Fatigueand

Residual Strengthof

Composite Aircraft Structures

Tonny Nyman

Department of Aeronautics

Kungliga Tekniska Högskolan

(Royal Institute of Technology)

SE-10044 Stockholm

Sweden

Report No. 99-26

ISSN 0280-4646

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PREFACE

Most of the work presented in this thesis was carried out within a national research prograSweden, “Life Prediction Certification and in Service Assessment of Composite Structubetween July 1995 to July 1998 at Saab AB and at the Royal Institute of Technology,Department of Aeronautics. It was financially supported by Saab and NFFP, the NationaAeronautical Research Program in Sweden, for which I am very grateful.

Most of the work carried out in this dissertation stems from research ideas formulated wthat program in which I had the privilege to act as a project leader. Participating parts wbesides Department of Structural Strength and Loads Technology at Saab AB, the AeronaResearch Institute of Sweden (FFA) and the Royal Institute of Technology, Department Aeronautics (KTH).

I would like to thank my supervisor Professor Anders Blom at the Royal Institute ofTechnology for his support and encouragement during the course of this study and ProJan Bäcklund for the opportunity to carry out this work at his department. Joakim Schön, aAeronautical Research Institute of Sweden, is also greatly acknowledged as a friend,companion in the project and as a valuable discussion partner. At the same institute I walso like to thank Dr. Börje Andersson for his enthusiasm and deep knowledge, especiacomputational mechanics, which enabled us to perform state-of-the-art calculations withdeveloped hp-adaptive FE-system STRIPE. I am also very grateful for the support recefrom colleagues at Saab AB, especially Tomas Ireman, Christina Altkvist, Magnus FribergAnders Bredberg at the composite group but also to Hans Ansell, previous manager of department and co-author of two of the papers. To my friends in the metallic group, althyou haven’t contributed to the appended papers I have benefited from your friendship for wI am very grateful. A special thanks to a skilful, experienced and observant technician at Cmaterialteknik, Bernt Forslind.

Finally I would like to thank my family, Ingrid, Torbjörn, Per and Henrik for their support anpatience during this period. To my very special dog, XITA, thank you for forcing me out long walks in the forest, thus clearing my thoughts and giving me new fresh angles to cproblems.

Linköping, August 1999

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ABSTRACT

This thesis consists of six papers within the area of composite fatigue and residual strenaircraft structures. The dissertation is a mixture of experimental, numerical and theoreticwork with focus on some structural hot-spots as defined in the certification process.

Basic fracture mechanical properties were investigated, both numerically and experimenfor the standard DCB-and ENF-specimens, with respect to different ply interfaces and mals. Toughness measurements on specimens with non-zero interfaces have a considerahigher toughness than specimens with zero interfaces, due to the fact that the initial crackates from the original symmetrical crack plane. A good prediction between FE-calculationbeam data analysis methods could be obtained if the non-homogenous characteristics imens with non-zero interfaces was taken into account. The cyclic delamination growth increased rapidly with increased∆G, raising doubts about the structural applicability of sucgrowth laws. The problem is rather a threshold problem. The importance of repeated loaon composite structures with artificial delaminations is also demonstrated.

The rule of interlaminar stresses with respect to damage initiation on notched and unnospecimens were numerically investigated with an state-of-the-art hp-adaptive FE-techniqucompared to some existing semi-analytical models. Though the number of initiation pointvarious quasiisotropic layups were found to be quite different the initiation strain, from aculation point of view, was almost the same. Investigated semi-analytical models as formutoday were not able to cope with the complex stress-state occurring at free edges.

A simplified approach for composite fatigue design is suggested. The methodology is basfatigue failure functions established on a laminate level and thus the calculated uniaxialerties will reflect the overall behaviour of the laminate. To establish the fatigue failure futions, a given SN-curve for a multi-directional laminate is considered as a series of failupoints where the failure in the short life-region essentially is assumed to be fibre fracturein the long life-region is assumed to be controlled by matrix properties.

An experimental investigation was also carried out regarding effects on resulting life for posite structures. Different elimination levels were considered on various load spectra aconstant amplitude block loading. Results revealed that the elimination level could be seapproximately 50% of the maximum load range. This means that considerable time and cstructural verification testing can be obtained. Spectrum fatigue testing accounting for aacteristic number of cycles could be mapped on constant amplitude data which means thaprediction methodology can be radically simplified.

The so called soft inclusion theory and delamination buckling theory was investigated oimpact damaged composite structures for residual strength prediction. The investigationsupported by FE-technique for determination of the stress distribution in the buckled statfor characterisation of the damaged region. It could be concluded that for low-energy imalthough with conservative assumptions on stiffness reduction, the soft inclusion theoryconservative for residual strength prediction. In contrast the delamination buckling theorshows good agreement for various impact energy levels, thicknesses and layups.

ii

CONTENTS

PREFACE ............................................................................................................... i

ABSTRACT ........................................................................................................... ii

CONTENTS ......................................................................................................... iii

DISSERTATION .................................................................................................. iv

INTRODUCTION ................................................................................................. 1

1 Background .................................................................................................. 1

1.1 Certification process ................................................................................ 2

1.2 Damage characterisation .......................................................................... 5

2 Interlaminar stresses ..................................................................................... 7

3 Basic fracture mechanical properties ............................................................ 9

4 Composite fatigue ....................................................................................... 14

4.1 Basic methodology ................................................................................ 14

4.2 Truncation and elimination .................................................................... 16

5 Residual strength ........................................................................................ 16

5.1 Equivalent damage methodology .......................................................... 18

5.2 Delamination buckling theory ............................................................... 20

5.3 Artificial delaminations ......................................................................... 21

SUMMARY OF APPENDED PAPERS ............................................................. 22

DIVISION OF WORK BETWEEN AUTHORS ................................................ 25

RECOMMENDATIONS FOR FUTURE WORK .............................................. 26

REFERENCES .................................................................................................... 27

PAPER A .....................................................................................................A1-A29

PAPER B ...................................................................................................... B1-B35

PAPER C ...................................................................................................... C1-C21

PAPER D ......................................................................................................D1-D23

PAPER E........................................................................................................E1-E24

PAPER F........................................................................................................F1-F22

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DISSERTATION

This dissertation contains the following appended papers:

PAPER A: Schön J., Nyman T., Blom A. and Ansell H., “A numerical and experimtal investigation of basic fracture mechanical properties in the DCB-spimen”. Accepted for publication in Composites Science and Technolo

PAPER B: Schön J., Nyman T., Blom A. and Ansell H, “A numerical and experimtal investigation of basic fracture mechanical properties in the ENF-simen”. Submitted to Engineering Fracture Mechanics.

