribution.

I t traditionally was assumed that al1 fibers have the same propenies. but in fact the

strength follows a statistical distribution. Contrary to the high arength of fibers. some

fibers are expccted to break at a very low stress state. The fact that the 90" has failed

causes stress concentration at the crack tip. Mien a crack initiates in 90' layer. it

propagates fast parallel to the fiber because the matrix is highly brittle. The Crack is

arr-rested when it reaches the 0". since crack does not travel across the 0" oriented fiber.

The stress concentration at the crack tip near the fiber causes the local debonding of

matriv and fiber interface. Since the fiber failure follows a statistical distribution, fiber

failure occurs locally. As the fibers are failing in a progressive marner. the local stress

grows larger and eventually the laminate failure occurs at a lower stress than its average

stress limit.

.As the figure 7.3 depicts. the laminate in tensile stress typically fails paraIlel to the fibers.

The fact that cracks grow parallel to the fibers for polyrneric matrut is a clear indication

of a high degree of orthotropy. The transverse strength of composites is very low in

cornparison to longitudinal strength that even in O" degree unidirectional laminate the

cracks run parallel to the fibers and cause a local debonding.

Figure 6.17 Failure of Multidirectional & unidirectional carbon fiber epoxy Laminate

Figure 6.1 8 Failure of Mult idirectional glass fiber epoxy laminate

The results of (0. +45. -45. O)? tests. AS4 & type 1003. see table 6.6 and 6.9. proved that

when any layers other than O" layer fail. they would not contribute to the strength and

stifiess of the laminate. The edge-replica process has also shown that the crack initiated

in matrix dominated area happened at approximately the same stress level as that of

found in cross-plied laminate. This was the case for both carbon fiber and glass fiber

Epoxy composite larninate. However. fkom the rate of change of strain after formation of

fvst crack. it also conciuded that the failure progress in matrix-dorninated area is a

function of orientation of the fiber and the orientation of the fiber in adjacent layer. In

othrr word. in cross-plied laminate the rate at which total failure of matn?< dominated

aren occurs is different than the failure rate of angle-plied laminate even thocgh the

initiation of first crack happens at the same stress values in both laminate.

If the layer 0" is the contributing layer to the ultirnate strength it c m also be said that the

strain of larninate at failwe would be the same as unidirectional strain. It tvas recorded

that cl = 1.1 5 % (.O 1 15 mdrnrn) for AS4 and EI = 2.5 % (0.025 mmimrn) for Glass fiber

1003 type. El . E:. and the maximum unidirectional strain in conjunction with Classical

Lamination Theory wrre used to determine the maximum strength of angle-plied

larninate. This calcularion was performed to fud the accuracy of the i n t ~ s i c effective

propenies found in this work. The value for membrane load was calculated as 37187

N/mm (8360 Iblin) and 34963 N / m (7860 Ibh) . The experimental applied failure load

was found to be approximately 25354 N/mrn (5700 I b h ) 20684 Nlmm (4650 Iblin) for

AS4 and Type 1003 respectively. But it was argued earlier that the O" is the only layer

contributhg to the strength of the laminate. Therefore. the effect of other layer has to be

set at zero in larninate theory formulation. Setting contribution of 45' plies to zero.

classical lamination gives 24467 N/mm (6175 l b h ) and 23620 N/mm (5310 l b h ) ,

which is within 7.7 and 12.5% for carbon and Glass fiber epoxy respectively. The

discrepancy resulted Born the fact that CLT assumes perfect bonding. ignores the inter-

laminar stresses and the edge effect of the laminate. Nevertheless. the failure stress of (O.

4 5 . -15. O)? is very close to that of cross-plied laminate.

A s the cross-plied and angle plied laminate results indicated. stress concentration f?ee

edge effect. and inter-Iamuiar stress did contnbute to the reduction of strength and

sri ffness in balanced symmetric laminates. There fore. quasi-isotropic whic h is balanced

and symmetric should not be an exception to this physical phenornenon. But. the

cxperimental result of quasi-iso tropic gave contradicting results to that found in other

experiments. The ultimate strength of both carbon and glass fiber epoxy were found to be

1830 W a and 955.6 W a respectively. The principal reason for inter-laminar stresses is

Poisson's mismatch v,. If the layers were not bonded. under uniaxial stress. each lamina

could under go different transverse strain. However. since a perfect bonding is assumed

the transverse strain must be the same. which causes inter-laminar stress. Knowing this

effect. one cm conclude that the stacking sequence can increase or decrease the inter-

laminar stresses such that the ultimate strength can be affected. Clearly the distributions

of stresses due to monotonic Ioading shows that stress components are constant within

each individual layer for in-plane stress in quasi-isotropie laminate. For an (8 / - 0 )

angle-plu laminate in uniaxial tension r, is the most significant inter-laminar stress at

the interface of the 8 and - 0 laminas. Its magnitude and direction is highly dependent on

the fiber orientation angle 0. Further more. r, has a higher value at the 0 and - 0

interface in a clustered (0. and - 0. ), laminate than in a altemathg [(O and - 8 ). 1,

laminate. For cross plied samples. significant inter-laminar stresses are a, ry,. The

stacking sequence of cross-plied larninate causes out-of-plane tende stress at rnid plane.

Therefore. delamination contributes to the failure of test specimen at lower stress. For a

eeneral larninate. different combination of r,,. r,. and o, c m be present between C

different layers. For the angle-plied. the inter-lamina shear stress is higher between + 45"

and - 45'' than between 0" and - 45". Also the maximum O, occurs at the laminate mid

plane. Thus. the stacking sequence arongly influenced the nature. magnitude. and the

location of inter-lamina stresses.

The use of CLT brings the attention to the ultimate transverse strength of iaminate

(ultimatr strength of 90" when laminated with other layers). In both cases of Carbon fiber

and Glass tiber r p o q the transverse strength is smaller that that of unidirectional. This

phenomenon can be esplained as that the unidirectional ultimate strength value is an

overail value away from fiber/matris interface. Therefore. locally the interface stress

must reach a stress value published for unidirectional at a lower applied stress at far field.

Allen Yu req-41 expiain this phenomenon as an inhomogeneous stress field that prevails

in quasi-static experiments as a result of mat* plasticity. Thus. failure happens in rnatrix

well before the intrinsic stress value of 55 MPa and 35 MPa for Carbon and Glas fiber

eposy is reached respectively.

In surnmary the inter-laminar stresses in two cases can be shown to be virtuaily the same

or identical. Therefore. the ody conclusion to be drawn is that the inter-larninar normal

stresses. a, may be the key to the difference in effective strength of a laminate. It is well

established that the cross-plied and the ahematively layered angle-plied laminate exhibits

this fact and therefore. delamination is the cause of premature failure if the laminifte.

However. in the case of quasi-isotropic laminate this apparent.

CHAPTER 7. CONCLUSION

In the analysis of data, three-dimensional constitutive equations have been reduced for

the case of plane stress. The engineering constants of a lamina oriented at any arbitrary

angle 0 to the global x-y coordinate have been developed. Testing balanced symmetric

laminate has eliminated the curvature effect of the laminate due to axial load. In addition,

the residual thermal stress has been excluded in the calcuIation of lamina stresses. The

Classical Lamination Theory has been utilized to find the elastic respond of laminate. The

CLT has been used to calculate the effective properties, X and Y, of each lamina. AIso

the mie of mixture has been used to determine the effective stiffness in transverse

direction knowing the stiffness of unidirectional O* and cross-plied laminate.

Manufacturing process that was adopted was recornmended by the manufacturer of pre-

preg carbon fiber epoxy by Hexel and glass fiber epoxy type 1003 by 3-M. The tensile

teaing was carried out in accordance with ASTM standard. Al1 the rnanufacturing and

testing equiprnent were designed and built in house or in few cases, they were upgraded

in house The experimental data obtained were teaed for validity using statistical means.

From the experiments it was concluded that the transverse modulus of a constrained

laminate is more than the modulus of unidirectional unconstrained laminate. This value

was found to be 16.2 GPa (2.35 Msi), which is approximately twice as much as the

published value. The transverse Young's modulus of glass fiber epoxy in a highly

constrained (cross-plied) was found to be 12.22 GPa (1.77 Msi). The longitudinal

ultimate strength X was measured to be 1.5 1 GPa (2 18.7 ksi) and 73 1 MPa (1 O6 ksi) for

carbon fiber and glass fiber epoxy respectively. And at last, the transverse ultimate

strength Y was found to be approximately 35.7 MPa (5.17 ksi) and 28.8 MPa (4.18 ksi).

The following table 7.1 summarizes the experimental values:

Material 1 Young's Modulus E2 ( Longitudinal Strength 1 Transverse Strength 1

Table 7.1 The Effective Properties of Carbon and Glass fiber epoxy

1 AS4 l

1 1003

For in-plane loading of specially orthotropic laminate, (Al6 = AZ6 = O) with only N

nonzero, stresses take the form:

Normalizing through thickness distribution of in-plane components of stress in cross-

plied. angle-plied and quasi-isotropie gives more information about inter-laminar

stresses. The stresses are constant between the layers and generally discontinuity occurs

at layer interface. The stresses are continuous only when the layers are in the sarne

orientation. In cross-plied os is different between 0" and 90. layers. However, q is the

same in both layers and only their direction is different (discontinuous). This reflects the

large difference in stiffness of each layer. Angle-plied on the other hand demonstrates

uniform axial stress and zero transverse stress $, Angle-plied laminate has also equal

16.2 Gpa

12.32 GPa

1550 MPa

73 1 MPa

35.7 MPa

28.8 MPa

and opposite shear stresses & in the f 0 layers. From physical stand point the transverse

stress is zero because the Poisson's ratio is the same for + 8 and - 9 layer. The fact that

shear stresses ?, is the same satisfies the equilibrium condition. It should me noted that

al1 angle-plied exhibits the same stress distribution independent of its orientation.

The quasi-isotropie laminate represents a combination of cross-pied and angle-plied

laminate. The axial stresses are proportional to their stiffness of the respective layer. The

transverse stresses are non zero in some layer because of Poisson's mismatch between

those layers. The shear stress is zero in 0" and 90" plies because of onhotropy and

equilibrium is satisfied by existence of? 45 layer.

Prediction of faiiure of these laminates is not simply a calculation of maximum stress in a

given layer. it rather requires more attention and calculation. From these expenments

usine balanced and symmetric laminate, it was found out that the mode of failure for

cross-plied was edge delamination. The contributing factor to the reduction of ultimate

strength of the laminate in angle-plied was found to be shear stresses. And quasi-isotropie

laminate demonstrated higher ultimate strength due to the distribution of these stresses

within each layer.

A basic assumption of CLT is that each layer is in a plane state of stress. The equilibrium

in each layer of finite width laminate can not be met under uniaxial if the plane stress is

the assumption. The nonzero transverse stress away 6om the free edge is acceptable but

as at the free edge transverse stress has to go to zero. If equilibnum is tme, there should

exist out-of-plane stress at the traction free edges. These stresses are called inter-laminar

stresses. These stresses are the results of staking sequence and fiber orientation. It was

previously mentioned that many researchers believe that properties of composites are

hnction of orientation as well as thickness of the layer. In this work most of the attention

was paid to the orientation of the fiber and the data obtained are valid for intersperse

configuration. However, to fully tabulate the properties of composite, more m d y must be

performed on the effect of the thickness of each layer (cluster sequence).

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COLOR 14.0. O: CLS L$="**********************************$*******************************" LIS = "* * N

COLOR 1 1 : LOCATE 5.6: PRMT U: FOR X = 6 TO 19: LOCATE X 6: PWNT LIS: NEXT X LOCATE 19.6: PRMT Lâ: LOCATE 18.25: COLOR 15: PRINT "RYERSON POLYTECHMC

u J I v E R S m " BEEP: COLOR 14 WS(1) = "A": WS(2) = "En: WS(3) = "Rn: WS(4) = " ": WS(5) = "LN: WS(6) = "An: WS(3 = "BN FOR .Y = f TO 8: FORT = 1 TO 1 0 : NEXT T: LOCATE 10.29 + 2 X: PRMT WS(m: NEXT X LOCATE 6.25: PRMT " STRUCTURES LABORATORY"; : COLOR 1 1 LOCATE 15.50: P M "Mr. J. KARPYNCZYK & H. GHAEMIN WS(I) = "A": WS(2) = "En: WS(31= "Rn: WS(4) = " ": WS(5) = "Ln: W$(6) = "A": WS(7) = "B" FOR ,Y = 1 TO 8: FOR T = 1 TO 200: NEXTT: LOCATE 10.29 + 2 X; P M WS(X) NEXT X KEY 36 = "COMPOSITE MANUFACTLTRMG SOFTWARE' ( 1 1/23/94)" LOCATE 12. (40 - LEN(KEY3S) / 2): COLOR 15: PRINT KEY3$ FOR T = 1 TO 5 0 : NEXT T: BEEP

460 CLS : SCREEN O: WlDTH 80: COLOR 7.0. O: CLS

LOCATE 1.25: PlUNT "STRUCIZTRES LABORATORY ": LOCATE 3.24: PRiNT "DATA ACQUISïïiON SELECTION MENU":

LOC ATE 6. 10: COLOR 14: PRMT " 1. ": : COLOR 15: LOCATE 6.11: PlUNT "MONITOR ": : COLOR 1 O: P M "40 start manufacturing composites coupon"; LOCATE 7. 10: COLOR 14: PRMT "2. ": : COLOR 15: LOCATE 7.14: PFUNT "DYNAMIC

PRESSüRE DATA MEASUREMENT ": : LûCATE 8,30: COLOR 10: PRINT "-automatic data logging to hard dnve":

LOCATE 9. 10: COLOR 14: PEUNT "3. ": : COLOR 15: LOCATE 9. 14: PEUNT "CONFIGURE SYSTEM"; : COLOR 10: PRMT " -to setup DataPAC": LOCATE 1 1. 10: COLOR 14: PRINT "1. ": : COLOR 15: LOCATE 11. 14: PFüNT "TERMINATE

PROGRAMME ": 530 LOCATE 22. 10: PRMT SPACES(65);

LOCATE 22. 10: COLOR 14: P W "Rase press numbered key to make your seldon. . ."; 550 KS = W Y $ : IF= = ""THEN 550 LOCATE 24. 10: PEUNT SPACES(65): iF K$' = "O" GDïO 1780' SYSTEM K = VAL(K$): IF K = MT(K) AND K > O AND K < 5 THEN GOTO 600 LOCATE 24. 10: COLOR 13: PRiM' "Somy, that was not a valid enm. Try again..."; : GOTO 530

