NASA/TM- 97-206317

da'& 3_B

The Impact Response of

Carbon/Epoxy Laminates

(MSFC Center Director's Discretionary Fund Final Report,

Project No. 94-13)

A. T. Nettles and A.J. Hodge

Marshall Space Flight Center, Marshall Space Flight Center, Alabama

November 1997

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NASA / TM-97-206317

The Impact Response of

Carbon/Epoxy Laminates

A. T. Nettles and A.J. Hodge

Marshall Space Flight Center, Marshall Space Flight Center, Alabama

National Aeronautics and

Space Administration

Marshall Space Flight Center

November 1997
Available from:

NASA Center for AeroSpace Information

800 Elkridge Landing Road

Linthicum Heights, MD 21090-2934

(301) 621-0390

National Technical Information Service

5285 Port Royal Road

Springfield, VA 22161

(703) 487-4650
TABLE OF CONTENTS

o

2.

INTRODUCTION ......................................................................................................................... 2

PREVIOUS WORK ...................................................................................................................... 3

2.1 History and Background .......................................................................................................... 3

2.2 Damage Resistance and Damage Tolerance ............................................................................ 6

2.3 Key Findings of Past Research ............................................................................................... 6

2.4 The Effects of In-Plane Loading, Nx ....................................................................................... 19

3. ANALYTICAL ............................................................................................................................... 25

3.1 Energy Method ....................................................................................................................... 25

3.2 Finite Element Technique ....................................................................................................... 30

3.3 Finite Element Results ............................................................................................................ 31

4. EXPERIMENTAL ........................................................................................................................ 41

4.1 Material ............................................................................................................................... 41

4.2 Impact Specimens ................................................................................................................... 43

4.3 Pretensioning Device .............................................................................................................. 43

4.4 Impact Testing ........................................................................................................................ 45

4.5 Postimpact Testing .................................................................................................................. 46

4.6 Static Indentation Testing ........................................................................................................ 48

5. RESULTS ..................................................................................................................................... 52

5.1 Instrumented Impact Testing .................................................................................................. 52

5.2 Visual Examination ............................................................................................................... .60

5.3 X-Ray Inspection .................................................................................................................... 61

5.4 Cross-Sectional Examination .................................................................................................. 63

5.5 Static Indentation Tests ........................................................................................................... 73

6. DISCUSSION OF RESULTS ...................................................................................................... 77

6.1 Instrumented Impact Testing .................................................................................................. 77

6.2 Visual Examination ................................................................................................................ 78

6.3 X-Ray Inspection .................................................................................................................... 79

6.4 Cross-Sectional Examination .................................................................................................. 81

6.5 Static Indentation Tests ........................................................................................................... 81

°°°

nl
TABLE OF CONTENTS (continued)

. CONCLUSIONS AND RECOMMENDATIONS ....................................................................

7.1 Overview .............................................................................................................................

7.2 Conclusions ..........................................................................................................................

7.3 Recommendations ................................................................................................................

8. SUMMARY

o,o**o.w.oooowooooooovolo,oo ooooooooo,o,loooooowoolw.loooo ..... °°°°=°°o°=°o.°°°o°°° .... ° o=.°-°o°°°o°°°°°,=°°°°.o°oo°°,l_olo°o

REFERENCES .............................................................................................................................

APPENDIX A--Load-Deflection Data From Instrumented Impact Tests ..........................................

APPENDIX B--X-Ray Signatures of Statically-Loaded and Impacted Specimens ............................

APPENDIX C--Load-Displacement Curves From Static Indentation Tests .......................................

APPENDIX D---Strain-Gauge Readings From Static Indentation Tests .............................................

86

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104

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144

iv
LIST OF FIGURES

1. Top ply cracking due to contact stresses .................................................................................. 7

2. Cracking through other plies due to contact stresses ................................................................ 8

3. Transverse cracks in adjacent 45 ° and 0 ° plies joined via delamination ................................... 8

4. Schematic representation of cross-sectional view of transverse cracks and delaminations

due to contact forces ................................................................................................................. 9

5. Schematic of cross-sectional view of bidirectional laminates ................................................... 10

6. Residual strength versus impact energy curve as predicted by Husman et al. 12 ...................... 13

7. Rapid reduction in residual strength as measured by Rhodes 50 .............................................. 14

8. Schematic of residual tensile strength versus impact energy for highand low-velocity impacts .......................................................................................................... 15

9. Example plot of preload versus impact energy curve to demonstrate concept

of a failure threshold curve ....................................................................................................... 20

10. Representative data from Avva 113 .......................................................................................... 22

11. Deflection, Wmax, versus preload for constant values of transverse load, P,

from equation (9) ...................................................................................................................... 29

12. Load versus displacement for constant values of prestrain from equation (9) ......................... 29

13. Plate geometry for the finite element model ............................................................................. 31

14. Finite element results for prestrain versus deflection at a constant transverse load .................. 34

15. Finite element results for displacement versus load for a constant value of prestrain .............. 34

16. Contour plot of transverse deflection for a transverse load of 223 N (50 lb)

and no prestrain ........................................................................................................................ 36

17. Contour plot of transverse deflection for a transverse load of 223 N (50 lb) and a prestrain

of 6,000 #e ............................................................................................................................... 37

18. Contour plot of transverse deflection for a transverse load of 2,003 N (450 lb) and

no prestrain ............................................................................................................................... 38

V
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40.

LIST OF FIGURES (continued)

Contour plot of transverse deflection for a transverse load of 2,003 N (450 lb) and a prestrain

of 6,000/_e ............................................................................................... 39

Specimens used for impact testing ........................................................................................... 43

Fixture to apply preloads .......................................................................................................... 44

Preload and impact apparatus ................................................................................................... 46

Dam of plumber's putty to hold zinc iodide solution over damage area .................................. 47

Maximum load of impact versus applied prestrain for 3.4 J (2.5 ft-lb) impacts ...................... 52

Maximum load of impact versus applied prestrain for 4.5 J (3.3 ft-lb) impacts ...................... 53

Maximum load of impact versus applied prestrain for 6.0 J (4.4 ft-lb) impacts ...................... 53

Total time of impact versus prestrain for 3.4 J (2.5 ft-lb) impacts ........................................... 54

Total time of impact versus prestrain for 4.5 J (3.3 fl-lb) impacts ........................................... 54

Total time of impact versus prestrain for 6.0 J (4.4 ft-lb) impacts ........................................... 55

Maximum center deflection versus prestrain for 3.4 J (2.5 ft-lb) impacts ............................... 55

Maximum center deflection versus prestrain for 4.5 J (3.3 ft-lb) impacts ............................... 56

>

Maximum center deflection versus prestrain for 6.0 J (4.4 ft-lb) impacts ............................... 56

Absorbed energy during impact versus prestrain for 3.4 J (2.5 ft-lb) impacts ......................... 57

Absorbed energy during impact versus prestrain for 4.5 J (3.3 ft-lb) impacts ......................... 57

Absorbed energy during impact versus prestrain for 6.0 J (4.4 ft-lb) impacts ......................... 58

Load/displacement curve for lowest level impact at lowest prestrain ....................................... 59

Load/displacement curve for highest level impact at highest prestrain ..................................... 59

Typical x-ray signature for a specimen that exhibits no visible damage ................................... 61

Typical x-ray signature for a specimen that contains back face matrix splitting ....................... 62

Typical x-ray signature for a specimen that exhibits visible damage on the impacted side ...... 63

vi
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59.

60.

LIST OF FIGURES (continued)

Cross-section of specimen # Static #5 .....................................................................................

Cross-section of specimen # 12/13B .......................................................................................

Cross-section of specimen # 12-14B, cut i .............................................................................

Cross-section of specimen # 12-14B, cut 2 .............................................................................

Cross-section of specimen # L6, cut 1 .....................................................................................

Cross-section of specimen # L6, cut 2 .....................................................................................

Cross-section of specimen # 12/12C, cut 1 ..............................................................................

Cross-section of specimen # 12/12C, cut 3 ..............................................................................

Cross-section of specimen # 12-14B, cut 2 .............................................................................

Load displacement from a statically loaded specimen at a low prestrain ..................................

Load displacement from a statically loaded specimen at a high prestrain .................................

Load displacement from a statically loaded specimen (loaded until failure) .............................

Load displacement from a statically loaded specimen (loaded until failure) .............................

Schematic of bottom ply splitting and associated delaminations ..............................................

Stresses associated with back face matrix crack .......................................................................

Load-deflection data comparison between equation (9), the finite element method,

and static indentation for a prestrain of 4,000/ae .......................................................................

Prestrain-deflection data for a constant transverse load of 223 N (50 lb) .................................

Comparison of load-deflection curves for a static indentation specimen and an impact

specimen both preloaded at -228 N/cm (130 lb/in) ................ .................................................

Comparison of load-deflection curves for a static indentation specimen and an impact

specimen both preloaded at -1,300 N/cm (742 lb/in) ..............................................................

Comparison of load-deflection curves for a static indentation specimen and an impact

specimen both preloaded at -2,600 N/cm (1,500 lb/in) ...........................................................

64

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1

2

3

4

5

LIST OF TABLES

Calculated data from equation (9) ............................................................................................ 28

Results from the finite element analysis ................................................................................... 33

Finite element results for stresses within plies ......................................................................... 40

Specimens for static indentation tests ....................................................................................... 49

Specimens tested with associated parameters and results ........................................................ 60

,°°

VII1
TECHNICAL MEMORANDUM

THE IMPACT RESPONSE OF CARBON/EPOXY LAMINATES

(MSFC Center Director's Discretionary Fund Final Report, Project No. 94-13)

1. INTRODUCTION

The term "composite materials" can be used to describe a host of different material forms used

for a vast variety of final applications. The one common denominator for all of these materials is that two

or more distinct phases of material are combined in a manner such that the resulting "composite"

properties are more desirable for the application at hand than any of the constituents alone. One of these

phases is usually called the matrix, which acts to hold together the composite material. The other major

phase is commonly called the reinforcement. Composites generally fall into three major categories defined

by the type of reinforcement phase: particulate, short fiber, and continuous fiber.

Particulate composites consist of particles (aspect ratio approximately less than 10) bound together

by the matrix phase. The particles usually do not act as reinforcements for the matrix, but are commonly

used as fillers (such as rocks in cement to make concrete) or to enhance properties other than strength

(such as rubber particles in a brittle plastic to enhance the plastic's toughness).

Short fiber composites contain fibers with an aspect ratio of approximately 10:10,000. With

the exception of most ceramic matrix composites (which use fibers for toughening), the fibers are the

major load-carrying component of the composite material. This class of composites represents the most

widely used and common composites, mainly due to what is commonly referred to as "fiberglass."

Strictly speaking, fiberglass is glass that has been melted and drawn into fiber form, although the term

"fiberglass" usually means chopped glass fiber in a polyester or other inexpensive resin matrix. This

material has been widely used since the 1950's for components such as boat hulls, automobile bodies

for sports cars, and recreational products such as surfboards, skis, and racket frames.

The use of continuous fiber-reinforced composites is growing at a rapid rate since the cost

of materials and processing is decreasing. These types of composites make the most efficient use

of the high stiffness and strength of the fibers used in them, and are usually what is being referred to when

the term "advanced composites" is used. Some of the more common fibers used in this category are

glass, carbon, and Kevlar TM. Glass fiber-reinforced plastics (FRP' s) have good tensile and compressive

strengths and modulii but are heavier than carbon FRP's, which offer very high specific properties but are

relatively weak in compression, limiting them to use in such tensile load-dominated structures as filament

wound pressure vessels.
High strengthcanbeachievedby aligningthefibersin thedirection(s)in whichthestructural

load(s)will beapplied.However,for laminated(layered)composites(whicharealmostalwaysplates

or shells),therearenofibersin thethrough-the-thicknessdirection.Thisis the"Achillesheel" of

laminatedadvancedcompositessince,on occasion,loadsmaygetratherhighin thisdirection.

Anotherproblemwith laminatedFRP'sis that thefailuremodesaremultipleandcomplex.

Thesemodesrely onbothtruematerialpropertiesaswell asstructuralproperties.A homogeneous

isotropicmaterialfailsby asingleflaw, propagatingto failure,whereascompositesfail asaresult

of a statisticalaccumulationof manyflaws.Unlikecomposites,traditionalhomogeneousiotropic

materials(metals,glass,unreinforcedplastics,etc.)havetheluxury of enjoyingamaturedatabase

developedovermanyyearsfor predictingtheeffectsof damage.In addition,thehomogeneousnature

of thesetraditionalengineeringmaterialsmakethemmucheasierto workwith thancomposites,both

analyticallyandexperimentally.Furthercomplicatingtheproblemfor compositesis thelargenumber

of ply orientationsthatareusedin compositestructures.A databasemustbegeneratedfor eachlayup

sequenceusedin a structurecomposedof laminatedcomposites,whichcanbecomequitecumbersome.

Foreignobjectimpactdamageto laminatedFRPis aloadingconditionin whichall of thecomplex

failuremodesmustbetakeninto account.Impactof laminatedcompositeshasbeenresearchedheavily

in thepasttwodecadesbutmuchremainsto beunderstood.Thisresearchis summarizedin thenext

chapter.Thethreatof foreignobjectimpacttocompositeshashinderedtheuseof laminatesin many

applications.Greatstrideshavebeenmadein developinghigherstrainfibersandtougherresinswhich

aremoredamage-resistant anddamage-tolerantthantheearlygenerationcarbonfiber/resinsystems.

However,progressin understandingtheeffectsof themanycomplexvariablesinvolvedin animpact

eventneedsto continuesothatlaminatedcompositescanbeusedwith evengreaterconfidence.Forthe

studypresentedhere,continuouscarbonfiber compositeswill bethefocus,sincetheconsequencesof

impactdamageto theseadvancedcompositesposethegreatesthreat.In addition,mostprimaryloadbearingapplicationsof advancedcompositesusecontinuouscarbonfiberasthereinforcementin

a polymermatrixresin,thustheemphasisonthisclassof material.

In chapter2 areviewof paststudiesonimpactdamageto compositelaminatesis presented. -

Most of thecitedliteratureisfrom experimentalstudiessothatageneralconsensusof theeffectsof

foreignobjectimpactdamageto compositelaminatescanbemade.Thischapterisratherbroadin range

andis intendedto givethereaderagoodbasisfor understandingmanyof theissuesinvolvedconcerning

impacttocompositelaminates.Chapter3presentsa simpleanalyticalassessmentof theeffectsof in-plane

loadsto changeimpactresponsesof orthotropicplateswith clamped-clamped/free-freeboundaryconditions.An energy-balancesolution,alongwith afinite elementassessment, is presented.Chapter4 contains

theexperimentalprocedureusedin this studyandchapter5 presentstheresults.A discussionof these

resultsis presentedin chapter6.In chapter7, somegeneralconclusionsfrom thisstudyarepresented

alongwith somerecommendationsfor futurework.Finally, asummaryis givenin chapter8.
2. PREVIOUS WORK

This chapter consists of three main parts and is intended to be a fairly thorough review of the

general problem of foreign object impact to composites. Section 2.1 is a chronological presentation

of some of the early work that laid the groundwork for the explosive growth of studies conducted in the

past decade. Section 2.2 covers impact damage resistance/tolerance to laminated composites. Section 2.3

summarizes the most important features of past research conducted to date. Section 2.4 deals specifically

with the thrust of this work and the response of impacted laminates due to a small change in boundary

conditions.

2.1 History and Background

The first experiments on impact damage to composites were performed (quite naturally) on the

first structural composite used--wood. In the early 19th century, ballistic-type impacts were studied

by several western European naval forces in an attempt to determine the size and speed of a projectile that

would maximize the damage to an oak ship hull. 1 As FRP came into use, the understanding of foreign

object impact to this class of material became critical. The first experiments in the 1940's were ballisticresistance tests since the major FRP of that time, glass reinforced polyester, was being used as secondary

load-bearing components for military aircraft. 2 With the advent of "advanced composites" came the need

to better understand the damage resistance due to a larger variety of types of impact, not just ballistic

impacts.

Some of the in'st research into the effects of foreign object impact on advanced composites

focused on the energy needed to inflict a certain level of damage, usually fracture or penetration. In the

mid-1960's, glass fiber woven into cloth form, impregnated with polyester resin, was tested to determine

the most efficient configuration to resist ballistic impacts. 3 Lower velocity impact testing was originally

performed using Charpy or Izod impact test machines, using beams made of laminated composites

to determine what type of composite possessed the best damage resistance and what variables were

responsible for energy absorption. 4 This was a natural extension of the type of impact testing being

conducted on metallic materials to determine a ductility index for brittleness. With this impact technique,

the relative importance of the fibers, matrix, and interfaces on impact resistance and energy absorption

could be determined. In using this type of test, it was found that the fiber strain to failure was of primary

importance in determining a composite beam's impact resistance. 5

As impact testing of composite laminates matured, test methods and analyses became more well

defined. In particular, for test methods, instrumented impact and noncatastrophic testing became widely

used and the analyses were broken down into dynamic response and quasi-static response impacts.

2.1.1 Instrumented Impact

By placing strain gauges, load cells, and/or accelerometers on the striker, load time, load deflection,

and other measurements of the dynamic impact event can be made using this instrumentation. Instru-
mentedimpacttestinghasprovento beavaluableresearchtool in studyingforeignobjectimpactof

compositelaminates.Thefirst instrumentedimpacttestswereconductedby theBritishWeldingResearch

Associationto accesstheimpactbehaviorof weldedsteeljoints.6Tolandwasamongthefirst to use

instrumentedimpactortechniquesto studytheprogressionof damagein acompositelaminatebeam.7

By utilizing theCharpytest,theimportancein distinguishingbetweeninitial elasticfailureenergyand

energyabsorbedduringtheresultingfailureprocesswasnoted.Aboutthesametime (early 1970's),

Phillips andTetelmanpublishedapaperdemonstratingthroughinstrumentedCharpytechniquesthat

linearelasticfracturemechanicscouldnotbereliedonfor compositeflaw analysisasit couldin the

analysisof homogeneousmaterials.8

2.1.2 Noncatastrophic Impact Testing

Through experience gained with composite laminates, it became evident that a major threat

is damage that cannot be seen from the outside surface. Most of this damage is due to a foreign object

impact event. Being a laminated material, the high shear stresses set up during an impact event can cause

the layers to delaminate in the vicinity of the impact site. In addition, with the advent of the first generation

carbon fibers (extremely brittle) as a reinforcement, FRP laminates no longer had the impact resistance

of glass-reinforced laminates. Thus, noncatastrophic impact testing became increasingly prevalent in the

mid-1970's. It is this type of testing that separated composite material impact studies from techniques

used on metallic materials. Some of the first work involving noncatastrophic impact was performed

on sandwich panels by Oplinger and Slepetz 9 in which a steel ball was dropped onto panels with carbon/

epoxy and glass/epoxy facesheets. The results clearly demonstrated the lower impact resistance

of the carbon/epoxy facesheets. Other papers presenting results of the study of noncatastrophic impact

to composites are found in reference. 10 Of particular interest is a study that simulated turbine blade

impacts with steel spheres (to represent rocks), ice (to represent hall), and gelatin (presumably to simulate

birds).l 1 This study classified the failure modes produced as transverse (matrix delamination and cracking), penetrative, or breakage of the specimen. When the damage was of the transverse type (noncatastrophic), the amount of damage was assessed using ultrasonic scans and cross-sectional photomicroscopy,

two techniques used extensively today in the study of impact to composite laminates. In another study

Husman et al.12 examined the postimpact strength-carrying capabilities (residual strength) of impacted

composite laminate tensile specimens. From the experimental data it was observed that the residual

strength does drop as the impactor kinetic energy is increased and that glass fiber composites retained

more of their initial undamaged strength than carbon fiber composites for a given impact level. Residual

strength of impacted composites will be discussed further in section 2.3.3.

2.1.3 Dynamic Analysis

Since an impact event is time-dependent, that is, a load inflicted over a short period of time,

it seems natural to use a time-dependent solution. This involves vibrations and stress waves set up

in the material due to the impact event. This dynamic aspect of impact was examined, mostly analytically,

fairly early among the impact studies. Moon 13 used a time-dependent solution to determine the stress

wave propagation due to impact in anisotropic fiber composite plates. The slowest (as well as fastest)

stress waves were associated with a unidirectional laynp and had a speed of slightly less than 2 mm/txsec

(0.08 in/lxsec), (the fastest wave was predicted to be almost 10 mm/[.tsec (0.4 in/lxsec)). Thus, if the

impact is of a short duration, the stress waves will not yet have reflected from the target boundaries

and will become the driving force in determining the resulting damage. Mortimer et al. 14 experimentally

4
measuredthewavevelocity inedge-impacted,cross-plylaminatesas7.38mm/lxsec(0.3in/_tsec),

sotheorderof magnitudeisthatpredictedby Moon.A previouslymentionedreference11showsthe

durationof impactfor ahigh-velocityevent,ontheorderof 100p.sec.This impliesthatfor a20-cm2

(7.9-in2)specimenimpactedatitscenter,theimpactingprojectilewill bemovingawayfromthetarget's

surfacejust asthereflectedstresswavereturnsto thecenterof theplate,assumingawavevelocityof

2 mm/txsec(0.08in/lxsec).In contrast,atypicallow-velocitydrop-weighttestwill beontheduration

of 10,000_tsecwhichgivesampletimefor thestresswavesto decayawaybeforethestrikerleaves

thetarget.

Inertialeffects,ontheotherhand,becomeimportantatmuchlowervelocitiesthanstresswave

propagationeffects.Inertialeffectspertainto theenergyneededto acceleratetheout-of-planedeformations

of theimpactedplate'smass.Duringlow-velocity/high-massimpact,the energy to do this is negligible

compared to the strain energies involved, but at higher velocities, the plate's kinetic energy becomes more

significant. Sun et al. 15 developed a continuum theory for the dynamic behavior of layered composites

which took into account the kinetic energy of the continuum. Greszczuk 16 incorporated the mass of the

target in his analytical solution of the impact of a sphere on an elastic half-plane. Results showed that

a higher target mass will give rise to higher impact forces and contact deformations (deformation of the

upper surface of the target right below the impactor). Global deflections were not accounted for in this

model. For thick laminates, it has been shown that global deflections can be ignored since they are so

small and the material response is governed almost solely by contact stresses. 17 Inertial effects are also

important in that the inertia of the plate and projectile can interact and cause multiple impacts.

2.1.4 Quasi-Static Analysis

For low-velocity impacts, many of the analyses are done as quasi-static, ignoring all timedependent behavior of the event and inertial effects of the material. In essence, this analysis is that

for a "static indentation" case in which the plate is loaded in increments of force over an essentially

infinite period of time. Practically, this is carried out in a material loading frame. In one of the early

studies, Oplinger and Slepetz 9 performed both low-velocity drop-weight and static indentation tests

on graphite/epoxy and glass/epoxy sandwich panels and found that the resulting damage was similar.

Similar results were obtained more recently by Kaczmarek and Maison 18 for quasi-isotropic and

bidirectional carbon/epoxy plates.

In the analytical approach discussed in reference 19, it was seen that a quasi-static event is defined

by the lowest natural frequency ratio of the impactor and target, regardless of their mass ratios and impactor velocity. However, for most practical purposes, high-velocity projectiles are small objects with high

natural frequencies which will produce a dynamic response; low-velocity impacts are really of interest

only when the impactor mass is large and its natural frequency is low, which will produce a quasi-static

response.

Swanson 20 has developed a procedure based on the impactor mass and the "equivalent lumped

mass" of the target which is not inconsistent with the work in reference 19. He showed that if the impactor

mass is more than 10 times greater than the target's "lumped mass," then the impact event can be

considered quasi-static. For various plates and beams of different boundary conditions, the "lumped

mass" was seen to be no more than one-half the entire plate or beam mass. Thus a crude, but effective,

rule of thumb is that the impactor must be more than 20 times the target mass for the impact event to be
consideredquasi-static.However,boundaryconditionscanaffectthisruleif thespecimenis supported

suchthatit is stiff in flexure.Highsmith21foundthatquasi-staticresultsproducedslightly moredamage

thanimpacteventsof similar maximum load. The specimens in his study were 20 plies thick and clamped

over a 6.35-cm- (2.5-in-) diameter opening.

Jackson and Poe 22 demonstrated that impact force alone can be used to quantify an impact event

for a given material and layup but only when the impactor mass is large and the velocity is low. In their

study it was found that the maximum delamination damage area was a linear function of the force

of impact whether the force was introduced by falling weight or static indentation tests. Furthermore,

the support diameter of the impacted plate had no effect on the delamination size when impact force

was used to quantify the event. However, for large plates, deviation from quasi-static behavior was noted

since the stress waves set up by the impact event took longer to lravel the length of the specimen and

reflect back from the boundaries, resulting in some dynamic effects.

2.2 Damage Resistance and Damage Tolerance

Damage resistance and damage tolerance are two distinct and completely different aspects

of impact damage to laminated composites. Damage resistance refers to the composite material's ability

to minimize matrix cracking, delamination, and fiber breakage due to an impact event. Damage tolerance

refers to a material's ability to retain its desired properties once damage has been incurred. This is usually

referred to as "residual" properties.

During an impact event, many parameters determine a material's damage resistance. Structural

parameters such as boundary conditions and the stress state during impact can greatly affect a laminated

composite's damage resistance. It is the intent of this study to assess the damage resistance effects of

a tensile stress state being applied to the target during an impact event. This will usually be referred to as

a "tensile preload" throughout the remainder of this paper. Tensile prestress and tensile prestrain will also

be used but implies the same physical phenomena.

2.3 Key Findings of Past Research

The following review concerns only fiber-reinforced laminated plastics being transversely

impacted by a foreign object. The types of damage incurred during an impact event will be examined

and the effect that the stacking sequence has on this damage will be investigated. The next extension

of the impact problem involves residual load-carrying capabilities of the damaged laminate. A discussion

of how different fiber/resin systems and boundary conditions can affect the amount and type of damage

induced in an impacted composite will then be examined. Finally, a very important topic to those working

with actual sta'uctures made of reinforced plastics, scaling from laboratory specimens to larger entities,

will be discussed.
2.3.1 Damage Modes

There are three major modes of damage that can be incurred during foreign object impact events.