PAPER C: Nyman T. and Friberg M., “Interlaminar stresses in composite notchedunnotched laminates”. Accepted for publication in Journal of ReinforcPlastics and Composites. Presented at the 13th International Conference ofthe American Society for Composites in Baltimore, September 21-231998.

PAPER D: Nyman T., " Composite Fatigue Design Methodology A SimplifiedApproach". Published in Composite Structures 35 (1996) 183-194. Cdensed version of paper presented at the Third International ConferencComposites Engineering, ICCE/3, in New Orleans, July 21-26, 1996.

PAPER E: Nyman T., Ansell H. and Blom A., “Effects of truncation and eliminatioon composite fatigue life”. Submitted to Composite Structures.

PAPER F: Nyman T., Bredberg A. and Schön J., “Equivalent damage and resistrength for impact damaged composite structures”. Accepted for pubtion in Journal of Reinforced Plastics and Composites. Presented at 13th International Conference of the American Society for CompositesBaltimore, September 21-23, 1998.

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INTRODUCTION

1 Background

New aircraft must comply with requirements for durability and damage tolerance. A modcombat aircraft today contains high percentages of advanced composite material in flighcal primary structures, e.g the B-2 stealh bomber, the F-22 advanced tactical fighter, the Fstealth fighter and the V-22 tilt-rotor, Ref. [1]. The new Swedish fighter JAS39 Gripen utiliabout 20%, by weight, carbon fibre/epoxy composite laminates in the primary structure,Fig. 1.

Figure 1 The extent of composite materials in JAS39 Gripen.

Specific damage tolerance requirements and design criteria for advanced composmaterials were not specified during design and development of various military aircraft iduced in the 1970’s and early 80’s such as the F-16, F-15 and F/A-18, Ref. [1]. Instead, datolerance were demonstrated by extensive structural testing and by the use of low designallowables in the design. An excellent review of lessons learned with the use of advancedposites can be found in Ref. [2]. Some conclusions are cited below:

o Significant weight savings have been achieved

o Outstanding fatigue and corrosion resistance have been demonstrated

o In-service performance has been very successful

o Durability of thin skin honeycomb structures has been less successful

A review of composite design philosophy/process can be found in Ref. [3] where it is em

Vertical fin

Nose landing

Fillet

Leading edge flaps

AccessDoors

Main landinggear door

gear doors

Main wing

Fin tip

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Glass-Fiber Composite (GFRP)

Aramid-Fiber Composite (AFRP)

Carbon-Fiber Composite (CFRP)

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sized that a good design is usually one where complexity and sophistication are kept toimum, as these often only serve to increase costs and reduce reliability. It is also pointethat joints frequently have been a problem, thus design which will reduce the number of jwill generally lead to weight and cost savings and an improvement in component integrity.siderations when looking at detailed design can be summarised as follows:

o Avoid offset loads which result in local bending

o Avoid peeling forces

o Consider galvanic corrosion effects between carbon composite and many met

o Look for local weaknesses and anticipated failure modes

o Remember that joints may be locally stiffened by the addition of extra laminati

Of special concern are also what type of loads that are significant/characteristic and howare applied on the geometrical and laminate configuration at hand, i.e whether they are tecompressive, shear etc., and how they appear in the sequence.

1.1 Certification process

Currently, the main design drivers for most composite structural components is either stiffstatic strength or low energy impact damage. In order to avoid structural problems curredesign procedures have led to rather low allowable strain levels. It is anticipated that thesues will be increased in future aircraft and then it may be necessary to further consider thlihood for fatigue damage. In the certification process, durability and damage tolerance crrequire composite structures containing undetected damage to be acceptable to fly.

Current design philosophy is based on no damage growth criteria and is realised bservative strain limitations (e.g assuming that a Ø6 mm hole can be present everywherestructure representing e.g impact damage, manufacturing defects etc.) and verification on components and built-up structures. Consequently, expensive testing, not only for certion of new structures but also for comparable structures with only minor alternations, isquently performed by the industry. The problem is usually represented by a test pyramid weach stage refers to an investigation level in terms of specimen category, see Fig. 2.

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Figure 2 Test pyramid to certify composite structures.

In Fig. 3 some of the structural verification tests performed on JAS39 are shown. In fact theing is on-going and extended with 39B the two-seater.

Figure 3 Structural verification testing on JAS39 Gripen.

Special attention must be made regarding coupon v.s structural testing. For examedge delamination occurs in unnotched coupons and significantly influences the result. ever, for a structural panel the effect is much less. Also of special interest is, depending ogeometry, the difference in static and cyclic strength between tension and compression lo

FULL SCALECOMPONENTS

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FATIGUE & DAMAGE TOLERANCE

DAMAGE TOLERANCE

FATIGUE

GUN FORWARDATTACHMENT

DEFLECTORFORWARDATTACHMENT

ENGINE AFTMOUNT

AFT ROOT RIB

WING-FUSELAGEJOINT

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Of special importance is also the scatter in structural response achieved in structural tethe structural response variability. Besides the scatter on a coupon level several other facontribute to the overall scatter, e.g structural geometry, loading conditions, manufacturinmaterial conditions. Due to this scatter unexpected failure can occur, i.e “hot-spot failure”Fig. 4. Very often these “hot-spot failures” are associated to out-of plane loading emphasthe need for analysis capability with respect to interlaminar stresses.

Figure 4 Schematic of “Hot-Spot” failure for a static test, from Ref. [4, 5].

In the process of structural verification testing usually both composite and metallic pare involved. In order to have reasonable test times, truncation and elimination of the losequence must be done. This process, load sequence reduction, usually leads to differeclusions for the composite structure and for the metallic part.

Lessons learned within the certification process are, see e.g. Ref. [2]:

o There is a need for analysis capability within the area of interlaminar stress cation since secondary loads (out-of-plane loading) are almost impossible to elimifrom complex built-up structures.

o The full scale static test is the key test for composites since it is the primary “hot sindicator.

o Lack of reliable life prediction tools makes it necessary to rely on a no-growth cept and makes it impossible to investigate the consequences of a load spectrupdate.

o There is an urgent need for an agreement what should be a detectable size of an

Hot-Spot Failure

Mean Strength

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tation, i.e there is a lack of appropriate definition of barely visible impact dama(BVID). There is also need for an agreement of cut-off energy, values from a feJoules to 140 J exists.

o There is also need for rational repair/reject criteria including a methodology fodetermining inspection intervals.

o No complete airframe test validation is required for composite damage tolerancetification. Extensive testing has shown that subcomponent validation tests accurrepresent fullscale composite damage tolerance behaviour.