600 ON K GOTO 780. 1780.620.9999 6LOCLS:PiUNTm ":PFUNTn ":PRiNTn " P M "Switch the EEPROM mitch ON and press the F5 FüNCTION key" DO DO KEms = INKEYS LOOP WHILE KEYINS = "" L û û P UNTIL MiDS(KEYM$. 2. 1) = CHRS(63 CLS REM TO SETUP THE DAYI'RONIC CHANNEL LOCATiON AM) CONSTANTS PRIM- # 1 . "UNLM PRINT #1. "LCTI=21" P m # 1. "LCT2=2ZW PRiNT # 1. "LCT3=23" P M #1. "TYP1=19" PRiNT # 1. 'WC= 19" PNNT # 1. "T(P3= 19"

P M " ": PRMT " ": PRMT " ": COLOR 20. O. O PRiNT "Tm OFF EEPROM switch and press the F5 FUNCTION kq"

DO DO liEmrs = INKEYE LOOP WHLLE r n E = "" LWP UNTIL MID$(KF(M$. 2.1) = cm(63)

780 CLS 880 COLOR IO. O. O

MPüï "MAKE SURE THE INITiAL PRESSURE IS SET AT 200 PSI. CAN WE START THE TEST (YIN)": IS

IF IS = "Y" THEN GOTO 890 END IF II; IS = "Nn THEN GOTO 460 END IF

890CLS: K=-1: Y=O:COLOR4.0.0 P W "STARTTNG TO RESET ALL VALVE AND ZERO TRANSDUCER": COLOR 2.0. O REM ZEROiNG ALL VALVE LOCATION FOR SETUP OPERATION

PRiNT #1. "BITO=lM SLEEP 2 PRiNT # 1. "BITO=OH

REM TO ZERO PRESSURE TRANSDUCER P W RI. "ZR09"

PRINT "NOW THE CURE CYCLE BEGMS:"

P R I N A B. C. X SLEEP 60

IF A < 345 OR B < 345 OR C < 345 THEN GOTO 891

895 P M #1. "BITO=lm SLEEP 2 PRINT #l. "BITO=OW SLEEP 2

900 PRMT #I . "BiT2=ln I F X > 183ïHENGOTO901 [F X < 183 THEN GOTO 900

90 1 P M # 1, "BIT2=OM SLEEP 2

902 PRINT #l. "CHN;"

W U T #l. A P M #1. "CHN2" MPUT #1. B P M #1. "CHN3" INPUT #1. C PRMT #1. " ~ m 9 ~ rM'm #1. X

PRMT A B. C. X SLEEP 60 IF X < 180 THEN GOTO 900

IF X >= 180 THEN GOTO 905 REM ( TO ENSURE CONSTANT PRESSURE THROUGH OUT THE CURING CYCLE).

1200 P M #I. "BIT1=ln SLEEP 4 PRNT al. "BITl=O" SLEEP 2

I2IO P M #I. "BITZ=la IF S . OTHEN GvrO 1210 IF'C=OTHENGOTO 1212

12 lZ P M #1. "BIT2=ON SLEEP 2 P W # 1. "BITO= I " SLEEP 2 P r n i t f l. nBno=on SLEEP 2

COLOR 4. O. 0 LOCATE 23.2: P W "REPEAT SCAN- 1" LOCATE 23.30: PRiNT "GO TO MAIN MENU-2": B E P : COLOR 3.0. O LOCATE 21.2: MPUT "YOUR CHOICE ": CHOICE Obi CHOICE GOSLB 890.460

P W "EWER FiLE NAME FOR DATA OUTWTw

MPUT "WOULD YOU LïKE TO START DATA SCAN 0"; IS IF IS = "Y" TrlEX GUTO 1880

END IF IF I$ = "N" THEN GOTO 460 END IF

1880 COLOR 10. O. O OPEN FLLES FOR APPEND AS #Z H5S = "##.# ##.#"

1885 CLS . COLOR 4.0. O pRR\TTn

*****************************************************************************~ PRINT" " PRMT " EXECUTING PRESSURE DATA LOGGING " PRINT" " P R M T "

***********************+*******************************************************n

1890 SOUND 60.4 PRiNT #1. "CHNI" INPUT #1. A P M #1. "CHNZ" INPUT #1. B PRINT f i l . "CHN3" MPUT $1. C P M il 1. "CHNO" rNPuT#l. 'i

CLOSE #Z

999 9 COLOR 7.0. O: CLS : COLOR 20.0. O P M " " : P M " " : P m " ":PRR\CT" " P m w

*****************************************************************************" PRiNT " THE TEST PROGRAMME HAS BEEN TERMMATED n

P W "

COLOR 7.0. O: CLOSE

The Autoclave program Consolidation Cornputer Program

DECLARE SU3 MENU ( InmingGrad DecreasingGrad AvTempûld AvTemp. D a i M e m p . TempA TempB. TempC. MaxTemp. RoomTemp. Ressure. RessureOld, -. CureTime. TirneElapsed SnoozeTime. X Sec. H5S) DECLARE SUB MITIALiZE (X) DECLARE SUB MANUFACTWECOMP (hwsingGrad. DeaeasingGrad. AvTernpûid AvTemp. DesiredTemp. TempA TempB. TempC. Ma..Temp RoomTemp. Ressure. PressureOid DesiredPressure. CureTime. TimeElapsed. SnoozeTime. X Sec. H5$) DECLARE SUB C O m G OC) DECLARE SUB TERMMATE ( X ) DECLARE SUB SENSORCHECK (X) DECLARE SUB DATAFILE (X) DECLARE SUB SETSCREEN (X) DECLARE SUB SCREENSTATUS (X) DECLARE SUB TEMPRNSE (IncreasUigGtad AvTempOld AvTemp. Desiredï'emp. TempA Te+. TcmpC. Mx~Tcmp. Ressure. PmmreûLd DairedRemire. CiireTime. TbneElapsed. SnoozeTime. X Sec. H5S) DECLARE SUB PRESSURUE (AvTempûld AvTernp. Deshfïemp. TempA TexnpB. TempC. MasTemp. Pressure. RessureOld DesiredRessure. CureTime. TimeElapsed SnoaeTime. X Sec. HSS) DECLARE SUB CURECYLCE (AvTempOld AvTemp. DesuedTernp. TempA TanpB. TexnpC. Mx~Ternp. Presmre. RessureOld Des-. CureTime. TimeElapsed SnoazeTime. X Sec. H5%) DECLARE SUB DEPRESSUREE (AvTempOld AvTemp. D e s M e m p . TempA TempB. TempC. MasTcmp. Pressure. PressureOld DesiredPressufe. CmTime. TimeElapsed. SnoozeTime. X Sec, H5S) DECLARE SUB TEMPLOWER (DecrcasingGnd AvTcmpOld. AvTemp. DesMemp. TempA TcmpB. TempC. MxrTemp. RoomTemp. Prcsme. PrcssureOld DesiredPressure. CureTime. TimcEhpscd SnoozeTimc. X Sec, H5S) DECLARE SUB EMERGENCYDEPRESS (Pressure)

CLS

RcadingsPcrMin = 3 'Number of radings h t the Datqtec v d e m is to takc p minute RcqurrcdRaisingGndient = 4 Rcquired Temperaturc RASiNG Gradicnt of 5 degrces F per min RcquiredLowcringGndient = 2.5 'Requ~rcd Tempemm LOWEFUNG Gradient of 2.5 degrces F per min CureTime = 120 'Duration of Cure CyIce (time at 260 Dcg,rees F) in Minutes RoomTcmp = 90 Xoom Temperature at Start MasTemp = 420 'Mauimum Operating Temperame in Degrees F DesifedTcmp = 260 'Sets the Temperature rqured for the cure cycle Desi rcdPressure = 50 'Sets Desird Pressure in psrg colour = 12 1 naasingGrad = RcqurredRaisingGradient / RcadingsPerMin DccrcasingGnd = RequircdLoweringGradient / RadingsPerMin SnoozeTime = 60 / ReadingsPerMin Zength of time between Reading in seconds CureTime = CureTime 60 'Coverts Cure time

'****************ktial Settings ************** TempA = RoomTemp: TempB = RoomTemp: TempC = RoomTemp AvgTemp = RoomTemp: AvgTempOtd = RoomTemp: OIdAvgTempOId = RoomTemp: Curing = O: n = O: Sec = O '***$********************************************

OPEN "ComZ:9600.1~7.2" FOR RAMXlM AS #1

OPEN "zglJ5-05" FOR OUTPUT AS #Z

DIM d(W) CLS KEY 4. "FILES" + CHRS(13) KEY 6. " " + C W ( 3 4) KEY 7. + cm(3 J) KEY 7. " " + CHR$(34) KEY 8. " " + CHR$(3 4) KEY 8. " " + CKRS(3-l) KEY 9. " " + CHRS(34) EY 10." " + c m ( 3 ) CLS

PEUNT # 1 . "UNL" PRiNT #1. "LCT1=21" PRINT #1. "LCT2=2tW PRINT #1. "LCT3=23"

P m #1. "TYPl=19" T h c d Couples PEUNT #1. "TYP2=19" PRiNT #1. "Tl'P3=19"

'********** Th1~sectionensutesth3tafIsolenoidstumedoff****** PrnT # 1 . "BrTO=O" P M # 1 . "BIT1=CW POT # 1 . "BITZ=On P W # 1. "B1T3=OU P M if 1 . "Brr4=O" PRINT # 1 . "BIT5=On PRINT # l . "B[T6=OU PRiNT if 1, "B1T7=On

CLS PRINT "Just checking if the RELEASE solenoid is openting poperly" PRINT "The RELEASE soicnoid shouid remain open for 5 seconds" P m PRMT PFUNT #1. "BIT3=ln PWNT # 1. "BIT5= 1 " SLEEP 1 PRNT # 1. "BIT5==OU SLEEP 1 PRMT # 1. "BIT3=OU SLEEP 5

P M "Just checking if the PRESSURIZE solenoids are working poperiyu PRiNT "The autciave shodd PRESSURiZE for 10 seconds" PRmT

PRINT #I. "BIT3=In PRMT #1. "BiT4=IU SLEEP 1 PRMT #l. "BIT5=Om SLEEP 10 PRiNT #1. "BIT3=On PRINT PRmT P M "PIease open the m u r i l release valve to releive the miduai pressure"

PRMT P M "Vacuum pump should be enmgized for 5 seconds" P M # I . "BIT6=ln SLEEP 5 PRINT #1. "BTT6=OM

PRn4-r P M "Wcating element should be on for 5 seconds" PRiNT # I . "BIT7=ln SLEEP 5 PRINT # l . *81T7=0"

PNNT # 1 . "ZRû9" 'Re-zero's the pressure transduccr P M # 1 . "TER=15" C.4LL MENU(Incrc3singGnd DccrmingGrad AvTempOid AvTcmp. DesircaTcmp. TempA TempB.

TempC. MrisTcmp. RcomTemp. Pressure. PrcsweûI d DesiredPressurc. CureTime. T I C E hpsed. SnoozcTimc. X Sec. H5S)

9999 CLS CALL SCREENSTANS(W P W # 1 . "BIT3= 1": LOCATE 10. 1: P M "ON II

PRINT # 1 . "BITS= 1 ": LOCATE 1 1. 1 : PRMT "Enerped " SLEEP 1 P W # 1. "BIT5=On: LOCATE 1 1 . 1 : P W "OPEN n

P R N T #1, "BiT3=ûn: LOCATE 10. 1: PRINT "OFF 9

P M #I. "%ITO=O" P M # 1, "%IT1=On P W # I . "BIT2=in PRNT # 1 . "BIT3=Om: LOCATE 10.35: PRRüï " OFF " 'Ensures that the power line IN is tunied

off PRMT#1."BITS=O":LOCATE11.35:PRMT"OFFn Ensutesthatthepwertorelease

solcnoid is tunied off P M #l. "BITS=ûw: LOCATE 1 1.35: P W " OFF " 'Ensures that the power to release

solenoid is turneci off P M # 1. "BIT+Ow: LOCATE 12.35: PRINT " OFF " Ensures that the Vacuum pump is

turned off P M #1. "BU7=û": LOCATE 13.35: PRINT " OFF " 'Ensures that the heating element is

turned off

LIS = "* ** LOCATE 1 . 1 : COLOR 4.0. O: PRINT LS: FOR n = 1 TO 6: PRMT L1S: NEXT n P M LS LOCATE 5 . 2 5 : COLOR 6: PRINT "Prognm has been Completed": COLOR 15

CLOSE #1 CLOSE #2 END

SUB CONFIG (X)

CLS PRiNTn ":PRMTU ":PRINTN "

PRMT " Suitch the EEPROM s\h.itch ON and press the F5 FUNCTION key" DO DO KEYTNS = N E Y $ L û û P WHlLE KEYIN$ = "" L O P W L MIDS(KEYMS. 2. 1) = CHRS(63) CLS 7'0 SETUP THE DAYTROMC CHANNEL LOCATION AND CONSTANTS PRihT #2. "UNLN PRINT #2. "LCT1=2 1 " P W #2. "LCT2=22" PRINT #2. "LCT3=23" PRiNT #2. m l = 19" PRMT # 2 . 7 Y F 2 = 19" PRihT #2. W 3 = 19"

PRiNT #2. "ASL 1O=In

P M #2. "LCT9=Jln P R I N #2. "pTp9=72" P M #2. "MVV9=3.08333.500" P m #2. "mm" PRINT #2. "TER=15"

PRlNT " ": P M " ": P M " ": COLOR 20.0.0 PFUNT " Tum OFF EEPROM mitch and press the FS FUNCTION kq"

DO DO K E m S = INKEY$ L û û P WHILE KEMNS = "" LOOP M I L MIDS(KEYTNE. 2, 1) = C M ( 6 3 )

CALL MEMQIncmsUigGrad DenasingGrad AvTempOld Avfernp. DesiredTemp. TempA TempB. TempC. MaxTernp. RoomTemp. Ressure. RessurrOld DesimPmwe. CureTime. TheElapred SnoozeTime. X Sec. H5S)

END SUB

SU3 CüRECYLCE (AvTempOld AvTemp. D c s ~ e m p . TempA TempB. Te@. MaxTemp. Pressure. PressureOld DesiredPressure. CureTime. TimeElapsed. SnoozeTime. X Sec. HS$)

CALL SCREENSTATUS(X) LOC ATE 3.20: COLOR 4: PRMT " Curc Cycle ": COLOR 15

P M # 1. "Now Beginning Cwe Cycle Portion of Program" P M # 1. TEMES; TimeEbpsed. Curing scc=o Curing = O DO WKILE (Curing <= CureThne)

' * * t**************f **TO be read from Da\?= *********** PFüNT #l . "CHNI": INPUT #l.TempA PRïhT fC 1. "CHNZn: M U T # 1. TempB PFUNT #1. "CHN3": NPOT #1. TempC PRINT # 1 . "CHN9": INPUT # l . Ressure '********************************************************