These are matrix cracking, including fiber/matrix debonding; delamination; and fiber breakage. The

formation and extent of these damage modes are heavily dependent upon specimen geometry, or more

precisely, the ability of the impacted structure to deform. For relatively stiff structures, contact forces will

dominate, and for flexible structures, in-plane forces will dominate due to global bending. The contact

force generated by the impactor, if sufficiently large, can cause transverse matrix failure in the impacted

ply and/or those beneath it parallel to the direction of fibers in each ply, due to the high shear stresses, Xyz,

that are set up (shown schematically in figure 1).

MatrixCracks

Z

Y

Directionof

OuterFibers

Figure 1. Top ply cracking due to contact stresses.

As the impactor drives itself into the specimen, the material from the top ply tends to spread

the load over a wider diameter, producing matrix cracks slightly away from the impact site in the next-tothe-top ply. As these transverse cracks develop, they become wider apart through the thickness as shown

schematically in figure 2. These transverse cracks usually join one another through the acute angle with the

adjacent ply via delamination, shown schematically in figure 3. When cross-sectioned, the transverse

cracks and delamination will form a characteristic cone shape as noted in references 23-27 (shown

schematically in figure 4). The through-the-thickness assimilation of the delaminations will give the

characteristic "stair-step" shape, noted in references 28 and 29.

For composite laminates that experience large deformations during an impact event, large in-plane

stresses can cause damage to form. The intrinsic weakness of a ply perpendicular to its fibers causes

the bottom plies of heavily deformed laminates to split along the fibers in that ply, due to the high tensile

strains from bending. The transverse shear in the plate will be at maximum near the center of the plate

and it has been observed that this region contains the planes of largest delaminations, except when spalling

occurs on the nonimpacted face, after an impact event. 29

7
_+45°

OO

Crack

Spacing

Becomes

-45 ° Larger

90o

Figure 2. Cracking through other plies due to contact stresses.

Delamination

45*

0° Ply

Figure 3. Transverse cracks in adjacent 45 ° and 0 ° plies joined via delamination.

8
+45°

0o

90°

-45 °

-45 °

90°

0o

/ \

(Splitsparallelto directionof cutandthuscannotbeseenin0° plies)

/ \

/ \

/ \

Figure 4. Schematic representation of cross-sectional view of transverse cracks

and delaminations due to contact forces.

By using simple bidirectional ]ayups (fibers aligned only in the 0 ° and 90 ° directions) and

examining the resulting matrix cracking and delamination, more insight is given into the prob]em

of damage development in impacted composite laminates. In tests by Chow et al.,30 for [On/9016_2n/On]

(n=6 and 7) stacking sequences, damage was seen to form as matrix cracks in the 90 ° plies with extensive

delamination between the 0 ° and 90 ° plies (when sectioned parallel to the 0 ° plies). Smaller matrix cracks

were observed in the 90 ° plies that did not join with the delamination, thus all matrix cracks do not end in

delamination propagation. In these tests, delamination was observed directly under the impact site,

indicating that flexural behavior is responsible for this growth. These experiments were unique in that the

laminates were supported as wide beams (plates clamped-clamped/flee-free) and impacted by a line nose

impactor rather than the conventional spherical-shaped nose impactor. Thus, cracking in the 0 ° plies was

limited since bending and deformation were only longitudinal to the fibers in these plies. A sphericalshaped nose impactor typically causes the laminate to deform in a "bowl" shape, causing matrix splitting

between fibers in all of the plies. By dispersing the plies more in a [03, 903, 03, 903, 03, 903, 03] layup, it

was found that matrix cracking still occurred in all of the 90 ° plies but the delaminations between the 0 °

and 90 ° plies became progressively smaller from bottom to top.

A finite element analysis of delamination growth from matrix cracks was given in reference 31.

It was demonstrated that normal stresses develop at the matrix cracks that are tensile on the upper interface

toward the impacted side and tensile on the lower face away from the impact site. Thus, delaminations are

predicted to grow toward the impact site on the upper interface and away from the impact site on the lower

face, producing larger delaminations near the bottom plies when bending is the governing stress producer.

These findings were verified for quasi-static loading of bidirectional laminates in three-point bends

while observing a polished edge in reference 32.

An experimental microstructural analysis of impacted [0/90/04] s and [0/90/0] 6 revealed that

the more 0/90 interfaces in a laminate, the smaller the delaminations. 33 In addition, delamination growth

appeared to grow only away from the impact site on the bottom interface of the 0/90 boundary. No trans9
verse cracking was observed in the 90 ° plies and the delaminations were observed only at every other

layer. The authors attributed this to the processing of the panel, but as demonstrated in figure 5, this

is due to the nature of how the specimen was sectioned coupled with how the delaminations form

as shown in figures 1-3.

For high-velocity impact loading, arguments have been given that the tensile stress waves

are responsible for matrix cracking on the upper surface of an impacted composite. 34 Evidence

was also given that matrix cracking and delaminations interact simultaneously since matrix cracking

was not observed in delaminated areas in interfaces away from the impacted side of a bidirectional

composite.

o

Cross-Sectional Cut

Cross-Sectional Views

0/90Plies 90/0Plies

DelaminationsVisible DelamlnationsNotVisible

Figure 5. Schematic of cross-sectional view of bidirectional laminates.

The process of delamination formation being the result of bending stiffness mismatch between

plies of different orientation was given by Liu. 35 He proposed a bending stiffness mismatching coefficient

that would qualitatively give the size and shape of the delamination, resulting from the differences in the

D11 terms of adjacent plies. The shape was seen to be of a "peanut" type, mathematically a lemniscate.

The size was seen to increase as the angle between plies approached 90 °. However, a host of empirical

data28,29, 36--42 shows that the characteristic shape of delaminations between plies is mostly governed

by matrix cracks and the lemniscate has a significant magnitude in the 45 ° direction, whereas it is

predicted by Liu to be zero in this direction.

Fiber failure occurs in laminates experience seeing such large contact stresses that surface fibers

are broken or such large bending stresses as to exceed the strength of a lamina in the direction of the

fibers. A detailed discussion with empirical data for fiber breakage in a laminate due to large deflections

is given in reference 43. This study showed that the onset of fiber breakage was in laminae near

the bottom of the plate.

10
2.3.2 Stacking Sequence

Stacking sequence of the laminate is very important in determining the mode and size of damage

due to foreign object impact; thus, the residual performance of the structure. For structures which

experience high contact stresses, it has been shown 44 that placing nonload-bearing fibers on the outer plies

of the laminate to protect the load-bearing 0 ° plies is advisable. However, as Cantwell et al. 45 suggested

outer 0 ° plies would increase the elastic energy absorption capacity by increasing the flexural stiffness.

When examined from a damage-tolerance point of view, Cantwell et al. 45 showed that the residual compressive strength of a composite with the load-bearing 0 ° fibers away from the surfaces and not next

to the outermost delamination was larger than specimens that had outer 0 ° fibers that delaminated

from the angled ply next to them. The delaminated 0 ° outer plies could easily buckle away from the rest

of the laminate and the strength effect of these fibers would then be lost.

From a damage-resistance point of view, Kwon and Sankar 27 examined 0/90 bidirectional and _4

and rd8 quasi-isotropic stacking sequences. Results showed that initial damage due to transverse loading

occurred relatively early in the rd8 laminates but this damage did not cause as drastic a load drop

(as measured by instrumented impact) as the other two layups. This is because transverse cracks between

plies of different orientation have a smaller angle between them in the rd8 laminates and the resulting

delaminations that form between these cracks between plies (see section 2.3.1) sweeps out a smaller area

which requires less energy. The bidirectional laminates have a full 90 ° between ply orientations and the

delaminations between cracks in adjacent plies in these laminates must sweep out the largest area of any

stacking sequence, therefore, the large load drop due to the formation of these large delaminations. In this

study, 27 cross-sectional examination of the damaged laminates revealed that the bidirectional laminates

contained the most massive and numerous delaminations followed by the n/4 and then the rd8 laminates.

By stacking a laminate such that the number of grouped plies of equal orientation is at a minimum,

increases the interfaces at which a delamination can occur (delaminations do not form between plies

of equal orientation). This was shown to reduce the overall damage area in references 46 and 38 since

the laminates with less ply interfaces at which delamination could occur produced larger delaminations

between the plies that could delaminate in order to dissipate the same amount of energy. For bidirectional

laminates it has been shown 47 that dispersing the plies of different orientation significantly increases

a laminate's impact resistance (as measured by impact energy to produce initial damage) and that

increasing the number of 00 plies on the outer surfaces increases a laminate' s compressive damage

tolerance by increasing the buckling resistance of the sublaminates on the critical outer plies. 48

For honeycomb sandwich' iganels, 49 it was found that the facings would experience less delamination than equivalent laminates and that contact forces dominate since bending is restricted on the facings

of sandwich panels due to the existence of the core. This was supported by data which showed that at low

impact, energies that did not cause large laminate deflections (and were thus contact stress-dominated),

facings and equivalent laminates had approximately the same total delamination areas, but as the impact

level increased to cause large deflections of the laminates, the facings would experience a smaller, total

delamination area since the large bending stresses in the laminate caused larger delaminations than those

caused by contact stresses. It was also found in this study that layups with large angular orientations

between plies (e.g., 90 °) tended to cause a larger total delamination area across all of the ply interfaces

(through-thickness) compared to layups that had 45 ° as the largest angular orientation between plies.

11
2.3.3 Residual Tensile Strength

In this and the following section, the effects of incurred impact damage on composite laminates

is assessed by examining residual strength (tension or compression), a measurement of the composite's

damage tolerance. While some papers will present results as plots of residual strength versus damage area

(thus isolating the damage tolerance response), many others show results on residual strength versus

impact energy plots which actually serves to effectively combine the impact resistance and damage

tolerance portions of the impact event.

Once damage has been introduced into a composite laminate, the critical question is how the damage will affect the desired properties of the material. Measurements and predictions of residual strength

have seen extensive research, especially since the early 1980' s. Much of this work has concentrated on

compression-after-impact (CAI) strength since impact events cause delaminations which can lead to a

lower compressive strength due to the localized instability caused by the delaminations (this will be discussed in the next section). However, many load-bearing composite structures will be dominated by tensile forces (such as pressure vessels), so this aspect of the impact problem also needs to be examined.

Husman et al. 12 were among the f'LrSt to investigate and attempt prediction of the residual tensile

strength of impacted coupons. The residual strength prediction was based on an analogy with an inserted

flaw of known dimensions and a stress intensity factor for an isotropic material. As a result it was predicted that the residual strength would drop off in proportion to the square root of the impact energy

of the projectile. It was noted that departure from this analysis will occur as the specimen experiences

complete penetration. Graphically, the residual tensile strength versus impact energy curve as predicted

by Husman et al. is given in figure 6. Limited data were given to support this analysis and since ballistictype impacts were used, the fibers on the impacted surface (which were 0 ° plies) may have broken at low

impact energies (but still high velocities), yielding a drop in residual at the lowest impact energies used

in this investigation.

12
ImpactEnergy

Figure 6. Residual strength versus impact energy curve as predicted by Husman et al. 12

This study also suggested that the width of the specimen be at least six times the diameter of the

projectile in order to minimize damage zone/edge interaction. This is an important factor to consider when

evaluating residual strength on laboratory specimens. For example, one researcher may impact 2.54-cmwide (1-in-wide) specimens while another performs identical impact tests on 15.24-cm- (6-in-) wide

specimens. If the damage zone is 2.54 cm (1 in) in both specimens, then the narrower specimens will

register a much lower residual strength (possibly zero; i.e., catastrophic failure) than the wider specimens

which possess undamaged material around the damage zone to help carry the load. Thus, if a structure

is larger than the coupons used to perform the impact damage characterization on it, then the coupons need

to be of sufficient width so that free-edge stress/damage zone interactions do not occur. In addition,

if the damage zone is a sizable percentage of the specimen width, then the cross-sectional area rendered

ineffective due to damage must be removed from the residual stress calculations. This is an argument for

measuring residual strain which eliminates needing to know the cross-sectional area rendered ineffective

to carry load.

Data by Rhodes, 50 also generated by ballistic-type impacts, seems to support the notion that both

residual tensile and compressive strengths drop rapidly from the undamaged strength even at the lowest

impact energies tested (fig. 7). It is worth noting that these data were from honeycomb sandwich panels

whose composite facesheets undergo smaller deflections (due to the backing rigidity of the core), resulting

in a more localized damage which includes more fiber breakage.

13
re"

ImpactEnergy

Figure 7. Rapid reduction in residual strength as measured by Rhodes. 50

Other studies have shown that for high-velocity impacts, the residual tensile strength does indeed

begin to drop quite rapidly at the lowest impact energy levels tested. 51-54 A study by Cantwell and

Morton 26 contained residual tensile strength data for eight-ply specimens in which the tensile strength did

not begin to drop off until a critical impactor kinetic energy was obtained. Cairns 55 also found that there

existed a plateau region of residual strength before degradation for 12-ply carbon/epoxy specimens. This

study also showed that for equivalent energy, higher velocity (smaller mass) projectiles, the residual

strength began dropping off much more rapidly, indicating that the plate's inertial effects become more

important at higher velocities, as pointed out in section 2.1.3.

Lower velocity impacts, such as drop-weight tests, generate a different material response than

high-velocity impacts, as mentioned in sections 2.1.3 and 2.1.4. Therefore, it may be expected that the

residual tensile strength response may differ. Comparisons of equal energy high- and low-velocity

impacts have been conducted by Adsit and Waszczak 56 and Dorey. 44, 57 In reference 56, carbon/epoxy

facesheet honeycomb panels were impacted at low velocities with drop weights and at high velocities

utilizing a compressed air gun to project small granite stones. A plot of residual tensile strength versus

impactor kinetic energy shows that the strength reduction begins at a lower level of impact energy for the

high-velocity stones and also drops off at a faster rate with increasing impactor kinetic energy for the

higher velocity tests. In reference 44, the residual strength is seen to drop at a smaller impact energy for

ballgun impacted specimens, but once a decrease in strength is seen, it is more rapid for drop-weight tests.

This trend is shown schematically in figure 8. This effect is much more pronounced for residual flexural

testing as seen in reference 56. The type of resin used can have an effect on these results, as discussed

in section 2.3.5.

14
I'--

"ID

rr /

DropWeight

Impacts

BallgunImpacts

ImpactEnergy

Figure 8. Schematic of residual tensile strength versus impact energy

for high- and low-velocity impacts.

Low-velocity impact tests with resulting residual tensile strength tests show that, for most cases,

a region of impact energies where strength degradation does not occur is readily apparent. This

phenomena is seen in data by Cantwell et al., 45, 58 Dorey et al., 59 Morton and Godwin, 60 Caprino, 61

Nettles, 62 Bishop, 63 Lal, 64 and poe.65

However, once strength degradation does occur, the residual strength usually decreases rapidly

with increasing impactor kinetic energy, depending upon the material and geometry of the impact event.

For example, specimens supported as short beams 60-61 or as small diaphragms 62 experience a high level

of contact stresses immediately below the impactor, thus breaking fibers and causing a rapid reduction

in residual tensile strength. Specimens supported over larger areas, such as in references 59 and 64,

demonstrate a more gradual decrease in residual tensile strength with increasing kinetic energy since

the large elastic deformations can absorb much of the incident impactor kinetic energy and local contact

forces under the impactor are less severe. The effects of boundary conditions will be examined further

in section 2.3.6.

2.3.4 Residual Compressive Strength

Residual compression strength after impact has drawn much more interest than residual tensile

strength since laminates have been shown to be much more susceptible to this type of loading. The impact

event can cause delaminations within the material which cannot be seen from the surface, yet can severely

reduce the CAI strength of the composite structure. While the tensile strength of a composite laminate is

fairly dependent on the test sample's geometry, compression testing of laminates is virtually a complete

structural test as opposed to a true material test. Thus extreme caution must be used when examining data

from CAI tests and direct comparisons between studies is virtually impossible. Therefore, only general

observations will be made based on the results of some of the findings of past research on this topic.

15
In low-velocityimpactstudieswherebothresidualtensileandcompressivestrengthswere

tested,theresidualcompressivestrengthalwaysfell atafasterratethantheresidualtensile

strength.44,45,57,60,63,66Rhodes50presenteddatafor projectile-typeimpactsthatshowedtheresidual

tensilestrengthdecreasingata slightlyfasterratethantheresidualcompressivestrength.Theseresults

arebecauseaspecimenwill losetensilestrengthonly whenload-beatingfibersarebroken,butcompressivestrengthcandropdueto delaminations.In mosttestprogramsthe specimensaresupported

suchthatmatrixcrackinganddelaminationwill occurbeforeanyfibersarebroken.Brokenfiberswithout

delaminationsonly occurin thick specimens,or in ballistic-typeimpacts,in whichthespecimendoesnot

havetimetodeformandcreatethehighinterlaminarshearstressesthatcausedelamination.

Forlow-velocityimpacts(dropweight),thecharacteristicshapeof theCAi versusimpactenergy

curveisheavilydependentuponthespecimenmaterialandgeometry.In astudyby PrichardandHogg,67

carbon/epoxyandcarbon/polyetheretherketone(PEEK)specimensweresupportedoverarelativelysmall,

clampedannularopeningto inducedelamination-typedamage.TheresultingCAI versusimpactenergy

curvesshowthatthecompressionstrengthof thecarbon/epoxydroppedoff extremelyrapidlywhereasthe

curvefor carbon/PEEKgraduallydroppedoff with increasingimpactorkinetic energy,almostin alinear

fashion.Thesameresultswerefoundin references60and68for specimensclampedwith alargerannular

diametersupport,althoughthecarbon/epoxyCAI strength did not drop off quite as rapidly. In general, it

has been shown that for a wide range of carbon/epoxy specimen geometry and CAI test methods, the CAI

strength falls rapidly for increasing impactor kinetic energies at the lower level of impact energies until

a plateau or modest linear decrease zone is reached. 44, 45, 61, 63, 66, 69-75

For high-velocity- (projectile-) type impacts the CAI strength versus impact kinetic energy curve

does not possess as drastic a drop in CAI strength with increasing impactor kinetic energy as shown

in references 50, 52, 76, and 77. This is mainly because the much more localized response of the highvelocity impacted specimens in which delamination tends to be extensive through the thickness, biat not

as widespread in areas such as lower velocity-impacted specimens. Thus, the lower velocity-impacted

specimens contain a larger delamination area, causing a larger reduction in structural stability of the test

coupon than the smaller delaminations formed by higher velocity impacts. This notion is supported

by data in reference 78 in which the CAI strength of specimens impacted with high- and low-velocity

impactors did not vary a great deal when plotted versus damage width, but were different when plotted

versus impact energy.

From drop-weight impact data, Prichard and Hogg 67 showed that when plotted as a function

of damage width, the residual compression strength of two very different laminated composite materials

(epoxy and PEEK matrices) were the same. Similar results were obtained by Teh and Morton 68 and

Dempsey and Horton. 73 However, other studies 79 have shown that a specimen of one type of material

may have a much larger damage zone than another type of material impacted under identical conditions,

yet have a higher retention of initial compressive strength. Indeed, if the residual strength versus damage

size plots were similar for all materials, then this would imply that there is no difference in the damage

tolerance of laminated composites, only a difference in damage resistance.

Of practical importance for composite structures that see multiple impacts at nearby sites (such

as during a hailstorm), Jones 80 showed that the CAI strength was unaffected by the multiple impacts

as long as they were approximately one damage zone diameter apart (i.e., the impact sites do not interact

when spaced apart by at least one diameter of damage zone).

16
2.3.5 Fiber/Resin Systems

Of the most interest to designers and end users of composite structures are which fiber/resin

systems are the most damage resistant, or more importantly, damage tolerant. Since the chosen fiber/

resin system is a variable that can be controlled from the onset of design, this information is vital to any

program utilizing laminated composites. In the past decade, a concentrated effort has taken place within

the industry to produce fibers with higher strain to failures and matrix resins that better resist cracking and

delaminations. More exotic techniques such as stitching or using three-dimensional preforms, 81-85

interleaving, 86-88 and using hybrids (more than one fiber type) 36, 89-95 have been investigated, although

they have not yet seen widespread usage.

The toughness of the matrix resin can severely affect the damage resistance/tolerance of laminated

structures. A direct comparison of thermoplastic and thermoset resins usually shows that the thermoplastic matrix material is more damage resistant. The most promising high-performance thermoplastic resin

is PEEK which has been studied extensively in damage-tolerance programs. A comparison of PEEK

with epoxy resin composites was made by Dorey et al.59 and Demuts and Sharpe 96 and showed that the

PEEK laminate was much more damage resistant (as measured by C-scanned damage area), resulting in

greater retention of CAI strength. In reference 59, for the highest impact energies used (8 J), the carbon

PEEK had about one-fourth of the damage area as the carbon/epoxy tested. When compared to another

thermoplastic such as polyphenylene sulfide (PPS), 97 PEEK was still more damage resistant with about

a 50-percent smaller delamination size for the highest impact energies tested.

By plotting CAI strength versus damage size for carbon/epoxy and carbon/PEEK laminates

and obtaining similar results, 67 it follows that the superior CAI resistance of the PEEK laminates is due

mainly to the smaller delamination area formed for a given impact energy level. This is further supported

by work performed on a number of toughened epoxies and bismaleimide (BMI) resins in which the CAI

strength was found to correlate strongly with mode 11 interlaminar fracture toughness. 98 Mode II fracture

resistance controls the amount of delamination that forms during an impact event whereas the CAI

strength is strongly controlled by mode I peeling as the delamination grows.

Although thermoplastics such as PEEK have distinct advantages over thermoset resins, these

materials require a high temperature and pressure to cure the plies into a laminate. Furthermore, lack

of any tack and little drape make further processing complicated. Until more advanced processing

methods are found, thermosets will continue to dominate the resins in use on composite structures; thus,

mechanisms for making existing thermosets tougher has seen much activity. The two principle methods

of toughening thermosets is by modifying these resins either with elastomeric particulates or by blending

in a controlled amount of thermoplastic resin. Both methods can substantially increase a laminate's

toughness and modified resins of both type are commercially available and in wide use. Damage

resistance/tolerance comparisons of some of these resins have been made in a number of

studies.53, 75, 79, 99

17
2.3.6 Boundary Conditions

Another critical factor that determines the amount of damage that a composite laminate will

experience due to a foreign object impact event is the laminate's geometry and boundary conditions.

For high-velocity impacts, the boundary conditions of the target are less important than the areal mass

of the target in determining the impact response 100 (unless stress waves account for damage). The

impactor has already rebounded from, or penetrated the target by the time the stress response reaches

the edges of the target specimen; thus, boundary conditions are much more critical for lower velocity

impacts. Cantwell and Morton 101 showed that for a given target thickness, the increase in energy required

to cause initial damage and to cause perforation for impacted beams, was roughly proportional to the

increase in beam length. It was also shown in this study, that for a given impact energy, lengthening

the beam caused the beam to experience larger deflections, thus storing more of the energy as elastic

energy and reducing the size of the damage zones. Kwon and Sankar 27 demonstrated that if the

delamination size is plotted as a function of maximum load of impact, then the specimen support size

effect is eliminated, since larger supports allow larger deflections and thus iower impact forces. In a study

by Verpoest et al., 102 the size of the support was shown to be more important than the shape. Another

aspect of the varying response of the target, due to support geometry, is where the majority of damage

is most likely to form. For specimens supported in a smaller geometry, the top surface will experience

higher contact stresses and thus more damage will form at this location than on a similar target, only

with larger geometrical boundaries, which will experience bottom surface splitting/fiber breakage due

to the larger bending stresses that are present. 103

Also of importance is how the specimen is supported at its boundaries (i.e., clamped, simply

supported, etc.). The same general rule of thumb applies here also. If the boundary conditions restrict

the specimen from deforming, then the damaging stresses will migrate more from bending type to contact

type. This was shown analytically by Shivakumar et al.104

Another critical aspect of the impact event, the shape of the impactor, has seen relatively little

attention. Various sizes of hemispherically ended impactors have been used, but other geometries

of impactor were studied by Poe. 105 This study was conducted on thick sections of graphite epoxy that

simulated a section of a rocket motor case. Impactors of two different hemispherically ended radii--a

sharp, bolt-like rod--were used. The comer-type impactor caused very high localized Hex'_zian contact

pressures, producing visible damage earlier than the hemispherical indentors. The cylindrical bolt-type

indentor acted like a punch, slicing through the upper layers of the laminate. The shape of the impactor

will have a much greater effect on laminates which are relatively stiff, since the Hertzian contact pressures

are more important for this type of impact event.

2.3.7 Scaling of Impact Damage

The ultimate goal of the vast amount of research that has been conducted on impact-to-laminated

advanced composites is to be able to predict the effects of a foreign object impact event to a full-scale

composite structure. Since this can turn into a very expensive program, laboratory-size specimens must

be used and the results scaled to those of larger structures. The problem of scaling impact damage

to composites has been investigated by Morton. 106 Morton used the principles of similitude and pointed

out, that for most practical cases, complete similarity between "model" (coupon) and "prototype" (full

scale) is rarely possible when rate- or notch-sensitive materials are involved. For composite laminates,

18
theproblemis furthercomplicatedsincethefiberdiameterandlamina_'cknessshouldideally bevariable

to accommodatescaling.In practicalapplications,thisisvirtually impossible.Testresultsonimpact

of compositebeamsshowedthat classical scaling laws apply for elastic behavior of the beams (i.e., before

damage is induced) but significant deviation was observed as damage was induced into the specimen with

the smaller specimens always stronger than the scaled-up ones. Swanson 107 has also examined the scaling

effects of impact damage to composite laminates and found that delamination damage was dependent

upon the absolute specimen size, whereas, fiber breakage was not. Thus, the effects of scaling must be

different for different failure modes.

For delamination damage, Bucinell et al.108 tested laminated plates of scaling factors 1, 2, and 4

and showed that the response followed an inverse square root of the scaling factor. The response measured was the percentage of laminate area delaminated as detected by ultrasonic C-scans. The impact

parameters and responses were scaled according to the development of Morton. 106 It was also demonstrated in this study that quasi-static and dynamic impact events of equal energy scale similarly.