1.2 Damage characterisation

During operating conditions damage of different nature, e.g service mechanical damagefatigue damage or environmental damage, will be observed/detected. Facing different dammanufacturing or service induced, delamination related damage is one of the far most impfailure mechanisms of aircraft-composite structures and can be detrimental for flight safConsequently delaminations must be considered in the design process but also in structuification testing. Examples are delamination growth due to local buckling, delamination grofrom free edges and notches such as holes, delamination growth from ply drops, and imdamaged composites containing considerable delamination. Delamination growth can aoccur due to interlaminar stresses which can arise from fuel pressure variations and in costructures due to unanticipated loading.

The complex nature of composite failure, involving different failure modes and theiinteractions, makes it necessary to characterise / identify the relevant parameter(s) for daresistance, accumulation and life prediction. It is not obvious that a major characteristic famechanism on a coupon level will be equally interesting for a real structure. It is well recnised that the most damaging state is impact damage at least for compression dominating. In the certification process different damage states must be identified to enable verificof the structure. For instance for barely visible impact damage (BVID), the structure mustain ultimate load before and after cyclic loading.

The three main types of defects observed in fiber composites are transverse cracksplitting, delamination between plies and fiber fractures which occur mainly in the last stathe life of the structure Refs. [6], [7]. Transverse matrix cracks are initiated very early in fatigue process and they multiply until a saturation state is reached, Characteristic DamState (CDS) Ref. [8]. The loading level has a large effect on the rate of crack multiplication[6], [9].

The matrix can not easily deform in a non-linear way which makes it difficult to relax astress concentrations. Instead, matrix cracking will serve to reduce any stress concentraand the consequence will be redistribution of load in the composite and increased stresstrain in the 0o plies. Most of the matrix cracks propagate through the thickness of the plythey can usually not propagate into the next ply if the difference in fiber orientation is suciently large. A stress singularity will occur at the crack tip of the matrix cracks which dur

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In the 0o plies matrix cracks will form parallel to the fibres, i.e. splitting, and the distanbetween the splitting cracks is larger than for the transverse matrix cracks and it is randdistributed throughout the specimen, Refs. [7], [9]. These dispersed cracks are observed mat low loads and increase with the number of cycles up to failure. At higher stresses, localongitudinal cracking appears to be the main damage mechanism prior to failure, Ref. [9]composite component will fail when the 0o plies fail, requiring fiber fracture. It has beenobserved that most of the fiber failure and damage localization occurs in the last 10% of thof the laminate, Ref. [9]. The last part of a specimen’s fatigue life has been reported to shoaccelerating rate of damage development characterized by the localization of damage inof increasing crack interaction, Ref. [10]. This involves the coalescence of microcracks adevelopment of macro-cracks.

It has been found that the growth of a delamination is dependent on the stacking sequDelaminations were found to propagate perpendicular to the loading direction in a (0/903)s lam-inate, as opposed to parallel to the loading direction in a (0/902)s laminate. The delaminationarea can become very large at some interfaces, values of 57% delamination have been reThis means that towards the end of the fatigue life of the composite the different plies areor less isolated from each other. It is also possible to have delaminations to form groups whigh density of delaminations Ref. [7].

Compression-induced damage is of a different nature than tension-induced damagecompression induced damage will depend on the interlaminar stresses which in turn strdepends on the stacking sequence and the thickness of the laminate. Tests were conduRatwani, Ref. [11], on AS/3501-6 graphite epoxy with four different stacking sequenceslaminates had the same percentage of 0o, 90o and ±45o degree plies. Constant amplitude teswere performed at a loading ratio of . Depending on stacking sequence diffefailure modes were obtained for the central circular notched specimens. It was also concthat the damage growth changed direction and location completely when the stacking seqwas changed. The location of delamination and matrix cracking for the different laminatescide with the highest interlaminar shear or normal stress calculated with FE-technique.

Test results presented, especially compressive test results, must be critically reviesince interlaminar stresses, depending on failure mode, can have played a dominant ro[12]. There is always a question to what extent they are invoked in the test results, i.e. theybe viewed in light of test methods, for example special fixtures used to prevent bucklinge.g. Highsmith, Ref. [13]. Compression-compression fatigue damage in notched ((±45/02)3)scomposite laminates, T800/924C, consists of matrix cracking, delamination, and at appcompression stresses approaching the laminate strength, fibre microbuckling, Ref. [14].loading tests were performed at constant-amplitude, R=10, with a frequency of 10Hz.

A failure prediction technique which takes microbuckling into account was introducedSoutis et al., Ref. [15]. Failure is due to 0ofibre microbuckling surrounded by delamination. Thmicrobuckling zone initiates at the hole boundary. It extends stably under increasing loa

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before becoming unstable at a critical microbuckling length of approximately 2-3 mm. Amicrobuckling failure criterion, in conjunction with Lekhnitskii’s complex variable techniqRef. [16], was introduced by Gürdal et al., Ref. [17]. The schematic kinking mode of failureshown in Fig. 5.

Figure 5 Kinking mode of failure, Ref. [17].

In-service detected damage must be correlated to life prediction methodology andservice inspection tools and repair/reject criteria must be developed. There must also brelation between what is inspectable/seen and life prediction- residual strength methodolomutual correspondence between methods and characterized damage must be sought.

2 Interlaminar stresses

Secondary loads are almost impossible to avoid in a built-up structure. These secondarywill produce out-of-plane loading resulting in interlaminar stresses in principal structuralments which could cause premature failure. It is believed that these interlaminar stressea dominant role in the process of delamination and thus are a limiting factor for the life ostructural component. Interlaminar stresses will arise when e.g a structural component jected to acoustic random loading, post buckling, fuel pressure variations or when impaThese stresses must thus be considered in e.g. structural verification testing and in the process.

Interlaminar stresses also provide important information regarding damage initiatiogrowth and could therefore be the key for better understanding of the failure process. Theage initiation will most likely depend on these interlaminar stresses and consequently ostacking sequence. It is also likely that delaminations start to develop where interlaminastresses are most severe.

To have analysis capability within this area, i.e to calculate the resulting interlaminstresses in a straightforward manner is important. Some out-of-plane sources are exemin Fig. 6 below. Coupled with a proper failure criterion the calculations could also be an im

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Among the first persons to identify and recognise the existence of interlaminar strewas Pagano, Ref. [18]. The three-dimensional problem was solved by applying finite differtechniques to the governing elasticity equations. The existence of a boundary layer and thsibility of a singular behaviour was launched. Since then many three-dimensional finite elesolutions have been published see e.g. Refs. [19, 20] and [21]. The singularity problem discussed in Refs. [22] and [23]. Boundary collaction methods, methods depending on datheory, specially developed continuous strain high order elements but also plasticity relamethods have been used for the solution of the problem, see Refs. [24, 25, 26, 27]. ThouFE-method provides relative accuracy to the problem, the need for simplified methods isous. In 1986 an article which fulfilled that need, Ref. [28] was published. By assuming sfunctions and applying traction continuity and minimising the complementary potential ftion the problem could be solved. The article was followed by a number of articles dealingother types of loading not only uni-axial loading but also combination of thermal loads amechanical effects, Refs. [29, 30, 31, 32]. Basically the same theory was adopted to studinfluence of a hole on the interlaminar stresses, i.e. a curved free edge Refs. [33, 34, 35, 3By turning the problem into a series of straight free edges an approximate solution coulfound. FE-based solutions to the hole problem can be found in e.g. Ref. [38].

