i
Distortions and Residual Stresses of GLARE
Induced by Manufacturing
iii
Distortions and Residual Stresses of GLARE Induced by Manufacturing
Proefschrift
ter verkrijging van de graad van doctor aan de Technische Universiteit Delft,
op gezag van de Rector Magnificus prof.ir. K.C.A.M. Luyben; voorzitter van het College voor Promoties,
in het openbaar te verdedigen op Woensdag 24 Februari 2016 om 12:30 uur
door
Morteza ABOUHAMZEH
Master of Science in Mechanical Engineering Amirkabir University of Technology (Tehran Polytechnic)
geboren te Tehran, Iran
This dissertation has been approved by the promotor: Prof.dr. ir. R. Benedictus
Copromotor: Ir. J. Sinke
Composition of the doctoral committee:
Rector Magnificus Chairman
Prof. dr. ir. R. Benedictus Delft University of Technology Ir. J. Sinke Delft University of Technology
Independent members:
Prof. dr. ir. A.H. van den Boogaard University of Twente, Netherlands Prof. dr. A.J.M. Ferreira University of Porto, Portugal
Prof. dr. ir. M. van Tooren University of South Carolina, United States Prof. dr. C. Bisagni Delft University of Technology
Prof. dr. ir. R. Marissen Delft University of Technology
Prof. dr. ir. K.M.B. Jansen Delft University of Technology, reserve member
Keywords: Orthotropic, Fibre Metal Laminates, Residual Stresses, Distortions, Cure, Thermo-Viscoelastic
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TO:
MY SPIRITUAL TEACHER, DR. A. GAVAHI, WHO TEACHES ME HOW TO LIVE MY MOTHER’S SOUL AND MY FATHER
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ONTENTS
1 Introduction and background ... 1 Introduction ... 1 1.1
Overview of development and features of fibre metal laminates and their manufacturing . 2 1.2
History of Innovation & Applicability ... 2 1.3
Special Features ... 3 1.4
Manufacturing Concepts and Procedures ... 4 1.5
Modelling development strategy ... 7 1.6
Experimental strategy ... 8 1.7
Thesis Objectives ... 9 1.8
Outline of the thesis ... 10 1.9
2 Literature review ... 13 Introduction ... 13 2.1
Manufacturing processes on metals ... 13 2.2
Manufacturing of precise full composites ... 14 2.3
Design for manufacturing accurate skin panels made of FMLs ... 24 2.4
3 Primary modelling and experiments ... 31 Introduction ... 31 3.1 Modelling ... 32 3.2 Experimental ... 34 3.3
Results and Discussion ... 36 3.4
Conclusion ... 42 3.5
4 Investigation of Curing Effects on Distortion of Fibre Metal Laminates ... 47 Introduction ... 47 4.1
Determination of Bending-Strain-Free Temperature (TBSF) ... 49
4.2
Specimens ... 51 4.3
Experimental results ... 52 4.4
Discussion on development of curvature during the cure cycle ... 54 4.5
Curing effects and future work ... 57 4.6
Conclusion ... 57 4.7
5 Kinetic and Thermo-Viscoelastic Characterization of the Epoxy Adhesive in GLARE ... 61 Introduction ... 61 5.1
Thermo-dynamic analysis ... 62 5.2
Conclusion and future work ... 76 5.3
6 Analytical model for Residual Stresses And Warpage During Cure of Composite Laminates 79
Introduction ... 79 6.1
Theoretical formulation ... 80 6.2
Resin shrinkage during cure ... 86 6.3
Material Characterization ... 87 6.4
Cure characterization of GLARE ... 87 6.5
Numerical results and discussion ... 92 6.6
Conclusions ... 95 6.7
Appendix 6-A: Self-consistent micromechanics equations [13,14] ... 97 7 A New Procedure For Thermo-Viscoelastic Modelling of Composites With General Orthotropy and Geometry ... 101
Introduction ... 101 7.1 Theoretical background ... 103 7.2 Implementation procedure ... 108 7.3
Numerical results and verification ... 110 7.4
Conclusion and future work ... 114 7.5
8 Analysis and Discussion on the distortions in GLARE products ... 117 Introduction ... 117 8.1 Curing stresses ... 118 8.2 Cooling stresses ... 118 8.3
Response of S2-glass/FM94-epoxy prepreg ... 119 8.4
Results ... 122 8.5
Conclusion ... 128 8.6
Contents ix 9 Conclusions ... 131 Concluding remarks ... 131 9.1 Future works ... 132 9.2
Summary ... Error! Bookmark not defined. List of publications ... 137
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HAPTER
1
1 Introduction and background
Introduction
1.1