Through finite element analysis, Chen et al. 109 showed that the stress concentration due to

a damage zone is concentrated around that zone and local stress distributions between test coupons and

stringer-stiffened panels were similar. Thus, the measured differences in CA/strength in which the larger

panels retained more strength 110 is due to the larger panels experiencing less damage for a given impact

energy per unit thickness.

2.4 The Effects of In-Plane Loading, N x

A limited amount of research into the effects of tensile or compressive preloads on the impact

response of composites has been performed. Olster and Roy 111 performed ballistic penetration tests on

tensile preloaded graphite/epoxy plates and found that the combination of impact and preload could cause

a 45-percent reduction in ultimate tensile strength whereas penetrated plates without a preload would fail

at an approximately 38-percent reduction in ultimate tensile strength. This suggests an interaction between

the preload stresses and the stresses set up by the impact event. Since all specimens were completely

penetrated by the projectiles, the residual tensile strength was found to be independent of the preload.

For nonpenetrating impacts on preloaded specimens, Awa (formerly Sharma) 112-115 has

produced the most experimental data on this topic. Typically, a "failure threshold curve" is obtained

by plotting specimen preload (tension or compression) versus impactor kinetic energy and drawing

a faired curve between the data in which the specimens broke upon impact (a "catastrophic" break)

and those which survive the impact. A sample curve (not from actual data) is presented in figure 9

for clarity on this subject. As will subsequently be seen, a weU-def'med area between the two zones

exists in most actual experimental data.

19
O

-d

t_

tX.

©

• = FailedOnImpact

0 = DidNot FailOn Impact

FailureThresholdCurve

• RegionOf Catastrophic Failures

O0

0

0

RegionOf Noncatastrophic Failures

• •

0

0

KineticEnergyof Impactors

Figure 9. Example plot of preload versus impact energy curve

to demonstrate concept of a failure threshold curve.

In a study on honeycomb sandwich panels impacted by projectiles from a fight gas gun 112

it was found that tensile preloads on the facings of sandwich panels could cause catastrophic failure

at stress ratios (ratio of preload stress to ultimate breaking stress) below 40 percent for 1 J (0.7 ft-lb)

of impactor kinetic energy. This was for facings of T300/5208 carbon/epoxy in a stacking sequence

of [90,+45,-45,0] s bonded to a 130 kg/m 3 (8.1 lbm/ft 3) aluminum honeycomb core. For stacking

sequences of [+45,-45, 04]s and [+45,-45,90,0]s, higher prestresses needed to be applied to cause

catastrophic failure for a given impact energy. This is understandable for the [+45,-45,04]s layup since

it is thicker and contains more 0 ° pries. Limited data were supplied for the [-1"45,--45,04] s and

[+45,-45,90,0] s specimens, so conclusions on why this may be are difficult to draw. Contact stresses

surely dominated in this study since the facesheets were backed by a relatively hefty honeycomb

of 130 kg/m 3 (8.1 lbm/ft 3) density. For specimens that did not catastrophically break upon impact,

the residual strength was obtained. Although this aspect of the study was not highlighted in the paper,

it is observed that the tensile residual strength is not much higher than the failure threshold curve

for the limited data presented. For compression, the residual strength could be much more above

the failure threshold curve. This is probably due to the stiff honeycomb backing preventing large

deflections; thus, the large delamination areas associated with them.

In a study that contained much more data, 113 the effects of a preload on the residual strength

of Specimens were noted. The specimens were made of T300/5208 carbon/epoxy with a stacking

sequence of [452,-452,02,902]s. The support conditions were clamped/clamped (perpendicular

to the 0°-direction)-free/free with a width of 7.62 cm (3 in) and length (0°-direction) of 15.24 cm

20
(6in) for tensionloadedspecimensand6.35cm (2.5in) for compression-loadedspecimens.An

aluminumsphereof 1.27cm (0.5in) diameterwaspropelledatvelocitiesbetween18.6m/sec(61ft/sec)

and106.7m/sec(350ft/sec)andstruckthespecimensattheirgeometricalmidpoints.Thistesting

geometrywill inducerelativelylarge,localcontactstresseswhilenotcausingmuchglobaldeflection

aswouldbepresentin ahigh-mass/low-velocityimpactevent,thusminimizing theincreasedstiffness

responsedueto thetensilepreloads(or decreasedstiffnessresponsedueto compressivepreloads).

For thedatapresentedontensileloading,afairedcurvecouldeasilybedrawnbetweenregions

of catastrophicandnoncatastrophicfailures.Forthe specimensthatdid notexperiencecatastrophic

failures,theresidualtensilestrengthsweremuchabovethecatastrophicfailurecurve.Specimenswhich

wereimpactedwhileundernopreloaddemonstratedaresidualstrengthversusimpactenergycurvehigher

thanthecatastrophicfailurecurve.Theresidualstrengthdatafrom thepreloadedspecimensthatdidnot

catastrophicallyfail fall closetothenopreloadresidualstrengthcurve.An additionalsetof testswere

conductedusingthecatastrophicfailurecurveasa guidefor settingpreloadsatthreedifferentimpact

energies.All of thesespecimenswerepreloadedin theregionof thecatastrophicfailurecurve.As

expected,mostof thespecimensfailedcatastrophicallyuponimpact,andthosethatdid nothadvery

littleresidualstrengthremaining.

Similarexperimentswereconductedfor specimensunderacompressivepreload.For the

specimensthatdid notfail catastrophicallyuponimpact,theresidualcompressionstrengthwasseen

notto risemuchabovethecatastrophicfailurecurve.In addition,thecatastrophicfailurecurvewasclose

totheresidualstrengthversusimpactenergycurveof specimensimpactedwith nopreloadwhereasfor

thetensilespecimens,thecatastrophiccurvewasapproximately30-percentlowerthantheresidual

strengthversusimpactenergycurvefor unloadedspecimens.Theseresultsaresketchedout graphicallyin

figure10for clarity andcomparison.Onlypertinentdatapointsareincludedto reduceclutterintheplots.

Thus,fromthisstudyit appearsthatatensilepreloadcanreducetheresidualstrength-carryingcapability

of theimpactedlaminate,butonlywhenpreloadednearthe"catastrophicfailure"curve.

21
Tensions Compression

• -- CatastrophicBreak -- •

O -- Survived Impact -- _7

-- ResidualStrength -- _'

[] _ ResidualStrength -- Z_

(No Preload)

o

-J

E

-E

13.

1.0

0.8

0.6

0.4

0.2

%

%

%

= Catastrophic Line ForTension

= CatastrophicLine ForCompression

....... ResidualStrengthWith No

Preload(Tension)

....... ResidualStrengthWith No

Preload(Compression)

I I I I 1 I I I I I

10 20 30 40 50 60 70 80 90

KineticEnergy/Thickness(J/cm)

Figure 10. Representative data from Avva. ! 13

In a very similar study, 114 specimens of T300/934 carbon/epoxy with a stacking sequence

of [+45,--45,0,9012s were impacted by an aluminum sphere of 1.27 cm (0.5 in) diameter while under

a tensile preload. The catastrophic failure curve for this study is very similar to that shown in figure 10.

The only noticeable difference between these two studies is that the residual strength of the preloaded

specimens that did not fail catastrophically fell closer to the catastrophic failure curve in this study as

compared to the study in reference 113. A summary of this study can be found in reference 115.

In an impact damage study by Starnes et al.76 the residual compressive strengths of 48-ply laminates

loaded with a compressive preload were evaluated. This represents an extremely thick test coupon

and global deflections should be negligible due to the laminates high bending stiffness. The projectile

used was a 1.27 cm (0.5 in) aluminum sphere propelled between velocities of 52 to 101 m/sec

(171 to 331 ft/sec). The material used was T300/5208 carbon/epoxy with a stacking sequence

of [+45,--45,02,+45,--45,02, +45,--45,0,9012s cut into specimens 11.4 cm (4.5 in) wide by 24.8 cm

(9.75 in) long. Catastrophic failure threshold curves were determined from the data and the two zones

between specimens that survived the impact and those that failed on impact were very distinguishable.

For those specimens that did survive impact the residual strengths were determined. Below 5 J (3.7 ft-lb)

22
of impactenergy,theresidualstrengthvalueswerewithin thescatterbandof specimensthathadnot been

impactdamaged.Forimpactenergyvaluesthatdidcauseareductionin residualcompressionstrength,

thepreloadappearsto decreasetheresidualcompressivestrengthto asmalldegree,althoughonly six data

pointsareusedforthis conclusion.In acontinuationof thisstudy,116additionaldataon40-ply laminates

of stackingsequence[+45,-45,04,-45,+45,+45,-45,04,-45,+45,+45,-45,02]s demonstratedthat

specimensimpactedwhile underacompressivepreloadexperiencedslightlylargerareasof damage,

asdetectedby brittlelacquerspallationonthenonimpactedsideof thetestcoupon.Specimensof aquasiisotropiclayup,[+45,-45,0,90,-45,+45,0,9013swerealsotested,with andwithoutcompressivepreloads,

with noreal discernibledropin CAI strengthdueto thecompressivepreloads.Theseresultsarealso

presentedin reference117.

Low-velocitydrop-weighttestsontensileprestressedthinnerspecimenswereperformed

in astudyby Park.118T300/914Ccarbon/epoxywith four- andeight-plybidirectionallayupswere

impactedby a 1.5-kg(3.3-Ibm)massimpactorwith a 6-mm(0.24-in)diametertup.The2.5-cm-(1-in-)

widespecimensweresupportedasclamped-clamped/free-freewith a spanof 5 cm(2 in). Catastrophic

thresholdcurvesweregenerated(termed"impactmaps"in thisstudy)andshowedthatballistic impacts

hadahigherthresholdcurvethandrop-weightimpactssincetheballistictestscausedlocalizedpenetration

andthedrop-weightestscauseddamageacrosstheentirespecimenarea.A limitednumberof testson

theresidualtensilestrengthof preloadedcoupons,impactedballistically,showedthattheresidualstrength

droppedsharplyasthepreloadnearedtheultimatetensilestrengthof thetestcoupon.

A C-scananalysisof impactedbeamsunderatensilepreloadwasconductedby Sankarand

Sun.119 Although only one initial preload value was used (about one-third of ultimate load) the results did

show differences in those beams impacted with an initial tensile stress and those impacted with no initial

stress. The beams were 20-ply graphite/epoxy with a stacking sequence of [02,902,02,902,02] s supported

as clamped-clamped/free-free. A 1.27-cm- (0.5-in-) diameter steel ball was used as a projectile at speeds

of 10-40 m/sec (33-131 ft/sec). The span of the impacted beams was set at 18 cm (7 in) and were 3.8 cm

(1.5 in) wide. Although ballistic projectiles were used, the relatively high aspect ratio for these plates could

possibly allow more global deflection than the specimens impacted in the previously cited studies in this

section. Results to note are that larger zones of delamination were found to exist for those specimens

impacted while preloaded. The only exception to this trend was a reduction in the delamination area near

the back plies on preloaded samples for larger energy impacts. Residual tensile strength was also

determined for these specimens. It is evident that the tensile preload causes a reduction in residual tensile

strength by approximately 10 percent for the impact velocities used in this study.

Data for thinner plates impacted at low-velocity (drop-weight) impacts have been presented

by Nettles and Lance. 120 Compressive preloads were applied to 16-ply test coupons of T300/934

and IM7/8551-7 carbon/epoxy with a quasi-isotropic stacking sequence of [0,45,90,-45]s2. Coupon

dimensions were 7.6 cm (3 in) wide by 10.2 cm (4 in) gauge length. The preloaded specimens were

supported over a 6.4-cm- (2.5-in-) diameter circular opening and clamped around its periphery. A design

of experiments approach was undertaken to try to identify significant variables without testing numerous

specimens. A total of 18 specimens (two replicates of 9) of each material type were tested. Preloads

ranged from 890 to 44,500 N (200 to 10,000 lb) and impact energies varied between 1 and 16 J

(0.7 and 11.8 ft-lb). In general, it was found that the compressive preloads tended to cause a decrease

in residual compression strength after impact as the preload was increased, although quantitative values

23
aresuspectdueto thesmall number of specimens tested and the variables not being allowed to go to their

extremes.

A recent study 121 examined the effect of impact response on biaxially loaded specimens

of chopped E-glass in a polyester resin matrix. The preloads could be introduced in any combination

of tension and compression in the two principal axes of the plate. Impacts were performed by a dropweight apparatus at an energy of 21.5 J (15.9 ft-lb) for all specimens. Results measured were indentation

depth, absorbed impact energy, maximum force of impact, and damage area. The largest level of damage

was observed to occur for specimens loaded in equal tension-compression (i.e., pure shear). The

maximum force of impact was largest for the highest values of tension-tension preloading and lowest

for the largest values of tension-compression (pure shear). A state of biaxial compression produced

maximum loads of impact between these two extremes. Very tittle energy was absorbed for specimens

loaded in tension-tension but as compressive preloads were introduced, the absorbed energy dramatically

increases. Care must be taken when interpreting these results to continuous fiber-reinforced, layered

laminates since the damage modes are so very different.

Analytical treatments of this problem have been undertaken in a few studies. 122-125 In all of these

studies it was shown that a tensile prestress will decrease the maximum deflection due to impact and will

raise the maximum load of impact. Compressive preloads will do the opposite.

24
3. ANALYTICAL

In this chapter, finite element solutions to the problem of transverse impact of a prestressed plate

are made. None of these solutions are detailed to the extent that heavy experimental testing need not be

performed. In keeping with the empirical nature of this work, the analyses presented will be simple and

produced more to explain observed physical phenomena rather than to predict extent of damage as a

function of impactor kinetic energy and applied preload. First, an approximate energy method model

of the problem is presented. A solution of this problem can only be performed numerically and many

assumptions have to be made, making the finite element technique much more informative and accurate.

However, some basic physical phenomena can be qualified from the energy method problem before

it is completely solved. These results can help explain some of the results from the finite element analysis.

3.1 Energy Method

The energy involved in the system is assumed to consist of the elastic potential energy stored

in the deflected laminate and the energy due to the external transverse load. The energy stored in a laminate

due to bending is given by

[.. (02w 2

1 y-bx-L [Ull_ -----_) 8 x

E bend" " "2 yfO xf 0 I

[+4(D16 02W

02w t_2W + D { 02w_ 2

+2 a =t-Ey

+

0 092W'_ 02w +4066(_W] 26"-_y2) o2r----_ _ 0y]

dxdy ,

2

(1)

where b is the width of the laminate, L is the length of the laminate, and w is the transverse displacement

of the laminate. The Dii are the components of the bending stiffness matrix, x is the cooridinate along the

length of the plate, and y is the coordinate along the width of the plate.

The energy stored by midplane (membrane) stretching is given by

lY'bx'L[All(O_Vl4 OWtW 2+

2 A66)(--_-x -_-- ) A22( °_v/4 ]

Emem"'2yfOxf 0 ['-'2-\-'_'x-x ] +(a12+ -"_--_-Y ] I dxdy '

(2)

where the A U are components of the extensional stiffness matrix.

25
Theenergystoreddueto theexternallyappliedpreload,N_, is given by

-_ - N x dxdy

E Nx 2 yf0xf 0

where Nx is the x-direction stress resultant.

The energy put into the system by the transverse load, P, is given by

y-bx-L 4p

E P "_ -yfO x_O--L-'_ wd.xdy

The total energy of the system is given by

Etot " Ebend + Emem + ENx + Ep

Assume the deflected surface to be of the form

(3)

(4)

(5)

,14,,.(7) ,6,

where Co is a constant, satisfiesthe boundary conditions at x=0 and x=L. The boundary conditions

at y=0 and y=b restrict the plate from deflecting along these edges when in actuality they are free to move.

(However, from experimental data to be given later, these edges deflect very little so this approximation

will suffice for qualitative purposes.)

Assume the typical values for carbon/epoxy as:

A 410o124o 124,000 410,000 00 1andD--22 i67221 ] 32

o 0 143,000 14 14

(units in pounds and inches)

and putting equation (6) into equations (1), (2), (3), and (4) and performing the integrations

(some numerically) with L=16 cm (6.3 in) and y=9.5 cm (3.75 in), then using equation (5) gives

7.10 x 1020 C04+ (8.3 x 107 Nx + 4.66 x 1010)C_ - 1.02 x 105 P CO - Etot , (7)

with units in pounds and inches.

From the principle of virtual work,

0 Etot

OCo

--_0 (8)

26
Applyingthisoperationto equation(7) gives:

2.84x 1021C 3 + (1.66 x 108Nx + 9.32 x 1010)C0 -- 1.02 x 105p

At the center of the plate, w=-Wmnx, x=- L/2 and y=b/2. Therefore,

CO = 1.032 x 10-4Wmax ,

(9)

(10)

and equation (9) becomes

P - 30, 600 W3max + (0.168Nx + 94.3) Wmax (11)

The cubic equation (9) can also be solved for Co and then these values placed in equation (6) to determine

the deflection at any point on the plate. Table 1 gives load (P)-preload (Nx) combinations with the

associated transverse displacement at the center of the plate (Wm_x). The same values are obtained using

equation (11). These values are plotted in figure 11 for constant values of P with varying Nx and in figure

12 with constant Nx and varying P.

From equation (7), if Etot represents the total impactor kinetic energy put into the system, then if Nx

increases, P increases and Co decreases in order to maintain the energy balance of the system Thus, from

an energy balance formulation of the problem of a transverse load acting on a prestressed plate, the higher

the prestress, the larger the maximum transverse load and the lower the maximum transverse deflection

for a given amount of energy put into the system.

27
Table 1. Calculated data from equation (9).

TransverseLoad, P

a (Ib)

223 (50)

223 (50)

223 (50)

223 (50)

445(100)

445(100)

445(190)

445 (100)

(150)

668 (150)

(15o)

890 (200)

800(200)

890 (200)

890(200)

1,113(250)

1,113(250)

1,113(250)

1,113(250)

1,335 (300)

1,335 (300)

.. 1,335 (300)

1,335 (300)

1,558(350)

1,558 (350)

1,558 (350)

1,558 (350)

1,780 (400) .....

1,780(400)

1,780(400)

1,780(400)

2,r003(450)

2,003 (450)

2.003 (450)

2,003 (450)

Preioad

N/cm fib/in)

o(o)

1,229(702)

2,459 (1,405)

3,687 (2,107)

o(o)

1,229(702)

2,459 (1,405)

3,687 (2,107)

o(o)

1,229{702)

2,459 (1,405)

3,687 (2,107)

o(o)

1,229 (702)..

2,459 (1,405)

3,687 (2,!0.7)

o(o)

1,229(702)

2,459 (1,405)

3,687 (2,107)

o(o)

1,229(702)

2,459 (1,405)

3,687 (2,107)

o(o)

......1,229(702)

2,459 (1,405)

3,687(2,107)

o(o)

1,229(702)

2,459 (1,405)

3,687 (2,107)

o(o)

1,229 (702)

2,459 (1,4o5)

3,687 (2,1o7)

Prestrain

me

0

2,000

4,000

6,000

2_000

4,000

6,0OO

0

2,000

4,000

6,000

0

2,000

4.000

6,000

0

2,000

4,000

6,000

0

2,000

4,000

..... 6'0 O0

0

2,000

4,000

6,000

0

2,000

4,000

6,000

0

2,000

4,000

6,000

Deflection, Wmax

mm (in)

2.82(0.111)

2.53 (0.0998)

2.27 (0.0892)

2.01 (0.0793)

3.66(0.144)

3.43 (0.135)

3.2 (0.126)

3.00 (0.118)

4.24 (0.167)

4.04 (0.159)

3.84(0.151)

3.66 (0.144)

4.70 (0.185) ,

4.52 (0.178)

4.34 (0.171)

4.17(0.164)

5.08 (0.200)

4.90 (0.193)

4.75 (0.187)

4:57 (0.180)

5.38 (0.212)

5.23 (0.206)

5.08 (0.200)

4.m (0.194)

5.69 (0.224)

5.54 (0.218)

5.38 (0.212)

5.23 (0.206)

594(0.234)

5.82 (0.229)

5.66 (0.223)

5.54 (0.218}

6.20 (0.244)

6.07 (0.239)

5.92 (0.233)

5.79 (0.228)

28
E

vE

CL_

P=2,003N (450 Ib)

P=1,558 N (350 Ib) -

........ ® ........ £).........

- -i

_- P=1,113N (250 Ib)

,- _-_ _ -

-(

0.24

_0.20

0.16

0.12

,0,08

O

"t3

=D

0 1,000 2,000 3,000 4,000 5,000 6,000

Prestrain (Microstrain)

Figure 11. Deflection, W_x, versus preload for constant values of transverse load, P from equation (9).

Load(Ib)

80 160 240 320 400 480 560

7

6 0.24

' 0.08

2

0 500 1,000 1,500 2,000 2,500

Load (N)

Figure 12. Load versus displacement for constant values of prestrain from equation (9).

29
3.1.1 Formation of Damage

The above analysis does not take into account any damage that forms in the laminate. From

equation (5), the energy due to formation of this damage must be included in the total energy. Since

this energy is lost to the system, it will have a negative value and would tend to decrease the terms

in equation (11), implying larger deflections will occur for a given transverse load if damage is present,

thus the "softening" effect.

3.2 Finite Element Technique

For this study, a commercially available finite element program called COSMOS/M was utilized.

The plate was modeled as being statically loaded with a transverse load at the geometric midpoint of the

plate. Two edges were clamped while the other two were free. Figure 13 shows the plate geometry used

for the finite element analysis. Large displacement formation was used where the structural stiffness

matrix was recalculated after each incremental loading of 44.5 N (10 Ib). This was necessary to account

for the changing geometry of the problem since the midplane strains could no longer be ignored; i.e., the

strain within the laminate midplane is given by:

2

 ÷ll w I

Ey "--_ _ "-_,] (12)

re+m+

exy- 0x 0y

where u is displacement in the x-direction and v is displacement in the y-direction. For small displacements

the final terms in equation (12) can be dropped. However, for displacements greater than approximately

one-half the plate thickness, these terms can become significant. The finite element large displacement

formulation performs calculations retaining these terms and recalculating them at each step in the loading

process.

30
Y"_r., _ __

lamped Free

_-" I- -% C f

9.144cm j (3.6 in) "_. / " ..

/

i"

Figure 13. Plate geometry for finite element model.

The plate was meshed with rectangular composite shell elements 1.016 cm (0.4 in) by 0.914 cm

(0.36 in) for a total of 160 elements. Each element contained eight layers of orthotropic material with

properties and directions ([+45,0,-45,90]s) similar to those of the actual specimens to be used. This mesh

provided adequate aspect ratios for the elements and also produced a grid that was not too course, yet not

so fine as to extend computational runtime and memory. The transverse load was modeled as a point load

at the top, centermost node in the mesh. The clamped boundary conditions were achieved by restricting all

displacements and rotations at all nodes along the sides that were to be clamped.

Preload was introduced into the laminate by initially displacing one of the clamped ends a certain

amount in the x-direction.

After each run, the desired output response (z-direction displacement, stress in the principal

material direction of any ply, etc.) could be loaded and plotted.

3.3 Finite Element Results

3.3.1 Transverse Displacement

The most easily compared result between the finite element (FE) analysis and experimental results

is the transverse load/displacement relationship. The FE runs were performed for transverse loads of 223;

445; 668; 890; 1,113; 1,335; 1,558; 1,760; and 2,003 N (50, 100, 150, 200, 250, 300, 350, 400 and

450 lb) with four prestrains of 0; 2,000; 4,000; and 6,000 ge at each transverse load value for a total

of 36 runs.

The output from each analysis was presented as a contour plot and the maximum transverse

displacement (the node under the transverse load) value was recorded. The results from the runs are

presented in table 2. Figure 14 shows results for a constant transverse load P with varying preload and

figure 15 for a constant preload Nx with a varying transverse load.

31
From the results it can be seen that the maximum deflection decreases with increasing prestrain

in a nonlinear fashion. The results from the energy balance formulation showed a linear relationship that

had a much smaller effect (fig. 11). Initially, the added prestrain causes a rapid drop in maximum

deflection and as the prestrain is increases, this drop becomes less pronounced. As the transverse load

increases, the absolute drop in maximum deflection is decreased by a very small amount. That is at a

transverse load of 1,335 N (300 lb), the maximum deflection decreases a total of 2.642 mm (0.104 in)

as the prestrain increases from 0 to 6,000/.re. For the same increase in prestrain at a transverse load of

2,003 N (450 lb), the maximum deflection decreases a total of 2.591 mm (0.102 in).

The load-displacement plots for given prestrain values show that as the prestrain increases, the

maximum load needed to achieve a given deflection increases by a substantial amount. As the prestrains

increase, the load-displacement plots become more linear. This was not the case for the energy balance

results as plotted in figure 2.

The finite element method predicts a larger influence of the preload on the impact response of the

laminate than does the energy balance method. Comparisons with experimental results will be made in

chapter 6.