For the prediction of damage initiation stress based criteria, see e.g. Ref. [39, 40], orture mechanics related criteria, see Ref. [41], have been adopted. Delamination and failurysis are also discussed in Refs. [42, 43, 44, 45]. An excellent review has been made in Rewhich also contains a number of numerical solutions to the free edge problem.

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Normal and or shear interlaminar stresses arise at the free edges depending on emismatch of properties between the layers. The magnitude of the stresses will depend oning sequence and on external loading. Interesting is also that these stresses will dependperature, as layers have different expansion coefficients, and also with moisture contenlayers expand to different extents on absorbing external moisture, Ref. [47].

In paper C interlaminar stresses for notched and unnotched plates, for various stasequences, are calculated with an hp-adaptive FE-analysis technique and compared toexisting semi-analytical methods. A delamination criterion is also examined and comparcalculated results. Fig. 7 shows, for one of the analysed cases, the interlaminar stressehole-edge.

Figure 7 Interlaminar stresses at an cut-out.

3 Basic fracture mechanical properties

It is important to improve the knowledge of delamination growth both theoretically and eximentally. The most straightforward way is by studying simple test specimens, such as the

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Several round robin tests of unidirectional composites have been performed, with on e.g. delamination onset, T-tab v.s hinge, insert type and thickness, Ref. [48]. These igations have been used for establishing the ASTM (American Society for Testing and Mals) DCB standard. An overview of experiments and analyses being performed at the colevel within the GARTEUR (Group of Aeronautical Research in Europe) action group AGcan be found in Ref. [49]. The crack initiation has been studied with microfocus radiograon DCB-specimens and it was found that the crack initiates in the interior of the specimresulting in a curved crack tip, and that the crack initiation corresponds to the non-lineadisplacement curve, Ref. [50]. For graphite/epoxy composite specimens the interfacial pentation do not have a significant effect on GIc values, Refs. [51], [61].

The simple beam theory has difficulties simulating the deformation near the crack More advanced models have been introduced by Olsson, Ref. [52], and Sun-Pandey, ReIn a comparison with a FEM solution the FEM results lie between the results obtained fromOlsson and Sun-Pandey models, see Fig. 8.

Figure 8 Comparison between Olsson’s model and classical beam theory.

In order to obtain consistent results when analysing experimental data it is necess

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Cracklength (mm)

Beam (0) 40 QDShow0 0 0 1 setcmykcolor/|______Times-Roman findfont 9 scalefont setfont44 -12 moveto(o) 5 QDShow0 0 0 1 setcmykcolor/|______Times-Roman findfont 12 scalefont setfont49 -16 moveto(/0) 9 QDShow0 0 0 1 setcmykcolor/|______Times-Roman findfont 9 scalefont setfont58 -12 moveto(o) 5 QDShow0 0 0 1 setcmykcolor/|______Times-Roman findfont 12 scalefont setfont63 -16 moveto()Olsson (0) 46 QDShow0 0 0 1 setcmykcolor/|______Times-Roman findfont 9 scalefont setfont50 -12 moveto(o) 5 QDShow0 0 0 1 setcmykcolor/|______Times-Roman findfont 12 scalefont setfont55 -16 moveto(/0) 9 QDShow0 0 0 1 setcmykcolor/|______Times-Roman findfont 9 scalefont setfont64 -12 moveto(o) 5 QDShow0 0 0 1 setcmykcolor/|______Times-Roman findfont 12 scalefont setfont69 -16 moveto()

Beam (90) 46 QDShow0 0 0 1 setcmykcolor/|______Times-Roman findfont 9 scalefont setfont50 -12 moveto(o) 5 QDShow0 0 0 1 setcmykcolor/|______Times-Roman findfont 12 scalefont setfont55 -16 moveto(/90) 15 QDShow0 0 0 1 setcmykcolor/|______Times-Roman findfont 9 scalefont setfont70 -12 moveto(o) 5 QDShow0 0 0 1 setcmykcolor/|______Times-Roman findfont 12 scalefont setfont75 -16 moveto() Olsson (90) 52 QDShow0 0 0 1 setcmykcolor/|______Times-Roman findfont 9 scalefont setfont56 -12 moveto(o) 5 QDShow0 0 0 1 setcmykcolor/|______Times-Roman findfont 12 scalefont setfont61 -16 moveto(/90) 15 QDShow0 0 0 1 setcmykcolor/|______Times-Roman findfont 9 scalefont setfont76 -12 moveto(o) 5 QDShow0 0 0 1 setcmykcolor/|______Times-Roman findfont 12 scalefont setfont81 -16 moveto()

Beam (45) 46 QDShow0 0 0 1 setcmykcolor/|______Times-Roman findfont 9 scalefont setfont50 -12 moveto(o) 5 QDShow0 0 0 1 setcmykcolor/|______Times-Roman findfont 12 scalefont setfont55 -16 moveto(/45) 15 QDShow0 0 0 1 setcmykcolor/|______Times-Roman findfont 9 scalefont setfont70 -12 moveto(o) 5 QDShow0 0 0 1 setcmykcolor/|______Times-Roman findfont 12 scalefont setfont75 -16 moveto()Olsson (45) 52 QDShow0 0 0 1 setcmykcolor/|______Times-Roman findfont 9 scalefont setfont56 -12 moveto(o) 5 QDShow0 0 0 1 setcmykcolor/|______Times-Roman findfont 12 scalefont setfont61 -16 moveto(/45) 15 QDShow0 0 0 1 setcmykcolor/|______Times-Roman findfont 9 scalefont setfont76 -12 moveto(o) 5 QDShow0 0 0 1 setcmykcolor/|______Times-Roman findfont 12 scalefont setfont81 -16 moveto()Olsson 450/450 Beam 450/450

Beam 900/900Olsson 900/900

Olsson 00/00 Beam 00/00

10

dueof the

ero or theith the

yhas

ate dis-e ratebeen

te dis-simen-, Ref.asingfreeis willith theinor

d infric-The

nd thatRus-ingation

HwuleaseininguphnessAS4/und in

with

consider end rotation and deflection of the crack tip, the effective shortening of the beamto large displacements of the arms, and the stiffening of the beam due to the presence blocks bonded to the specimens, Ref. [54].