IF ( AvTcmp > 420) OR (Pressure > 100) THEN CALL EMERGENCYDEPRESS(Pressure)

DcltaTcmp = AvTcmp - AvTempOld LOCATE 6. 1 : P R I W USING "#####": X: : P m " ": : PRMT USMG ":Wu; (TimeElapsed ',

7600): (TimcElripscd \ 60); (TimeEhpscd MOD 60); : PRMT USMG "###.###.#": TcmpA; TcmpB; TcmpC; AvTemp: DcltaTernp IncrasingGrad; Pressue

LOCATE 2. 10. PRiNT TIMES: LOCATE 2.65: PMNT DATES LOCATE 8. 1: P M " Curing Time Elripsed ......" : : COLOR 8: P M USMG ":##": (Curing \

3 600): (Cunng \ 60): (Curing MOD 60): COLOR 15

LKATE 9. 37: PFUNT " Deskd Curc T h e (Min)..."; : P W USMG "####IV: (CurcTimc) 160

iF ( AvTcmp >= DcsiredTemp + 5 ) THEN Pmi # 1 . "BITI=O": L K A T E 13.35: PFüM' " OFF " ' T m clement Off END IF

IF ( AvTcmp < ûesiredTemp - 2) THEN P W #I. "BlT7=1": LOCAE 13-35: PRMT"0N " Tunis heating

Elcmcnt

E r n IF IF (TcmpA > MasTemp) OR (TempB >= MrisTernp) OR (TernpC >= MrisTemp) THEN

BEEP P M #1. "BIT7=Ou: LOCATE 13-35: PRMT " OFF "; : COLOR 4: L O C A E 14.35: P M "

Tm HOT ". COLOR 15 Turns heating element off Sec = Sec .25 EMI IF

IF (Pressure < D c s i r c d P r ~ ~ ~ ~ ~ e - 5 ) THEN BuniCounter = O PM#l."BIT3=l":LOCATE10.35:PRR\TT"OPEN " ' Opens Pressure M

soicnoiod P M #1. "Bïï4=lU: LOCATE 11.35: PRMT " Energîzed "

SLEEP 1 P M #1. "BITI=O": LOCATE 11.35: PFUNT "OPEN "

DO WHlLE (Pressure C DesiredPressure)

SLEEP 1 s e c = s e c + 1 BurnCounter = BuniCounter + 1 IF BuCounter > 500 THEN PRINT "You are going to burn the line in solenoid " P R N ï #1. "CHNI": iNPüT #1. TempA P W #1. "CHN2": MPUT #1. TempB PRR\IT #l. "CHN3": iNPUT #l. TempC PRMT #I, "CHN9": MPUT #1. Pressure LOCATE 6. 1: P W USiNG "#####": X : P M " "; : PRMT USMG ":##";

(TimeEiapsed \ 3600); (TimeElapsed \ 60): (TimeElapsed MOD 60); : PRMT USING "###.W.#"; TempA; TempB: TcmpC; AvTemp: DeltaTemp; Incre;isingGnd: Pressure

LOCATE 2. 10: PRMT TIME$: LOCATE 2.65: P M DATES LOCATE 8. 1 : PRMT " Curing Timc Eiap sed......." ; : P W USMG ":##"; ( C u ~ g \

3600); (Curing \ 60): (Curing MOD 60) LOOP PIUNT #1. "BlT3=0": LOCATE 10.35: P W " CLOSED"

END IF

IF (Pressure > DesireciPressure + 5 ) TKEN BurnCountcr = O P M #1. "BtT3=On: LOCATE 10.35: P W " CLOSED "

Solcnoid PFüN'T#l."BITS=I":LOCATE 11.35: PEüNTnEnergized " SLEEP 1 PRMT # 1. "BITJ=O" LOCATE 1 1.35: PRMT "OPEN n

'Closes linc in solenoid

'Closses linc ui

DO WHILE (Prcssurc > DesircdPressure) SLEEP 1 S c t = S c c + l P M #1. 'CHNI'': M U T #1. TempA PRINT # l . "CHNZ": INPUT #1. TcmpB P M #l. "CHN3": INPUT # 1 . TcmpC P M # 1. "CHN9": INPUT # 1. Pmsure LOCATE 6. 1: P R I N USiNG "###Mn: X; : PRiNT " "; : PEUNT USiNG ":##":

(TirncElapscd \ 3600): (TimeElapscd \ 60); (TimcElripscd MOD 60): : P M USING "###.##*.ffn: Temp.4; TcrnpB; TcmpC; AvTcrnp: DeltaTcmp; incre3singGnd: Pressure

LOCATE 2.10: PRINT T m : LOCATE 2.65: PFüNT DATES LOCATE 8. 1 : PFüNT " Curing Timc Elapscd ......." : : P M USING ":##": (Curing \

3600); (Curing \ 60): (Curïng MOD 60) LOOP

P M # 1. "BIT* 1 ": LOC.4TE 1 1.35: PEUNT " Energized " SLEEP 1 P M #l. "BIT=i=OW: LOCATE Il. 35: PRMT " Closed "

Sec=Sec+ l END IF

SLEEP SnoozeTirne

Curing = (Curuig + SnoozeTime + Sec) \ 1 X = X + SnoozeTime + Sec TirneElapscd = (TimeElapsed + SnoozcTirnc + Sec) \ 1

Sec = O

PRINT #2. TIMES: : P W #2, USMG H5$: TimeEhpsed: Curing: P M #2, USING H5S; Tempk TempB: TempC: AvTcmp: DcltaTemp: [ncrcasuigGrad: Pressure A\.TempOld = AvTcmp PressurcOf d = Ressure

LOOP

P M # 1 . "BIT3=OW: LOCATE PRMT # 1. "BIT4=On: LOCATE solenoid PRïNT #1. "BIT5=OW: LOCATE solenoid PEüNT 1. "BIT6=01': LOCATE PumP P R N i #1 . "BITI=On: LOCATE clcmcnt

10. 35: PRINT " OFF " 11.35: PKiNT " OFF "

12. 35: PEUNT " OFF "

13. 35: F'F4.INT " OFF "

'Ensures that iN line is closed 'Ensures that power is tunid off to release

'Ensures that pwer is tumed off to release

'Ensures that power is m e d off to vacuum

'ENUT~S that powcr is tumed off to Heating

PRIhT If 2. "End of Cure cycle" 0998 EMISUB

SUB DATAFILE (XI

CLS CALL SETSCREEN(X) LOCATE 13.10: PRiNT "ENTER FILE NAME FOR DATA OUTUT" IWUT "(c.g. C:TESTl.DATl ". FILES PRIM-" " P M " "

ILKATE 22. 10: W U T "WOüLD YOU LIKE TO START DATA SCAN Con; iS IF (iS = "Y") OR (iS = "y") THEN CALL ~ A C T U R E C O ~ ( I n c r c a s i n g G n d DecreasingGrad

A\.TcmpOld AvTcmp. DcsiredTemp. TempA TempB. TernpC. W T e m p . RoomTemp. Prcssurc. PrcssurcOld DcsircdPrcssurc. CurcTimc. TimcEhpsed SnoozcTime. X Sec. HSS)

IF (i$ = "N") OR ( i $ = "n") THEN CALL MENU(lncrc3singGnd ûecrcasingGnd AvTcmpOld A1.Temp. Desudcmp. TcmpA TcmpB. TcmpC. MasTemp. RoomTemp. Rcsnuc. Ressudld DcsircdPrcssurc. CureTime. TimeEiapsed SnoozeTimc. .X Sec. H5$)

END SL'B

SUB DEPRESSURIZE ( AvTernpOld AvTemp. kiredTemp. TempA TempB. TcrnpC. MasTemp. Prcsnrrc. PressureOld DesiredPressurr. CureTime. TimeEiapsed SnoozeTirne. X Sec. H5S)

C.4.L.L SCREENSTATUS(.V LOCATE 3.20: COLOR 4: PRiNT "Depresnirization": COLOR 15

P M fC1. "BIT3=ln: LOCATE 10. 35: P W " ON "

PRMT # 1. "BITS= 1 ": LOCATE 1 1. 35: PRINT " Energized " SLEEP 1 P M iC1. "BIT5=On: LOCATE 11.35: P M "OPEN " 'Opens R e l a x Solenoid

PRNT it 1. "BIT3=On: LOCATE 10. 35: PRMT " OFF " 'Closes line M Solenoid P M # 1. "BIT4=OW: LOCATE 1 1-35: P M " OFF " PRIW #1, "BIT5=On: LOCATE 11.35: PRMT " OFF " PRiNT Y 1. "BIT6=On: LOCATE 12.35: PRMT " OFF "

P W # l . "BIT7=Ow; LOCATE 13.35: PRMT " OFF "

DO WHILE (Pressure > 5 ) P~T#1."BIT3=1":LOCATE10.35:PWnOPEN " PRNT # 1. "BIT5= 1": LOCATE 11.35: PEUKT " Energized " 'Opens Release Solenoid SLEEP 1 P~#l."BfT5=O":LOCATEll.35:PRINT"OPEN " PRNT # 1. "BIT3=OW: LOCATE 10.35: P W " CLOSED "

.4\.Tcrnp = (TcmpA TempC) / 2 'Correction for TC-B

IF ( AvTcmp > 420) OR (Pressure > 100) THEN CALL EMERGENCYDEPRESS(kssure)

DcltriTcmp = AvTcmp - AvTcmpOtd LOCATE 6. 1 : P W USMG "#####": X: : PFUNT " ": : PRMT USING ":##"; (TieEbpsed \

3600); (TimcElapscd \ 60): (TimcElapsed MOD 60): : PRINT USiNG "###.###.#": TempA. TempB: Tcm pC ; A\.Tcrnp: DeitriTcmp: IncmsingGnd: Rcssurc

LOCATE 2. 10: PRINT TTMES: LOCATE 2-65: PFüNT DATE5

SLEEP 3 0 Sec = Sec ' 10 'i = S + Sec TimcElrtpsed = TimeElripsed + Sec

PFWT fc2. TTMES: : P M #2. USING H5$; TirneEhpsed; Curmg; P W #2. USWG H5$: TcmpA: TempB: TempC; AvTemp; DeltaTemp: IncrcsisingGrad: Pressure

scc = O LOOP PlUTT # l . "BIT3=On: LOCATE 10.35: P M " OFF " 'Ensures that Iine M soienoid is closed PRiNT # l . "BIT4=OW: LOCATE 1 1.35: P M " OFF " 'Ensures that BIT4 is de-energized P W # l . "BIT5=OW: LOCATE 11.35: PRMT " OFF " 'Ensures that BIT5 is deenergized P M #1. "BIT6=O": LOCATE 12.35: PRMT " OFF " 'Ensures th31 vacul~lll pump 1s turneci off PRMT # 1 . "BIT7=On: LOCATE 13.35: PRMT " OFF " Ensures that heating element is tumed off

PFüNT #2. "End of Depressurization cycle"

END SUB

S U 3 EMERGENCYDEPRESS (Pressure)

CLS

L$="***************************************************************************" L13 = "* * n

LOCATE 1 . 1 : COLOR 4. O. O: PRMT U: FOR n = 1 TO 2: P M LI$: NEXT n LOCATE 2.25: COLOR 6: P W "EMERGENCY DEPRESSUREIZATION": COLOR 15

PRiNT # l . "BIT7=On: LOCATE 13.35: P M " OFF " Turns element off PW#l."BIT3=1":LOCATE10.35:PW"OPEN " P W #1. "BIT5=ln: LOCATE 1 1 . 35: PRMT " Energized " 'ûpens solenoid for emergenq dcpressuriza tion SLEEP 1 P R N ï trl. "BIT5=On: LOCATE 1 1 . 35: PRINT " Open "

PRINTCfl."BIT3=ln:LOCATE1O.35:P~"OPEN "

DO %'HILE Prcssure > O PFühT c: 1. "CHN 1 " : INPUT # 1 . TcmpA P M # 1 . "CHNY: MPüT #1. TempB PRlNT #1. "CHN3": MUT # l . TempC P M # l . "CHN9": INPUT M. Prcssure LOCATE 6. 1 : PRMT USMG "#####": X: : PFUNT " ": : PRMT USNG ":##": (TimcEhpsed 1

3600): (TimcElapscd i 60); (TimcEiripscd N D 60): : PRMT USiNG "###.W.#": TcmpA: TcmpBi TcmpC ; A\.Tcmp; DclmTcmg: IncreasingGnd: Pressure

LOCATE 2. 10: P M TIMES: LOCATE 2.65: PRNT DATE$

PNNT #2. TiMES. : P M zf2, USMG (TimeEIrtpscd): (Curing); PRIhT f(2. USING HSS: TcmpA: TcmpB; TcmpC: AvTemp; DeltriTcrnp; IncmsingGnd: Pressure SOUND 130. 1 SLEEP 10 scc = SCC - 10

LOOP

1000 KS = N E Y $ K = V.U(K$) coiour = colour + 1 S O W 330. 1 SLEEP 1 IF KE = "" THEN 1000

END SUB

SUB INTTULiïE (X) OPEN "COM I:96OO.nï.2" FOR RAMXlM AS #2 CLS LOCATE 10. 10: P W "htializrition Sut, routine"

BIT0 = 1: PRMT $2. "BiT(Z= 1": SLEEP 1 'Opcns Pressure Pl Vdve BIT0 = O- PRMT #2. "BKû=On: SLEEP 1 'Closes F b s w e M Vaive

W #2. "BIT 1= 1 ": SLEEP 1 UNT #2. "BIT1=On: SLEEP 1 UNT #t. "BIT2= 1 ": SLEEP i 'ûpens Pressure Solenoid UNT #2. "BiT2=0": SLEEP 1 'Closes Pressure SoIenoid UNT $2. "BiT3= 1": SLEEP 1 UNT #2. "BIT3=On: SLEEP 1 UNT #2. "BIT+ 1 ": SLEEP 1 UNT #7. "BITS=O": SLEEP 1 iINT #2. "BIT5= 1 ": SLEEP 1 W #2. "BiT5=ON: SLEEP 1

KEY 4. "FILESw + CHR!R 13) W Y 6. " " + CHRS(34) KEY 7. " " + C W ( 3 4 ) KEY 7. " " + C W ( 3 4 ) KEY 8. " " + CHRS(33) KEY 8. " " + CHRS(34) E Y Y. " " - CHRS(3.S) uk' 10. " " 7 C W ( 3 ,