32
Table 2. Results from the finite element analysis.

TransverseLoad, P

N(tb)

223 (50)

223(50)

223 (50)

223 (50)

445 (100)

445(1oo)

445(ioo)

445 (100)

_8 (15o)

_8 (15o)

(150)

668 (150)

890(200)

890(200)

89o(200)

89o(200)

.1,113 (250)

lr,113(250)

1,..113(250)

1,113(250)

1,335 (300)

1,335 (300)

1,335 (300)

1,335 (300)

1,558 (350)

1_1,558(350)

1,558 {350)

1,558 (350)

1,760 (400)

1,760 (400)

1,760 (400)

1,760 (400)

2,003(450)

2,003 (450)

2,003 (450)

2,OO3(450)

Preload

N/cm (Ibfin)

o(o)

1,229(702)

2.459 (1,405)

3,687(2,107)

o(o)

1,229(702)

2_459(11405)

3,687 (2,107)

o(o)

1,229(702)

2,459 (1,405)

3,687 (2,!0_7)

o(o)

lr229 (702)

2,459 (1,4.0_5)

3,687 (2,107)

o(o)

1,229(702)

2_459(1,405)

3,687 (2,1..0.7)

0(0)

1,229(702)

2,459(1,405)

31687(2r107)

o(o)

1,229(702)

2r459(1_405)

3,687 (2,107)

o(o)

1,229(702)

.2,459 (1,405)

3,687 (2,107)

o(o)

1,229(702)

2,459 (1,405)

3,687 (2,107)

Prestrain

2_000

4,000

6,000

0

2,000

4,ooo

6tO00

0

2,000

4,000

6,000

0

2_000

4,000

61000

0

2,000

4,000

6_000

0

2_000

3.66

Deflection, Wmax

mm (in)

2.82 (0.111)

1.25 (0.0493)

0.833 (0.0328)

0.638 (0.0251)

3.78 (0.149)

2.21 (0.0870)

1.57(o.o618)

1.23 (0.0486)

4.47 (0.176)

2.95 (0.116)

2.20 (0.0866)

1.77 (0.0698)

5.00(0.197)

3.56(0.140)

2.74 (0.108)

2.26(0.0888)

5.46 (0.215)

4.09(0.161)

3.23 (0.127)

2.69 (0.106)

5.87 (0.231)

4.55 (0.179)

4tO00

6_000

0

2,000

4,000

6_000

0

2,000

4_000

6,000

0

3.23

6.22

4.95

4.04

3.61

6.55

5.31

4.42

3.94

6.86

2,000 5.64

41000 4.75

6,000 4.27

(o.144)

(o.127)

(0.245)

(o.195)

(o.159)

(o.142)

(0.258)

(0.209)

(0.174)

(0.155)

(0.270)

(0.222)

(0187)

(0.168)

33
7

•, _ P=2003 N (450 Ib) 0.25

6 "

"- ,,, P=1558 N (350 Ib)

0.20

E

'- 0.15 = 0 0m

== 3 ==

P=1113 N (250 Ib) 0.10

2 (I50 Ib)

1 0,05

P=223 N 150Ib)

0 0

0 1,000 2,000 3,000 4,000 5,000 6,000

Prestrain(Microstrain)

Figure 14. Finite element results for prestrain versus deflection at constant transverse load.

Load(Ib)

0 80 I60 240 320 400 480 560

=

0

1I

,,_I_,,,I,,,

.est=o1oo

o

0 500 1,000 1,500 2,000 2,500

Load(N)

Figure 15. Finiteelement results fordisplacement versus load for constant values ofprestrain.

34
3.3.2 Shape of Deflected Surface

From the contour plots of transverse displacements output by the FE program, the shape of the

deflected surface could be predicted. The surface tended to form a "bowl" shape around the loading point

rather than deflect as a cylindrical surface (as would be expected for long spans). This precluded any

closed form analysis of the problem as one of cylindrical bending which would have greatly simplified

the problem.

Contour plots for four combinations of transverse load-prestrain values are given in figures 16-19.

These values represent both no prestrains and high prestrains at a low transverse load and a high

transverse load.

As the prestrain increased, the edges along the free sides of the laminate deflected less and even

began to deflect upward at the highest prestrain, as demonstrated in figures 7 and 9. With no preload, the

specimen's free edge did deflect, but at a relatively small amount compared to the maximum deflection,

as shown in figures 6 and 8.

3.3.3 Maximum Stresses Within Plies

The maximum stresses within each ply were also calculated by the finite element method.

Table 3 gives results (in principal material directions) for the four highest transverse loads for which

the maximum stresses were calculated. These values are presented to demonstrate general trends

in the maximum ply stresses with increasing transverse load and prestrain. The maximum stresses

presented are those at the center of the respective ply. This value was used since most of the stress arises

from membrane stretching and the values are only an approximation to begin with.

For the lower plies, the maximum stress increases with both increasing transverse load and

increasing prestrain, an expected result. However, this trend also holds for the upper plies as well. This

is to be expected with increasing prestrain but with increasing transverse loads, one might expect the upper

plies to go more into compression due to the higher bending involved. This demonstrates the significance

of the large deflection theory as the higher transverse loads add more to the overall membrane-type

stresses than to the bending stresses. The maximum stresses in the plies in both principal material

directions decrease from the bottom to the upper plies. Note that only the top two plies ever go into

compression in the data presented, with these values being very small. Thus, compressive-type failures are

not expected unless they are due to some mechanism not accounted for in the finite element analysis such

as contact stresses or dynamic effects.

35
Y

P= 223 N(50 Ib)

Prestraln= 0 p.E

2.54 mm

i

/

Figure 16. Contour plot of transverse deflection for a transverse load

of 223 N (50 lb) and no prestrain.

36
i

°

K _

Figure 17. Contour plot of transverse deflection for a transverse load

of 223 N (50 lb) and a prestrain of 6,000/_E.

1.

37
P = 2,003 N (450 Ib)

Prestraln= 0 p._

. 'IT

2.05 mm

5.4!

Figure 18. Contour plot of transverse deflection for a transverse load

of 2,003 N (450 lb) with no prestrain.

38

Table 3. Finite element results for stresses within plies.

Prestrain

0

2,000

4,000

6,000

P=1,335N (300 I))

Ply8 Ply7 Ply6 Ply5 Ply4 Ply3

o2 Ol 02 Ol 02 _ 02 ol _ Ol o2

758 72 641 59 544 45 365 34 227 20 143 10

799 76 675 62 577 48 402 37 258 24 183 13

847 80 717 66 624 53 449 41 323 28 240 19

882 83 737 69 664 57 493 44 411 35 300 22

{ValuesinMPa)

Ply2 Ply1

Ol o'2 Ol o2

49 -1.6 37 -11.0

58 5.2 42 -7.3

114 7.0 54 -1.3

156 12.0 66 4.2

P=1,558N (35(

Ply8 Ply7

Prestrain _ 02 01 02

0 841 80 717 65

2,000 875 83 751 69

4,000 923 88 785 72

6,000 965 9I 834 75

Ib) {ValuesinMPa)

Ply 6 P_ 5 Ply 4 Ply3 Ply2 P_ 1

02 o_ 02 ol 1_. °1 _ 01 02 01 02

612 50 411 38 251 23 201 8.9 54 -.8 41 "-10.0

644 53 448 41 320 27 217 14 66 3.1 45 -6.5

689 58 665 45 464 31 283 21 123 7.5 57 -1.1

730 62 677 48 420 36 318 24 161 13 68 4.0

Prestrain

0

2,000

4,000

6,000

Ply8

Prestrain ol

0 978

2,000 lr013

4,000 1,061

6,000 1,102

P=1,780N (400

PI_/8 Ply7

01 02 Ol 02

909 88 779 72

951 91 813 75

992 g5 854 79

1,034 99 889 61

P=2,003N (4501

Ply 7

o2 01 a,_

94 847 77

98 882 81

102 916 85

105 951 87

b) (ValuesinMPa)

Ply2 PV1

al 02 Ol 02

59 2.4 45 -8.9

76 3.5 49 -5.6

134 7.9 60 -0.6

165 13 70 4.2

Ply6 Ply5 PI_/4 Ply3

01 .oz 01 02 01 02 Ol oz

677 54 456 42 314 26 226 12

710 58 492 45 336 30 259 17

751 62 544 49 379 34 312 23

792 66 601 52 430 38 341 26

)) {ValuesinMPa)

Ply6 Ply5 Ply4 Ply3 Ply2 Ply1

Ol 02 Ol 02 Ol 02 01 c_ 01_... 02 Ol 02

737 59 500 46 340 29 263 16 63 6.1 48 -7.6

772 62 535 49 350 33 294 17 84 7.0 52 -4.5

813 67 557 53 431 37 347 21 143 8.3 62 0.2

847 70 613 56 440 40 368 26 169 14.0 72 4.7

The strength value in the second principal material direction (-,35 Mpa (5,100 psi) as determined

by 90 ° tension tests) in the bottom four plies is always exceeded. However, synergistic effects with

neighboring plies can help keep the ply from failing in this direction at the values obtained by the 90 °

tension tests. It is still expected that much matrix cracking parallel to the fibers in these plies should occur,

especially in the bottom ply which has the highest stresses and also has only one neighboring ply to help

"hold together" this ply.

The breaking stress (--2,700 Mpa (400,000 psi)) in the f'trst principal material direction (along

the fibers) is never exceeded in any of the plies, thus fiber breakage is not expected to occur for the loadprestrain values presented. In fact the stresses calculated are never even one-half that of the breaking stress

of a ply in the fiber direction. Thus, if fiber breakage is found, it must be due to another mechanism other

than the stresses set up solely by the transverse load-prestrain combination.

40
4. EXPERIMENTAL

Before the specifics of this particular research program are given, a few general comments about

the empirical study of impact to composites are in order; however, direct comparisons between studies are

very difficult due to a lack of standardization, allowing whatever convenient specimen geometry or support system best suits the researcher. However, a hasty rush for a standard can, and would, do more harm

than good at this point in the evolution of composites, since much remains to be understood about the

many variables involved in impact testing. Even a standard meant to compare composite materials for

damage resistance and damage tolerance (such as the Boeing 48-ply CAI standard) can be very misleading. For example, in this standard, a 48-ply specimen of a quasi-isotropic stacking sequence, 8 inches

by 10 inches in size is impacted with 1,500 in-lb of energy per inch of specimen thickness. Assuming

all impact events are conducted with the same boundary conditions, impactor size, shape, and velocity

(which they are not), the only response measured is damage tolerance in a structurally compressive test,

which is inconclusive for assessing a fiber/resin system for impact damage tolerance. It should be noted

that compression testing a panel of this size is a structural, not material, property test. The most important

parametermvelocity of the impactor--is not specified, thus high-velocity/low-mass impacts could be used

to obtain the required energy level, resulting in very localized damage and a smaller reduction in CAI

strength than would have been found from a high-mass/low-velocity test. In addition, 48 plies is a

relatively thick laminate and would not be encountered in many engineering applications, thus bringing

the issue of scaling into the problem. The thick laminate is called for to minimize the global buckling

of the plate during compression testing but during the impact event, the flexural stiffness is so high

that the large interlaminar shear stresses that normally develop in an impact event are minimized. Thus,

a material with a higher mode II shear strength would not register as being more damage resistant

than a material with a lower mode 1I shear strength, although it would for thinner test panels.

Any standard developed is going to have to examine a range of impact energies. The 1,500 in-lb/in

called for on the Boeing CAI tests represents just one datum. A material may perforate at this level, giving

a CAI strength equal to the compression strength of a panel with a hole the size of the impactor drilled

through it. Another, more damage-resistant/tolerant material may develop a large area of fiber breakage

(not penetration) associated with delaminations that could produce a much lower CAI strength than the

less damage-resistant material. At lower impact energies this would have been detected, since perforation

would not have occurred in the less damage-resistant material and a much larger area of damage would

have developed whereas the tougher material would probably show a smaller region of damage and

a higher CAI strength. In other words, it is entirely possible to have material A be more damageresistant/tolerant at one impact energy level than material B, but at a different impact energy level,

material B can be more damage-resistant/tolerant than materialA.

4.1 Material

The fiber/resin system chosen for study was IM7 fiber in 977-2 toughened epoxy. This material

was readily available and represents a very common type of system (high-strength carbon fibers in a

toughened epoxy) being used on many programs for NASA. The layup configuration chosen for the

41
impacttestswasaquasi-isotropiconeof [+45,0,-45,90]s.Thisconfigurationistypicalof thebasiclayup

unit beingusedfor manyprimarystructuresonair andspacecraft.Thematerialwasprocessedfrom unidirectionalprepregtape as flat panels, using a hot press. The cure cycle consisted of vacuum bagging, then

heating the panel at 1.7 °C (3 °F) per minute up to 177 °C (350 °F) and holding for 2 hr. Cooling was

then ramped back down to room temperature at a rate of 2.8 °C (5 °F) per minute. A constant pressure

of 550 kPa (80 psi) was applied during the entire cycle. The resulting panels had nominal ply thicknesses

of approximately 0.12 mm (0.0047 in).

A select series of tests were performed on specimens to determine some key material property

values. These specimens were manufactured from the same prepreg roll as the impact specimens. The

cure cycle also remained the same for these specimens.

4.1.1 0 ° and 90 ° Tensile Testing

Unidirectional laminates of IM7/977-2 were tested to obtain load/strain data so the modulus

and Poisson's ratio of the laminae in the fiber direction and perpendicular to the fiber direction could

be determined. Strain-gauge rosettes were attached to five different specimens of each orientation and

readings were taken at 445-N (100-1b) increments for the 0 ° specimens and at 4.45-N (10-1b) increments

for the 90 ° specimens. These data were then plotted and a least-squares regression line fit determined the

measured modulus of elasticity and Poisson's ratio in the respective direction. The averages were

El1=1,867 Gpa (27.1 msi), v12=0.37, E22- 9.85 Gpa (1.43 msi), and v21--0.0187.

4.1.2 -+45 ° Tensile Testing

Tensile testing was performed on -+45 ° specimens to determine the shear modulus of the laminae.

Twenty-two-ply specimens were tested with strain-gauge rosettes with readings being taken at 223-N

(50-1b) increments. The stress divided by two versus lell + le21 was plotted and a least-squares regression

fit on the linear portion of the plot gave an average G12 of approximately 7.03 Gpa (1.02 msi).

4.1.3 Laminate Testing

Tensile specimens of the eight-ply [+45,0,-45,90]s laminates were also tested with strain-gauge

rosettes. Readings were taken at 445-N (100-1b) increments and the data were plotted and a least-squares

fit used to determine the tensile modulus and Poisson's ratio. The average values were Ex=-Ey=60.5 Gpa

(8.78 msi) and v12--0.32.

Strength measurements were also taken on these laminates. It should be noted that edge delamination occurred before ultimate fracture. This result is expected as high tensile normal stresses can be set up

in the region of a free edge on a tensile specimen as outlined by Pipes and Pagano. 126 The ultimate average failure strain was 12,875/_e which corresponds to an average stress of 779 Mpa (113 ksi). The edges

began to noticeably delaminate at approximately 689 Mpa (100 ksi). The delamination extended far into

the specimen width before fracture of the 0 ° plies. A full analysis and discussion of this phenomenon

is beyond the scope of this text and the interested reader is referred to other references for a more detailed

treatment of the subject.127,128

42
4.2 Impact Specimens

The specimens manufactured for impact testing consisted of eight-ply quasi-isotropic samples

with a stacking sequence of [+45,0,-45,90]s, cut into coupons 9.5 cm (3.75 in) wide and 29.8 cm

(11.75 in) long. The coupons were cut from the cured panel with a diamond-wheel table saw. Fiberglass

end tabs 11.4 cm (4.5 in) wide were bonded onto the ends of the specimen with epoxy adhesive to

facilitate loading into the pretensioning device. The specimen dimensions are given in figure 20.

< 29,8 cm (11.75 in)

......

9.5 cm (3.75 in) Gauges

1......

(1.25in)

Figure 20. Specimens used for impact testing.

4.2.1 Specimen Strain-Gauging

Each specimen had two strain gauges bonded to it so that a uniform prestrain was ensured. The

gauges were 1.3 cm (0.5 in) from the specimen's edges to detect strain gradients across the width (see

fig. 20). This instrumentation was determined from a series of preliminary tests on samples that contained

numerous strain gauges on the surfaces of the specimen. From these tests it was found that bending was

never introduced into the specimen so gauges need only be placed on the same surface. In addition, it was

found that the strain remained constant along the longitudinal axis of the specimen and the only variations

in strain were across the width of the specimen. As long as the outermost gauges (1.3 cm (0.5 in) in from

each edge) had the same strain, the ones between them also had this same strain, thus it was determined

that only the two gauges nearest the edges and on the same side need be measured to ensure a uniform

strain field was being set up in the specimen.

4.3 Pretensioning Device

The device to apply tensile preloads to the specimens basically consisted of two wedge grips connected to a load cell and a hydraulic piston. As the piston was pumped with hydraulic fluid it would cause

the grips to move apart, putting the composite specimen in a state of tension. A sketch of the apparatus

with the various parts is given in figure 21.

43
SideView

Load Cell--"--3 Wedge Grlp

Upper \ (4 total)

Support

Plate

! I

I •

I I

LI

BottomClamp _ 'i Specimen

Plate - Base Plate

Top Clamp

F- Plate(2 total) /

/ I

I!!1_--

I I

L!

Clamp Bolt

(12 total

_-- Wedge

Adjustment

Block

(2 total)

(Tierods joining2 wedgeadjustmentblocks not shown)

TopView y BasePlate

,/ --Top Clamp Plate

OpperSupport--_ __j-_//////////////////////_ "__ --_-,

Plate

I°1Fspec'men Iql

i |l I'--"I I II II

, | |'

i' - "I |i

........ I i

I i

I ¢ I' I L =I I

! I

I.. _ I Ii

L_ __ ..-Z

Wedge

Adjustment /

Block

i i, Ii

, " 'I

I II II

I II II

ill ,,

o/

Figure 21. Fixture to apply preloads.

44
Thewedgeadjustmentblockswerefirst tightenedtogetherto lodgethespecimenin thegrips.

Theseblocksalsopreventedthewedgesfrom movingapart(andstopgripping) during the application

of the preload. As the preload was applied and the strain gauges were monitored, when one gauge deviated

from the other, that side reading the lower strain could have its adjustment blocks loosened, allowing the

specimen to expand slightly along that side, causing the strain to rise on that side. These adjustments were

made until the final level of strain was reached and the adjustment blocks could be fine-tuned to give

a nearly uniform prestrain across the width of the specimen.

4.3.1 Boundary Conditions

The boundary conditions chosen for this study were clamped-clamped/free-free with the clamped

edges perpendicular to the applied preload. This geometry was chosen since it was demonstrated in a previous program 129 that this condition most closely represented a pressure vessel being impacted. Pressure

vessels are of vital interest to this study since they are tension-carrying structural members.

Once the level of prestrain was achieved, the top clamp plates were tightened down with the

12 bolts, each at a torque of 28 N-m (250 in-lb), to obtain the desired boundary condition. A true clamped

boundary condition would mean that the hydraulic piston could be released and the specimen would continue to carry the prestrain between the region of the clamps. However, in practice this was not the case

and was never achieved despite trying a number of techniques. The methods that came closest to achieving

this tended to crush the specimen in the region of clamping; therefore, it was determined that the hydraulic

piston would be held in place during the impact event. Measurements of the strain before and after impact

indicated that the specimen did not move significantly in the clamps. The high torque on the clamp-down

bolts ensured that no bending moments could occur under the clamps.

4.4 Impact Testing

Once the desired level of prestrain was achieved, the preloading device with the specimen was

positioned under a Dynatup 8200 drop tower as shown in figure 22. The height of the crosshead was

adjusted to give the desired impact energy. In this study, the mass of the crosshead with tup was 2.4 kg

(5.22 lb). Since the crosshead/guideposts interface does not contain linear bearings or another such

friction-reducing device, an exact prediction of impact energy from the drop height was impossible.

Although this friction is present, the measured impact energy is quite precise since the velocity of the

falling crosshead is measured just prior to impact. Once a desired drop height was decided upon, all

impact tests at that corresponding impact energy level were conducted before the cross head was moved

again. This ensured uniformity in the impact energy levels used. A total of three impact energies were

used in this study: 3.4, 4.5, and 6.0 J (2.5, 3.3, and 4.4 ft-lb). These levels induced damage from visible

to nonvisible.

45
Guideposts ,2 •

Weighted

Crosshead

--- PCData

Acquisition

[r lr

--J

Figure 22. Preload and impact apparatus.

4.5 Postimpact Testing

In order to assess the effects of tensile preloads on the damage resistance of the specimens tested,

three forms of postimpact inspection were used. The first type of inspection was a visual one to see if any

damage could be detected from either surface of the plate. This was followed by a nondestructive x-ray

inspection and then a destructive cross-sectional examination of the specimen through the damage zone.

4.5.1 Visual Inspection

After each specimen was impacted, it was carefully removed from the preload device and

inspected for visual damage. No magnifying devices were used; however, if damage could be felt and not

seen, this was recorded as "visual" damage since no special tools were needed to identify this damage.

The extent and type of visual damage was recorded for each impacted specimen. The type of damage that

could be detected consisted of back face (nonimpacted side) matrix splitting, front face (impacted side)

indentation damage, and small matrix splits being felt on the back surface.

4.5.2 X-ray Inspection

For specimens that did contain observable damage on one of the two faces, a small circular dam

of plumber' s putty was placed around the damage on one of the faces to hold in a zinc iodide (ZnI) dye

46
penetrantsolutionasshownin figure23.Thedyepenetrantis opaqueto x raysandwasallowedto soak

onthespecimen(atadamagesite)for atleast24hr.Thesolutionconsistedof 100g of ZnI powder

in 100ml of waterand100ml of KodakPhotoflow_' which breaksthesurfacetensionof thesolution

to allowfull penetrationintoall cracksanddelaminations.

ZincIodide

Solution

CircularDamof

Plumber'sPutty

Figure 23. Dam of plumber' s putty to hold zinc iodide solution over damage area.

For specimens that did not contain damage on one of the two faces, a small hole of -0.5 mm

(0.02 in) was drilled through the specimen at the impact site to allow the dye penetrant to be exposed

to any through-the-thickness damage that may have occurred within the laminate.

The specimens were wiped clean of the plumber' s putty and excess penetrant after at least a 24-hr

soak and exposed to x rays with a Torrex 150D x-ray inspection system manufactured by EG&G Astrophysics Research. Polaroid type 55 film was placed directly under the specimen to obtain a permanent

copy of the x-ray signature of each specimen. The dye penetrant was seen to seep into all matrix delaminations and cracks and was easily detectable on the developed film.

4.5.3 Cross-Sectional Examination

Some of the specimens were sectioned through a part of the damage zone in an attempt to gather

more detailed information about the induced damage. This was particularly helpful in identifying areas

where fibers were broken, and in identifying the interface of a delamination.

A technique was developed during this study whereby the location of the cross-sectional cut

relative to the damage as seen by the x rays could be determined. The x rays, which corresponded to

the specimens at a 1:1 ratio, were photocopied onto film transparency. The clear film now had the image

of the x ray at a 1:1 ratio with the specimen. Strain gauges that would register on the x rays provided

accurate guides as to the location that the film needed to be placed to be a surface marker of the internal

damage. The area of impact on the specimen was painted white to enhance the visibility of the x ray

on the clear film transparency which was matched to the specimen and applied with double-sided tape.

The damage zone was then cut from the specimen and two clear acrylic plates were used to sandwich

47
thecut specimenwith thex-raypatternonthesurface.Theseclearplasticplateswerebondedto the

specimenwith cyanoacrylatesterglue.Thex-raypatterncouldbeseenclearlythroughtheacrylicplate.

Any cutsmadeonthesandwichedspecimencouldnowbeassociatedwith thedamage,asseenby

thex-rayinspectiontechnique.

A microscopeequippedwith aPolaroidcamerawasusedto examinethecross-sectionedspecimens.Magnificationsbetweenx 7.25andx 160werepossible.

4.6 Static Indentation Testing

In order to better understand the load/deflection and in-plane strain response of the transversely

loaded specimens, a series of static indentation tests were performed using a dial indicator to measure

transverse displacement and various strain gauges to measure the surface strains at given locations. The

pretensioning device was placed in an Instron 1125 load frame and the indentor was attached to the upper

crosshead/load cell. After applying the desired amount of prestrain, the specimen was incrementally

loaded with the indentor at the center of the specimen (as in the impacts) and transverse displacement

and strain-gauge readings were taken. This step process was continued until a cracking noise was heard,

indicating specimen damage.

The specimens were unloaded in the same incremental fashion to obtain a complete loaddisplacement history of the entire process. This could provide valuable information on how much energy

is being lost during the process.

In addition, two static indentation tests were performed in which the specimen was loaded to failure. Surface strains were not monitored but load-displacement data were. These two tests were performed

to gain insight on the load-deflection behavior of specimens with a medium level of prestrain, as well as

gather information on the static behavior once damage becomes severe in the specimen.

4.6.1 Static Indentation Strain Gauging

Strain gauges were placed on the specimens so that information could be obtained about the surface strains at areas of interest. For practical reasons, strain gauges could not be placed at the center of the

specimen since the indentor would interfere with it from the top and the dial indicator from the bottom.

A number of strain-gauge setups were used on the 10 samples tested. Table 4 lists the strain-gauge placement on each sample along with the rationale behind choosing these locations.

4.6.2 Loading/Unloading Curves

As mentioned previously, the transverse load was applied incrementally. After initial readings

were taken at zero-applied load, the specimen was loaded with 44.5 N (10 Ib) and a second reading taken.

The third reading was taken at 111.25 N (25 lb) and then readings were taken at 111.25-N (25-1b)

increments thereafter until a popping or cracking noise was heard, at which point the specimen was then

unloaded in 222.5-N (50-1b) increments. The transverse load was applied at a rate of 1.27 mm/min

(0.05 irgmin) between stops.

48
Table 4. Specimens for static indentation tests.

Specimen StrainGaugePlacement

4

2F-q

IO

11--I

1_6

4[---'1

3[Z]

IO

21--1

1E3

5EZ]

10[_l (9)

4[_1 (5)

3 !---] (6)

IO

2 [_ (7)

1I_l (8)

2 J_l (5)

Gauge x1 yl

1 67 13

2 67 82

]2 79 47

Objective

To test setupandobtain

load/deflectiondata.

Gauge

Gauge

1

2

3

4

(5)3

(6)

(7)

(8)

9

(10)

Gauge

1

2

3

4

5

6

!

79

79

79

79

140

13

79

x

79

79

79

79

79

79

79

79

10

19

x

7a--

73

146

73

73

146

79

13

38

58

83

48

48

48

Y

13

38

57

82

82

57

38

13

47

47

Y

82

13

47

82

13

47

47

Toexamineacross-the-width strain

fieldand strainnear centerendson

top surfaceof a low prestrained

(500/_e) specimen.

To examineacross-the-widthstrain

field on top and bottom sidesof

specimenof a low prestralned

(500/_c) specimen.

To examinetop and bottomstrains

nearedgesand ends at a moderate

(4,500/_).

Note: Specimen not clamped.