The energy release rate along a straight crack will not be constant, but approaches zinfinity, depending on interface ply orientation, at the vertices where the crack intersectsfree surface. This will cause the crack to become curved and the angle the crack makes wsurface is determined by the interface ply orientation, Ref. [69]. A ratiohas been introduced, Refs. [55], [56], and found to correlate with the difference in energrelease rate at a specimen’s centre versus that at its edges, Ref. [57]. A ratioalso been introduced and found to correlate with the asymmetry in the energy release rtribution about the centre of the specimen’s width, Refs. [55], [56]. Since the energy releasis not evenly distributed along a straight crack the crack tip will become curved. This hasstudied analytically, Ref. [55], [57] and with FEM, Ref. [58].

For an ENF-specimen it has been observed numerically that the energy release ratribution is uneven with a straight delamination front, with maxima occurring at the pointwhere the delamination front intersects the free surface. To predict this behaviour a non-dsional ratioDc has been defined where conventional laminate plate theory notation is used[59]. The amount of mode III energy release rate for ENF specimens increases with increDc ratio, Ref. [72]. The vertex singularity at the points where the crack intersects with thesurface causes the energy release rate to approach infinity for a straight crack front. Thcause subcritical crack advance at the specimen edges which qualitatively correlates wDc ratio, Ref. [60]. During toughness testing of specimens with different interfaces only minfluence on interface type has been found, Ref. [61].

An analytical equation based on beam theory taking friction into account can be founRefs. [62, 63]. Mall et al., Ref. [64], have studied the influence on energy release rate fromtion, using a two-dimensional finite element model which assumed plain strain condition.energy release rate was computed with a virtual crack closure technique and it was fouthe friction effect is large for short crack lengths and negligible for longer crack lengths. sell, Ref. [65], estimated the effect of friction from the extent of hysteresis observed durloading and unloading of ENF specimens and friction was found to result in an overestimof the energy release rate of less than 2%.

The effect of fiber orientation at the interface for glass/epoxy have been studied byet al., Ref. [66], and no clear trend was found. It has been found that the critical energy rerate increases with a thicker matrix layer at the interface. Unidirectional specimens contaa resin rich interface of 50µm and 200µm showed an increase in delamination resistance ofto 200% when compared to control specimens of the same material. The increase in tougobserved was attributed to the relief of plastic constraint at the crack tip, Ref. [67]. For the3502 composite a damage zone, characterized by extensive microcracking, has been fofront of the delamination. The damage zone was at least 180µm long and extended about 5µmabove and below the plane of the delamination, Ref. [68].

The distribution of the total energy release rate along the crack tip for various layups,

DC D122 D11D22( )⁄=

Bt D16 D11⁄=

11

ke rate.ivatedith

gy vertexnevenfromr to.

the interface crack at the 0o/0o, 90o/90o, and 45o/45o interfaces are shown in Fig. 9. The craclength is 25 mm and the coefficient of friction is assumed to be zero. The energy releashas been normalized with the total energy release rate (GII+GIII ) at the centre of the specimensIn the figure the energy release rate for mode II and III have been added. This can be motby the similarities in fracture which might cause , Ref. [66]. The specimen wa 90o/90o interface has the flattest distribution followed by the specimen with a 0o/0o interfacewhereas the specimen with the 45o/45o interface has the most uneven distribution. The enerrelease rate approaches infinity as the free surfaces are approached. This is due to thesingularity located at those positions and has been described elsewhere, Ref. [69]. The udistribution of the energy release rate would suggest that subcritical crack growth, startingthe free edges, where the energy release rate approaches infinite values, will occur priounstable crack growth. This has also been observed experimentally, Refs. [70] and [71]

Figure 9 Distribution of G along the crack front for an ENF-specimen, different layupsconsidered.

GIIC GIIIC=

12

. Ford thisintro-f

ropic. Asason-creas-

n theith an

The classical data analysis methods are strictly valid only for homogenous beamsmultidirectional laminates the introduced error will depend on the stacking sequence anhas been discussed in detail in Ref. [72]. A more general expression for the compliance isduced with an alternative expression for GII, i.e. a modified beam theory, where the ratio, R, oflexural rigidities of the cracked and uncracked region is involved. For a homogenous isotmaterial R=8 but will alter for a multi-directional laminate depending on stacking sequencecan be seen in Fig. 10 the modified direct beam theory, according to that equation give reable predictions as compared to the FE-solutions. For a/L=0.5 the error is less than 5%. Ining the crack length to 35mm (a/L=0.7) good predictions are made except for the 45o/45o

interface where approximately the same error remains, i.e 5%.

Figure 10 Comparison FE-model with modified direct beam theory

In paper A and B results from thorough investigations, both numerical and experimental, oDCB- and ENF-specimens are discussed. The numerical investigation has been done whp-adaptive FE-analysis technique.

13

r real-g that

nsld beg on

ere isecific

con-te is yieldbased

4 Composite fatigue

4.1 Basic methodology

There is a strong need to be able to predict the fatigue life for a structural component undeistic loading conditions in aircraft design. These conditions imply general in-plane loadinwhere the load sequence produced will contain numerous loading conditions, R-values,must be handled in a convenient way in a design methodology, see Fig. 11.

Figure 11 Schematic of loading conditions.

Structural-verification testing with similar components but with different loading conditioand repair/reject criteria are areas where an improved methodology for life prediction woubeneficial. Composite fatigue, i.e life prediction, is not a straightforward matter dependinvarious failure modes and their interactions. Scanning the literature, it is obvious that thno complete, “ready to use”, model today. What exists are different approaches to solve spproblems, regarding e.g different R-values, layups, notched / unnotched behaviour etc.

One of the problems encountered with composite materials is that every new lay-upstitutes a new material, thus making the need for testing infinite. Besides from constraineffects, causing differences between unidirectional and multi-directional laminates, theralso a scale-up problem, i.e. going from coupons to a structural part will not necessarilythe same test results with respect to strength and fatigue life. The current philosophy is

INTERCEPTION

AIR COMBAT

LANDING

DESCENT

CLIMBCRUISERETURN

CRUISE

No of exceedences

Load

fact

or -

n z

14

tionxten-d cur-

tz et.[74].Ref.the

[76-d [81-asedl lam-uents theion. load-eralsug-

countehav-

ge)

on no damage growth and this is achieved by conservative strain limitations and verificatesting on components and built-up structures, i.e. basically a threshold formulation. Any esion of that philosophy must also take into account current static design methodology anrent NDT-techniques.