END SLB

SUB MANUFACTURECOMP (IncmsingGnd DecmsingGnd AvTempOld AvTemp. DenreûTemp. TcrnpiL TcrnpB. TcrnpC. MsisTcmp. RoomTemp. Prcssurc. Ressu~OId DesiredPressure. CureTime, TimcEIapscd SnoozcTimc. X. Sec. HS$)

CLS COLOR 9. O. O

P R I M " Makc sure that the door and al1 finings arc tight and sccure. " INPLT " Crin the qclc bcgîn? (YM)": jS IF (j$ = "Y") OR CjS = "y") THEN

K = - 1 : Y = O : COLOR4.0. O PRINT "Strtrting to rcset ail valves and zcro the prcssure transduccr": COLOR 2. O. O '*** ZEROING ALL VALVE LOCATION FOR SETUP OPERATION

PNNT # 1 . "BIT3= ln: SLEEP 1: P M #1. "BIT3=ON PRINT # i . "BIT4=In: SLEEP 1: PRMT#1. "BITI=On PRINT # 1. "BITJ= 1 ": SLEEP 1: PRINT # 1. "BIT5=OW PRINT # 1 . "BIT+ ln: SLEEP 1: P M # 1. "BIT6=On PRIW #l . "BIT7=ln: SLEEP 1: PRINT #1. "BIT7=0"

PRINT # 1. "ZRO9" 'Zcro's Pressure Transduccr PWNT "NOW THE CYCLE BEGINS:"

PRMT # 1. "CHN 1 ": INPUT # 1. TempA P M ti 1 , "CHNZn: MPUT # 1, TempB PRMT #I . "CHN3": INPUT #1. TempC PNNT #1. "CHN9": MUT # 1. Pressure

CALL TEMPRAISE(1ncreasingGnd AvTempOld AvTemp. Desiredïemp. TempA TempB. TcrnpC. MasTcmp. Ressure, i3xswcûld DesiredPressure. CureTime. TimeElapd. SnoozeTirne. X Sec. H5S)

CALL PRESSURIZE(AvTemp0Id AvTemp- DesMemp. TempA TempB. TempC. MaxTemp. Prcssurc. PrcsnrreOld DesircdPressure. CureTime, TimcE hpsed. SnoozeTimc. X Sec. H5 S)

CALL CURECYLCE(.AvTempûld AvTemp. lkskdïemp. TempA TempB. TanpC. Ma~Temp. Pressure. PressureO1d CureTime. TUneEIripscd Snoozeïiie. h Sec. WS)

CALL TEMPLOWER(DecreasingGrad AvTempûld AvTemp. Desirecfïemp. TempA TmpB. TempC. MaxTemp. RoomTemp. Ressure. RessureOld DesiredResnrre. CureTime. TimeEhpsed. SnoozeTirne. X Sec. WS)

CALL DEPRESSUlUZE(AvTempOld AvTemp DesMemp. TernpA TempB. TempC. MasTemp. Pressure. RessweOld DesiredRessure. CureTime. T imeE1apse-d. SnoozeTime. X Sec. H5S)

END IF iF (if = "Y") OR (jS = "y") THEN CALL MENU(IncreasingGnd DecreasingGrad AvTempOld

A\.Temp. Desiredïcmp. TempA TempB. TempC. Ma-xTemp. RoomTernp. Resnire. ResnireOld DcsiredPresswc. CureTime. TimcElapsed SnoozeTime, X Sec. H5â)

EhD SUB

SUB M E N (InmsingGnd DecrcasingGnd AvTempOld A\-Temp. DesiredTemp. TempA TcmpB. Tc m pC. MasTcmp. RoomTemp. Pressure. PressumOid DcsidFVtxmc. CureTime. TimcEla psed. SnoozcTime, N. Sec. H5$)

CLS CALL SETSCREEN(X)

LOCATE 7 . 3 : COLOR 6 : PRMT" Data Acquisition Selcction Mcnuw: COLOR 15 LûCATE 14.3 : COLOR 14: P W " 1. ": : COLOR 15: LOCATE 14.7: PRMT "Scnsor Check ";

: COLOR 10: PRiNT " ": L K A T E 1 5.3: COLOR 14: PWNT "2. ": : COLOR 15: LOCATE 1 S. 7: PRïNT "Configure Synem

": : COLOR 10: PRDIT "40 setup DataPACW LOCATE 16. 3: COLOR 14: P M "3. ": : COLOR 15: LOCATE 16. 7: PRINT "Composite Cure

Cycle ": : COLOR 10: P M "-[O manufrtcturc composite coupon" LOCATE 17. 3: COLOR 14: P M "4. ": : COLOR 15: LOCATE 17.7: PFUNT "Terminate Program

"; . COLOR 10: PRINT "40 tcnninatc the program" 570 LOCATE 18. 3: COLOR 14: PRMT "Plcase enter thc numbcr to m k c your seleaion.. ";

550 KS = IMiEYS: iF KS = "" THEN 550 K = V.AL(KS) coiour = colour + 1 iF (K > 0) .4ND (K <= 4) THEN IF KS = " 1 " THEN CALL SENSORCHECK(X) IF KS = "2" THEN CALL CONFIG(X) IF KE = "3" THEN CALL MANUFACTURECOMP(1ncrerisingGrad DecmingGnd AvTempOld

AvTcmp. D e s d e m p . TempA TempB. TernpC, MasTemp. RoomTemp. Prcssurc. PressureOld DcsiredPressure. CureTime. TimeElapsed SnoozeTime. X Sec. H5f)

IF KS = "4" THEN CALL TERMMATE(X) ELSE LOCATE 24. 10: COLOR (colour): P M "Som. th1 was not a vaiid e n g . Try again..": :

GOTO 530 Eh?> IF

END SLTB

SUB PRESSLTRIZE (AvTempOId AvTemp, DesiMemp. TempA TempB. TempC, MasTemp. Preswe. PressureOld DesiredPressufe. CureTime, TimeElapsed. SnoozeTime. X Sec. H5$)

LOCATE 3.20: COLOR 4: PRMT " Pressurization ": COLOR 15

BEEP

SLEEP 2 BEEP SLEEP 7 BEEP

CALL SCREENSTANS(.X) LOCATE 3.20: COLOR 4: P M "Ressurization ": COLOR 15

PRlNT#1."BIT3=In:LOCATE10.35:PRMT"ON " PRlNT # 1 . "BIT4= 1 ": LOCATE 1 1. 35: P M " Energized ": Solcnoid is Closed SLEEP 1 PRINT#1."B1T5=On:LOCATE 1 1 . 35:PRINTUCLOSED ":

SLEEP 2: Sec = Sec + 2

' Ensures Emergency Release

'Opens Solenoid B ï ï Z = 1

AvTcmp = (TcmpA + TcmpC) / Z 'Comaion for TC-B

IF ( AvTcmp > 420) OR (Pressure > 100) THEN CALL EMERGENCYDEPRESS(Prcsmrc)

LOCATE 6. 1: PRINT USiNG "#####"; X; : PRINT " "; : PRINT USMG ":##"; (TimcEIripsed \ 3600): (TimeEhpscd \ 60): (TimeEhpscd MOD 60): : P m USMG "W.###.#": TempA, TempB; TcmpC . AvTc mp: DeltaTc mp: Incrcasin gGrsid: Pressure

LOCATE 2.10: P M TIMES. LOCATE 2.65: PRMT DATES

' IF rDclt3Tcmp > incrc3singGrrid) THEN l PRlNT #l. "BIT7=On: LOCATE 15.35: PRMT " OFF " I LOCATE 14.35: COLOR 4: PRMT " Tcmp Gndicnt too High ": COLOR 15 ' Turns clement

OFF Bit lo=O" I END IF

IF ( DeltaTemp <= LncrasingGnd) THEN P M #1. "BIT7=In LOCATE 13.35: PWNT " ON ": LOCATE 14-35: COLOR 6: PiUNT " Temp Gradient OK

": COLOR 15 ' Turns ctcrncnt Off Bit lû=On E M I IF

IF (TcmpA > MasTernp) OR (TempB > MasTcmp) OR (TempC > MauTemp) THEN BEEP PRMT # 1. "BIT7=On: LOCATE i7.35: PRlNT " OFF ": COLOR 1: LOCATE 14.35:

P M " Hccituig Element is Too HOT": COLOR 15 ' Turns element OFF BiT10=O Sec = Sec + .E END IF

A\-TcmpOld = A~vTemp SLEEP 1 Sec = Scc - 1

x=x+scc TimcElripscd = TimeElapsed + Sec

P M #2. TIMES: : P M #2. USING H5S: TimeElapsed Curing PRMT #2. USMG H S : TempA: TempB: Te'empC: AvTemp DeliaTemp: IncreasùigGrad: Ressure S e c = O

LOOP

PRDT #1. "BIT3=Qn: LOCATE 10.35: PRMT " OFF " ' C l o s IN solenoid P M #2. "End of Presswke qcIen

SUB SCREENSTATUS (m CLS

LOCATE 2. 1 : COLOR 15: PRINT "Timc .... Date.. ..... " LOC.4TE 2. ?O: COLOR 9: P M " G W Fiber Mmufactunng Rogram": COLOR 15 L K A T E 5. 1 : PRMT "Currcntly in the portion of the C m Cycle" LOCATE 5. 1 : P M " X TirncElapsed TcmpA TcrnpB TcmpC AvTemp DclwT ReqDT Press"

...... LOCATE 10. 1. P M " Pressure Linc ïN BIT? ....... " LOC ATE 1 1 . 1 : PRINT " Rclasc Solenoid. ..... BIT4 & 5 . . . "

........ ..... LOC.4TE II. 1 : P M " Vacuum Pump.. .BIT6.. " ....... LOC.4TE 13. 1 : PfUNT " Hcating Element ....... BIT7 "

END SUB

SUB SENSORCHECK (X)

' This sub rouunc has not bccn dcbuggcd !et

CLS LOCATE 1 . 5 : COLOR 4: P M "Scnsor check": COLOR 6: PRINT

PFUNT # 1. "BITO= 1" INPUT "1s BIT0 currentl> activateci? (Y/N)": b$ IF (bS = "Y") OR (bS = "y") THEN PRiNT #1. "BIT* 1": ELSE GOTO 666

PRiNT #1, "BiTl=I" INPUT "Did BIT0 go off. and is BIT1 nirrentiy actisated ? (Y N": CS IF (CS = "Y") OR (d = "y") THEN PRDIT #1. "BIT1=Om: ELSE GOTO 666

P R [ N # 1. "BIT?= 1" NPUT "Did BIT 1 go off. md is 6 IT2 currendy activateci ? (Y M)"; dS IF (cl$ = "Yn) OR (dS = "y") THEN PRMT #I. "B1T2=Ow: ELSE GOTO 666

PRINT #1. "BIT3=L1' MPUT "Did BIT;! go off, and is BiT3 currentiy activated ? (Y /N)"; e$ IF (CE = "Yw) OR (e$ = " j ") THEN PRMT # 1. "Bïï3=0": ELSE GOTO 666

P R N ï #1. "BIT4=ln MPUT "Did BIT: go off. and is BIT4 cuneritiy activateci ? (Y M)": 6 IF ( fS = "Y ") OR (fs = "y") THEN PFUNT # 1. "BiTS=On: ELSE GUTO 666

P M #1. "BrrS=l" INPUT "Did BIT4 go off. and is BIT5 currentiy activated ? (Y M)": g$ IF (g$ = "Y") OR (g$ = "v") THEN P W #l. "BIT5=Om: ELSE GOTO 666

PRMT #l. "BiT6=lW INPüT "Did BIT5 go OR. and is BIT6 a m a î i y activated ? (Y M)": HS IF (HS = "Y") OR (HS = "y") TKEN PRMT #l. "BIT6=Ow: ELSE GOTO 666

P M $1. "BIT7=ln INPUT "Did BIT6 go off. and is BIT7 currentiy activated ? N Nu: iS IF (iS = "Y ") OR (iS = "y") THEN P W # 1. "BIT7=Ow: ELSE GOTO 646

IWüT "Did B I T go off3 (Y /N)": jS IF (jS = "Y") OR US = "f"' THEN P M "Gd": ELSE GOTO 666

P M " Sensor checked completed successfdly"

GOTO 777

666 CLS LS = rn** * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * * *n

LI5 = "* " LOCATE 1. 1: COLOR 4. O. O: PRMT LS: FOR n = t TO 6: PRMT LIS: NE,YT n PRINT LS LOCATE 5 . 2 5 : COLOR 6: PRiNT "FAiLED SENSOR CHECK": COLOR 15

777 PMNT 81. "BITO=OH: PRMT #1. "BITI=OH: PRiNT # 1. "SIT2=On P M P 1. "BIT3=OW: P W #I. "BIT4=ûU: PRiFJT #1. "BIT5=Om PRINT # 1, "BIT6=01': PRiNT # 1. "BfT7=Ow SLEEP 2

CALL MENU(lncrwsingGrad DecrcasingGrad AvTempûid AvTemp. D e s M c m p . TcmpA TempB. TcmpC. MrisTcrnp. RoomTemp. Pressure. PressureOld DesiredPrcssure. CureTime. TimcEIapsed SnoozcTimc. X. Sec. H5$) END SUB

SUE3 SETSCREEN ( X )

CLS L$ = "***************************************************************************** LI$ = "* *" COLOR I l : PRINTLS:FORn= 1TO 11: PRiBTL1E:NEXh LOCATE 2 . 5 : COLOR 6: P M "Timc: ": LOCATE 2-60: : PRiNT "Date: "

LOCATE 3.10: COLOR 9: P M " Ryerson Polytechnic Universityn LOCATE 5. 10: P W " Aircraft Stnictures Laburatoryn LOCATE5,10:PRiNTn Mr.J-ncqkH.Ghaemi&K.MackyN LOCATE 6.10: PRiNT " Composite Manufaauring Software" L K A E 8. 10: P W * Glass Fiber Composite Program"

COLOR 15 P W PRiNT PRMT P m L$

END SL'B

SUB TEMPLOWER (DecreasingGnd AvTempOld AvTemp. DesiredTemp. TempA TernpB. TempC. MasTcmp. RoomTemp. Pressure. RessureOld DeskdPmmue. CureTime. TimeElapsed. SnoozeTime. X Sec. H5S) Sec = O CALL SCREENSTAWS(X) LOCATE 3.20: COLOR 4: P M "Temp Lowering ": COLOR 15 DO WHiLE ( .AvTcmp > ( 1 rio))

.A\.Tcmp = (TempA + Tc@) / 2 'Correction for TC-B

IF ( AvTemp > 420) OR (Pressure > 100) THEN CALL EMERGENCM)EPRESS(Rcssure)

DeltriTemp = .4vTernp - AvTernpOld LOCATE 6. 1: PWNT USING "#####": X; : PRMT " ": : PRMT USING ":##": (TimcEhpsed \