1. Dimensionsinram.

2. ! denotesIndentorlocation,

3. Numbers In parenthesisIndicatestraingauge on oppositesideof specimen,

49
Table 4. Specimens for static indentation tests (continued).

Specimen Strain GaugePlacement

2r---I (4)

io 3 [_] (6)

1l--1 (5) X

2 [_1 (4)

5 [_1(6) I0

1 [_] (3)

(8)

E_

4

2r--l (6)

(7)

[_1 i0

3

1F-'] (5) X

2 r--I (6)

4 3

I--1 I--] io

(8) (7)

1 [_1 (5) X

p,

Gauge x y

1 86 13

2 86 78

3 137 45

4 86 78

5 86 13

6 137 45

I 79 45

Gauge x y

1 79 13

2 79 78

3 79 13

4 79 78

5 17 45

6 17 45

! 79 45

Gauge x y

1 92 13

2 92 82

3 57 47

4 13 47

(5) 92 13

(6) 92 82

(7) 57 47

(8) t3 47

! 79 47

Gauge x y

1 79 13

2 79 83

3 57 48

4 13 48

5 79 13

6 79 83

7 57 48

8 13 48

] 79 48

Objective

To replicate specimen#4 with

plates clamped.

To replicate #5 witha low

prestraln (300 pc).

To examinethe changeIn strainfrom

nearan end toward theindentor

locationat a small prestrain(300 _.s).

To replicate #7 at a higher prestrain

(4,100 pc).

1. DimensionsIn mm.

2. I denotesindentorlocation.

3. Numbersin parenthesisIndicatestraingaugeon oppositesideof specimen.

5O
Table 4. Specimens for static indentation tests (continued).

Specimen

9

10

7

[--1

(4)

8

I--I

(3)

5

EEl

(2)

6

(1)

StrainGaugePlacement

4 2

(8) (6)

[23

3 1

IO

x

IO

Gauge x y

1 51 13

2 51 78

3 13 13

4 13 78

(5) 51 78

(6) 51 13

(7) 13 78

(8) 13 13

I 79 46

Gauge x y

1 13

2 82

3 13

4 82

(5) 82

(6) 13

(7) 82

(8) 13

[ 47

1. Dimensionsin ram.

2. [ denotesindentorlocation.

3. Numbersinparenthesisindicatestraingaugeon oppositesideof specimen.

Objective

Examinestrainnearcenterendof

specimen.

To replicate#9ata low

prestrain(400/_).

51
5. RESULTS

This chapter presents the results of the experimental portion of the work. Since this study is

empirically based, the material presented is quite detailed, and most of the raw data are presented in the

appendices. Every effort has been made to make this chapter as complete and thorough as possible. Discussion of the particulars of the results and noteworthy comparisons are not made in this chapter but rather

will be made in chapter 6 which is devoted entirely to a discussion of the results.

5.1 Instrumented Impact Testing

The amount of tensile preload on the impacted specimens had a measurable effect on the maximum load of impact, duration of impact, maximum deflection, and absorbed energy of the specimen.

The following sections will present the results from these data.

5.1.1 Maximum Load of Impact

As calculated in chapter 3, the higher the tensile prestrain, the higher the impact force should

be for given impact conditions. This was clearly the case for the specimens examined in this study.

The results for each of the three impact levels used are shown in figures 24-26. As damage forms

in the specimen (not accounted for in the analysis in chapter 3), the maximum load of impact tends

to level off, and even drop some with increasing prestrain (which tends to increase the damage).

z

-d

E

E

=[

1,650

1,600

1,550

1,500

1,450

1,400

1,350

1,300

_ _ 370

i

m

m

n

I

B

B

m

D

n

B

m

7__ • --

m

Illllllllllll II Vm

360

z

350

- 34O

330

320

310

300

0 1,000 2,000 3,000 4,000 5,000 6,000 7,000

--.I

E

E

Prestrain(Microstrain)

Figure 24. Maximum load of impact versus applied prestrain for 3.4 J (2.5 ft-lb) impacts.

52
1,850

1,800

1,750

Z

v

1,700

..1

_lilillIIililillllJ111111111111=llilil'_

_-- • :

m m

D

m

w

B

m

m

u

I

I

408

400

392

384

376 ..u

368

360

1,650 "

"-C1,600 _--- • __

1,550 3j=,l,_=_lli,,l,],,Im,,=l,_,,l,,,,l,,_ 352

0 1,000 2,000 3,000 4,000 5,000 6,000 7,000 8,000

Prestrain (Microstrain)

Figure 25. Maximum load of impact versus applied prestrain for 4.5 J (3.3 ft-lb) impacts.

2,150

2,100

z 2,050

C=

-J

2,000

1,950

1,900

_lll'l'l II ll'lll''ll lllll Illl Illl Ill I'lL__

• • -:

m

i

i

B

m

7,,,Itii,11,,,l,,lilillil,,,,111,,l,,,_

0 1,000 2,000 3,000 4,000 5,000 6,000 7,000 8,000

480

472

464

456

448

440

432

8

.--I

Prestrain (Microstrain)

Figure 26. Maximum load of impact versus applied prestrain for 6.0 J (4.4 ft-lb) impacts.

5.1.2 Duration of Impact

Figures 27-29 show that as the tensile prestrain increased, the duration of the impact event

decreased. This variable was not examined in the analysis, but as the maximum deflection decreases

53
with increasing prestrain (as will be seen in the next section), it is expected that the duration of impact

should follow the same trend. Note the nonlinearity in this relationship.

12

11

_" 10

E

G)

E 9

Prestrain (Microstrain)

Figure 27. Total time of impact versus prestrain for 3.4 J (2.5 ft-lb) impacts.

E

v

E

I---

11.0

10.5

10.0

9.5

9.0

8.5

8.0

7.5

7.0

I ' llllllll l''"l lllll lllllallllt•e'

--- • "-s

- • -

- • • -

- • • • -

-ll,illllJlI,IIl,,,,llllJlllll I,i,JI,IIF

0 1,000 2,000 3,000 4,000 5,000 6,000 7,000 8,000

Prestrain (Microstrain)

Figure 28. Total time of impact versus prestrain for 4.5 J (3.3 ft-lb) impacts.

54
_D

tt_

E

v

I--

10.5

10.0

9.5

9.0

8.5

8.0

- •

- •

•

0 1,000 2,000 3,000 4,000 5,000 6,000 7,000 8,000

Prestratn (Microstrain)

Figure 29. Total time of impact versus prestrain for 6.0 J (4.4 ft-lb) impacts.

5.1.3 Maximum Deflection

The maximum deflection of impact was predicted to decrease with increasing tensile prestrain as

determined in chapter 3. Plots of the measured data are shown in figures 30-32. Th_ese plots are nonlinear

in a similar fashion to those from the previous section. The maximum deflection data begins to drop rather

rapidly with increasing prestrain, then level off at the higher end of prestrains.

E

E

v

.e,

s=

E

7.0

6.5

6.0

5.5

5.0

4.5

jb, ,i ,,,,l,l ,,i,,,,i,, ,,i,,,,i,,,,_

B

,, ,I,, ,, II ,, ,I ,, ,,I,, ,,ll, i ,_,,

0

0.26

0.24 =O

0.22 E

E

0.20 _

0.18

1,000 2,000 3,000 4,000 5,000 6,000 7,000

Prestrain(Microstrain)

Figure 30. Maximum center deflection versus prestrain for 3.4 J (2.5 ft-lb) impacts.

55
E

E

v

rr'_

E

E

8.0

7.5

7.0

6.5

6.0

5.5

5.0

4.5

-Illll,l''llllll',,,lllllllllllZlllll,lL -- m

-00

D

m

B m

B m

- O0 z

- O0 -

- • "I

-_,,,l,,,,I,,,,I,l,,I,,,,I [ijlljlllll I_-

0 1,000 2,000 3,000 4,000 5,000 6,000 7,000 8,000

0.30

0.28

0.26

0.24

0.22

0.20

0.18

O

G3

E

Prestrain(Microstrain)

Figure 31. Maximum center deflection versus prestrain for 4.5 J (3.3 ft-lb) impacts.

8.5

8.0

E_ 7.5

7.0

t'_

_E 6.5

_: 6.0

5.5

5.0

_i_, I" " I" " I" " I" " I" " I" " I" "

- I -

-,, ,, 1,, ,, I,, ,, I,, ,,I,, i, I;, ,, l,l,l I,, ,--_

0.32

0.30

0.28

0.26

0.24 _

0.22

0.20

0 1,000 2,000 3,000 4,000 5,000 6,000 7,000 8,000

Prestrain(Microstrain)

Figure 32. Maximum center deflection versus prestrain for 6.0 J (4.4 ft-lb) impacts.

5.1.4 Absorbed Energy of Impact

The amount of energy lost during the impact event is termed "absorbed energy." The absorbed

energy versus prestrain plots are given in figures 33-35. For low prestrains the absorbed energy is seen

to decrease with increasing prestrain up to a critical level, at which point it begins to rise. The smaller the

56
impactenergythelargertheinitial dropin absorbedenergywithincreasingprestrain.Thisis thoughtto be

aresultof compliancein thepreloadingdevice.Sincea trueclampedconditionwasnotpractical,at low

prestrainsthespecimencoulddisplaceagivenamountinthefixture,thuslosingagivenamountof

energy.As theimpactenergiesincreased,thisamountof energybecamealowerpercentageandwasless

noticed.As theprestrainincreased,thegripstendedto helpholdthespecimenin atrueclampedconfigurationandtherewaslessdisplacementof thespecimen,thuslessenergylostdueto thismovementof the

laminatebeingtested.

v

,,=,

0

.o

2.35

2.30

2.25

2.20

2.15

2.10

2.05

2.00

-'" 'I''"I "' 'I'"'I' '" I" "I '"±

-o

•

-- • --

• -

- • --

1.72

1.68

1.64

1.60

1.56

1.52

1.48

u

==

nlu

"8

_3

,<

1.95 1.44

0 1,000 2,000 3,000 4,000 5,000 6,000 7,000

Prestrain(Micr0strain)

Figure 33. Absorbed energy during impact versus prestrain for 3.4 J (2.5 ft-lb) impacts.

4.0

- O-- 2.8

-_ • Z _-

- 2.6 "" 3.5 _ -- =

v

ILl __ C

"o - -- 2.4 "'

_- Oo -

< 3.0 -- -- 2.2 ,<

-0 • -

- • • -- 2.0

-- •

2.5 n,,,l=nn,l,_,l,,,,l=,,,l,,,,l,,,,l,,,,

0 1,000 2,000 3,000 4,000 5,000 6,000 7,000 8,000

Prestrain(Mtcrostrain)

Figure 34. Absorbed energy during impact versus prestrain for 4.5 J (3.3 ft-lb) impacts.

57
kid

O

t_

5.0

4.8

4.6

4.4

4.2

4.0

3.8

3.6

_l ii I I, ,, _11ii ill nl Ill ii lil I, ii_lJ _1' '_ L

-_ • •

- •

-O

7,,,I,,e,,io,,,ll,,,I,,,,n,,,,I,,,,i,,,r

0 1,000 2,000 3,000 4,000 5,000 6,000 7,000 8,000

3.6

3.4

Prestrain (Microstrain)

i

!

3.2 >,

eLM

a.0 0

2,8

Figure 35. Absorbed energy during impact versus prestrain for 6.0 J (4.4 ft-lb) impacts.

5.1.5 Load-Deflection Curves

For all of the specimens tested, the load-deflection plots show that a significant amount of energy

is lost during the impact event, even at the lowest impact energy level used. A typical load/deflection curve

is shown in figure 36 for the lowest level impact at the smallest prestrain and in figure 37 for the highest

level impact at the highest prestrain. These two plots represent the two extreme cases of impact tested in

this study. All of the load/deflection curves are given in appendix A.

58
Deflection (mm)

700

60O

5OO

4OO

300

2OO

100

0

1 2 3 4

I I I I

5 6 7 8 9 10

I t I I I I

Specimen # 12/17A

I I I I I I

0 0.1 0.2 0.3 0.4

3,000

1,780

50O

w

v

8

._1

Deflection (in)

Figure 36. Load/displacement curve for lowest level impact at lowest prestrain.

Deflection(mm)

0

.-,I

0 1 2 3 4 5 6 7 8 9 10

,oof, 600

5O0 ,,,,,

Specimen

,

12/12A

,,l ,ooo

400 - 1,780

30O

20O

- 500

IO0

1 I I

0 0.1 0.2 0.3 0.4

I 1 I I

Deflection(in)

Figure 37. Load/displacement curve for highest level impact at highest prestrain.

59
5.2 Visual Examination

In general, as the preload increased, the amount of visual damage also increased. This corresponds

to the higher impact force induced into the specimen. The first sign of visible damage was always splitting

of the matrix between fibers on the back 45 ° face. Contact stresses produced visible damage at the impact

site in some of the specimens. Table 5 gives values for all of the specimens tested, along with a description of the visual damage.

60

Table 5. Specimens tested with associated parameters and results.

Specimen Impact

No. Energy

.... J)

L1 4.5

1_2 4.5

L3 4.5

L4 4.5

L5 4.5

L6 4.5

11/'13A 4.5

11/1313 4.5

11/14A 4.5

Prestraln M_mum Maximum Total Absorbed

(me) Load(N) Deflection .time Energy(J)

(n'm) (rrsec)

5,182 1_11 5.3 .....8;06 3.16

3,229 1,682 6.1 8.70 2.73

452 1,615 7.6 10.79 2.93

833 1,598 7.6 10.73 2.73

7,399 1,664 4.8 7.88 3.81

5,564 1,793 5.3 8.31 3.11

2,430 1,700 6.4 8.95 2.52

1 960 lr660 6.4 9.31 2.66

7,970 1,736 4.8 7.36

11/14B 4.5 4,205

11/25A 4.5 4,268

11/27A 4.5 6,698

11/28A 6.0 5,384

11/268 6.0 498

12/12A 6.0 7,173

12/12B 6.0 2,132

12/12C 6.0 4,156

12/12D 6.0 3,053

12/13A 6.0 1,594

12/13B 6.0 634

12/14B 3.4 6,972

12/14C 3.4 987

12/14D 3.4 3,387

12/14E 3.4 5,012

12/17A 3.4 400

12/17B 3.4 4,334

12/17C 3.4 6,060

17./17D 3.4 2,045

1,825 5.6 8.01

1,829 5.6 8.25

1,727 5.1 8.14

2,069 5.8 8.13

2,025 8.4 10.21

2,074 5.3 8.29

2;100 7.1 8.86

2,123 6.4 8.06

2,096 6.6 8.59

1_967 7.6 9.20

1,931 ....... 8.4 10.49

1,633 4.6 7.44

1,335 6.6 10.66

1,477 5.3 8.69

1,562 4.8 8.09

1,308 6.9 1134

1,531 5.1 8.38

1,615 4.6 7.85

1,406 5.8 9.50

3.35

2.51

2.92

3.65

4.80

3.89

4.66

3.70

4.26

4.22

3.70

4.04

2.32

2.14

2.07 None

2.28

2.24

2.09

2.12

1.98

Visual

Damage

22-mm spliton backface

None

18-mm spliton backface

17-mm spliton backface

116-mm spliton backface

Dentw/short split, front

47-mm spliton back face

17-ram split onbackface

18-mm split onbackface

65-mm split on backface

Dentw/smallsplit, front

None

30-mm spliton backface

60-mm spliton backface

Dent w/small split, front

Split across width

Dentwith short split onfront

21-mm split onback face

Splitacrosswidth

Dentwithsplitson front

18-ram spliton backface

45-ram spliton backface

Dentw/smallsplit,front

53-ramspliton backface.Dentonfront

20-ramsplitonbackface

19-ram spliton backface

27-mm spliton back face

None

26-mm spliton back face

None

19-mm spliton backface

24-mm spliton backface

None
5.3 X-ray Inspection _J:

As expected, the specimens that contained more visible damage also contained more internal damage, mostly in the form of matrix splitting and delaminations. The specimens that had no externally visible

damage demonstrated that some delaminating occurred (the small drilled hole allowed the penetrant to

reach these areas). These delamination zones were relatively small compared to the specimens that had

visual damage associated with them. A typical x ray of this type is shown in figure 38. Note that matrix

cracking is associated with the delaminations, a feature seen in all of the specimens.

O

0 0.1 0.2 0.3 0.4

1 I

o 5 lO

I I II I

mm

Figure 38. Typical x-ray signature for a specimen that exhibits no visible damage.

For the specimens that did contain externally visible damage, the matrix crack on the back 45 ° face

was clearly visible and always had delaminations emanating from the crack. These delaminations tended

to form a lobe shape and end at a crack in the ply above the 45 ° ply (the 0 ° ply). A typical x ray with this

form of damage is shown in figure 39.

61
0

I I I

0 0.1 0.2 0.3 0.4

I I I I I

in.

5 10 15 20

I I II I1 I II I !1 I I I I II

Figure 39. Typical x-ray signature for a specimen that contains back face matrix splitting.

As the damage in the specimen became more severe due to higher preloads, damage on the

impacted side of the specimen would begin to form. For these specimens, the dye penetrant would tend

to "pool up" and form very dark areas on the x-ray signatures as shown in figure 40. This made detection

of broken fibers difficult since these breaks are hidden by the large amount of dye penetrant. Crosssectional examination was the method used in this study to detect broken fibers. All of the x-ray signatures

of damage are shown in appendix B. The static indentation test samples are also included.

62
0

0

II IIll Illl II Ill

Figure 40. Typical x-ray signature for a specimen that exhibits visible damage on the impacted side.

5.4 Cross-Sectional Examination

For specimens in which the x rays did not provide sufficient enough detail (usually associated

with fiber breakage), cross-sectional examinations proved useful. The x rays gave an accurate description

of the damage formed due to the impact event, but fiber breakage was difficult to detect. When specimens

were cross-sectioned through the dark areas of the x ray, fiber breakage could sometimes be seen. Crosssectional photographs of selected specimens, along with their x rays showing where the cuts were made,

are shown in figures 41-49. Note that the 0 ° fiber breakage areas are always in very close proximity to the

matrix cracks both above and below this area because these areas are subject to the highest stresses, and

also, as a matrix crack forms, the component of stress perpendicular to the fibers in that layer is now

transferred to the layer next to it, in the vicinity of the crack.

63
Q

®

v

®

Oo

• ....d.._ .... _.._ ¢_.

Figure 41. Cross-section of specimen # Static # 5.

64
® ® @

®

®

• 0°

......

® ®

Figure 42. Crossisection of specimen # 12/13B.

65
o

@

@

©

Figure 43. Cross-section of spec_en # !2/12A, cut 1.

66
©

®

+45

0

-45

90

90

-45

@ o

+45 @

Figure 44. Cross-section of specimen # 12/12A, cut 2.

67
• 0 °

©

90

-45

0

+45

Figure 45. Cross-section of specimen # L6, cut 1.

68
p 0 °

©

+45

0

--45

90

Figure 46. Cross-section of specimen # L6, cut 2.

69
+45

0

--45

90

Figure 47. Cross-section of specimen # 12/12C, cut 1.

70
0 °

Figure 48. Cross-section of specimen # 12/12C, cut 3.

+45

0

-45

90

90

-45-

0

+45

71
+45

90

90

-45

Figure 49. Cross-section of specimen #12/12C, cut 2.

0

+45

Figure 41 is a cross-sectional view of a statically loaded specimen that experienced a maximum

load of 1,780 N (400 lb). The cut on this specimen was made perpendicular to the 0 ° fibers in order to

demonstrate that the long dark areas running in the 0 ° direction consists of heavy matrix cracking in

the 0 ° ply. Note that between delamination zone #3 and matrix cracks #4 and #5, there is much matrix

damage in the lower 0 ° ply, which shows up as a dark stripe in the 0 ° direction on the x-ray signature.

For the remainder of the cross-sectioned specimens presented, the cross-sectional cut was made

parallel to the 0 ° fibers in order to check for fiber breakage in the critical 0 ° plies. Figure 42 demonstrates

72
thatmuchmatrixcrackingcanexistwith nofiberbreakage.Whenfiberbreakageis seen,asin figure43,

amatrix crackin oneof thetwo adjacentpliesisaccompanyingthe0° fiberbreaks.The0° fiber breaks

(crack#1)arealsoseento bein relativelycloseproximityto thebottom+45° split (crack #3), that is the

first and most visible form of damage noted. When the same specimen is sectioned at a point where the

main +45 ° and -45 ° cracks intersect, as shown in figure 44, the bottom 0 ° fiber breakage is closer to the

bottom +45 ° split. For more heavily damaged specimens, fiber breakage in the lower 0 ° ply is directly

above the +45 ° ply split as shown in figure 45.

Broken 0 ° fibers in the upper part of the laminate are always seen to be a part of a crack running

either from the ply above it (the top +45 ° ply) or below it (the first -45 ° ply) or both as shown in

figure 46. The crack in the upper +45 ° ply (crack #2) and the first 45 ° crack (crack #1) are joined by

broken fibers in the 0 ° ply between them. Figure 47 shows a feature that was seen in many of the

sectioned specimens. A crack will form from a ply adjacent to the upper 0 ° ply and transverse part of the

ply until the crack begins to grow parallel to the fibers, causing the 0 ° ply to be only partly fractured. This

crack

is highlighted by white arrows in the figure. Figure 49 shows a -45 ° ply crack that starts into the 0 ° ply

above it but only breaks fibers a short distance into the 0 ° ply when the crack is deflected parallel to the

0 ° fibers. Note the severe damage in all plies in this cross-sectional view.

Figure 49 demonstrated the difficulty in trying to identify broken fibers from the x-ray signatures.

From the cross-sectional view, bottom 0 ° fibers are clearly broken, yet between cracks #1 and #3 on the

x-ray signature, there is nothing obvious to indicate that fibers are broken (area #2). This was the case

for the vast majority of specimens.

5.5 Static Indentation Tests

Data from the static indentation tests aided in indicating the shape of the deflected surface as well

as showing the different strain fields associated with a low versus a high prestrain. The lower preloaded

specimen demonstrated more nonlinearity at the lower end of loading than the heavily preloaded specimen

as shown in figures 50 and 51. This is to be expected from the analysis in chapter 3. From the unloading

data it is seen that a hysteresis is associated with all of the specimens tested at the lower prestrain levels.

This indicates that there is either more damage being formed in these specimens or more compliance

in the preloading device at the lower prestrains. Results from the x-ray analysis indicates that more

damage was not formed, indicating compliance in the fixture at low levels of preload. Complete data

from the 10 specimens tested are given in appendix C (specimen #2 is not included since the dial indicator

was not set up properly).

73
_8°

4OO

350

300

250

200

150

IO0

5O

0

Deflection (mm)

0 1 2 3 4 5 6 7

_--""1 ""l""["'"l ""["'ll e'r,_,-_

-- O0 --

Nx=203 N/cm (116 Ib/in) •

.- •© -

= • --

=.

_-- • © -

- • = Loading • --

- • 0 :

- 0 = Unloading • --E

_-- • 0 --

• 0

- • Z

_-- • 0 -

? I ° I

,I ,,, ,,,, ,,, ,I,,, ,I,,,,

0.05 0.1 0.15 0.2 0.25

1,600

1,400

1,200

1,ooo

800 o._1

6OO

40O

20O

0

0.3

Deflection(in)

Figure 50. Load displacement from a statically loaded specimen at a low prestrain.

O

.J

400

350

300

250

20O

150

100

50

0

Deflection(mm)

0 1 2 3 4 5

_" '' '1' '' ' I' '' ' I' '' '_' '' ':1

- Nx=2,847 N/cm (1,625 Ib/in)

• = Loading

© = Unloading

- •

- e©

_-- e©

- •

- eO

eO

tD

00

O0

00

0 0.05 0.1 0.15

1,600

1,400

1,200

1,ooo

"1o

800 8,--I

6OO

Deflection(in)

Figure 51. Load displacement from a statically loaded specimen at a high prestrain.

74
From the in-plane strain measurements as determined from the gauges placed on the specimens,

as shown in table 4, it was clear that little strain was set up along the edges of the specimen due to transverse loading, especially on the highly prestrained specimens. Compressive strains due to transverse

loading were always found in the vicinity of the applied load on the top surface and the strains on the

opposite side (back face) would be tensile unless the preload was high, in which case they became negative. Plots of strain versus applied transverse load for the 10 specimens tested are given in appendix D.

Note that at regions near the clamped edges (up to 57 mm from a clamped edge for specimens #7 and #8),

both the top and bottom surface are put into tension due to the transverse load. At high preloads, the

bottom surface sometimes went into compression from the transverse load (static #5), but this appears

to be an isolated case.

5.5.1 Static Indentation Tests to Failure

Two specimens were prestrained and statically loaded until they could no longer hold a load.

Only load-deflection data were generated from these tests. The results are shown in figures 52 and 53.

w

O

-.1

600

500

400

300

200

100

Deflection(mm)

0 1 2 3 4 5 6 7

Nx=1,354N/cm(773Ib/in)

2,500

2,000

g

1,500

o,

1,000

500

O_ 0

0 0.05 0.1 0.15 0.2 0.25 0.3

Deflection(in)

Figure 52. Load displacement from a statically loaded specimen (loaded until failure).

75
Deflection(mm)

m

O

0 1 2 3 4 5 6

6O0

5OO

400

300

2O0

100

Nx=2,847 N/cm

(1,625 Ib/in)

O( 0

0 0.05 0.1 0.15 0.2 0.25

2,500

2,000

1,500

1,000

5OO

v

"o

¢¢1

0

..J

Deflection(in)

Figure 53. Load displacement from a statically loaded specimen (loaded until failure).

?6
6. DISCUSSION OF RESULTS

This chapter is a discussion of the experimental results obtained in the previous chapter, along with

results from the analytical analysis performed in chapter 3.

6.1 Instrumented Impact Testing

6.1.1 Maximum Load of Impact

For impact conditions in which the specimen did not sustain heavy damage, the maximum load

of impact increases linearly with increasing preload. This can be seen for all of the data at the 3.4 J

(2.5 ft-lb) impact energy level, data up to -5,000/ue prestrain for the 4.5 J (3.3 fl-lb) impact energy level

and up to -4,000 bte for the 6.0 J (4.4 ft-lb) impact energy level. More scatter exists for the higher two

impact energy levels due to the development of more severe damage. From equation (7) in chapter 3, it

was predicted that the maximum load of impact was expected to increase at a linear rate with increasing

prestrain for a given amount of impact energy.