Pioneering work in the fatigue area of composite materials has been done by Schüal., Ref. [73]. A general overview directed towards aircraft structures can be found in Ref.For classic fatigue and especially on life prediction methodology an overview was made in[75]. Life prediction of CFRP (Carbon Fibre Reinforced Plastics) has been investigated inpast from different points of view. This would typically imply, degradation models, Refs. 80], concepts with damage modelling including residual strength concepts, Refs. [78] an83] or semi-empirical models, Refs. [73, 74, 84-91]. Also typical is that many models are bon micro-mechanical phenomena observed under unidirectional loading for unidirectionainates. There is, however, little hope that damage/fatigue models built on laminae constitwith associated micro-mechanical phenomena, e.g. micro-cracking, but not consideringmulti-directional nature of real structures would solve the problem of structural life predictThe nature of fracture is essentially ruled by, at least for in-plane loaded structures withcarrying fibres in all major directions, when essential fibre fracture starts to occur. A genoverview of the problem can be found in Refs. [11, 13, 47, 74, 98-104]. In Refs. [105-107]gested methodologies for constant-amplitude, block and variable loading can be found.

In paper D a methodology is suggested which in a simplified manner takes into acthe laminate behaviour. That is the lamina characteristics are deduced from the global biour, see Fig. 12.

Figure 12 Schematic determination of fatigue failure functions, constant (averastress/strain.

εx

εx

εx

15

a

ichionise

onvesti-f lowSim-ancean 50%nsivemero-ion of. It is thanA-18rposehechedld be

spec-

E. Ittesting. The 50%

nly ines andd byrmina-pli-

4.2 Truncation and elimination

Durability and damage tolerance criteria are and become more and more important. Fromcertification point of view there is a strict demand to validate the long-term behaviour ofcomposite structures during repeated loading. It is also absolutely necessary to know whstates of loads in the sequence that can be truncated or eliminated in structural verificattesting and under operational conditions. That is, which load states dominate/characterthe life of a structural component.

An extensive investigation was conducted by Badaliance and Hill, Ref. [108, 109], effects of fighter attack spectrum on composite fatigue life. Typical wing spectra were ingated, where e.g clipping to 90% test limit stress, addition of stress overloads, addition oloads, elimination to 70% test limit load and clipping of tension loads, were investigated.ple coupons with two different layups, zero- and matrix dominated were used. The importof retaining high loads in the spectra is emphasized, and cycles that have peaks less thof limit load can be eliminated since they do not contribute to the degradation. A compreheinvestigation has also been made by Shütz and Gerharz, Ref. [73], with basically the saresearch approach, i.e effect of truncation, overload and elimination on life. A typical zedominated layup was used with un-notched specimens. The conclusion was that eliminatlow loads, roughly 70% was eliminated, was the prime effect of investigated parametersalso well known that testing with basically un-notched specimens produce larger scatterspecimens with e.g an open hole, Ref. [110]. In the certification process for the Navy F/E/F upgraded model MacDonnel Douglas shortened their composite spectra with the puto reduce time and cost. Up to 60% of Test Limit Load (TLL) were eliminated, Ref. [111]. Teffect of elimination on a compression dominated spectrum, TWIST, was studied on notquasi-isotropic specimens by Phillips, Ref. [112], The results suggest that high loads shouretained and that it is likely to have large reductions in the low-load end when defining testtra.

A limited experimental investigation has been carried out and is discussed in paperwas demonstrated that it is possible to map calculated and tested results from spectrumon constant-amplitude data if a characteristic number of cycles were taken into accountexperimental results also indicated that the elimination level can be set to approximatelyof maximum range for in-plane loaded structures.

5 Residual strength

Structural durability, damage resistance and damage tolerance are important tasks, not othe design process but also when the aircraft is in operational use. There are many sourctypes of damage e.g: fatigue cracking, environmental degradation, or damage introduceforeign objects. When occurring, all types of damage need immediate attention for detetion of the effect on aircraft performance or functionality. There is therefore a need for simfied predictive methods for rapid assessment of occurring damage.

16

. Butlami-e andcan

ig. 14,cklingsting

We are still some way from being able to predict the true effects of service damageas pointed out in Ref. [113], if damage could be reduced to an idealised notch and/or denation then it might be possible to relate these to local stress fields in a realistic structurtherefore be able to predict structural failure loads. A schematic view of the actual problembe seen in Fig. 13.

Figure 13 Schematic flow chart for assessing in-service damage.

In paper F, three different ways to characterise impact damage are suggested, see Fand compared to two residual strength prediction methods. These are a delamination butheory and the so called “soft inclusion theory” and they are applied to the strength of exiimpact damaged panels. The importance of artificial delaminations are also discussed.

stress calculationDelaminationbuckling

Fatigue lifeprediction

Interlaminar

Damage criticality

Hole damagemethodology

Operationalconditions

Damage size

Structuralconfiguration

Diameter

Diameter

Impact

Hole

Delamination

Dπ4---ab=

Equivalent damagemodel

Damage Resistance

17

alentation,ivalentenous

ostf the

e platedimen-thate “doesspects

be

f ahetotal

Figure 14 Characterisation of damage, technique 1 and 2.

5.1 Equivalent damage methodology

A number of methods for different types of damage is suggested in Ref. [114]. An equivdamage size concept was introduced in Ref. [76], treating such different topics as delaminimpact and ordinary holes. The residual strength model was formulated based on the equdamage concept and a fracture criterion, i.e a model assuming macroscopically homogbehaviour.

Different approaches have been adopted to quantify the damaged area, which in mcases is the delaminated region due to impact. In Ref. [115] it is postulated that the ratio odamaged area to some reference area is proportional to the ratio of the compliances of thmaterial and the average inclusion material. The model uses experimentally measured sions of the damaged region to reduce the effective average stiffness of the material in region and to establish the size and shape of the affected area. The nature of the damagnot matter” except that it causes changes in the stiffness in the damaged region. General aregarding shape and size of individual delaminations in impact damaged structures canfound in e.g Refs. [116-118].

In Ref. [119] a model is presented for predicting the effective anisotropic modulii omicrocracked damaged material. By taking into account the energy dissipated through topening of microcracks the model can predict the effective response of the material. The

Observed damage

Equivalent damage

Each delamination have the samesize and shape

45o

2a

2b=a

2a

2b=0.76a

Relative angle:90o

Relative angle:45o

’Real’ delamination

Assumed delamination

T1

T2

(a)

(b)

18

e voidRef.eiontingective

lami-rentarger[123]hich

anged

pres-harac-. Theh-

nt ofible tomod-Fig.

strain is composed of the strain in the base material and the change in size and shape of thvolume (microcracks). An alternative is to use the micromehanical model introduced in [120] to study the effect of broken fibres, cracked matrix and fiber-matrix debonding. Thmodel is based on an integrated micro-mechanical and macromechanical fracture criter(IMMFC). Other micro-mechanical models can be found in e.g Ref. [121] where a repeacell philosophy is used. It is stated that the model can calculate elastic or non-elastic effmodulii of composites with or without including micro-mechanical damage.