3600): (TimcEbpscd \ 60): (TimcElapsed MOD 60): : PRRuT USiNG "##.###.Mn: TcmpA Tcrn~B: TcmpC; A\.Tcmp: DcltaTemp; DecreasuigGrad: Ressure

LOCATE 2.10: PWNT T I M E S : LOCATE 2.65: PRINT DATE$

IF (DcltaTemp > DecreasingGnd) THEN PWNT # I . "BIT7=IU: LOCATE 13.35. P W " ON " LOCATE 14.35: COLOR 4: PRiNT " Cooling Gndicnt too Kgh ": COLOR 15

' Turns clcment OFF " END IF

IF (DclbTemp <= DecreasingGrad) THEN PRINT #1. "BiT74": LOCATE 13-35: P W " OFF " LOCATE t 4.3 5: COLOR 6: PRMT " Temp Gradient OK ": COLOR 15 ' Tunis

dcmcnt Off Bit 10=OR END IF

ff (TempA > Ma~Temp) OR (TempB >= MaxTemp) OR (TempC >= Ma~Temp) THEN BEEP P W #1. "BIV=O": LOCATZ 13-35: PRiNT " O f f " LOCATE 14.35: COLOR 4: PEUNT " Element is Too HOT": COLOR 15

Turns element OFF BiTlO=O Sec = Sec + .X END IF

IF (Pressure > DesvedPressure + 5) THEN P W #1. "E3iT3=Om: LOCATE 10.35: PRMT " OFF " P M #1. "BITS= 1": LOCATE 1 1-35: PRïNT " Energized " SLEEP 1 PRMT #1. "BIT5=On: LOCATE 1 1.35: PEUNT " D e p r d i n g " S e c = S e c + l Dû UNTU, (Pressure <= Desirahssm)

P W #l. "CHN9": WUT #1. Ressure 'Reads Rcssure from the DAYTEC Unit LOCATE 6. 1: PRMT USMG "#####": X: : PRiNT " "; : PRiNT USMG ":##":

(TimeEIapsed \ 3600): (TirneElapsed \ 60): (TimeElapsed MOD 60); : PRMT USING "###,###.#": TcmpA; TcmpB: Tc*: AvTemp; DcltaTemp; incfe;isingGnd: Ressure

LOCATE 2.10: PRINT TIMES: LOCATE 2.65: PFUNT DATE$ LOOP PRMT #1. "BIT4=ln: LOCATE 11.35: PRiNT " Energized " SLEEP 1 PRINT # 1. "BITS=O": LOCATE 1 1.35: P W " Open "

END tF

IF (Prcssurc < DesiredPrcssurc - 5 ) TfIEN PRINT fll. "BIT3=lN: LOCATE 10.35: PRMT " OPEN" PRINT#l. "BIT4=lW: LûCATE 11.35: P M " Encrgized " SLEEP 1 PRlNT#t."BITG-O":LOCATE11.35:PRMT"OPEN " Scc = Sec - I

DO UNTTL (Ressurc >= DcsircdPresswe) PfüNT #l. "CHN9": N'PUT #1. Rcssure 'Rads Pressure from the D AYTEC Unit LOCATE 6. 1: PFUNT USMG "#####": X : PRMT " ": : PRINT USMG ":#":

(TimcElaspcd '\ 36iK)); (TimcElapsed \ 60): (TimcEhpscd MOD 60): : PRMT USING "###.###.#": TcrnpA. TcmpB: TempC; .4vTcmp: DeliaTernp Pfessurc

LOCATE 2. 1 0: PRIM TI=: L W ATE 2.65: P W DATES SLEEP 1 Scc = Sec + 1

LOOP PRiNT # 1. "BIT3=OU: LOCATE IO. 35: PRiNT " OFF " 'Closes linc M solcnoid

EYD IF

SLEEP SnoozcTime S = X A SnoozeTime + Sec TimeElapscd = TimcElapscd + SnoozeTimc + Sec

PRIPUT #2. TIMES: : P M #2. USING H5S; T i r n e E l a m Curing; PEUNT # l USING H5S: TempA; TempB: TempC; AvTemp DeliaTemp; IncreasingGnd: Pressure

scc = O OldAvTemo01d = AvTempOld AvTempOld = AvTemp OldPressurc = Pressure

LOOP

PRlNT #2. "End of Temperature Lower cyclen EM> SUB

SUB TEMPRAiSE (LncreasingGrad AvTempOld AvTemp. Deskâïemp. TempA TempB. TempC. MasTcrnp. Pressure. hsureOld. Des-, CureTime. TimeElapsed SnoozeTirne. X Sec. H5E)

PRMT # 1 . "BIT6= I n Tums on the vacuum pump

CALL SCREENSTANS(X)

Sec=O LOCATE 3.20: COLOR 4: P W "Tempentufe Raise": COLOR 15 P W # l . "BIT7=l*: LOCATE 13. 35: PRMT " ON "

AvTcmp = (TcmpA + TcmpC) / 2 'Conccrion for TC-B

DcltaTcmp = AvTcmp - AvTempOId

IF (AvTemp > 420) OR (Pressure > 100) THEN CALL EMERGENCYDEPRESS(Rcssurc)

LOCATE 6. 1 : P M USING "X####": X; : PRINT " "; : P W USiNG ":##"; (TicEhpsed \ 3600); (TimcElapscd \ 60): (T imeElapsed MOD 60); : P M USMG "###.###.#"; TcmpA: TempB: TcmpC. Ai.Temp: DeltriTcmp: tncrcasingGnd: Pressure

LOCATE 2. 10: P W TIMES: LOCATE 2.65: P M DATE.

LOCATE 7. 5 : P M "DcltriTemp ........" LOCATE 7.20: PFUNT USING "##.###.##": DelbTemp; In~sïngGrrid; MmTemp

SLEEP 3

P M #1. "BIT7=Ow: LûCAïE 13.35: PRiNT " OFF " Twns heating clemcnt off

LOCATE 14.35: COLOR 4: P W " Temp Gndient too High ": COLOR 15 ' Tunis elemcnt OFF Bit lû=OW

END IF

iF (DeitaTemp < IncreasingGrad) THEN PR[NT#I. "BIT7=In: LOCATE 13.35: PRINTn ON ": LOCATE 1435: COLOR6: PRMT"

Temp Gndicnt OK ": COLOR 15 ' Turns eIement ON ENDIF

IF (TempA > Mz-Temp) OR (TcmpB >= MiiuTemp) OR (TempC >= MxxTemp) T'KEN

BEEP PEUNT #1. "BII7=Ou: LOCATE 13.35: PRMT " OFF "; : LOCATE 14.35: COLOR 4: PRDE

" One of the tcrmal coupies is too HOT": COLOR 15 ' Tunis heamg element OfF sec= Sec+ 2 5 END IF

SLEEP SnoozeTime

P R N ï #2. TIMES; : P W #2. U S N G H5S; TimeElapscd. Curing P M #2. USMG H56: TempA TempB: TempC: AvTemp: DeltaTemp: IncreasingGrad: Ressut

AvTempûld = AvTemp OldPrcssurc = Ressure

N = .Y + SnoozcTime + Sec TimcElripscd = TimeElapd + SnoozeTime + Sec Scc = 0

LOOP

LOCATE 16, 1: P M #2. "End of Tcmpcraturc Raise cycle" END S U B

SUB TERMINATE {,Y) CLS COLOR 7. O. 0: CLS : COLOR 70. O. O

P M " ": PFUNT" ": P m " ":PFUNTn " P R r ? w "

******************************************** PRIhT " THE PROGRAM HAS BEEN TERMMATED n

PRNI-" ******************************************************************************-

COLOR 7. O . O END END Sm

Testing Cornputer Program

Tensile Testing

DECLARE SUB DCREATE (NAMES. TL!. AREA! ) DECLARE SUB STRAIN (ADD2!. U!. L!. TL!. EX!. S m ! ) DECL ARE SUI3 EXT (ADD2!. U!. L! . TL!. DL! ) DECLARE SUB OPTO (ADDl!. ER!. FIBYTE!. MSYTE!. LBYTE!) DECLARE SUE LCELL (ADDZ!. SF!. OFS!, W!) DECLARE SUB SERVO (ADD2!. V!) DECLARE SUB RELAI' (ADD2!. R!) DECLARE SUB TESTSYS (ADDI!. ADDL!) REM VARIABLES FOR TENSiLE TESTER REM OPTICAL ENCODER READER: ADDl = 220h = 544 REM DATA AQUISITfON BOARD: ADD2 = 300h = 768 REM SERVO = SERVO CONTROL VOLTAGE REM LCELL = LOADCELL READING (SCALED) REM POSIT = OPTICAL ENCODER RE ADMG (SCALED) REM E.'YTEN = EXTENSIOMETER READMG REM DLD = DELTA LOAD OR LOAD INCREMENT REM DT=DWELLTIMEFORSAMPLE M AUTO = RELAY (R) = 1 REM STOP = RELAY (RI = 2 REM START = RELAY (R) = 4 REM RETURN = RELAY (R) = 8

REM CONSTANT COEFFICENTS DFS = "TEST.DATn MAXLD = 3000 DLD = 50 n=1 x = .25 Y = l D T = 8 M.4YSTIW = 15

5 ADDI = S U ADDZ = 768 SPD = 0 m = o E k X N = O SF = 13.46 12 'SCALING FACTOR FOR LOADCELL ALGORYTHM OFS = 25245.94 'OFFSET VALUE FOR LOADCELL ALGORYTHM ZER = O AREA = 1 U=.19 'UPPER L M T FOR EXTENSIOMETER L = -.O9 'LOWER L M FOR EXENSIOMETER n= I 'DEFAULT GAGE LENGTH = 1 .O INCHES UL = .5 'DEFAULT MAX POSITIVE SERVO CONTROL VALUE LL = -.5 'DEFAULT MAX NEGATiVE SERVO CONTROL VALUE Iif = -.O08 'PROPORTIONAL COEFFICIENT I P = O 'MTEGR4L COEFFiCENT KD = .O0 1 'DERIVATIVE COEFFICIENT KV = .O04 VELAV = O

DWAV = O TSAMP = .25 m'IT=o V=-.13 z = O P = O

CALL SERVQADDZ. O ) 'SET SERVO CONTROL VOLTAGE TO ZERO CALL RELAY(ADD2. O) ' S i 3 ALL OUTPUT RELAYS TO ZERO = MANUAL OPERATION

D M STRN(255) 'DIMENSION ARRAY FOR STRAM DATA READMGS DIM STRS(Z55) ' STRESS DATA READMGS DIM E.?CR+l(255) ' EXTENSION DATA READMGS Dihl LDC(Z55) ' LOADCELL DATA READiNGS

REM AbTOMATED TESTING OF SAMPLE

REM PROGRAM TO ENTER DATA FOR TEST CLS ?RINT "THIS PROGRAM OPERATES THE TESTER UNDER A CONSTANT LEAD SCREW VELOCITY1' PRINT "SAMPLE READlNGS ARE T.4KEN EVERY 0.25 SECONDS UNTU. THE SAMPLE BREAKS" PNNT "OR THE W h l U M LOAD IS REACHED" PRINT

10 CALL REL.QY(ADD2. O ) CALL SERVQADDZ. O) N'UT "DO YOU WANT TO START A NEW TEST (YIN OR REPEAT A TEST (RI' ". D$ IF DS = "N" Go 85 IF D$ = "R" GOTO 50 INPUT "USE DEFALTLT DATA ??? Y/N ". AN$ IF AN$ = "Y" T'KEN GOTO 50 INPUT "ENTER FiLE NAME FOR RAW DATA ". DFS INPUT "ENTER FILE NAME FOR SELECTED DATA ". SDF$ INPUT "ENTER DATE DD/MM/YY ". DA- INPUT "ENTER GAGE LENGTH ( INCES) ". TL INPUT "ENTER SAMPLE WIDTH (MCHES) ". X INPUT "ENTER SAMPLE DEITH (MCHES) ". Y CLS P W "RAW DATA FlLE NAME = ". DFS P M "SELECT DATA F L E NAME = ". SDFS PRINT "CURRENT DATE = ". DAT$ P M "GAGE LENGTH = ". TL P m "SAMPLE WiDTH = ". X PRINT "SAMPLE DEITH = ". Y PiUNT "SAMPLE CROSS AREA = ". X * Y NPLT "IS THIS DATA C O W C T ? YfN ". G$ iF G$ = "Nu THEN GOTO 10

15 WUT "ENTER MA- LOAD ". MAXLD MPUT "EN?ER MAXIMUM EXPECTED STRN% ", MAXSTRN INPUT "ARE YOU SURE '??? YM ". Q$ IFQ$="NnGOT015

REM PROGRAM FOR WRITMG DATA TO FILE

50 CALL SERVO(ADD2. O) OPEN DFS FOR OUTPUT AS # 1 PlUNT # 1, "FILE NAME = ". TAB( 17): DFS P M # 1 . "CJJRRENT DATE = ": TAN 17); DAT$ P M # 1. "GAGE LENGTH = "; TM( 17): TL PRMT #1. "SAMPLE WIDTH = "; TAB(17): X PRiNT # 1. "SAMPLE DEITH = ": TAB(17); Y PRiNT #1. "SAMPLE CROSS AREA = ": TAB(21); X Y PFt..INT # 1. PRlNT # 1 . PRINT #1. "LOAD (LBs)". "STRESS ". "EXTENSION (in)"; TAB(J8): "STWJN": TAB(6O):

" S M (O/*)"

CLOSE # 1 OPEN SDFS FOR OUTPUT AS #2 PRïNT #2. "FILE NAME = ". TAB(17): SDFS P M #2. "CURENT DATE = ": TAB(1T): DAI'S P M #2. "GAGE LENGTH = "; TAB(17): TL P M #2. "SAMPLE WIDTH = "; TAB( 17): X P M #2. "SAMPLE DEPTH = ": TAB( 17); Y P M #2. "SAMPLE CROSS AREA = ": TAB(Z1): X Y P M K. P M #2. P M #2, "LOAD (LBs)". "STRESS ". "EXENSION (in)"; TAB(48); "STRAM": TAB(60):

" S W (?/O)" CLOSE #3

R = O CLS

PRINT "MAMMtiN LOAD = ". MAXLD CALL LCELL(ADD2. SF. OFS. W WUT "1s SAMPLE READY FORTESTMG??? YM ". PS IF P$ = "N" GOTO 50 PRMT "DO YOU WANT T O START TEST ?? LOADCELL IS READMG ". W INPUT "1s THIS OC; ? Y M ". PS IF P$ = "N" GOTO 50 PRINT TESTMG HAS STARTED: PRESS (X) TO STOP TESTING "