6.1.2 Duration of Impact

The total time of impact is seen to decrease in a nonlinear fashion with increasing tensile prestrain.

As damage is formed in the specimen, the duration tends to be a little longer due to local "softening" of

the specimen. Thus, the nonlinearity is due mostly to damage. Note that the total duration of impact seems

independent of impact energy. This should be the case if the transverse load is approximately proportional

to the transverse deflection. For a spring-mass system (which can approximate the impact problem), the

total time of impact (contact of mass on spring) is given by:

t = j2 x (13)

where F is the force exerted by the laminate (spring), x is the displacement, and m is the impactor mass.

Assuming F 0cx (or F=kx, k=-"spring constant"), then equation (1) can be written as:

(14)

which is independent of deflection and therefore independent of impact energy since m and k are fixed

constants.

77
6.1.3Maximum Deflection

The maximum deflection of impact decreases in an almost linear fashion with increasing prestrain

for a given impact level. This linear relationship is not obvious from equation (7). This tends to hold true

of samples in which incurred damage caused departure from linearity in maximum load and duration data.

This is due to the "softening" which caused departure from linearity in the maximum load and duration

data, increasing the deflection, yet the load drops which cancels it out.

6.1.4 Maximum Absorbed Energy

These data are seen to decrease with increasing prestrain for the lower values of prestrain, then rise

with increasing prestrain for the rest of the data. It is predicted that as the preload increases, more energy

would be lost due to the impact event since higher prestrains tended to cause more damage in the specimen. An explanation for this can be gained from the results of the static indentation tests in which more

hysterises were formed at the lower prestrains, due to fixture compliance.

Thus at the low preloads, fixture compliance attributes to energy lost in the system, but as the

preload increases, this compliance tends to become less severe. At higher prestrain levels, damage begins

to form in the specimens and the energy lost in the system begins to rise again.

6.1.5 Load-Deflection Curves

For higher preloads, the load-displacement plots become more linear up until the point of damage.

At this point on the curve, the load begins to fluctuate at a high frequency as matrix cracks, delaminations,

and possibly fiber breakage occurs.

A large amount of hysteresis is observed for all of the impacted specimens regardless of transverse load-preload combinations. For specimens that contain significant damage, this result is expected,

but for s oecimens with little or no damage, little hysterisis would be expected. This implies that energy

is being lost to the system in modes other than specimen damage. The particular impact tester used in this

study does not have linear bearings or other such friction-reducing devices between the falling weight

crosshead and the guideposts. This has always been suspected of causing the system to lose a fair amount

of energy in the reaction between the falling crosshead and the guideposts during the impact event. This

suspicion has never formally been conf'Lrmed and documented.

6.2 Visual Examination

The first form of noticeable damage that formed was splitting of the matrix along the fibers in the

bottom +45 ° ply. This is to be expected from the finite element results, presented in table 3, which show

these plies develop the highest stresses in the second principal material direction, which are almost twice

as high as the breaking stress for a lamina.

The next sign of visual damage was usually associated with contact stresses on the impacted face

of the specimen. A small dent would become noticeable and as either the prestrain or the transverse impact

load increased, matrix cracking and fiber breakage was noticed inside the indentation.

78
At relativelyhighimpactlevels,brokenfiberswouldbeginto protrudefrom thebackfaceof the

impactedspecimenalongacrackin thebottom+45° ply.

This type of damage progression is typical of fibrous laminates that are allowed to undergo large

deflections during impact, thus setting up high tensile strains on the back surface of the impacted laminate.

6.3 X-ray Inspection

The most notable feature of the x-ray signatures for both the statically loaded and impacted specimens is that the damage tended to form along the direction of the bottom +45 ° ply. Delaminations, emanating from a matrix crack in this ply, were always noted on the x-ray results (unless no damage at all

formed). This feature is shown schematically in figure 54. The mechanics thought to be behind these

delaminations are given in the next section.

6.3.1 Initial Damage Formed

The initial damage consisted of a matrix crack in the bottom +45" ply with delaminations forming

as shown in figure 54. On either side of the impact site, the delaminations consisted of two zones, one

usually smaller than the other. As a matrix crack develops in the bottom +45 ° ply, a free edge is created

and the stresses must redistribute themselves so that equilibrium is maintained. This "free edge effect" is

the same mechanism outlined by Pipes and Pagano. 126Figure 55 shows schematically how the geometry

of the problem influences the delamination growth, mainly due to the transverse interlaminar stress, O Z,

that is set up because of the free edge (called face 1 in the figure).

ImpactSite

00- Direction

Figure 54. Schematic of bottom ply splitting and associated delaminations.

79
Txy

Bottomof Laminate

"¢xy

--Face 3

Oz

Face1

xzy y-Direction

Face4

x-Direction

Figure 55. Stresses associated with back face matrix crack.

Face 3 of the ply is carrying the x-direction stresses, just as it was before the crack was formed.

The stress distribution is such that the closer a point on the face is to the impact point (I), the higher the

x-direction and y-direction (face 2) stresses are. This is known from the finite element results. The

moment about a line through point A running in the y-direction must be zero for equilibrium conditions

to be satisfied. The face 3 stress, crx, creates a moment about this line. To balance this moment, stresses in

the z-direction, az, are set up along the crack boundary. Since Orx is largest at the point closest to the

impact point, I, tYz is also largest at this point. As the crack moves farther away from the line through A

(i.e., as y increases), the z-direction stress rapidly decays since O'x gets smaller in this direction and also the

"moment arm" is getting longer, so to maintain a given moment, the z-direction force gets smaller. The

opposite should hold true as the crack gets closer to face 3. The "moment arm" becomes smaller, so the

stresses need to be larger to maintain a constant moment. However, Ox gets smaller, so the moment

needed for equilibrium also gets smaller at points further away from the impact point I.

The above also holds true for tyy which also creates a moment about a line through point B

running along the x-direction. This will tend to cause symmetry of tYz along the crack for equal values

of (7x and try. However, since try is generally smaller than crx, especially for heavily preloaded specimens,

the overall effect is to have slightly larger tTz values near the face 4 end of the crack than at the face 2 end.

80
Whenface1is created,theshearstressesalsoredistributethemselvesto maintainequilibrium.

Theshearstresseson faces2 and3, "r_-y, are no longer balanced by shear stresses on face 1, since it is free,

or on face 4, since it can be considered infinitely small. New interlaminar shear stresses lrzy and "rzx are

created to balance the moments due to the shear stresses. Since z:ty is largest nearest the point of impact

(as it was for O'x and try), "lrzx and "t'zy will also be largest nearest the impact point, I.

6.4 Cross-Sectional Examination

Large matrix cracks and delaminations could easily be detected by x-ray analysis but fiber

breakage was difficult to detect except by cross-sectional examination. Cross-sectional examination

showed that fiber breakage could occur in a laminate at a point that was not obvious from the x-ray

signature. A large amount of matrix damage was always associated with neighboring plies of a ply with

broken fibers. These areas of high crack density tended to cause the dye penetrant to "pool-up" and create

dark areas on the x-ray signatures which would obscure the broken fiber damage.

The cross-sectional analysis also revealed that fiber breakage in the bottom 0 ° ply differed from

that in the upper 0 ° ply. The fiber breakage in the bottom 0 ° ply always ran completely through the ply

but fiber breakage in the upper 0 ° ply did not. In many instances the fiber fracture would only travel onehalf or one-third the thickness of the lamina and then deflect parallel to the fibers, running through the

matrix in the upper 0 ° ply. The bottom 0 ° ply fiber breakage was always associated with large delaminations between this ply and the bottom +45 ° ply. The fiber breakage in the upper 0 ° ply was characterized

by the propagation of the crack containing the fiber breakage into one, or if the ply was completely fractured, both adjacent plies through the matrix material in these adjacent plies. Delamination was sometimes,

but not necessarily, associated with these cracks.

The failure modes of the two 0 ° plies can be expected to differ, based on the maximum ply stress

values given in table 3. The upper 0 ° plies are subjected to small stress values that are not even within an

order of magnitude of the fiber-breaking stress for these plies. Thus, the fiber breakage must come from

another stress mechanism setup, other than that by preload or membrane and bending stresses arising

from the transverse load. Since dents were observed on the impacted surface of many of these laminates,

it is assumed that contact stresses are responsible for fiber breakage in the upper 0 ° ply. The stresses

predicted for the lower 0 ° plies are much larger (although they never exceed one-half the breaking stress)

and with added stress from cracked plies transferring stress to the main load-bearing 0 ° fibers, it is likely

that these fibers simply break in tension (causing the complete fracture of the ply).

6.5 Static Indentation Tests

The static indentation tests showed that a tensile preload increases the flexural stiffness of the

laminate, especially at lower levels of transverse load, at which point the induced membrane stresses due

to large deflection are not as large.

81
6.5.1 Comparison With Energy Balance and Finite Element Models

Figure 56 shows load-deflection curves for specimens prestrained at 4,000/.re as predicted

by equation (9) and the finite element analysis. Data from a static indentation test are included for

comparison. It is clear that the finite element method comes closer to predicting the actual loaddisplacement behavior of the laminate.

Figure 57 shows prestrain-deflection data as calculated by the two models, along with data

from the static indentation tests at a constant transverse load of 223 N (50 lb). The finite element method

matches up extremely well with the experimental data. The curve as predicted by the energy balance model

(equation (9)), greatly underestimates the effect of prestrain on deflection and also predicts a linear relationship which does not agree with experimental data.

Displacement(mm)

._1

0

500 ,

400 --

300 --

200 --

100 --

0 ,J

0

1 2 3 4 5

'''1''''1''''1''''1''''1'

•©

FiniteElement

Static • ©

Indentation_, • C) []

4k

• © []

e(3

1[3

(3

0 []

(9 []

,,, l,,,, I,LI

0.05 0.1

[]

[]

_Equation

If,,,, I,

0.15 0.2

6

I I I [ -

[]-- 2,000

[]

-- 1,500

-- 1,000

- 500

(g)

I I I 0

0.25

0

Displacement(in)

Figure 56. Load-deflection data comparison between equation (9), the finite element method, and static

indentation data for a prestrain of 4,000/_E.

82
0.12

0.1

O.O8

o

C_ 0.06

0.04

0.02 I

0

tE]S Finite Element

m

3.0

- 1.0

D i

1,000 2,000 3,000 4,000 5,000 6,000 7,000

2.5

2.0 _"E

v

1.5

Prestrain(Microstrain)

Figure 57. Prestrain-deflection data for a constant transverse load of 223 N (50 lb).

The data from the in-plane strain measurements agreed qualitatively with the finite element results

and showed that at high preloads, little deflection or strain is induced into the laminate near the free edges

due to the transverse load.

A result that was found experimentally that was not predicted was the compressive strains that

appeared on the top surface of the laminates due to the transverse load during the static indentation tests.

From table 3 it was predicted that as the transverse load increased, the stress (or strain) on the top surface

would become more positive due to the large membrane stresses set up in the laminate. One cause for this

could be the previously mentioned contact stresses which are not accounted for in the model.

6.5.2 Comparisons of Impact and Static Indentation Curves

From all accounts, the impacts in this study were of a quasi-static nature with no contributions

from material inertia or material vibrations. The conditions for a quasi-static impact event, low target

mass, and a large impactor mass were certainly met in this study. Thus, it is expected that the static

indentation data should be very similar to the impact data.

Figures 58-60 show load-deflection data for three different levels of preload for both the static

indentation and impact tests. These data are seen to match fairly well, especially at the lower end

of transverse deflection and at lower preloads. The static indentation data show a larger load for a given

deflection at the higher end of deflection. Note that the data from the impact tests show a small drop

in stiffness along the load-deflection curve whereas this drop is not seen in the data from the static

indentation tests. In addition, figure 59, which has static indentation data for a specimen with heavy

damage, shows that the static indentation specimens can sustain a slightly higher load before the fLrst

sudden drop in load is observed. In figure 60, no damage has occurred on the static indentation specimen

at 1,780 N (400 lb), but obvious damage has occurred in the impacted specimen.

83
Deflection(mm)

0 1 2 3 4 5 6 7 8 9 10

700 = , i i i l i i i i

-- 3,000

600 -- Nx-228 N/cm (1301b/in)

500 • = Static Indentation Data

_" 400- - 1,780

300 .J

r J

100° _ 500

0 0.1 0.2 0.3 0.4

Deflection(in)

Figure 58. Comparison of load-deflection curves for a static indentation specimen

and an impact specimen both preloaded at -228 N/cm (130 ]b/in).

8

..d

0

700

60O

5OO

400

300

20O

100

0

Deflection(mm)

1 2 3 4 5 6 7

I I I I I I I

_ Nx-1,300 N/cm (742 Ib/in)

• = Static IndentationData e=

_

8 9

I I

I I • I I 1 1

0 0.1 0.2 0.3 0.4

Deflection(in)

10

I

-- 3,000

- 1,780E

-- 500

Figure 59. Comparison of load-deflection curves for a static indentation specimen

and an impact specimen both preloaded at -1,300 N/cm (742 lb/in).

84
0

700

600 -

50o_-

Deflection(turn)

1 2 3 4 5 6 7 8 9 10

I I I ! I I I I I I

Nx-2,600N/cm(1,500Ib/in)

• = StaticIndentationData

3,000

- ,ool- --#j 1 1,78°

...?'"l /

F _..j / J

lOI _ ] 500

0 0.1 0.2 0.3 0.4

Deflection(in)

Figure 60. Comparison of load-deflection curves for a static indentation specimen

and an impact specimen both pre]oaded at -2,600 N/cm (1,500 ]b/in).

It was noticed during the static indentation tests that when a specimen was at a load that could

cause damage formation, damage did not always immediately occur in the specimen. As the load was held

constant and a displacement reading was being taken, the specimen would sometimes produce a loud pop

which indicated the formation of damage after a time interval at a given load.

From x-ray results on specimens that experienced similar maximum transverse load-preload

combinations, some impacted specimens showed more damage than the specimens that were statically

loaded, although the differences were not great. Most of these specimens contained very similar x-ray

signatures.

Thus some subtle differences can be noted for impacted and statically loaded specimens in this

study.

85
7. CONCLUSIONS AND RECOMMENDATIONS

7.1 Overview

This chapter contains conclusions that can be drawn from this study and recommendations

for future studies of relevance.

7.2 Conclusions

The following are the major conclusions of this study:

For all other impact parameters which are the same, a tensile prestress tends to increase the maximum

force of impact and the amount of damage formed. A decrease is seen in the duration of impact and in

the maximum transverse deflection due to impact.

The types of damage formed are matrix cracking (splitting), delaminations, and fiber breakage. The

first sign of damage is matrix splitting in the bottom +45 ° ply. Large delaminations emanate from

the matrix split and tend to run between this split to a split in a neighboring ply. Smaller delaminations

form directly under the impact zone due to the high contact stresses set up in this area.

Large deflection plate theory must be taken into account for analysis in which the plate deflects more

than one-half its thickness. The induced midplane stretching from the large deflections rapidly

dominate the load/displacement response of the plate.

Finite element techniques do a good job of predicting the elastic response of a transversely loaded

plate, but the introduction 0fdamage greatly complicates the problem_ :=

• X-ray inspection can easily detect matrix cracks and delaminations; however, fiber breakage

is difficult to detect.

Cross-sectional examination revealed that fiber breakage did occur, even though the finite element

analysis predicted that the maximum stresses along the fibers in a ply would not even be one-half the

breaking strength of the ply.

• Fiber breakage in the lower 0 ° ply was dominated by tensile stresses whereas fiber breakage

in the upper 0 ° ply was dominated by contact stresses and showed a complex failure mode.

Statically loaded specimens and impact loaded specimens show the same general behavior

if the target mass is small and the impactor mass high, but subtle differences do exist, especially at

high tensile prestrain values.

86
7.3 Recommendations

The following recommendations are for related programs that may be conducted in the future

concerning impact damage of composite laminates:

• The gray area between "quasi-static" and dynamic cases must be accounted for since many impacts

may fall into this regime.

• Finite element analysis must be handled with care once damage is incorporated into the model.

• The inclusion of the x-displacement values, u, in the energy balance solution should be incorporated

for more accuracy involving the influence of the preload.

• The effects of compressive preloads should be studied, since by the arguments given in section 6.3.1,

if a crack forms, crz should be compressive and help to suppress delamination growth.

• Nondestructive evaluation (NDE) techniques need to be examined extensively since, in the field, this

will ultimately be the tool that determines the amount of damage present due to an impact event.

87
8. SUMMARY

Tensile preloads tend to increase the maximum force of impact and decrease the duration

of the impact event and the maximum transverse deflection for low-velocity impacts. Higher tensile

preloads also tend to cause more damage to form in the specimen due to the higher impact forces and

normal tensile stresses, oz, that are set up along a back face split that helps to form a delamination.

Finite element analysis can predict the material response well but once damage is induced into

the specimen, the problem becomes much more complex.

Fiber breakage is difficult to detect from x-ray signatures of the damage zone due to the high

matrix crack density associated with such a specimen, causing the dye penetrant to "pool-up" and hide

the fiber breaks.

Contact stresses are very important, even when the specimen is flexible and allowed to deform

many times its thickness. Static indentation and "quasi-static" impact tests show the same basic behavior

and damage formation, but subtle differences do exists.

88
APPENDIX A--LOAD-DEFLECTION DATA FROM INSTRUMENTED IMPACT TESTS

89
0

700

60O

5OO

400

300

2OO

IO0

0

Deflection(mm)

1 2 3 4 5 6 7 8 9

I I I I I I ! I I

I 1 1

0 0.1

Specimen#L1

I.E. = 4.5 J (3.3 ft-lb)

Nx =3,185 N/cm (1,820 Ib/in)

I I I I

0.2 0.3

Deflection(in)

10

- 3,000

- 1,780

- 500

0.4

Z

o,

0

700

6O0

5O0

400

o 300

2O0

100

Deflection(mm)

1 2 3 4 5 6 7 8 9 10

I I I I I I I I I I

w

m

Specimen#L2

I.E. = 4.5J (3.3 ft-lb)

Nx = 1,985 N/cm (1,134 Ib/in)

I 1 I I I I I

0.1 0.2 0.3

Deflection(in)

3,000

1,780

50O

0.4

v

o,

90
v

8

.--I

0

700

600 -

50O

400

300

200

100

0

Deflection(ram)

1 2 3 4 5 6 7 8

I I I I I I I I

Specimen#L3

I.E.= 4.5J (3.3 ft-lb)

Nx = 278 N/cm (159 lb/in)

I I I I I

0.1 0.2

Deflection(in)

I I

0.3

9 10

I I

3,000

1,780

- 5OO

0.4

...9°

,£_

v

,...3

0 1

700

600 -

500 -

400 -

300 -

200 -

100

0

!

0

Deflection(mm)

2 3 4 5 6 7

I I I I I I

8 9 10

I I I

I.E.= 4.5J (3.3 ft-lb)

Nx = 513 N/cm (293 Ib/in)

I I I

0.1 0.2

Deflection(in)

Specimen#L4

I

0.3

3,000

1,780

500

0.4

8

,-J

91
Deflection (mm)

0 1 2 3 4 5 6 7 8 9

700 i I i i , , l i i

600 - Specimen#L5

500 -

400-

e.s

"_ 300

._J

2OO

IO0

0 I.E.= 4.5 J (3.3 ft-lb)

Nx= 4,548 N/cm (2,599 Ib/in)

I I I I I I

0.1 0.2 0.3

Deflection(in)

10

- 3,000

1,780

5OO

0.4

z

¢'0

3

J_

v

o

0

700

600

500

40O

300

200

100

0

Deflection (mm)

1 2 3 4 5 6 7 8

I I I I I I I I

Specimen #L6

-- I.E.=4.5J(3.3ft-lb)

Nx= 3,420N/cm(1,954II_n)

I I I I I 1

0.1 0.2 0.3

Deflection(in)

10

I

3,000

1,780

500

0.4

92
.¢3

_9°

0

,o: t

5oor

4OO

3OO

2OO

IO0

0

Deflection(mm)

1 2 3 4 5 6 7 8 9

I I I I I I I I I

Specimen#11/13A

_.5 J (3.3 ft-lb)

Nx = 1,493 N/cm (853 Ib/in)

I I I I I I

0 0.1 0.2 0.3

Deflection(in)

10

I

3,000

1,780

- 500

0.4

-9

v

8

,.J

0

400--

300

200

100

0

Deflection (mm)

1 2 3 4 5 6 7 8 9 10

I I I I I I l I I i i

Specimen #11/13B

I.E. = 4.5 J (3.3 ft-lb)

Nx = 1,204 N/cm (688 Ib/in)

I I I I I I

0.1 0.2 0.3

Deflection(in)

3,000

1,780

Z

_9

5OO

0.4

93
0

Deflection (mm)

0 1 2 3 4 5 6 7 8 9

700 i i I l i , , , i

600 Specimen #11/14A

500

400

300 ==_

20O

100

0 I.E. = 4.5 J (3.3 ft-lb)

Nx = 4,898 N/cm (2,7991b/in)

I I I I I I

0 .1 .2 .3

Deflection(in)

10

I

3000

1780

500

.4

A

o,

m

v

'1o

0

..J

0

700

600 5OOt

40O

300

200

100

Deflection(mm)

1 2 3 4 5 6 7 8 9 10

I I I I I I I I I !

Specimen#I 1/14B

_ . . = 4.5 J (3.3 ft-lb)

Nx = 2,585 N/cm (1,4771b/in)

I I 1 I I I I

0.1 0.2 0.3

Deflection(in)

3,000

1,780

- 5O0

0.4

Z

v

0

.-J

94
..9°

0

700

600 500 t

4O0

30O

20O

100

Deflection(mm)

1 2 3 4 5 6 7 8 9

I I I I I I I I I

Specimen#11/25A

_ . . = 4.5 J (3.3 ft-lb)

Nx = 2,623 N/cm (2,3521b/in)

1 1 I I I I I

0.1 0.2 0.3

Deflection (in)

10

'1 3,000

- 1,780

- 500

0.4

Z

v

0

-.I

700

6OO

5OO

400

'_0 300

._1

20O

100

Deflection(ram)

1 2 3 4 5 6 7 8 9

I I I I I I I I I

Specimen#11/27A

m

_E. = 4.5 J (3.3 ft-lb)

Nx = 4,116 N/cm (2,3521b/in)

I I I I I I I

0.1 0.2 0.3

Deflection(in)

10

I

3,000

1,780

50O

0.4

A

Z

..J

95
0

....I

0

700

600 t

5O0

400

300

200

100

0

Deflection(mm)

1 2 3 4 5 6 7 8 9

I i I I I I I I I

Specimen#11/28A

....... 0J (4.4ft-lb)

Nx = 3,309N/cm (1,8911bAn)

1 I I I I I 1

0.1 0.2 0.3

Deflection(in)

10

I

- 3,000

- 1,780

- 500

0.4

,..J

J_m

o

0

700

600 t

500

400

300

200

100

Deflection(ram)

1 2 3 4 5 6 7 8 9

I I I I I I I I I

Specimen#11/28B

I.E.= 6.0J (4.4ft-lb)

Nx = 306 N/cm (175Ib/in)

I I I I I I I

0.1 0.2 0.3

Deflection(in)

10

i

3,000

1,780

500

0.4

z.

96
8

..,,t

0

:t

5OO?

4OO

3OO

2OO

tO0

0

1 2 3

! ! I

Deflection(mm)

4 5 6 7 8 9 10

I I I I I I I

3,000

Speclmen #12/12A

I.E. = 6.0 J (4.4 ft-lb)

Nx = 4,408 N/cm (2,519 Ib/in)

I I I I I I

0 0.1 0.2 0.3 0.4

1,780

50O

Deflection (in)

,--I

0

ml

0

700 /

OOoOoE

400

300

200

100

0

Deflection(ram)

1 2 3 4 5 6 7 8 9 10

I I I I I I I I I L_

Specimen#12/12B

..... J (4.4 ft-lb)

Nx = 1,312 N/cm (749 Ib/in)

I I I I I I I

0 0.1 0.2 0.3

Oeflection (in)

3,000

- 1,780._.

z

- 500

0.4

9"7
Deflection (ram)

0 1 2 3 4 5 6 7 8 9

700 , I i , e t J j j

6O0

5O0

_. 400

_--_ 300

,..J

2OO

100

Speclmen#12/12C

I.E.= 6.0 J (4.4 ft-lb)

Nx = 2,555 N/cm (1,460/bAn)

I I I I ] I I

0.1 0.2 0.3

Deflection (in)

10

J1 3,000

1,780

5OO

0.4

Z

._1

0 1

700 t s

600

5O0

Deflection (mm)

2 3 4 5 6 7 8 9

I I I I I I I I

_ooF / /

'oi_-_ _/

L i--- I

Specimen#12/12D

I,E.= 6.0J (4.4 ff-lb)

Nx = 1,876 N/cm (1,072 Ib/in)

I I,, I I .... I

0.1 0.2 0.3

Deflection(in)

10

1

3,000

1,780

5OO

0.4

0

,.-I

98
J_

]

0

700

600 -

500

400

30O

200

100

0

Deflection(mm)

1 2 3 4 5 6 7 8 9

I I I I I 1 I I I

Specimen#12/13A

I.E.= 6.0 J (4.4 ft-lb)

Nx= 980 N/cm (560 Ib/in)

1 I I I I I I

0.1 0.2 0.3

Deflection (in)

10

I1 3,000

0.4

8

0

7OO

600 t

50O

4O0

300

200

100

0

Deflection(mm)

1 2 3 4 5 6 7 8 9

I I I I I I I I I

Specimen#12/13B

IE 60J (44ft Ib)

Nx = 390 N/cm (223 Ib/in)

I I I I I 1 I

0.1 0.2 0.3

Deflection(in)

10

=1 3,000

- 1,780

- 500

0.4

0

,_1

99
0

7O0

BO0_

500_

400"--

300

200

100

0

Deflection(ram)

1 2 3 4 5 6 7 8 9

I I l l I I I I I

Specimen#12/14B

__----_'_J I.E. = 3.4 J (2.5 ft-lb)

Nx = 4286 N/cm (2,44g Ib/in)

I I I I I I I

0 0.1 0.2 0.3

Deflection (in)

10

I

- 3,000

1,780

5OO

0.4

0

,--I

0

..J

0

700

600

50O

400

300

200

100

0

Deflection(mm)

1 2 3 4 5 6 7 8 g 10

I I I I I I I I I I

Specimen#12/140

12.5.-,bl

Nx = 607 N/cm (347 Ib/in)

I - I I I I I

0 0.1 0.2 0.3

Deflection(in)

3,0O0

1,780 .._.