A quantitative index to describe different damages due to low-velocity damage isdescribed in Ref. [122]. The relations between damage and modulus degradation of thenate, and between the damage and buckling of the laminate are preliminary given. Diffeimpact damages (low-velocity) are divided into four classes: slight, small, medium and lcorresponding to different damage parameters. A more direct approach was used in Ref.where the impact damage zone was varied from 0%, which is similar to a hole, to 100% wis no damage of the virgin material properties. Poisson’s ratio is assumed to remain unchfor the damaged area.

An analytical method has been suggested in Ref. [124] for the prediction of the comsive strength of composite laminates containing impact damage. The impact damage is cterised by ultrasonic C-scanning and the damage is modelled by an elliptical soft inclusiontechnique employed for stress and strain calculation is Lekhniskii’s complex variable tecnique. The stiffness of the inclusion is quantified by sublaminate buckling analysis.

Lekhnitskii’s complex variable technique is an elegant technique for rapid assessmestress and strain of notched composite structures, Ref. [16]. This technique makes it posscalculate not only notched plates but also plates with regions comprising different elasticuli, so called soft inclusions. Consider an infinite plate during general in-plane loading, see15.

Figure 15 General in-plane loading for a plate with a soft inclusion.

q

p

t

x

y x’y’

α

19

n cancatedfor

metricbasedork inUB-

ationiatedrgyte, G,y in theturedbleareacatedeory,

ment solu-thstem,e validf the

Using the complex mapping technique by Lekhnitskii, stresses around and in the inclusiobe obtained. As an alternative the delamination buckling theory, with more or less sophistiassumptions on blister displacement, or a combination of both, Ref. [124], can be used determination of residual strength.

5.2 Delamination buckling theory

Calculation methods that are cheap, fast, accurate and provides an opportunity for parastudies are most desirable from an engineering point of view. Semi-analytical methods on e.g a Rayleigh-Ritz approach are just that kind of methods. Based on the pioneering wRef. [125], the theory for delamination buckling was incorporated into a computer code (SLAM) for damage tolerance analysis in Ref. [126]. Loads and stacking sequence informare taken from the load distribution analysis. A conservatively assumed flaw size assocwith visible impact damage is selected. Through differentiation of the total potential enewith respect to the minor and major axes of the ellipse the total strain energy release racan be calculated and compared to experimental data. The authors also used the theordesign and development stage for quality control as a accept/reject criteria for manufacparts. In a recent aircraft program all composite parts were “zoned” for maximum allowadelamination sizes at ultimate load. Flaw depth was not specified in the criteria, only flawfor the sake of conservatism, inspectability and brevity. It was assumed that the flaw was loat the worst position (critical interface). Various aspects on the delamination buckling thcould be found in e.g. Refs. [125, 127, 128].

For the current approach an elliptical delamination is assumed to be present at sodepth, i.e oriented somewhere in the stacking sequence and an admissible displacemetion is assumed that fulfils the boundary conditions. Orientation and size are optional wirespect to loading directions. The problem is solved in a local coordinate system, x,y-syand transferred back to a user defined system, see Fig. 16. The solution is assumed to bif the so called thin film approach is present, i.e sublaminate is approximately one tenth ototal thickness.

20

ld be

age of

d.g. com-UR

.

nnereep-

loadificial orthick-

Figure 16 Problem statement.

An approximative way to calculate the strain-energy release rate for the delamination coufound in Ref. [125].

5.3 Artificial delaminations

When artificial flaws are considered it is of vital importance to know the extent of the dambut also to know where to locate the damage. What is lacking today is an understandingwhere the initial delamination(s) will occur, i.e a position criterion.

Artificial delaminations, located at the right place in structural verification testing ancoupled to a proper theory, could be an interesting way to characterise the strength of eimpact damaged structures or to demonstrate damage tolerance. An extensive work onpression loaded plates with artificial delaminations has been performed within the GARTEAction Group (AG16) on “Damage Propagation in Composite Structural Elements”, Ref[129].

Depending on size, shape and depth a delamination can first buckle in a stable mabefore it becomes critical or it can fail directly in an unstable manner which is the case for dlying and/or small delaminations. If subjected to successive loading, the critical bucklingwill decrease. A probable cause is that adhesive forces between base-laminate and artdelamination will relax. The question is if artificial delaminations reflect a real behaviournot. The problem is of course closely related to the on-going discussions of the size andness of the insert for DCB and ENF-specimens.

Xref

Yref

x

y

x’y’

βγ

NxNxy

Nxy

Nxy

Nxy

Nx

Ny

Ny

Base region

Delaminatedregion

Nx’b

Nx

Nx’d

-10-5

05

10

-10

-5

0

5

10-0.1

0

0.1

0.2

0.3

0.4

0.5

21

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er thebeam

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hey didof they-zero

everall finite-exist-

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SUMMARY OF APPENDED PAPERS

PAPER A

The static and cyclic fracture mechanical properties for the DCB-specimen are presentediscussed. Numerical and experimental results are compared to standard data collectioods but also to state-of-the-art finite element analysis (FEA). The investigation covers smaterials, IM7/8552, HTA7/6376 and T300/914, and different interfaces, 0o/0o, 45o/45o and90o/90o. It turned out that the material systems IM7/8552 and HTA7/6376 have similar critenergy release rate whereas T300/914 has a significantly lower critical energy release ranon-zero interfaces intralaminar growth occurred through the 90o or 45o plies, giving raise to ahardening R-curve behaviour and a wavy appearance of the fracture surface. Numericadicted curved crack front for zero-interface could be verified experimentally. The cyclic dmination growth rates increased rapidly with increased∆G. Comparison between specimenswith different layups showed that the specimens with a 90o/90o interface had the flattest energyrelease rate distribution along the crack tip and the specimens with a 45o/45o interface had themost uneven one. The calculated compliance compared well with the measured one aftcurved crack front had been accounted for. Energy release rate calculated with a modifiedmodel has a smaller error than that calculated with the simple beam model.

PAPER B

An extensive numerical and experimental investigation regarding fracture mechanical prties for ENF-specimens is presented. For different interfaces, 0o/0o, 45o/45o and 90o/90o andfor several graphite/epoxy composites, IM7/8552, HTA7/6376 and T300/914, static and cproperties are presented. State of the art numerical analyses are described and experimresults are interpreted in terms of usefulness for applications to real structures. The FEM rare compared with beam models and it is concluded that for the specimen with a zero intethe beam models were acceptable whereas for the specimens with non-zero interface tshow a large error. This could be compensated for by taking the non-homogenous naturebeams into account. Numerically, it was found that the effect of friction lowers the energrelease rate with a few percent. Experimental results showed that for specimens with noninterfaces the delamination crack changed interface.