REM SET RELAY S FOR AUTOMATED OPERATION BEEP CALL SERVO(ADD2. O) 'SET SERVO VOLTAGE TO ZERO CALL RELAY(ADD2.1) 'SET RELAYS FOR AUTOMATED OPEFUTION TO = TIMER

55 IF (TIMER - TO) < 1 THEN GOTO 55 'WAIT 1 SECOND CALL RELAY(ADDZ.3) 'SET RELAY FOR RESET TO = TIMICR BEEP

56 IF (TIMER - TO) <= 3 THEN GOTO 56 'WAIT 1 SECOND

CALL RELAY(ADD2.5) 'START SYSTEM TO = TIMER BEEP

57 IF (TTMER - TO) C= 1 THEN GOTO 57 'MOMENTARILY HOLD START RELAY CALL RELAY(ADD2.1)

CLS SCREEN 9 LOCATE 1.1 PRINT "LOADn: TAB(I0): "STRAiNn LINE ( 5 . 50)-(505.320). . B

REM AUTOMATED TESTTNG OF SAMPLE

CALL OPTQADD 1.1. HBYTE. MEYTE. LBYTE) 'RESET ENCODER TO ZERO T3 = TIMER

REM DATA RECORDING L O P : SAMPLE READMGS EVERY TSAMP = 0.25 SECONDS VAV = 0 VCON-r= - 1 TST = TIMER CALL SERC'QADDZ. 3)

58 IF(TIMER-TSc13<=.jTHENGOTO58

60 Tl = TIMER IF (Tl - TZ) <= TSAMP THEN GOTO 79

C U L STRAiN(ADD2. U. L. TL. EX. STRM CALL LCELL(ADD2. SF. OFS. W)

STRS = W / ( X * Y)

OPEN DFE FOR APPEND AS # 1 PFUNT #1. W. STRS. EX; TAB(J8): STRN: TAB(60); STRN 100 CLOSE # 1 T2 = TLMER IF c W > WO) AND (STRN > STRNO) THEN GOTO 61 ELSE GOTO 62

6 1 OPEN SDFS FOR APPEND AS #Z PRiNT #2. W. STRS. EX TAB(48): STRN: TAB(60): STRN * 100 CLOSE #2 W O = W STRNO = STRN

62 REM DRAWDATATOGRAPHX GSTRN = iNT(SïRN * 10000) 5 1 MAXSTRN

GLOAD = - M ( W / 10) 2700 / MAXLD PSET (GSTRN + 10. GLOAD + 3 IO)

REM CLOSED LOOP CONTROL OF LEAD-SCREW VELOCITY

REM TESTS FOR EXITMG CONTROL LOOP

79 CALL STRAiN(ADD2. U. L, TL. E X STRN)

CALL LCELL(ADD2. SF. OFS. W)

ff W >= blAXLD THEN GOTO 98 iF t STRN * 100) >= MAXSTRN THEN GOTO 98 IF W > 3 0 T H E N Z = 1 BS = W Y S IF B$ = "X" THEN GOTO 85 iF B$ = "Sn THEN GOTO 97 C.4LL OFTO(ADD 1. O. HBYTE. MBYTE. LBYTE) DX = MBYTE * 256 + LBYTE

IF ( W <= 50) AND (2 = 1 ) THEN GOTO 98

REM CELOCINCONTROL

CALL OPTO(PJ)D 1. O. HBYTE. MBYTE. LBYTE) Tl =TIMER D X 1 = MBYTE * 256 + LBYTE

81 LF(T1MER-T1)<=.02THENGOTOSl T5 = TIMER C U L OPTO(.4DD1. O. HBk"ïE. MBk'TE. LBYTE) DXO = MBYTE * 256 + LBYTE VEL=(D,YO-DXl)I(T3 - T l ) VAV = (VAV + VEL) / Z \ E R R = ( V E R R + ( W - V . 4 V ) * W ) / Z

VCONT = (VCONT -+ V + VERR) / 2 IF W < 3 O T K € N VCOhT=-25 IF VCONT >= UL THEN VCONT = UL IF VCONT <= LL THEN VCONT = LL CALL SERVQADDZ. -VCONT)

LOCATE 3. 1 P M USiNG "##ff#.##": W: Tm( 10): STRN * IO0

GOTO 60 97 CALL SERVQADDZ, O)

PRihT "TESTING STOPPED" M U T "DO YOU WANT TO C O m (Cl OR RETURN HEADER (R) ". AS IFAS="C"GOTO60 ff AS = "Rn GOTO 84

98 CALL SERVO(ADD2, O) WUT "MAX LOAD REACHED: PRESS (R) TO RETURN HEADER (X) TO EXIT". AS IF .a = "RN THEN GOTO 84 iF A$ = "X" THEN GOTO 85

84 REM RETURN KEAD TO ZERO

REM ETWRN HEAD TO ORIGINAL POSITION

83 C.4LL OPTO(ADD1. O. HBYTE. MBYTE, LBYTE) MS = W Y S E M S = "Y THENGOTO85 rFMS="SNTHENP= 1 iFM$="CmTHENP=O IFP= i THENGOTO83 DX = MBYTE 256 + LBYTE VERROR = DX * K IF VERROR >= UL THEN VERROR = UL IF VERROR <= LL THEN VERROR = LL CALL SERVO(ADD2. -VERROR) PRINT DX VERROR IF DX >= 10 THEN GOTO 83 CALL SERVqADDZ. O ) PRiNT "HEPJ) AT ZERO INITAL POSITION"

CALL RELAY(ADD2. O) GOTO 5

'SHUT OFF SYSTEM

85 CALL SERVO(ADD2. O ) CALL RELAY(ADD2. O)

END

SUB LCELL (ADD2, SF. OFS. W) REM THIS PROGRAM READ THE N D CONVERTER C H A ! ! L = O REM SF = SCALiNG FACTOR REM OFS = OFFSET FACTOR FOR ZERO REM LOAD = SF*(AID READING) + OFS REM SAMP = NUMBER OF SAMPLE READINGS REM AVG = AVERAGE READWG FROM LOADCELL

S M = - ! W=O C - O OUT ADDZ + 2. O

860 OUT ADDZ + 1. O 'CONFIGURE AD FOR C H W L O 'START MD CONVERSION

880 IF W(ADD2 + 2) >= 128 THEN GCYïO 880 'CHECK FOR COMPLETION

HBYTE = iNP(ADD2 + 1) * 16 LBYTE = iNT(INP(ADD2) / 16) WGKT = HBYTE + LBYTE W = INT(SF * WGHT - OFS) iFABS(W-TW)>=5TKENC=O C = C + I

IF C >= SAME THEN GOTO 890 T W = W GOrO 860

890 END SUB

SUI3 OPTO (ADD 1. ZER HBYTE, MBYTE, LBYTE)

IF ZER = O THEN GOTO 900

REM MASTER CONTROL REGISTER: PRESET THE OPllCAL ENCODER COUNTER TO ZERO OüTADDl + 1-32

REM OUTPUT CONTROL REGISTER 900 OüT ADDI + 1. 128

REM INPUT CONTROL REGISTER: NORMAL OPERATION AND A&B W U T S ENABLED OUT ADDl + 1. 104

REM QUADRATURE CONTROL REGISTER = MODE X 1 OUT ADD1 + 1.193

REM MASTER CONTROL REGISTER OUT ADDl + 1.3

REM COUNTER OüTPUT LATCH REM Y = MP(ADD1 + 8) LBk'TE = INP(ADD1) hiB\rTE = PP( ADD 1 1 HBYTE = INP(ADD 1 )

END SUB

SUI3 RELAY (ACDL R)

OUT ADD2 + 3. R

SUB SERVO (ADDL V) S = INT((V + 1) * 20-17) HEYTE = iNT(S / 256) LBYTE = S - HBYTE * 2% OUT ADDZ + 4. LBYTE OLT ADDZ A 5. HBYTE z = rNP( .4DDZ + 5 )

E N û SUB

SUB STRAiN (ADD2. U. L. TL. EX STRN)

REM THIS SU8ROUTiNE READS THE OUTPUT VOLTAGE FROM THE EXTENSIOMETER REM THROUGH A D Ch. 1

REM CONFIGURE A/D CHANNEL & START CONVERSION OUT ADD2 + 2.1 'CONFIGURE A D FOR CHANNEL #1 OUT ADD2 + 1, O 'START CONVERSION

REM TEST FOR COMPLETION OF A/D CONVERSION CYCLE 2000 IF W ( A D D 2 + 2) >= 128 THEN GOTO ZOO0

HBYTE = iNP(ADD2 + 1) * 16 LBYTE = INT(MP(ADD2) / 16)

EX = -((U - L) 1 4007) (HBYTE + LBYTE) + -19 EX = M I E X * 10000) 1 1000 'ROUND OFF NUMBER TO NEAREST 1/ 10000 INCH STRN = EX / TL

SUB TESTSYS (ADD 1, ADD2)

PRINT "THIS PROGRAM WILL TEST THE VARIOUS CONTROLS FOR THE TENSILE TESTER"

20 WüT "SYSTEM TO TEST: RELAYS. LOADCELL. SERVO. OPTO. END ". BS IF BS = "RELAYS" THEN CiOTO 100 IF BS = "L0.4DCELLw THEN GOTO 200 IF BS = "SERVO" THEN GOTO 300 IF ES = "OPTO" THEN GOTO 400 IF BS = "EM>" THEN GOTO 500

GOTO 20

REM TEST RELAY S 100 CLS INPUT "ENTER RELAY NUMBER: O , 1.2.3.4 (666 ENDS RELAY TEST) *. R IF R = 666 THEN GOTO 20 R = NI-(? " (R - 1)) OUT ADD2 +- 3, R GOTO 1 O0

REM READ VALUES FROM MAiN LOAD CELL 200 CLS

PFUhT " THlS TEST READS THE OUTPUT FROM LOAD CELL" PRINT " OPERATE TENSILE MACHINE MANUALLY AM) SUBJECT " PRlNT " TEST SAMFLE TO A LOAD: PRESS S=START. Z=ZERO. C=END TEST "

210 KS = INKEYS N = 0 IF KS = "S" THEN GOTO 250 IF KS = "C" THEN GOTO 20 IF KS = "ZN THEN N = i m o 2 10

250 IF I?KEY$ = "CM THEN GOTO 20 REM S A M P - O REM AVG=O REM FOR SAMP = 1 TO 100 REM OUTADD2+2.0 REM OüT ADD2 + 1. O REM 260 IF INP(ADD2 + 2) >= 128 THEN GOTO 260 REM HBYTE=MP(ADD2+ 1 ) * 16 REM LBYTE = N(lNP(ADD2 + 2) / 16) REM LCELL = HBYTE + LBYTE REM AVG = AVG + LCELL REM NE.- S M REM WEIGHT = AVG / 100 - NUL

REM GEhERATE VOLTAGE FOR SERVO AMPLIFIER 300 NPUT "ENTER A VALUE -1 TO +l . 666 = EXITS ". E IFE=666THENGOTO20 V = IhT((E + 1) * 2037) HBYTE = N ( V 1 256) LBYTE = V - HBYTE * 256 OUT ADDZ + 4. LBYTE OLT ADDZ + 5. HBYTE

REM UPDATE DIGiTAUA PRINT m m . E Z = W ( A D D 2 + 3 )

LOG CONVERTER

GOTO 300

REM TEST OPTICAL ENCODER READER JO0 OUT ADD 1 -. 1.64

OUTADDl + 1. 193 OUT ADDI + 1. 1 OUT ADDI + 1.3 BYTE0 = INP(ADD 1 ) BYTE 1 = INP(ADD I ) BYTE? = MP(ADDI)

PRiYT BYTEO. BYTEI. BYTE3 [F i h W Y S = "C" GOTO 20 m o 400

Step by Step testing

DECLARE SUE? BLANK @F$. SDFS) DECL.ARE SUB DCREATE (NAME.$. TL!. AREA!) DECLARE SL! S T R M N (ADD2!. U!. L!. TL!. EX!. STRN!) DECLARE SUB EXT (ADDZ!. U!. L!. TL!. DL!) DECLARE SUB OPTO (ADDI!. ZER!. HBYTE!, MBYTE!. LBYTE!) DECLARE SUB LCELL (ADDZ!. SF!. OFS!. W!) DECLARE SUB SERVO (ADDZ!. V!) DECLARE SUB RELAY IADDL!. R!) DECLARE SUB TESTSYS (ADDI!, ADD2!) REM VARIABLES FOR T'ENSILE TESTER REM OFTICAL ENCODER READER: ADDI= 220h = 544 REM DATA AQUlSITION BOARD: ADD2 = 3ûûh = 768 REM SERVO = SERVO CONTROL VOLTAGE REM LCELL = LOADCELL READmG (SCALED) REM POSIT = OPllCAL ENCODER READING (SCALED) REM EdYTEN = EXTENSIOhETER READMG REM DLD = DELTA LOAD OR LOAD INCREMENT REM DT=DWELLTiMEFORSAMPLE REM AUTO = RELAY (R) = 1 REM STOP = RELAY (R) = 2 REM START = RELAY (R) = 4 REM RETüW = RELAY (R) = 8

REM CONSTANT COEFFICENTS DFS = "TEST.DATU h.1.4XLD = 5000 DLD = 1(W SDLD = 500 TL= 1 5 = .25 Y = 1 D T = 8 bf AYSTRN = i 5

5 ADDi =54J ADD2 = 768 SPD = O t u I l = O E.XTEN = O SF = 13.46 12 'SCALiNG FACTOR FOR LOADCELL ALGORYTHM OFS = 25245.94 'OFFSET VALUE FOR LOADCELL ALGORYTHM ZER = O AREA = 1 U = .19 'UPPER LMlT FOR EXTENSIOMETER L = -.O9 'LOWER L M T FOR EXTENSIOMETER TL= 1 'DEFALLT GAGE LENGTH = 1.0 MCHES UL = .5 'DEFAULT MAX POSïilVE SERVO CONTROL VALUE LL = -.5 'DEFALJLT MAX NEGATlVE SERVO CONTROL VALUE W = -.NI8 'PROPORTIONAL COEFFICIENT IP = O 'INTEGRAL COEFFICENT KD = .O0 1 'DERIVATIVE COEFFICIENT KV = ,004 V'ELAV = O

DWAV = O TSAh4P = 2 5 WNIT = 0 v = - . l ? z = O P=O C N T = O 'INITIAL VALUE FOR REPEAT CYCLE COUNTER (SEE SUB HEAD)

REhI MITIALIZE SYSTEM

CALL SERVO(ADDZ.0) 'SET SERVO CONTROL VOLTAGE TO ZERO CALL RELAY(ADD2. O) 'SET ALL OUTPUT RELAYS TO ZERO = MANUAL OPERATION