Z

500

0.4

IO0
m

-.J

0

700;

600 -

500 -

400 -

30O

20O

100

0

Deflection(ram)

1 2 3 4 5 6 7 8 9

t I I I I I I I I

Specimen#12/14D

.. .. = 3.4 J (2.5 ft-lb)

Nx = 2,083 N/cm (1,190 Ib/in)

I I I 1 I I

0.1 0.2 0,3

Deflection(in)

10

I

- 3,000

1,780

5O0

0.4

Z

0

,--I

i

.--I

700

6O0

500

400

30O

200

100

0

Deflection(mm)

1 2 3 4 5 6 7 8 9 10

I I I I I I I I I 1

Specimen #12/14E

I.E, = 3.4 J (2.5 ft-lb)

Nx = 3,081 N/cm (1,760 Ib/in)

I [ I I I I I,

0 0.1 0.2 0.3 0.4

3,000

1,780 A

Z

-.I

5O0

Deflection (in)

101
0

7OO

600 -

500 -

400 -

300

200

100

0

Deflection(ram)

1 2 3 4 5 6 7 8 9

I I I I I ! I I I

Specimen #12/17A

Nx = 245 N/cm (140 Ib/in)

] I I 1 1 I

0 0.1 0.2 0.3

Deflection(in)

10

I

- 3,000

1,780

5OO

0.4

..J

J_

.9

0

700,

600

50O

4O0

30O

2OO

100

0

Deflection(ram)

1 2 3 4 5 6 7 8 9 10

I I I I I I I I I I

Specimen#12/17B

I.E. = 3.4 J (2.5 ft-lb)

Nx = 2,664 N/cm (1,522 Ib/in)

I I I I I I

0 0.1 0.2 0.3

Deflection (in)

3,000

1,780

S

500

0.4

102
0

0

7OO

6OO

5OO

40O

3OO

2OO

IO0

0

Deflection(mm)

1 2 3 4 5 6 7 8 9 10

I I I I I I I I I I

m

Specimen #12/170

- 500

- I.E.= 3.4 J (2.5 ft-lb)

Nx = 3,520 N/cm (2,009 Ib/in)

I I I I 1 I I

0.1 0.2 0.3 0.4

Deflection (in)

3,000

1,780

"0

0

,_1

Deflection(mm)

0 1 2 3 4 5 6 7 8 9

700 i i , i i i i i i

600

5OO

400

300

2O0

100

0

Specimen#12/17D

I.E. = 3.4 J (2.5 ft-lb)

Nx = 1,186 N/cm (678 lb/in)

I I I I i I I

0.1 0.2 0.3

10

I

- 3,000

- 1,780

- 500

Deflection(in)

0.4

Z

103
APPENDIX BmX-RAY SIGNATURES OF STATIC INDENTATION

AND IMPACTED SPECIMENS

104
Staticindentationspecimen# 1.

.............................................................. _::_::::_:_::::::_:i:i|_i_i iii_iii_i ̧i_? _

Static indentation specimen # 3.

105
F

Static indentation specimen # 4.

Static indentation specimen # 4.

106
Static indentation specimen # 7.

Static indentation specimen # 9.

107
Staticindentationspecimen# 10.

108
0 1

I I

2 3 4 5 6 7

Illlll

0 5 .10 15 20

tltillilttltlttillt ri_

Specimen # L1.

109
0 1 2 3 4 5 6 7

I I I I I I 1 I

cm

0

I l l l

2 _ ̧ ' r

4 8

I I I I I

mm

Specimen # L2

110
0 1 2 3 4 5 6 7

10

rrl'n

Specimen # L3

111
0 1 2 3 4 5 6 7

I I t I I 1 I I

cm

0 5 10

I I I I I I II I I I I I

mm

Specimen # L4\

112
0 1 2 3 4 5 6 7

I I I I I I I I

cm

0

I

5

III

10 15 20

II IIIIIII11

mm

Specimen # L5

113
0 1 2 3 4 5 6 7

I1111111

0 5 10

I I I t_l I i I I I I

mm

Specimen # L6

114
0 1 2 3 4 5 6 7

I I I I I I I I

cm

o 5 10

I I1 III II

mm

Specimen # 11/13A

115
0 1 2 3 4 5 6 7

11111111

0 5

II

mm

Specimen # 11/13B

10

I

116
0 1 2 3 4 5 6 7

IIIIIIII

0 5 10

In

mm

Specimen # 11/14A

15

li I!

20

i

117
0 1 2 3 4 5 6 7

I I I I I I I I

cm

• _ ,,1" •

r ._p

r

I I

mm

Specimen # 11/14B

118
0 1 2 3 4 5 6 7

I I 1 I I I I I

cm

i

0 5 10 15 20

II I II III II II II I I II I II

mm

Specimen # 11/25A

119
0 1 2 3 4 5 6 7

I I I I I I I 1

cm

0

I IIII

5

i I11

10 15

IJlilll

mm

Specimen # 11/27A

20

lllJl

120
0 1 2 3 4 5 6 7

I I I I I I I I

cm

ITITI

Specimen # 11/28A

121
.:," J!,:i-_Z_-:-,-"i

.' ;,d,._':_ _ _ _:'. " " " . _.i_,: • _..: . _.. - i

0 1 2 3 4 5 6 7

I I I I I I I I

CITI

0

1 I I I

5

I I I

film

lO

I I I I I I I

Specimen # 11/28B

122
0

1

10

mm

Specimen # 12/12A

20

III

123
0 1 2 3 4 5 6 7

I I I I I I I I

cm

0

II

5

IIII

10

IIIII

mm

15

I1[I

Specimen # 12/12B

124
0 1 2 3 4 5 6 7

III11111

0 5 10 15 20

III11 IIIII IIIII II I I11

mm

Specimen # 12/12C

125
0 1

I I

2 3 4 5 6 7

II1111

o 5

IIIIliil

10

IIIII

mm

15 20

Ji Illil

Specimen # 12/12D

126
0 1 2 3 4 5 6 7

IIIIIIII

5 10

III

mm

15 20

I IJl I

Specimen # 12/14A

127
0 1 2 3 4 5 6 7

i I 1 I l I l I

cm

0 5 10

III I II I

mm

Specimen # 12/14B

15

I

128
0 1 2 3 4 5 6 7

IIIIII11

4,

0 5

I I I I I I I

mm

10

I

Specimen # 12/14C

129
0 1 2 3 4 5 6 7

I I I I I I I I

cm

0 5

t I I

I'T'R3

10

I

Specimen # 12/14D

130
0 1 2 3 4 5 6 7

I I I I I I 1 I

cm

Specimen 12/14E

131
0 1 2 3 4 5 6 7

I I I I I I I I

cm

0 5

I I I I I I

r'r'lrrl

Specimen # 12/17A

132
0 1 2 3 4 5 6 7

I I I I I I I 1

cm

0 5 10

I I I I I I I I I I I 1 I I I

mm

15

I

Specimen # 12/17B

133
0 1 2 3 4 5 6 7

I I I I I I I I

crn

0 5 10 15

III II III I1111 111

n'lm

Specimen # 12/17C

134
5 10

I III I

n'lm

Specimen# 12/13B

15

135
0 1 2 3 4 5 6 7

I I I I I I I 1

cm

I I I

5

I TTT

n'lm

10

I

15

Specimen # 12/13A

136
0 1 2 3 4 5 6 7

IIIIIII1

0 5 10 15

III II 111 IIIII II1

rrl'n

Specimen # 12/17D

137
APPENDIX CmLOAD-DISPLACEMENT CURVES FROM STATIC INDENTATION

TESTS

138
30O

250

"_ 200

"o

¢¢1

0

.._1

150

100

50

Deflection(ram)

0 0.5 1 1.5 2 2.5 3 3.5

-""1""1""1""1'"'1'"'1'"'t

- Static Indentation#1

_ Nx=1,290 Ib/in

- • = Loading

- © = Unloading KZ)

-- (3

- ([3

-- K}

-- C

-- ¢}

- ¢3

5,Q, I,, ,I,, ,I ,, ,I ,, ,I

0 0.02 0.04 0.06 0.08 0.1

1,200

¢}

([_ - 1,000

(3

- 800

- 600

- 400

- 200

,,,1,,, 0

0.12 0.14

Deflection(in)

¢1=

0

-J

Deflection(mm)

0 1 2 3 4 5 6 7

400 " " I''"1 '"' I'"'1''"1'"'1"" I

350

300

250

'10

0

-- 200

150

100

5O

-- Static Indentation#3

- Nx=127 Ib/in

- £)

_ •

- • = Loading ([_

- 0 = Unloading •

- •

-

- •

-- @0

- •

-- 0

o_,I, 1 ,,,I,jll ii,,l,,,,I,,,

0 0.05 0.1 0.15 0.2 0.25

Deflection(in)

II--

1,600

-- 1,400

-- 1,200

-- 1,000 "-"

-- 800

-- 600

-- 400

-- 200

i- 0

0.3

139
Deflection(mm)

0 1 2 3 4 5

'°°L,,, ,,,

Static Indentation #4 2,000

400p Nx=1,550 ,b/in (IIgQ (130([_Q 1

_" I- • = Loading 1,500

-_ 300 0 = Unloading "_

3 _9o

• 1,000

200___-- • • 000 00 i

5OO

100 F • •

Did Not Clamp

oek_=,•,, I = = = j l, = =, I,,, ,-I o

0 0.05 0.1 0.15 0.2

Deflection(in)

.--I

0

5OO

400

300

200

100

Deflection (mm)

1 2 3 4 5

'''' I'''' I'''' I'''' I '' ''

Static Indentation#5

Nx=1,625Ib/in

m

• = Loading

0 = Unloading

(1(3

_,,I,,,,I,

0 0.05 0.1

(1(3

I

I I I I I I I I

0.15 0.2

Deflection(in)

2,000

1,500

1,000

5O0

E

3

140
v

o,

0

400

350

300

250

200

150

100

5O

0

0

Deflection(mm)

1 2 3 4 5 6 7

-' "' I"" I" "1 ""1' " '1' " '1' ' '_L"

Static Indentation#6 O0 --

Nx:l 16 tb/in

B

• = Loading

- 0 = Unloading

.

00

•

• 0

•

• 0

0

• 0

• 0

0.05 0.1 0.15 0.2 0.25 0.3

Deflection(in)

1,600

1,400

1,200

1,000

800

6OO

5OO

2OO

0

....I

5OO

400

300

2O0

100

Deflection(mm)

0 1 2 3 4 5 6 7 8

-""1'"'1'"'1'"'1""1 ""1'"'1'"'1'"

StaticIndentation#7

- Nx=107 tb/in

B

- • = Loading

B 0 = Unloading

m

0_

0

00

• 0

•

0

• 0

• 0

0

-, ,=, I,e, ,, I,,,, I ,_91,, ,, I,,,, I,,,,

2,000

1,500

1,000

500

0.05 0.1 0.15 0.2 0.25 0.3

0

0.35

Deflection(in)

0

-J

141
Deflection(mm)

0 I 2 3 4 5

5OO

-"" I"" I"" J"" I"i'1

Static Indentation #8 •

400 Nx=1,476 Ib/in

• = Loading _([_)

300 0 = Unloading

-- 200 (E)Q

I;3 °

100

,,,,

0 0.05 0.1 0.15 0.2

2,000

1,500

1,000 "_

5O0

0

Deflection(in)

Deflection(mm)

5O00

400

300

.._1

200

100

1 2 3 4 5

'''' I''' '1'' ''1''' '1'' '"

• --

StaticIndentation #9 •

- Nx=1,504 Ib/in • ©

- •

•©

• = Loading •

_ C) = Unloading • C)

•©

•

- • ©

•©

o_- ''' '-'' ' '

0 0.05 0.1 0.15

500

0

0.2

Deflection(in)

2,000

1,500

1,000 _= .9°

142
Deflection(mm)

5O00

400

300

J_

"o

0

._.1

200

100

1 2 3 4 5 6 7

-"" I"" I"" I""1"" I'"' I"" I"

- StaticIndentation#10 •

- Nx=125Ib/in •

- •

- • = Loading • ©_

- © = Unloading •

-- • ©

- •

- • ©

-- • --

-- • ©

- •

" • ©

- • _

-- • ©

- •

- • © =.

o_:_.._L-_,_ I,c_,,°l,,,,I ,,,,I ,,,,

0 0.05 0.1 0.15 0.2 0.25

2,000

1,500

1,000

5OO

0

0.3

Deflection(in)

...I

143
APPENDIX D---STRAIN GAUGE READINGS FROM STATIC INDENTATION TESTS

144
Clamped

Y

k Free

Free

Clamped

r X

3,800

3,600

._ 3,400

o

3,200

3,000

2,800

Load(N)

DO •

200 400 600 800 1,000 1,200

'' ' I '' ' I '' ' I '' ' I '' ' I '' ' I ' '

©

©

©

w

n

StaticIndentation#1

Nx=2,260Ib/inN/cm(1,290Ib/in)

©

©

©

©

©

m

m

u

©

0

,,,,I,,,, I,,,,I,,,,I,, ,, I,,,,-

50 100 150 200 250 300

Load(Ib)

145
Y

k Free

Clamped

4_

3_

2I0 I

I IX I

Free

Clamped

r X

¢=

2

¢J

Load(N)

0 500 1,000 1,500 2,000

3,800 - ' "' " I ' ' ' ' I ' ' ' ' ' ' ' ' I '

2,000 _ 6 6 6 6 6 6 6 6 • 6 6

1,000 6 6 6 v

_ Static Indentation#2

Nx--312 N/cm (178 Ib/in)

o

ooo "__

_,,ooo _ OOooo,, _,,_ ,,,, I,, 11_ I I I, , ,,,, -2,000

0 100 200 300 400 500

Load (ib)

146
Clamped

Free

10[+"_ (9)

Free

Clamped

_x

Load (N)

5,000

I

- StaticIndentation#3

° -- i . + +

._ 1,OOOFim,i i / i / i i i Ill

_i+ ,i * 'i

1'°°°I, _x o _ i

-_,ooo,,,I,,,,I,,,,I, ,,i,, ,i, ,_,,x,

0 200 400 600 800 1,000 1,200 1,400 1,600

'''1'''1'''1 '''1'''1'''1'''1'''1'''

_ - I I I I I

0 50 100 150 200 250 300 350 400

Load (Ib)

147
Clamped

Free

i _ (4)

3 _ (6)

2 _ (5)

Free

X

Clamped

2

.2

=E

Load(N)

0 200 400 600 800 1,000 1,200 1,400 1,600

5,000 i' ' ID ' 'D ' I D' ' D' I D' ' '_ ''D '1''' D _ I'_ ''_ I'''_j W I ii' 't

, ool ii" StaticIndentation #4

_ _ . NX= 2,716 N/cm= (1,550 Ib/in)

,,oo0 o o _o_ _ _ _,

Z_

3,500 Did Not Clamp

- 0

_ 0

_ 0

0

- I -2,000 Ilnl ii,,n,,,,I,,,,I,,,, I,,,,I,,,,1,,,,

0 50 100 150 200 250 300 350 400

Load(Ib)

148
Clamped

Y

k Free

1 _ (3)

5 _ (6)

2 _ (4)

Free

Clamped

X

.m.__

2

=Z

Load(N)

0 200 400 600 800 1,000 1,200 1,400 1,600

5'500 t' ' ' I' ' ' I' ' ' I '' '1''' I''' I' '' I ' ' ' I 'z_'

- A __X

Static Indentation#5 A A

Nx = 2,847 N/cm = (1,625 Ib/in) A

A A

s,ooo [] [] ix _

• I_ _ D D D D D

00_ _ _ _ D O D

4,500 • o _ • • n _

- • O X • • • = I_1"_

• O x

4,000

• o X X

- • o x -

- • o x x

• O

• O --

• O -

• o

• o Z

,,,,I,,,,I,, ,,ll, ,,I,, ,, 1,,,, I ,,,,I , #',,

-3,500

0 50 100 150 200 250 300 350 400

Load (Ib)

149
Clamped

Y

Free

2 _ (4)

5 _ (6)

1

1 _ (3)

Free

Clamped

r X

Load (N)

0 200 400 600 800 1,000 1,200 1,400 1,600

3,000 ' ' 'itstatic I 'Indentation ' ' I ' #6' ' I ' ' ' I ' ' ' I ' ' ' I ' ' ' I ' ' ' I ' '

Nx = 203 N/cm = (116 Ib/in) _,

2,000 - /% Z_ Z_ A /_ /k /_ /k /k Z_

1,000 "--- _ _' i_I/k []X []X _ _ .X X, X •X • .X

)

_ 1_ o o • e

_ o ;

° i

- ,ooo io, Ooo -2,000

-3,ooo , ,I,,,,I,,,,I, ,1,, ,1,,,,I,,,,I,,,

0 50 100 150 200 250 300 350 400

Load(Ib)

150
Y

Free

2 _ (6)

(8) 4 (7) 3

Clamped _

1 _ (5)

Clamped

_' x

Free

o

Load(N)

0 500 1,000 1,500 2,000

5,000 I<>' <>' <>' o I '

4,000

3,000

'''' I'''' I ' ' ' '

StaticIndentation#7

Nx = 187 N/cm= (107Ib/in)

<>

0

0

0

0

<>

<>

o []

o []

n nn

[]

DDD D

2,000 - 0 []

-1,000

'& A A A A A ,_

0 0 0 0 0 0 °

_- 0 0 0 0 -

0 100 200 300 400 500

Load(Ib)

151
Clamped

Y

k Free

(8) 4 (7) 3

2 _ (6)

1 _ (5)

Clamped

r X

Free

Load(N)

0 500 1,000 1,500 2,000

8,500 ' ' ' ' I ' ' ' ' I ' ' ' ' I ' ' ' ' I ' '

StaticIndentation#8

6,000 Nx = 2,586 N/cm = (1,476 Ib/in)

,,,oo [] [] []o __x _xx

5,ooo [] [] [] 0 0 °O X

nODv Y* _ox x x x x

,,,oo =_._x;o

_I |lI II IIi I II • • • II

3,ooo_ __6 _,_,

2,500='''' '' ''

0 100 200 300 400 500

Load (Ib)

!

1

! 152
Clamped

Y

b, Free

Ix ol

(4) 7 5 (2)

8 (3) (1) 6

li---51 I" el

Clamped

r X

Free

c:

P

¢J

Load(N)

7,000

6,000

0

m

m

500 1,000 1,500 2,000

''' ' I'''' I '''' I ''' ' I

StaticIndentation#9

Nx = 2,635 N/cm = (1,504 Ib/in)

4,ooo li t " i I

3,000

2,000

'1 I

u

w

m

i

m

m

m

w

m

N

m

,,,, I,,,, I ,,,, I ,, [, I ,, ,,-

100 200 300 400 500

Load(lb)

153
Clamped

Free

Ix "1

4 (7) (5) 2

Gauges1 and 8 Bad

3 (6)

V---d _

Clamped

F X

Free

f2

Load(N)

0 500 1,000 1,500 2,000

3'5°°1-''' ' I ' ' ' ' I ' ' ' ' I ' ' ' I I

I q

J_ Static Indentation#10

3,000 b Nx=219N/cm= (1251b/in) [] [] __

2,5oo _-- D D D D D

 ,oool ooooi ooo1

1,500 J_- [] 0 0 X X X X X X X X

I- D o x x x -I

1,ooo E- O x _ _ x z_ z_ z_ z_ z_ z_ z_ 'x z_ ---J-1

qP_n X © ,, A A J

soo • 6 • t • = = = • 6 6 6 6 •

0 100 200 300 400 500

Load(lb)

154
,

o

°

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REFERENCES

Johnson, W.: Oak Targets and Navel Gunnery in the Early 19th Century," Proceedings of

International Symposium on Composite Materials and Structures, Beijing, China, pp. 667-675,

June 10-13, 1986.

Paul, B; and Zaid, M.: "Penetration of a Thin Plate," Journal of the Franklin Institute, Vol. 265,

No. 4, pp. 317-335, 1948.

Gupta, B.P.; and Davids, N.: "Penetration Experiments With Fiberglass-Reinforced Plastics,"

Experimental Mechanics, pp. 445-450, September 1966.

Chamis, C.C.; Hanson, M.P.; and Serafini, T.T.: "Impact Resistance of Unidirectional Fiber

Composites," Composite Materials: Testing and Design (Second Conference), ASTM STP 497,

pp. 324-349, 1972.

Novak, R.C.; and DeCrescente, M.A.: "Impact Behavior of Unidirectional Resin Matrix

Composites Tested in the Fiber Direction," Composite Materials: Testing and Design (Second

Conference), ASTMSTP 497, pp. 311-323, 1972.

Cotterell, B.: "Fracture Toughness and the Charpy V-Notch Impact Test," British Welding

Journal, pp. 83-89, February 1962.

Toland, R.H.: "Instrumented Impact Testing of Composite Materials," Instrumented Impact

Testing, ASTM STP 563, pp. 133-145, 1974.

Phillips, D.C.; and Tetelman, A.S.: "The Fracture Toughness of Fibre Composites," Composites,

Vol. 3, pp. 216--223, September 1972.

Oplinger, D.W.; and Slepetz, J.M.: "Impact Damage Tolerance of Graphite/Epoxy Sandwich

Panels," Foreign Object Impact Damage to Composites, ASTMSTP 568, pp. 30-48, 1975.

Foreign Object Impact Damage to Composites, ASTM STP 568, American Society for Testing

and Materials, Philadelphia, PA, 1975.

Preston, J.L.; Jr., and Cook, T.S., "Impact Response of Graphite-Epoxy Flat Laminates Using

Projectiles That Simulate Aircraft Engine Encounters," Foreign Object Impact Damage to

Composites, ASTMSTP 568, pp. 49-71, 1975.

Husman, G.E.; Whitney, J.M.; and Halpin, J.C.: "Residual Strength Characterization of

Laminated Composites Subjected to Impact Loading," Foreign Object Impact Damage to

Composites, ASTMSTP 568, pp. 92-113, 1975.

155
13. Moon,F.C.;"Wave SurfacesDueto Impactof AnisotropicPlates,"Journal of Composite

Materials, Vol. 6, pp. 306-319, 1972.

14. Mortimer, R.W.; Chou, P.C.; and Carleone, J.: "Behavior of Laminated Composite Plates

Subjected to Impact," Foreign Object Impact Damage to Composites, ASTM STP 568, pp. 173-

182, 1975.

15. Sun, C.T.; Achenbach, J.D.; and Hermaann, G.; "Continuum Theory for a Laminated Medium,"

Journal of Applied Mechanics, pp. 467-475, September 1968.

16. Greszcuk, L.B.: "Response of Isotropic and Composite Materials to Particle Impact," Foreign

Object Impact Damage to Composites, ASTM STP 568, pp. 183-21 l, 1975.

17. Pintado, P.; Vogler, T.J.; and Morton, J.: "Impact Damage Development in Thick Composite

Laminates," Composites Engineering, Vol. 1, No. 4, pp. 195-210, 1991.

18. Kaczmarek, H.; and Maison, S.: "Comparative Ultrasonic Analysis of Damage in CFRP Under

Static Indentation and Low-Velocity Impact," Composites Science and Technology, Vol. 51, pp.

11-26, 1994.

19. Madsen, C.B.; Morgan, M.E.; and Nuismer, R.J.: "Scaling Impact Response and Damage in

Composites. Damage Assessment for Composite Cases," Aeronautical Laboratories--Technical

Report 90-03 7, 1990.

20. Swanson, S.R.: "Limits of Quasi-Static Solutions in Impact of Composite Structures,"

Composites Engineering, Vol. 2, No. 4, pp. 261-267, 1992.

21. Highsmith, A.L.: "A Study of the Use of Contact Loading to Simulate Low Velocity Impact,"

Final Report, NASA Grant No. NASA NCC8-23, 1993.

22. Jackson, W.C.; and Poe, C.C., Jr.: "The Use of Impact Force as a Scale Parameter for the Impact

Response of Composite Laminates," Journal of Composites Technology and Research, Vol. 15,

No. 4, pp. 282-289, Winter 1993.

23. Joshi, S.P.; and Sun, C.T.: "Impact Induced Fracture in a Laminated Composite," Journal of

Composite Materials, Vol. 19, pp. 51-66, January 1985.

24. Srinivasan, K.; Jackson, W.C., Smith, B.T.; and Hinkley, J.A.: "Characterization of Damage

Modes in Impacted Thermoset and Thermoplastic Composites," Journal of Reinforced Plastics

and Composites, Vol. 11, pp. 1111-1126, October 1992.

156
25.

26.

27.

28.

29.

30.

31.

32.

33.

34.

35.

36.

Salpekar,S.A.:"Analysis of Delaminationin Cross-PlyLaminatesInitiating from Impact

InducedMatrix Cracking,"Journal of Composites Technology and Research, Vol. 15, No. 2, pp.

88-94, 1993.

Cantwell, W.J.; and Morton, J.: "Detection of Impact Damage in CFRP Laminates." Composite

Structures, Vol. 3, pp. 241-257, 1985.

Kwon, Y.S.; and Sankar, B.V.: "Indentation-Flexure and Low-Velocity Impact Damage in

Graphite Epoxy Laminates," Journal of Composites Technology and Research, Vol. 15, No. 2,

pp. 101-111, 1993.

Dost, E.F.; Avery, W.B.; Ilcewicz, L.B.; Grande, D.H.; and Coxon, B.R.: "Impact Damage

Resistance of Composite Fuselage Structure, Part I," Proceedings of the Ninth DoD/NASA/FAA

Conference on Fibrous Composites in Structural Design, pp. 1037-1069, 1991.

Hull, D.; and Shi, Y.B.; "Damage Mechanism Characterization in Composite Damage Tolerance

Investigations," Composite Structures, Vol. 23, pp. 99-120, 1993.