PAPER C

Interlaminar stress calculations are presented for straight and curved free edges and for sstacking sequences. The interlaminar stresses are calculated using a three-dimensionaelement code with error control. For straight free edges the results are compared to someing semi-analytical models. A quadratic failure criterion for prediction of initial load for dmination and the most likely position for the delamination are also discussed. It can beconcluded that interlaminar stresses exhibit a thickness effect for repeating bundles and thdominant component for curved free edges are the high interlaminar shear stresses andnormal stress. The singularity of the interlaminar components is concentrated to a narroregion, approximately 1-2 fiber diameters in thickness and radial direction and it is concluthat since the interlaminar stresses are so localised, a continuum approach will not provid

22

ibleodels

sedlatedav-curvere in isnceiffer-ailurewhich

nec-times

ositectra

spec- thetaff”.pprox-sider-

alsoingtio oftigues elim-

atelye, i.e

d dela-mpos-ferentstresse con-

ful solutions in this area. The investigation also shows that the number of±45°/0° or ±45°/90°alterations should be kept to a minimum to minimise the interlaminar stresses and possdamage initiation points. Finally, it is demonstrated that the investigated semi-analytical mdo not correctly describe the interlaminar stress distribution.

PAPER D

A simplified methodology for composite life prediction is presented. The methodology is baon fatigue failure functions which are established on a laminate level and thus the calcuuniaxial properties, in this case with the Tsai-Hill failure criterion, will reflect the overall behiour of the laminate. To establish the fatigue failure functions, a documented/reported SN-for a multi-directional laminate is considered as a series of failure points where the failuthe short life-region essentially is assumed to be fibre fracture and in the long life-regionassumed to be controlled by matrix properties. For fibre failure it is also assumed that oessential fibre failure starts to occur the remaining time to total failure is very short. The dence between first and last ply failure, in a fatigue sense, is of less importance. Once ply fdata are established they are used in a semi-log representation to calculate the fatigue lifeis a straight forward matter since normalised, calculated, ply failure data collapse into ocurve. Different stress ratios (R=σmin/σmax) are presently handled with the Goodman-corretion approach. For combined loading a series of failure envelopes, representing differentin a log scale constitutes the behaviour.

PAPER E

An experimental investigation regarding load-sequence effects on the fatigue life of compstructures is presented. Different elimination levels were considered for various load speand for constant amplitude block loading. The spectra considered were an early designtrum for the Gripen fighter aircraft, spectra associated with the aft fitting of the fin and toupper fitting of the wing, and the compression dominated spectrum “Short Inverted FalsBoth experimental and calculated results show, that the elimination level can be set to aimately 50% of the maximum range occurring in the load sequence. This means that conable time and cost reductions in structural verification testing can be achieved. This willhave impact on a life prediction methodology since only a characteristic number of loadstates needs to be considered. The resulting life for block testing, especially for a load raR=-1, is highly influenced by elimination of high load ranges in the sequence. Spectrum fatest results accounting for ranges up to 90% of maximum range, the remaining load stateinated, can be mapped on constant amplitude data. At chosen elimination levels approxim80 to 90% of loading states were eliminated. Calculation results with the same techniquaccounting for a characteristic number of cycles, show promising results.

PAPER F

In this paper two residual strength models are presented, the so called soft inclusion anmination buckling theories, and compared to experimental results on impact damaged coite structures. Those experiments span a variety of impact events, from 8J to 55J and diflayups. The investigation has been supported by FE-technique for determination of the distribution in the buckled state and for characterisation of the damaged region. It can b

23

educ-trastls,th arti-ishat arti-

cluded that for low-energy impact, although with conservative assumptions on stiffness rtion, that the soft inclusion theory is un-conservative for residual strength prediction. In conthe delamination buckling theory shows good agreement for various impact energy levethicknesses and layups. The importance of repeated loading for composite structures wificial delaminations is also demonstrated. If repeatedly loaded, the local buckling strain strongly reduced as compared to a non-repeated loaded structure. It can also be shown tficial delaminations, although deep-lying, can grow in a stable manner if pre-buckled.

24

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DIVISION OF WORK BETWEEN AUTHORS

PAPER A: “A Numerical and Experimental Investigation of Delamination Behavioin the DCB-specimen”

by J. Schön, T. Nyman, A. Blom and H. Ansell

T. Nyman outlined the initial work, made the initial calculations and was also responsiblthe experimental investigation which was conducted in cooperation with BAe. J. Schön all the subsequent calculations and also most of the fractographic investigations. H. Ansmade the crack growth calculations and contributed to the damage tolerance aspects opaper. The paper was written by T. Nyman and J. Schön under the supervision of A. Blomalso helped with interpretation of the results.

PAPER B: “A Numerical and Experimental Investigation of a Composite ENF-spimen”

by J. Schön, T. Nyman, A. Blom and H. Ansell

Same division as in paper A.

PAPER C: “Interlaminar Stresses in Composite Notched and Unnotched Lamin

by T. Nyman and M. Friberg

T. Nyman outlined the work and made the initial calculations. M. Friberg made all the suquent calculations under the guidance of T. Nyman. The paper was written by T. Nymanthe assistance of M. Friberg.

PAPER E: “Effects of Truncation and Elimination on Composite Fatigue Life”

by T. Nyman, A. Blom and H. Ansell

T. Nyman initiated the experimental work. The theory was jointly developed by T. NymanBlom and H. Ansell. T. Nyman and A. Blom wrote the paper.

PAPER F: “Equivalent Damage and Residual Strength for Impact DamagedComposite Structures”

by T. Nyman. A. Bredberg and J. Schön

The basic theories were developed by T. Nyman although A. Bredberg introduced the dacharacterisation methods. The experimental work was conducted by J. Schön. T. Nymanthe first draft of the paper and finalized it together with J. Schön.

25

strial

RECOMMENDATIONS FOR FUTURE WORK

This work has dealt with some basic research topics of which some are of immediate induimportance. Below follow a number of points for possible research in the future.

o Different threats and the associated damage need to be investigated,especially in terms of size and position of the damage.

o The difference between structural and coupon level on residualstrength and life, i.e the scale-up effect, needs to be quantified to high-er degree.

o Truncation and elimination criteria must be further looked into, espe-cially the sorting principles for the load sequence leading to e.g. “con-stant-amplitude events”.

o Investigate the fatigue behaviour of impact damaged panels, "Fatigueafter Impact" (FAI), under realistic threats and loading conditions, in-cluding both constant-amplitude and spectrum loading

o Multiaxial static and fatigue loading of laminates with different cut-outs, impact damaged and bolted joints.

o Basic fracture mechanical properties, both static and cyclic, needs fur-ther penetration. Modus III testing should be increased and interactionformulas established. Threshold values are also of great interest.

o Simplified methods or forms for screening purposes, i.e engineeringtools with a relatively high accuracy, should be sought for residualstrength, life prediction and for interlaminar stress calculations.

o Environmental effects on static and fatigue loading should be furtherinvestigated.

26

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