DiM STRN(Z55) 'DIMENSION ARRAY FOR S W DATA READMGS DiM STRS(255) ' STRESS DATA READiNGS DIM E.XïK(255) ' EXTENSION DATA READMGS DIM LDC(Zj5) ' LOADCELL DATA READiNGS

REM AüTOMATED TESTING OF SAMPLE

REM PROGRAM TO ENTER DATA FOR TEST CLS PRiNT "THIS PROGRAM OPERATES THE ESTER LINDER A C 0 N S T . W LEAD SCREW VELoCI-IYW PRINT "SAMPLE READINGS ARE TAKEN EVERY 0.25 SECONDS." f FUNT "THE SAMPLE IS REPEATEDLY CYCLED UNDER UNFORMLY INCREASING LOADS" P M "UNTiL THE SAMPLE BEAKS" PWNT

10 CALL RELAY( ADDZ. O ) CALL SERVO(ADD2. O ) WUT "DO YOU W M T O REPEPLT A TEST '! (R) OR START A NEW TEST '' (Y/N) ". D$ IF DS = "3" GOTO 85 IF DS = "RN GOTO 50 INPUT "USE DEFAULT DATA ?1'? Y/N ". AN$ IF ANS = "Y" THEN GOTO 50 INPUT "ENTER FILE NAME FOR RAW DATA ". DFS M U T "ENTER FELE NAME FOR SELECTED DATA ". SDFG M U T "ENTER DATE DD/MM/YY ". DATS MUT "ENTER GAGE LENGTH (MCHES) *. n W L T "ENTER SAMPLE WtDTH (INCHES) ". X WUT "ENTER SAMPLE DEPTH (MCKES) ". Y CLS PRNT "R4W DATA FILE NAME = ". DFS PRJW "SELECT DATA FlLE NAME = ", SDF$ P W "CURRENT DATE = ". D A I 3 PEUNT "GAGE LENGTH = ". TL PRINT "SAMPLE WIDTH = ". X P M "SAMPLE DEFTi-i = ". Y PRNT "SAMPLE CROSS AREA = ". ,Y * Y MPUT "1s THIS DATA CORRECT ? YM ". GS IF GS = "Nu THEN GOTO 10

15 MPUT "ENTER MAXIMUM LOAD ". MAXLD INPUT "ENTER LOAD [NCREMENTS FOR CYCLES ". DLD rwuT * E ~ R LOAD m w FOR CYCLE START m. SDLD MPUT "ENTER MAXiMUM EXPECTED STRN% ". MAXSTRN M U T "ARE YOU SURE ??? YIN ", Q$ IF QS = "Nn GOTO 15

REM PROGRAM FOR WRITTNG DATA TO FILE

50 CALL SERVO(ADD2. O ) OPEN DF$ FOR OUTPUT AS # 1 PRINT # 1. "FILE NAME = ". TAB( 19); DF6 P M #1. "CURRENT DATE = "; TAB(19); DATE PRiNT #L. "DELTA LOAD = ". TAB(19); DLD P R N ï # 1. "GAGE LENGTH = ": TAB( 19); TL P M #1. "SAMPLE WZDTH = ": TAB(19): X PRINT # 1. " S.WLE DEPTH = ": TAB( 19): Y PMNT # 1. "S.4MPLE CROSS AREA = ": TAB(2 1): X * Y PRlNT $1. P W S l . P M # 1. "LOAD (LBs)". "STRESS ". "EXTENSION (in)": TAB(48): "STRAIN": TAB(60):

"STRAIN (%)"

CLOSE # 1 OPEN SDFS FOR OUTPUT AS #1! PFüNT #2. "FILE NAME = ". TAB( 19): SDFS PRINT it2. "CURRENT DATE = ": TAB( 19): DATS P M #2. "DELTA LOAD = ". TAB( 19): DLD P R l W #2. "GAGE LENGTH = ": TAB( 19); TL PRiNT #2. "SAMPLE WTDTH = ": TA& 19): X P M #2. "SAMPLE D E m = ": TM( 19): Y PRIW #2. "SAMPLE CROSS AREA = ": TAB(2 1); X * Y PRINT #2. P N N T #2. P N N T #2. "LOAD (LBs)". "STRESS ". "EXTENSION (in)": TAB(J8): "STRAM": TAB(60):

" STRAIN (%) " CLOSE #Z

R = O CLS

PRINT "M LOAD = ". MAXLD PIUNT "LOAD ICREMENTS = ". DLD P W "CYCLIC START LOAD = ". SDLD CALL LCELL(ADD2. SF. OFS. W INPUT "1s SAMPLE READY FOR TESTMG ??? YM ". P$ IF P$ = "N" GOïO 50 P M "Dû YOU WANT TO START TEST ?? LOADCELL IS READiNG ". W WUT "IS THIS OK ? Y/N ". P$ IFPS="NM GOTOSO PiüXT 'TESTMG H M STARTED: PRESS (X) TO STOP TESTMG "

REM SET RELAYS FOR AUTOMATED OPERATION BEEP CALL SERVO(ADD2, O) 'SET SERVO VOLTAGE TO ZERO

CALL RELAY(ADD2.1) 'SET RELAYS FOR AUTOMATED OPERATION TO = TIMER

5 5 LF (TUIIER - TO) < 1 THEN GOTO 55 'WAIT 1 SECOND CALL RELAY(ADD2.3) 'SET RELAY FOR RESET TO = T7MER BEEP

56 IF (TIMER - TO) <= 2 THEN GOTO 56 'WAIT 1 SECOND CALL RELAY(ADD2.5) 'START SY STEM TO = TiMER BEEP

57 IF (TLkER - TO) <= 1 THEN GOTO 57 'MOMENTARICY HOLD START RELAY CALL RELAY(ADD2.1)

CLS SCREEN 9 LOCATE 1 . 1 PRINT "LOAD"; TAB( IO): "STRAM": TAB(20): "CYCLE LOAD L M T " LINE ( 5 , 51))-(505. 320). . B

REM AWOMATEDTESTMGOFSAMPLE

CALL OPTO( A D D 1 . 1 . HBtTE. MBYTE. LBYTE) 'RESET ENCODER TO ZERO T2 = TïMER

REM DATA RECORDiNG LOOP: SAMPLE READWGS EVERY TSAMP = 0.25 SECONDS v.4v = O VCONT = - t TST = TlMER CALL SERVû(ADD3. . S ) F (TIMER - TST) <= .5 THEN GOTO 58

C = O Tl = TTMER IF (Tl - T2) <= TSAMP THEN GOTO 79

CALL STRATN(ADD2. U. L. TL. EX STR!! CALL LCELL(ADD2. SF. OFS. W)

OPEN DF$ FOR APPEND AS # 1 PRINT # 1. W. STRS. E4X TAB(J8); S M : TAB(60); STRN * 100 CLOSE #l T2 = TIMER rr: c W > WO) AM) (STRN > STRNO) THEN GOTO 6 1 ELSE GOTO 62 OPEN SDFS FOR APPEND AS #2

PWNT #2. W. SfRS. EX: TAB(48); STRN; TAB(60): STRN * 100 CLOSE #2 W O = W s r n o = S m

REM DRAW DATA TO GRAPH GSTRN = INT(STRN * 10000) * 5 / MAXSTRN GLOAD = - M ( W / 10) 2700 1 MAXLD PSET (GSTRN + 10, GLOAD + 3 10)

REM CLOSED LOOP CONTROL OF LEAD-SCREW VELOClTY

REM TESTS FOR EXïiTNG CONTROL LOOP

79 CALL STRAiN(ADD2, U. L. TL. EX STRN) CALL LCELL(ADD2. SF. OFS. W)

S T R S = W / ( X * Y ) IF W >= (Cm * DLD + SDLD) THEN C = 1 IF C = lTHENZ=O [FC= l T H E N G a r O 8 4 IF W >= MAXLD THEN GOTO 98 IF (STRN * 100) >= MAXSTRN THEN GOTO 98 IF W>3OOTHENZ= 1 ES = MKEYS IF £33 = "X" T'KEN GOTO 85 IF Bf = "S" THEN GOTO 97 CALL OPTO(ADD1- O. BYTE. MBYTE. LBYTE) DN = MBYTE 256 + LBYTE

IF ( W <= 50) AND (Z = 1 ) THEN GCYïO 98

REM VELOClTY CONTROL

CALL OPTO(ADD1. O. KBYTE. M B m . L B m ) Tl =TiMER DX1= MBYTE * 256 + LBYTE

81 IF(TIMER-T1)<=.02THENGOTO81 T3 = TIMER CALL OPTO(ADD 1. O. HBYTE. MBYTE. LBYTE) DXO = m Y T E 256 + LBYTE VEL=(DXO-DXl)/(T3 -Tl) VAV = (VAV + E L ) / 2 VERR = (VERR + (VMIT - VAV) * KP) / 2

VCONT = (VCONT + V + VERR) / 2 IF W < 3 0 THEN VCONT = -.25 IF VCONT >= UL THEN VCOElT = UL IF VCONT <= LL THEN VCONT = Li, CALL SERVO(ADD2, -VCONT)

LOCATE 3.1 PRINT USNG "####.##": W; TAB(I0); STRN 100: TAB(20): CNT DLD + SDLD

GûTO 60 97 CALL SERVO(ADD2. O)

P M "TESTING STOPPED" M U T "Dû YOU WANT TO CONTrNUE (C) OR RETURN HEADER (R) ". AS F A $ = "C"GOTO6O IF A$ = "RaTHENC=O IF A$="R"THENP=O E A$ = "R* GOTO 83

98 CALL SERVû(ADD2. O) MPüT "MAX LOAD REACHED: PRESS (R) TO RETURN HEADER ( X ) TO EXIT". AS iF .AS = "Rn THEN GCKO û4 iF AS = "Xn THEN GOTO 85

84 REM RETURN HEAD TO ZERO CALL SERVO(ADD2. O ) P = 1 LOCATE 6. 1 P M "CYCLIC LOAD REACHED. PRESS C TO CONTIMIE" P R I M ' "PRESS X TO EXIT"

REM RETURN iEAD TO ORIGINAL POSITION K = .O02

83 CALL OPTQADDI. O. BYTE. MBYTE. LBYTE) CALL LCELL(ADD2. SF. OFS. W) CALL STRAIN(ADD2. U. L. TL. EX. STRN) LOCATE 3. 1 P M USING "###.##": W: TAB(10): STRN "100: TAB(2O); CNT * DLD + SDLD

MS = W Y S IF MS = "X" THEN GOTO 85 IFMS="S"THENP= 1 iF MS = "C" THEN P = 0 IFM$="R"THENC=O

I F P = YrHENGOTTO83 DS = MBYTE * 256 + LBYTE VERROR = DX K IF VERROR >= UL THEN VERROR = LJL IF VERROR <= LL THEN VERROR = LL C ALL SERVO(ADD2. -VERROR) IF (C = 1 ) AND(W >25)THEN GûTO83 IF (C = 0) AND (DX >= 10) THEN GOTO 83 CALL SERVO(ADD2. O ) TW = TIMER

87 IF (TIMER - TW) < 2 THEN GOTO 87 IFC= lTHENCNT=CNT+ 1 IF C = 1 THEN CALL BLANK(DF$. SDFS) iFC= 1 THEN W O = O F C = lTHENSTRNO=O IFC= lTKENGOïO6û

P M "HEAD AT ZERO NITAL POSITION"

CALL RELAY(ADD2, O) GOTO 5

'SHUT OFF SY STEM

85 CALL SERVO(ADD2.O) CALL RELAY(ADD2. O)

rn

SUB BLANK (DFS. SDFS) OPEN DFS FOR APPEND AS # 1 PWNT #1* P W #1. CLOSE #I

OPEN SDFS FOR APPEM) AS #2 PRMT #2. P m K. CLOSE #2

END SUB

SUB LCELL (ADDL SF. OFS. W REM THIS PROGRAM READ THE AID CONVERTER C H M L = O REM SF = SCALING FACTOR REM OFS = OFFSET FACTOR FOR ZERO REM LOAD = SFS(.4/D READiNG) + OFS REM SAMP = NüMBER OF SAMPLE READINGS REM AVG = AVERAGE READMG FROM LOADCELL

SAME = 4 T W = O C = O O W ADDZ + 2. o

86(i OUT ADDZ + 1. O 'CONFIGURE A/D FOR CHANNEL O

'START A D CONVERSION

880 IF INP(ADD2 + 2) >= 128 THEN W O 880 'CHECK FOR COMPLETION

HBYTE = W ( A D D 2 + 1) 16 LBYTE = MT(W(ADD2) / 16) WGHT = KBYTE + LBYTE W = N ( S F * WGHT - OFS) lFABS(W-TW)>=STHENC=O c = c - i

890 END SüB

SUB OPTO (ADDI. ZER HBYTE, MBYTE. LBYTE) IF ZER = O THEN GOTO 900

REM MASTER CONTROL REGISTER: PRESET THE OFTICAL ENCODER COUNTER TO ZERO OUT .4DD1+ 1.32

REM OUTPL! CONTROL REGISTER 900 OUT ADDl4 1.128

REM INPUT CONTROL REGISTER: NORMAL OPERATION AND A&EJ INPUTS ENABLED OUT ADDI+ 1.104

REM QUADRATURE CONTROL REGISTER = MODE X1

OUT ADDl + 1. 193

REM MASTER CONTROL REGISTER OüTADDl+ 1.3

REM COUNTER OUTPUT LATCH REM Y = MP(ADD1+8) LBYTE = MP(ADD 1 ) MBYTE = INP(ADD1) HBYTE = PP(ADD1)

END SUB

SUB RELAY (ADDZ. R)

OUT ADD2 + 3. R

END SUE3

SUB SERVO (ADDZ. V) S = IW((V + 1) * 2047) HBYTE = N U S / 256) LBYTE = S - HBYTE * 256 OLT .OD2 + 4. LBYTE OUT ADDt + 5 . KBITE Z = INP(ADD2 + 3)

END Sm

REM ' I W S SLIBROUTINE READS THE OUTPUT VOLTAGE FROM THE EXTENSIOMETER REM THROUGH N D Ch. 1

REM CONFIGURE N D C M W L & START CONVERSION OUT ADDZ A 2 . 1 'CONFIGURE A/D FOR CHANNEL # 1 OLT ADD2 + 1. O 'START CONVERSION

REM TEST FOR COMPLETION OF An> CONVERSION CYCLE 2000 IF W(ADD2 + 2) >= 128 THEN GOTO 2000

EX = -((U - L) / 4097) (HBYTE + LBYTE) + .19 EX = M ( E X * 100001 / 10000 'ROUND OFF NUMBER TO NEAREST 1/10000 iNCH S m = EX/TL