Choi, H.Y.; Downs, R.J.; and Chang, F.K.: "A New Approach Toward Understanding Damage

Mechanisms and Mechanics of Laminated Composites Due to Low-Velocity Impact: Part IExperiments," Journal of Composite Materials, Vol. 25, pp. 992-1011, August 1991.

Choi, H.Y.; Wu, H.Y.T.; and Chang, F.K.; "A New Approach Toward Understanding Damage

Mechanisms and Mechanics of Laminated Composites Due to Low-Velocity Impact: Part II--

Analysis," Journal of Composite Materials, Vol. 25, pp. 1012-1038, August 1991.

Lammerant, L.; and Verpoest, I.: "The Interaction Between Matrix Cracks and Delaminations

During Quasi-Static Impact of Composites," Composites Science and Technology, Vol. 51, pp.

505-516, 1994.

Doucet, A.B.; and Qin, J.: "Microstructural Analysis of Impact Damage in Two Glass/Epoxy

Composite Laminates," Composites Engineering, Vol. 3, No. 1, pp. 55-68, 1993.

Liu, D.; and Malvern, L.E.: "Matrix Cracking in Impacted Glass/Epoxy Plates," Journal of

Composite Materials, Vol. 21, pp. 594--609, July 1987.

Liu, D.: "Impact-Induced Delamination--A View of Bending Stiffness Mismatching," Journal of

Composite Materials, Vol. 22, pp. 674-692, July 1988.

Kumar, P.; and Rai, B.: "Reduction of Impact Damage in KFRP through Replacement of Surface

Plies with Glass Fabric Plies," Journal of Composite Materials, Vol. 25, pp. 694-702, June 1991.

157
37. Clarke,M.P.; andPavier,M.J.; "Artificial DamageTechniquesfor Low Velocity Impactin

CarbonFibreComposites,"Composite Structures, Vol. 25, pp. 113-120, 1993.

38. Hitchen, S.A.; and Kemp, R.M.: "The Effect of Stacking Sequence on Impact Damage in a

Carbon Fiber/Epoxy Composite," Composites, Vol. 26, No. 3, pp. 207-214, 1995.

39. Choi, H.Y.; and Chang, F.K.: "A Model for Predicting Damage in Graphite/Epoxy Laminated

Composites Resulting from Low-Velocity Point Impact," Journal of Composite Materials, Vol.

26, No. 14, pp. 2134-2169, 1992.

40. Cantwell, W.J., "The Influence of Target Geometry on the High Velocity Impact Response of

CFRP," Composite Structures, Vol. 10, pp. 247-265, 1988.

41. Wyfick, D.A.; and Adams, D.F.: "Damage Sustained by a Carbon/Epoxy Composite Material

Subjected to Repeated Impact," Composites, Vol. 19, No. 1, pp. 19-27, January 1988.

42. Wang, H.; and Vu-Khanh, T.: "Damage Extension in Carbon Fiber/PEEK Crossply Laminates

Under Low Velocity Impact," Journal of Composite Materials, Vol. 28, No. 8, pp. 684-707,

1994.

43. Elber, W., "Failure Mechanics in Low-Velocity Impacts on Thin Composite Plates," NASA

Technical Paper 2152, 1983.

44, Dorey, G.: "Impact and Crashworthiness of Composite Structures." Structural Impact and

Crashworthiness, Vol. 1, J. Morton, ed., Elsevier Applied Science Publishers, London and New

York, 1984.

45. Cantwell,W.; Curtis, P.; and Morton, J., "Post-Impact Fatigue Performance of Carbon Fibre

Laminates With Non-Woven and Mixed-Woven Layers," Composites, Vol. 14, No. 3, pp. 301-

305, July 1983.

46. Avery, W.B.; and Grande, D.H.: "Influence of Materials and Lay-up Parameters on Impact

Damage Mechanisms, "Proceedings of the 22nd International SAMPE Technical Conference, pp.

470--483, November 1990.

47. Choi, H.Y.; Wang, H.S.; and Chang, F.K.: "Effect of Laminate Configuration and Impactor's

Mass on the Initial Impact Damage of Graphite/Epoxy Composite Plates Due to Line-Loading

Impact," Journal of Composite Materials, Vol. 26, No. 6, pp. 804-827, 1992.

48. Finn, S.R.; He, Y.F.; Springer, G.S.; and Lee, H.J.: "Compressive Strength of Damaged and

Repaired Composite Plates," Journal of Composite Materials, Vol. 26, No. 12, pp. 1796-1825,

1992.

158
49.

50.

51.

52.

53.

54.

55.

56.

57.

58.

59.

60.

Kim, C.G.; and Jun, E.J.: "Impact Resistance of Composite Laminated Sandwich Plates," Journal

of Composite Materials, Vol. 26, No. 15, pp. 2247-2261, 1992.

Rhodes, M.D.: Impact Tests on Fibrous Composite Sandwich Structures," NASA Technical

Memorandum 78719, 1978.

Awa, V.S.; Vala, J.R.; and Jeyaseelan, M.: "Effect of Impact and Fatigue Loads on the Strength

of Graphite/Epoxy Composites," Composite Materials: Testing and Design (Seventh

Conference), ASTM STP 893, pp. 196-206, 1986.

Jegley, D.: "Effect of Low-Speed Impact Damage and Damage Location on Behavior of

Composite Panels," Proceedings of the Ninth DoD/NASA/FAA Conference On Fibrous

Composites In Structural Design, pp. 1013-1036, November 1991.

Dow, M.B.; and Smith, D.L.: "Properties of Two Composite Materials Made of Toughened

Epoxy Resin and High-Strain Graphite Fiber," NASA Technical Paper 2826, 1988.

Jenq, S.T.; Wang, S.B.; and Sheu, L.T.: "A Model for Predicting the Residual Strength of GFRP

Laminates Subject to Ballistic Impact," Journal of Reinforced Plastics and Composites, Vol. 11,

pp. 1127-1141, October 1992.

Cairns, D.S.: "Impact and Post-Impact Response of Graphite/Epoxy and Kevlar/Epoxy

Structures," Ph.D. Dissertation, M.I.T., 1987.

Adsit, N.R.; and Waszczak, J.P.: "Effect of Near-Visual Damage on the Properties of

Graphite/Epoxy," Composite Materials: Testing and Design (Fifth Conference), ASTM STP 674,

pp. 101-117, 1979.

Dorey, G.: "Impact Damage in Composites--Development, Consequences and Prevention,"

Proceedings of the Sixth International Conference on Composite Materials, Second European

Conference on Composite Materials, Vol. 3, pp. 3.1-3.26, 1986.

Cantwell, W.J.; Morton, J.; and Curtis, P.: "A Study of the Impact Resistance and Subsequent 0-

Compression Fatigue Performance of Non-Woven and Mixed-Woven Composites," in Structural

Impact and Crashworthiness, Vol. 2 Conference Papers, J. Morton Ed., Elsevier Applied Science

Publishers, London and New York, pp. 521-531, 1985.

Dorey, G.; Bishop, S.M.; and Curtis, P.T.: "On the Impact Performance of Carbon Fiber

Laminates with Epoxy and PEEK Matrices," Composites Science and Technology, Vol. 23, pp.

221-237, 1985.

Morton, J., and Godwin, E.W.: "Impact Response of Tough Carbon Fibre Composites,"

Composite Structures, Vol. 13, pp. 1-19, 1989.

159
61.

62.

63.

4.

65.

66.

67.

68.

69.

70.

71.

72.

Caprino. G.: "Residual Strength Prediction of Impacted CFRP Laminates," Journal of Composite

Materials," Vol. 18, pp. 508-518, November 1984.

Nettles, A.T.: "Instrumented Impact and Residual Tensile Strength Testing of Eight-Ply

Carbon/Epoxy Specimens," NASA Technical Paper 2981, 1990.

Bishop, S.M.: "The Mechanical Performance and Impact Behavior of Carbon-Fibre Reinforced

PEEK," Composite Structures, Vol. 3, pp. 295-318, 1985.

Lal, K.M.: "Residual Strength Assessment of Low Velocity Impact Damage of Graphite-Epoxy

Laminates." Journal of Reinforced Plastics and Composites, Vol. 2, pp. 226-238, October 1983.

Poe, C.C., Jr.: "Summary of a Study to Determine Low-Velocity Impact Damage and Residual

Tension Strength for a Thick Graphite/Epoxy Motor Case," Proceedings of the 17th Congress of

the International Council of the Aeronautical Sciences, pp. 1005-1016, 1990.

Cantwell, W.J.; Curtis, P.T.; and Morton, J.: "An Assessment of the Impact Performance of

CFRP Reinforced with High-Strain Carbon Fibers," Composites Science and Technology, Vol. 25,

pp. 133-148, 1986.

Pilchard, J.C.; and Hogg, P.J.: "The Role of Impact Damage in Post-Impact Compression

Testing," Composites, Vol. 21, No. 6, pp. 503-511, November 1990.

Teh, K.T.; and Morton, J.: "Impact Damage Development and Residual Compression

Performance of Advanced Composite Material Systems," American Institute for Aaeronautics

and Astrionics Paper 93-1401-CP, 1993.

Silnivasan, K.; Jackson, W.C.; and Hinkley, J.A.: "Response of Composite Materials to Low

Velocity Impact, "Proceedings of the 36th International SAMPE Symposium, pp. 850-862, 1991.

Sjoblom, P.; and Hwang, B.: "Compression-After-Impact: The $5,000 Data Point!," Proceedings

of the 34th International SAMPE Symposium, pp. 1411-1421, 1989.

Nettles, A.T.; and Hodge, A.J.: "Compression-After-Impact Testing of Thin Composite

Materials, "Proceedings of the 23rd International SAMPE Technical Conference, pp. 177-183,

1991.

Manders, P.W.; and Harris, W.C.: "A Parametric Study of Composite Performance in

Compression-After-Impact Testing," SAMPE Journal, Vol. 22, No. 6, pp. 47-51,

November/December 1986.

160
73.

74.

75.

76.

77.

78.

79.

80.

81.

82.

83.

84.

85.

Tsai,X.; andTang,J.:"The ImpactBehaviorof LaminatedGlassFiber Composites by Weight

Dropping Testing Method," Proceedings of the 36th International SAMPE Symposium, pp.

1118-1127, 1991.

Kinsey, A., Saunders, D.E.J.; and Soutis, C.: "Post-Impact Compressive Behavior of Low

Temperature Curing Woven CFRP Laminates," Composites, Vol. 26, No. 9, pp. 661-667, 1995.

Dempsey, R.L.; and Horton, R.E.: "Damage Tolerance Evaluation of Several Elevated

Temperature Graphite Composite Materials," Proceedings of the 35th International SAMPE

Symposium, pp. 1292-1305, 1990.

Starnes, J.H., Jr.; Rhodes, M.D.; and Williams, J.G.: "The Effect of Impact Damage and Circular

Holes on the Compressive Strength of a Graphite-Epoxy Laminate," NASA Technical

Memorandum 78796, 1978.

Byers, B.A.; "Behavior of Damaged Graphite/Epoxy Laminates Under Compression Loading,"

NASA CR 159293, 1980.

Moon, D.; and Shively, J.H.; "Towards a Unified Method of Causing Impact Damage in Thick

Laminated Composites," Proceedings of the 35th International SAMPE Symposium, pp. 1466-

1478, 1990.

Nettles, A.T.; Lance, D.G.; and Hodge, A.J.: "Damage Tolerance of Candidate Thermoset

Composites for Use on Single Stage to Orbit Vehicles," NASA Technical Paper 3506, 1994.

Jones, R., "Residual Strength of Composites With Multiple Impact Damage," Composite

Structures, Vol. 28, pp. 347-356, 1994.

Gong, J.C.; and Sankar, B.V.: "Impact Properties of Three-Dimensional Braided Graphite/Epoxy

Composites," Journal of Composite Materials, Vol. 25, pp. 715-731, June 1991.

Evans, D.A.; and Boyce, J.S.: "Transverse Reinforcement Methods for Improved Delamination

Resistance, "Proceedings of the 34th International SAMPE Symposium, pp. 271-282, 1989.

Yamashita, S.; Hatta, H.; Takei, T.; and Sugano, T.: "Interlaminar Reinforcement of Laminated

Composites by Addition of Oriented Whiskers in the Matrix," Journal of Composite Materials,

Vol. 26, No. 9, pp. 1254-1268, 1992.

Ko, F.K.; and Hartman, D.: "Impact Behavior of 2-D and 3-D Glass/Epoxy Composites," Sampe

Journal, Vol. 22, No. 4, pp. 26-30, July/August 1986.

Dransfield, K.; Baillie, C.; and Mai, Y.W.: "Improving the Delamination Resistance of CFRP by

Stitching--A Review," Composites Science and Technology, Vol. 50, pp. 305-317, 1994.

161
86. Lu, W.H.; Liao, F.S.;Su,A.C.; Kao,P.W.;andHsu,T.J.:"Effect of InterleavingontheImpact

Responseof a UnidirectionalCarbon/EpoxyComposite,"Composites, Vol. 26, No. 3, pp. 215-

222, 1995.

87. Elber, W.: "Toughening of Graphite/Epoxy Composites by Interlarninar Perforated Mylar

Films," NASA Technical Memorandum 78643, 1978.

88. Norman, T.L.; and Sun, C.T.: "Delamination Growth in Composite Laminates with Adhesive

Strips Subjected to Static and Impact Loading," Composites Science and Technology, Vol. 46, pp.

203-211, 1993.

89. Marom, G.; Drukker, E.; Weinberg, A.; and Banbaji, J.: "Impact Behavior of Carbort/Kevlar

Hybrid Composites." Composites, Vol. 17, No. 2, pp. 150-153, April 1986.

90. Adams, D.F.; and Zimmerman, R.S.: "Static and Impact Performance of Polyethylene

Fiber/Graphite Fiber Hybrid Composites," SAMPE Journal, Vol. 22, No. 6, pp. 10-16,

November/December 1986.

91. Phillips, M.G.: "Fracture and Fatigue of Hybrid Composites," in Fibre Composite Hybrid

Materials, N.L. Hancox, ed., MacMillan Publishing Co., Inc., NY, pp. 119-162, 1981.

92. Peijs, A.A.J.M.; Venderbosch, R.W.; and Lemstra, P.J.: "Hybrid Composites Based on

Polyethylene and Carbon Fibers, Part 3: Impact Resistant Structural Composites Through

Damage Management," Composites, Vol. 21, No. 6, pp. 522-530, November 1990.

93. Busgen, A.W.; Effing, M.; and Scholle, M.: "Improved Damage Tolerance of Carbon Fibre

Composites by Hybridization with Polyethylene Fibre, Dyneema SK60," Proceedings of the

American Society for Composites Fourth Technical Conference, pp. 418-423, 1989.

94. Nettles, A.T.; and Lance, D.G.: "On the Enhancement of Impact Damage Tolerance

of Composite Laminates," Composites Engineering, Vol. 3, No. 5, pp. 383-394, 1993.

95. Jang, B.Z.; Chen, L.C.; Wang, C.Z.; Lin, H.T.; and Zee, R.H.: "Impact Resistance and Energy

Absorption Mechanisms in Hybrid Composites," Composites Science and Technology, Vol. 34,

pp. 305-335, 1989.

96. Demuts, E.; and Sharpe, P.: "Tougher Advanced Composite Structures," American Institute for

Aeronautics and Astrionics Paper 87-0794.

97. Ghasemi Nejhad, M.N.; and Parvizi-Majidi, A.: "Impact Behavior and Damage Tolerance of

Woven Carbon Fibre-Reinforced Thermoplastic Composites," Composites, Vol. 21, No. 2, pp.

155-I68, March 1990.

i62
98.

99.

100.

101.

102.

103.

104.

105.

106.

107.

108.

109.

Recker, H.G.; Altstadt, V.; Eberle, W.; Folda, T.; Gerth, D.; Heckmann, W.; Ittemann, P.; Tesch,

H.; and Weber, T.: "Toughened Thermosets for Damage Tolerant Carbon Fiber Reinforced

Composites," SAMPEJournal, Vol. 26, No. 2, pp. 73-78, March/April 1990.

Lance, D.G.; and Nettles, A.T.: "Low Velocity Instrumented Impact Testing of Four New

Damage Tolerant Carbon/Epoxy Composite Systems," NASA Technical Paper 3029, 1990.

Cantwell, W.J.; and Morton, J.: "Comparison of the Low and High Velocity Impact Response of

CFRP," Composites, Vol. 20, No. 6, pp. 545-551, November 1989.

Cantwell, W.J.; and Morton, J.: "Geometrical Effects in the Low Velocity Impact Response of

CFRP," Composite Structures, Vol. 12, pp. 39-59, 1989.

Verpoest, I.; Marien, J.; Devoes, J.; and Wevers, M.: "Is Absorbed Energy a Test Independent

Parameter for Damage and Residual Strength After Impact?," Proceedings of the Second

International Conference on Testing, Evaluation and Quality Control of Composites--TEQC 87,

pp. 69-77, September 1987.

Robinson, P.; and Davies, G.A.O.; "Impactor Mass and Specimen Geometry Effects in Low

Velocity Impact of Laminated Composites," International Journal of Impact Engineering, Vol.

12, No. 2, pp. 189-207, 1992.

Shivakumar, K.N.; Elber, W.; and Illg, W.: "Prediction of Impact Force and Duration Due to

Low-Velocity Impact on Circular Composite Laminates," Transactions of the ASME, Vol. 52, pp.

674-680, September 1985.

Poe, C.C., Jr.: "Relevance of Impacter Shape to Nonvisible Damage and Residual Tensile

Strength of a Thick Graphite/Epoxy Laminate," NASA Technical Memorandum 102599, 1990.

Morton, J.: "Scaling of Impact Loaded Carbon Fiber Composites," AIAA Journal, Vol. 26, No. 8,

pp. 989-994, 1988.

Swanson, S.R.: "Scaling of Impact Damage in Fiber Composites from Laboratory Specimens to

Structures," Composite Structures, Vol. 25, pp. 249-255, 1993.

Bucinell, R.B.; Madsen, C.B.; Nuismer, R.J.; Benzinger, S.T.; and Morgan, M.E.: "Experimental

Investigation of Scaling Impact Response and Damage in Composite Rocket Motor Cases,"

Proceedings, JANNAF Composite Motor Case and Structures and Mechanical Behavior Meeting,

Jet Propulsion Laboratory, Anaheim, CA, November 7-9, 1989.

Chen, V.L.; Wu, H.Y.T.; and Yeh, H.Y.: "A Parametric Study of Residual Strength and Stiffness

for Impact Damaged Composites," Composite Structures, Vol. 25, pp. 267-275, 1993.

163
110.

111.

112.

113.

114.

115.

116.

117.

118.

119.

120.

121.

122.

Horton,R.E.; andMcCarty,J.E.:"DamageToleranceof Composites,"in Engineered Materials

Handbook, Vol. 1--Composites. ASME International, Metals Park, OH, pp. 264-270, 1988.

Olster, E.F.; and Roy, P.A.: "Tolerance of Advanced Composites to Ballistic Damage,"

Composite Materials: Testing and Design (Third Conference), ASTMSTP 546, pp. 583--603,

1974.

Sharma, A.V.: "Low-Velocity Impact Tests on Fibrous Composite Sandwich Structures," Test

Methods and Design Allowables for Fibrous Composites, ASTM STP 734, pp. 54-70, 1981.

Avva, V.S.: "Impact Initiated Damage in Laminated Composites," AFOSR TR 82-1038, 1982.

Avva, V.S.: "Fatigue/Impact Studies on Laminated Composites," AFWAL TR 83-3060, May

1983.

Avva, V.S.; Vala, J.R.; and Jeyaseelan, M.: "Effect of Impact and Fatigue Loads on the Strength

of Graphite/Epoxy Composites," Composite Materials: Testing and Design (Seventh

Conference), ASTM STP 893, pp. 196-206, 1986.

Rhodes, M.D.; Williams, J.G.; and Starnes, J.H., Jr.: "Low-Velocity Impact Damage in GraphiteFiber Reinforced Epoxy Laminates," Polymer Composites, Vol. 2, No. 1, pp. 36-44, January

1981.

Starnes, J.H.; Rhodes, M.D.; and Williams, J.G.: "Effect of Impact Damage and Holes on the

Compressive Strength of a Graphite/Epoxy Laminate," Nondestructive Evaluation and Flaw

Criticality for Composite Materials, ASTM STP 696, pp. 145-171, 1979.

Park, N.: "The Impact Response of Composites Under Stress," IMechE Paper No. C400/019,

pp. 137-142, 1990.

Sankar, B.V.; and Sun, C.T.: "Low-Velocity Impact Damage in Graphite-Epoxy Laminates

Subjected to Tensile Initial Stresses," A/AA Journal, Vol. 24, No. 3, pp. 470--471, 1985.

Nettles, A.T.; and Lance, D.G.: "The Effects of Compressive Preloads on the CompressionAfter-Impact Strength of Carbon Epoxy," NASA Technical Paper 3303, 1992.

Robb, M.D.; Arnold, W.S.; and Marshall, I.H.: "The Damage Tolerance of GRP Laminates

Under Biaxial Prestress," Composite Structures, Vol. 32, pp. 141-149, 1995.

Sun, C.T.; and Chen, J.K.: "On the Impact of Initially Stressed Composite Laminates," Journal

of Composite Materials, Vol. 19, pp. 490-504, November 1985.

164
123.

124.

125.

126.

127.

128.

129.

Chen,J.K.; andSun,C.T.:"DynamicLargeDeflectionResponseof CompositeLaminates

Subjectedto Impact,"Composite Structures, Vol. 4, pp. 59-73, 1985.

Sun, C.T.; and Chattopadhyay, S.: "Dynamic Response of Anisotropic Laminated Plates Under

Initial Stress to Impact of a Mass," Journal of Applied Mechanics, pp. 693-698, September

1975.

Bert, C.W.; and Birman, V.: "Response of Prestressed Cylindrically Curved Composite

Structures Subjected to Low-Velocity Impact," Proceedings of the Fourth Japan-U.S.

Conference on Composite Materials, American Society for Composites, pp. 43-52, 1988.

Pipes, R.B.; and Pagano, N.J.: "Interlaminar Stresses in Composite Laminates Under Uniform

Axial Extension," Journal of Composite Materials, Vol. 4, pp. 538-548, October 1970.

Pipes, R.B.; and Daniel, I.M.: "Moire Analysis of the Interlaminar Shear Edge Effect in

Laminated Composites," Journal of Composite Materials, Vol. 5, pp. 255-259, April 1971.

Pagano, N.J.; and Pipes, R.B.: "The Influence of Stacking Sequence on Laminate Strength,"

Journal of Composite Materials, Vol. 5, pp. 50-57, January 1971.

Madsen, C.B.; Morgan, M.E.; and Nuismer, R.J.: "Subelement Validation for Damage

Assessment for Composite Cases," Phillips Laboratory Technical Report #93-3015, April 1993.

165

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1. AGENCY USE ONLY (Leave Blank) 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED

November 1997 Technical Memorandum

4. TITLE AND SUBTITLE 5. FUNDING NUMBERS

The Impact Response of Carbon/Epoxy laminates

MSFC Center Director's Discretionary Fund Final Report, Project No. 94-13

6. AUTHORS

A.T. Nettles and A.J. Hodge

7. PERFORMING ORGANIZATION NAMES(S) AND ADDRESS(ES)

George C. Marshall Space Flight Center

Marshall Space Flight Center, Alabama 35812

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

Washington, DC 20546-0001

8. PERFORMING ORGANIZATION

REPORT NUMBER

M-844

10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

NASA/TM-97-206317

11. SUPPLEMENTARY NOTES

Prepared by Materials and Processes Laboratory, Science and Engineering Directorate

1211. DISTRIBUTION/AVAILABILrrY STATEMENT

Subject Category 39

Unclassified-Unlimited

Nonstandard Distribution

12b. DISTRIBUTION CODE

13. ABSTRACT (Maximum200 words)

Low velocity dropweight impact tests were conducted on carbon/epoxy laminates under various boundary conditions. The

composite plates were 8-ply (+45,0,-45,90)s laminates supported in a clamped-clamped/free-free configuration with varying

amounts of in-plane load, Nx, applied. Specimens were impacted at energies of 3.4, 4.5, and 6 Joules (2.5, 3.3, and 4.4 ft-lb).

The amount of damage induced into the specimen was evaluated using instrumented impact techniques, x-ray inspection, and

cross-sectional photomicroscopy. Some static identation tests were performed to examine if the impact events utilized in

this study were of a quasi-static nature and also to gain insight into the shape of the deflected surface at various impact load

combinations. Load-displacement curves from these tests were compared to those of the impact tests, as was damage

determined from x-ray inspection. The finite element technique was used to model the impact event and determine the stress

field within the laminae.

Results showed that for a given impact energy level, more damage was induced into the specimen as the external in-plane

load, Nx, was increased. The majority of damage observed consisted of back face splitting of the matrix parallel to the

fibers in that ply, associated with delaminations emanating from these splits. The analysis showed qualitatively the results of

impact conditions on maximum load of impact, maximum transverse deflection, and first failure mode and location.

14. SUBJECT TERMS

composite materials, impact, damage tolerance

17. SECURITY CLASSIRCATION

OF REPORT

Unclassified

NSN 7540-01-280-5500

15. SECURITY CLASSIRCATION

OF THIS PAGE

Unclassified

19. SECURITY CLASSIRCATION

OF ABSTRACT

Unclassified

15. NUMBER OF PAGES

176

!16. PRIDE COOE

A09

20. UMITATION OF ABSTRACT

Unlimited

Standard Form 298 (Rev. 2-89)

Pnmc_bed by ANgel S_l. 230-16

290"102
APPROVAL

THE IMPACT RESPONSE OF CARBON/EPOXY LAMINATES

(MSFC Center Director's Discretionary Fund Final Report, Project No. 94-13)

A.T. Nettles and A.J. Hodge

The information in this report has been reviewed for technical content. Review of any

information concerning Department of Defense or nuclear energy activities or programs has been made

by the MSFC Security ClaSsification Officer. This report, in its entirety, has been determined to be

unclassified.

Dr. A.F. Whitaker

DIRECTOR, MATERIALS and PROCESSES LABORATORY