SUB TESTSYS (ADDI. ADD2)

P W "TKiS PROGRAM WILL TEST THE VARIOUS COECTEIOLS FOR THE l'ENSILE TESTER"

20 INPUT "SYSTEM TO TEST: RELAYS. LOADCELL. SERVO. OPTO. END ". BS IF BE = "RELAYS" THEN GûTO 100 IF BS = "LOADCELL" THEN GOTO 200 IF B$ = "SERVO" THEN GOTO 300 IF BS = "OPTO" THEN GOTO JO0 IF BS = "END" THEN GOTO 500

GUTO 20

REM TEST RELAYS 100 CLS INPUT "ENTER RELAY NUMBER: O. 1.2.3.4 (666 ENDS RELAY TEST) ". R IFR=666THENGOTO20 R = MT(2 A (R - 1)) OUT ADDZ + 3. R GOTO 100

REM READ VALLES FROM MAiN LOAD CELL zoo CLS

PRlNT " THIS TEST READS THE OUTPUT FROM LOAD CELL" P M " OPERATE T'ENSILE MACHME MANUALLY AND SUBJECT " PRiNT " TEST SAMPLE TO A LOAD: PRESS S=START. Z=ZERO. C = W TEST "

210 KS = INEiEYS N = o iF KS = "Su THEN GOTO 250 1F KS = "CU THEN GOTO ZO LF KS = "2" THEN N = 1 m o 210

250 IF I N E Y S = "CM THEN GUïO 20 REM SAMP=O REM AVG=O REM FOR SAMP = 1 TO 100 REM OUT ADDZ + 2. O REM OUT ADDZ + 1. 0 REM 260 IF MP(ADDZ + 2) >= 128 TKEN GOTO 260 REM HBYTE = MP(ADDZ + 1) 16 REM LBYTE = ïNT(MP(ADD2 + 2 ) 1 16) REM LCELL = HBYTE + LBYTE REM AVG = AVG + LCELL REM NEXTSAMP REM WEIGHT = AVG / 100 - NUL

PRINT WEIGHT W T O 250

REM GENERATE VOLTAGE FOR SERVO AMPLiFlER 300 INPUT "ENTER A VALUE - 1 TO + 1.466 = EXrrS ". E IF E = 666 THEN GOTO 20 V = N ( ( E + 1) * 2047) HBYTE = N ( V / 256) LBYTE = V - HBYTE * 236 OLT ADDZ + 4. LBYTE OUT ADD2 + 5. HBYTE

REM WDATE DIGiTAUANALOG CONVERTER PFUNT HBYTE. E Z = INP(ADD2 + 3)

GOTO 300

REM TEST OPTICAL ENCODER READER 400 OUT ADDl + 1-64

OUT ADD1 + 1, 193 OUT ADDl + 1 . 1 OLTADDl+ 1.3 BYTE0 = MP(ADD1) BYTE l = iNP(ADD1) BYTE2 = iNP(ADD 1)

PFUNT BYTEO. BYTE1. BYTE3 IF INKEYS = "Cm GOTO 20 m o 400

500 END S U E

APPENDIX B

INTAKE

/

1 RYERSON POLYTECHNK UNIVERSITY

DRVN BYi H, GHAEHl 1 DRAVlNG ND 1

FILM

~A~CXJ~~ELINE RELEASE \

COMPOSITE PLIES

1 / BLEEDER

BLEEDER ! 1

// sEPARATING

1 BREATHER

SEALANT

TIlL FRACTIONAL 1/16 3 PLCS 0,001 4 PLCS 40002 ANGULAR 1 /2'

POLYTECHNIC UNIVERSITY

146

References

A.A. Poukov "A Self-consistent and Statisticai Mechanical Approach for

Determining Effective Elastic Properties of Composites7', Theoretical and Applied

fracture Mechanics. (1999) 1 157-16 1.

P.K. Gotsis*, C.C. Chamis and L. Minnetyan, "Prediction of composite Laminate

Fracture: Micro-mechanical and Progressive Fracture7' Composite Science and

Technology, 1998, 1 137.

Peter A. Zinoviev and Andrey A. Smerdov*, " Ultimate properties of unidirectional

fiber composite ", Composite science and technology, 1999 , P. 625-634.

Dostal. Cyril A. " Engineering material handbook Volume 2, Composites " ASM.

International, Metals Park, Ohio, USA. 1987.

.Hart-Smith, L.J. " Simplified Estimation of Stifiess and Biaxial Strengths for

Design of Lam inated Carbon-epoxy Composite structure. Douglas Aircrafi paper

7548, ( 1985).

Mc Laughlin. P.V. and, Y.W. Chun. '%ILAM - A Composite Laminate Failure

Analysis code Using Bilinear Stress Strain Approximation." Lawrence Livermore

Laboratory. UCRL- 1 53 7 1 1 -0ctober 1980..

Puck. A. and Schurmann, H.. 'Sailure Analysis of FRP Laminates by Means of

Physically Based Phenomenological Models." Compos. Sci. Technol., 1998, 58(7)

1 083.

Wolfe. W.E.and Butalia. T.S., "A Strain -Energy Based Failure Criterion for Non-

linear .4nalysis of Composite Laminates Subjected to Biaxial Loading" Composite

Sci., 1988, 58 (7) 1 107.

Hashin . 2. " Analysis of composite rnaterials: A survey", Joumal of applied

mechanics. 1983, P. 481

10. Tsai. S.W. and Wu, E.M. " A general theory of arength for anisotmpic materials ':

Journal of Composite Material (1971) 5, 55-80.

11. Carl, T. Herkovich, 'Wechanics of Fibrous Composites7' , John Wiley & Sons Inc.,

New York U.S.A., 1998.

12. Siboni, G. " Prediction of the effective elastic properties of multiphase composite

media with orthotropic constituents ': Composite Science and Technology (1994), 50,

293 -298.

13. P. S. Theocaries and C. A. Stassinakies. " Crack propagation in fibrous composite

materids" Journal of composite materials, vol 15 p. 133 Tehchnomic publishing.

1 4 Tsai, Stephen and Hahn, Thomas H. " Introduction to composite materiais "

Technomatic Publishing Company, West port, Connecticut, USA. 1980.

15. Tsai. Stephen " Theory of composite materials " Think Composite Publishing,

Day-tona, Ohio, USA. 1992.

16. Naik, G.N. and Murty, A. V., "A failure Mechanism Based Approach to Design of

Composite Laminate." Composite Structure, 1999, P. 71-80.

17. Bader. M. G. and P. T. Curtis, "the micro-mechanics of fiber composites - Carbon

fiber Program", First report agreements A93BQ5, Dep of Metallurgy and Matenal

Science. Univ of Surrey, U.K. Juiy 1978.

18. A. K. Bledzki, A. Kessler , '7Ietermination of Elastic Constants of GlasdEpoxy

Lrnidirectional Laminate by the Vibration Testing of Plates" , Composite Science and

Technology, April 9,1999, p. 20 1 5-2024.

19. Mc Cartney. L.N., Tredicting Transverse Crack Formation in Cross-Ply Laminates."

Compos, Sci. Technol., 1998, 58 (7), 1069.

20. Donald Flaggs, " Experimental detemination of In Situ transverse lamina strength in

graphitdepoxy laminates ", Lockheed Pa10 Alto Research Laboratory and Lockheed

Missiles and Space Company Inc. March 1982

2 1. J. L. Thomson, " The influence of fiber properties of the performance of glass fiber-

reinforced polyamide " Composite science and technology, 1999, P. 23 15 - 2328.

22. Schwartq Mel M. " Composite material handbook " McGraw Hill, Toronto, Canada,

1984

23. Doaal, Cyril A. " Engineering material handbook, Volume 1, Composites " ASM.

International, Metais Park, Ohio, USA. 1987.

24. Allen Yu and Vijay Gupta *, " Measurements of in situ fibedmatrix interface strength

in graphite / epoxy composites ", Composite science and technology, Febniary 1998,

P. 48.

15. Ramesh Talreja, " transverse cracking and stifiess reduction in composite laminates"

, Journal of composite matenals, vol 19, July 1985, P. 355

26. Siboni, G. " Prediction of the effective elastic properties of multiphase composite

media with orthotropic constituents ", Composite Science and Technology (1994), 50,

293 -298.

27. Pendleton, Richard L. " manual on experimental methods for mechanical testing of

composites " Elsevier Science Publisher LTD. Bethei, CT 06801, USA 1989.

28. F. P. Gerstle, Jr. and E. D. Reedy Ir. " On the application of the maximum stress

failure criterion to off-auis and angle-plied laminate ", Journal of composite

materials, vol 19, Nov 1985, P. 505

39. S. C. Tan. " Effective stress fracture models for un-notched and notched

multidirectional laminates ", U. S. Air force Laboratory. Wright Patterson AFB,

AFW.U/MLBM, Jan 1987.

30. S. G. Lim and C.. S. Hong, " Prediction of transverse cracking and stiffness reduction

in cross-plied laminated composites ", journal of composite matenals. vol 23, July

1989, P. 695.

3 1 B. Gumrnners * , 1. Verpoa and Van Houtte, ' Determination of the mechanicd

propenies of composite materials by tensile tests. Part 1 elastic properties ", Journal of

composite materials. vol 32 Nov 1998, P. 3 10.

32. Rolfe, Stanley T. " fracture and fatigue control in structures " Prentice-Hall, Inc.,

Englewood Cliffs, New Jersey, USA. 1977. Daley, James and william Riley. "

Experirnental stress analysis " 2nd ed. John Wiley & Sons iNC, 1994

33. Peterson, D.A. and Hart-smith, L.J. " A rational development of lamina to laminate

analysis method for fibrous composites ", Testing and Design, ASTM STP 1059,

(1990), 121-164

34. Car1 T. Herakovich, " On the relationship between Engineering Properties and

delamination of composite matenals" Journal of composite materials vol 15, July

1981 P. 336

35. Narayanaswami, R. and H.M. Adelman, " Evaluation of the tensor polynomial and

Hoffman strength theones for composite materials", Journal of composite materials.

Vol 1 1, oct 1977, P. 366.

36. T. D. Papathanasiou and M. S. Ingber, A. L. graham, " The effective elastic modulus

of fiber-reinforced composites ", Joumal of composite materials, vol. 28, Nov 1994,

P. 288

37. Pinchas Bar-Yoseph and Theodore H.H. Pian," Calculation of inter laminar stress

concentration in composite laminates" Joumal of composite materials, vol 1 5, May

198 1, P. 225

3 8 . Luo-Yo Xu*, " Study on the charactenstic curve of stiffness degradation caused by

transverse matrix cracking in multidirectional composite laminates ': Joumal of

composite materials, vol. 30, Nov 1996, P. 820

P. K. Gotsis*, C. C. Chamis and L. Minnetyan, " Prediction of composite laminate

Fracture: Micrornechanics and progressive kacture ", Composite science and

technology, 1998 P. 1 137.

Gabriel O. Shonaike and T. matsuo, " Fabrication and mechanical properties of fiber

glas reinforced thermoplastic composite ", Composite Structure 32 (1 995), Elsevier

Science Limited.

Daley. James and william Riley. " Experimental stress analysis " 2nd ed. John Wiley

& Sons MC, 1994 G. Wacker *, A. K. ledzki and A. Chate, " Effect of inter-phase on

the transverse Young's modulus of glass / epoxy composites". Composite Part A 29A

. 1998_ P. 6 19 - 626, Elsevier Science Lirnited.

T. E. Tay and E. H. Lim, " Analysis of composite laminate with transverse cracks ",

Composite structure 34 ( 1999), P. 4 19 - 426, Elsevier science limited.

Wilfiied Becker *, Pei Peng Jin and Petra Neuser, " Inter-laminar stresses at the fiee

corners of a laminate ", composite structure, (1 999), P. 1 5 5 - 1 62, Elsevier science

limited.

14. Michael R. Wisnom and M. 1. Jones, " Delamination due to interaction between

overall interlaminar shear and stresses at terminating plies ", Composite structure

( 1995), P. 39 - 47, Elsevier science lirnited.

45. 1. Sevostianov *, V. E. Verijenko and C. J. Von Klemperer? " Overall properties of

composite with physically non-linear discrete phase ", Composite mcture, (2000), P.

187 -1 95, Elsevier science limited.

46. Michael R. Wisnom, Rachael Dkon and Graham Hill, " Delamination in

syrnrnetncally tapered composites loaded in tension ", Composite structure, (1 996), P.

309 - 322, EIsevier science limited.

47. F Ducept *, D. garnby and P. Davies, " A mixed-mode failure criterion derived fiom

test on symmetric and asymmetric specimens ", Composite science and technology,

May 1998, P. 609 - 619

48. M. J. Hinton and P. D. Soden, " Prediction failure in composite laminates: The

background to the exercise ", Composite science and technology, March 1998, P. 74

49. M. H. Alibadi * and P. Sollero, " Crack growth analysis in homogeneous orthotropic

laminate ", Composite science and technology, December 1997, P. 240.

50. Tennyson, R. C and D. MacDonald and A. P. Nanyaro, " Evaluation of tensor

polynomial failure critenon for composite materials", Journal of composite materials,

vol 12, Jan. 1978, P. 63

5 1. J . R. M. D'Almeida, " an analysis of the effect of the diameter of glass micropheres

on the mechanical behavior of glass-microphere 1 epoxy-matnx ", Composite science

and technology, April 1999, P. 59.

52. C. Filious * and C. Galiotis ," In situ monitoring of fiber arain distribution in carbon

fiber thennoplastic composites: 1. Application to tensile stress ", Composite science

and technology, May 1999. P. 74.

53. Shiniji Ogihara *, Nobouo Takeda, Satoshi Kobayashi, and Akira Kobayashi , "

Effects of stacking sequence on rnicroscopic fatigue darnage development in quasi-

isotropic laminate with interlarninar toughened layer ", Composite material

technology (1 999), P. 13 87 - 1398, Elsevier science limited.

54. J. Varna, R. Joffe. "Damage in Composite Laminate with oEaxis Plies" Composite

Science and Technology, May 1999, p. 2139-2147.

55. Joffe. , J Varna, '"Analytical Modeling of Stiffness Reduction in Symmetric and

Balanced Laminate due to Crack in 90 degree Layers.", Composite Science and

Technology, Jan. 1999, P. 16 14-1652.

56. Euginia de Azerdo Sariano, Sergio Frascina Mulleer de Aalmeida, " Notch Sensitivity

of CarbonEpoxy Fabnc Laminate" Composite Science and Technology, Sep. 1998,

P. 1143-1151.

57. Robert A. Jurf and R. R. Byron Pipes, Inter Larninar Fracture of Composite

Materials", Vol. 16, Sept. 1982, P. 386.