Fibre Metal Laminates (FMLs) are hybrid materials consisting of alternating metal and composite layers. FMLs have found applications in structural parts of aircraft like fuselage panels and leading edges of tail planes. Integrated fuselage panels made of Fibre metal laminates (FMLs) encounter shape deviations and residual stresses after cure and post-cure processes due to different mechanisms which are the main objectives of this study. In this chapter, first an introduction to the whole concept of design and manufacturing processes of FMLs and the relevant subjects are presented. After that, the methodology of research is introduced. All of the modelling, characterisation and measurement methods are described to investigate different manufacturing processes on the final distortions and residual stresses of FMLs. The proposed procedure should be followed in order to adapt the model to become capable of predicting the final geometry of a fuselage panel made of FMLs. The steps carried out in this research and the recommended steps for future work are clarified.
Although, the FML concept is further developed by researchers in the field [1-7], production of integrated panels made of FML has not been analysed and the elastic responses upon manufacturing processes are not yet studied in detail. Investigation of residual stresses and shape deviations induced by manufacturing on the geometry of glass fibre reinforced aluminium laminates (GLARE) and the methodology to aim for a predictive model for that is the main subject of this chapter. Manufacturing result in deviations from designed dimensions and these inaccuracies hamper the assembly. The residual stresses generated at the same time, reduce the material load capacity of the composite structure made from FMLs and can even cause premature failure.
Manufacturing processes on FMLs are described in this chapter and the factors influencing the final geometry and the stress state of the panels are discussed. In this chapter, an introduction is made of the whole research subject. In other words, the methodology and the necessary modelling and experimental approaches are described which are required to predict the manufacturing response of an integrated fuselage panel made of GLARE. The predictive model will eventually help to revise the mould to produce more accurate panels from GLARE. Further steps are presented to obtain a predictive model for residual stresses and distortions in a complete integrated fuselage panel made of GLARE. The full experimental measurements and modelling (simulations) needed are also discussed.
Overview of development and features of fibre metal laminates and
1.2
their manufacturing
First, an introduction to FMLs is presented, reviewing the history of its development, its (dis)advantages and its characteristics. The material is made from thin layers of metal sheet and prepreg, i.e. unidirectional fibres embedded in an adhesive. FMLs with aramid fibres are named ARALL and with glass fibres are named GLARE. The data presented in this section is mainly from [1].
History of Innovation & Applicability
1.3
A history of development of different types of FMLs (ARALL and GLARE) is given in [1, 2] as a unique story in the history of aircraft materials. To give a brief story of the development, it was in the late 1970’s that the material was developed at TU Delft and Fokker. Fokker had to be rebuilt after the second world war. In 1955 it presented its new aircraft, the Fokker F-27. One of its features was the extensive application of adhesive bonding. This was a useful method for manufacturing thick panels from thinner ones avoiding milling thick panels with expensive milling machines that were at the time unavailable for Fokker.
So it is not surprising that Fokker investigated the first Fiber Metal Laminates in the 1970s. Because of a lack of industrial potential (no suitable new aircraft under development), TU Delft, with academic curiosity, took over and in 1979 the first FML was developed by the group of prof. Vogelesang, His enthusiasm brought it to a success.
It was from 1978 until 1980 that Prof. Schijve together with prof. Vogelesang and Marissen presented the unique characteristic of FMLs, the fiber bridging, resulting in slow crack growth and high residual strength in ARALL and GLARE. The patent on FML by TU Delft was filed in 1982 by these three inventors. In the following years, the first commercial products named as ARALL-1 with AL-7075 aluminum alloy and ARALL-2 with Al-2024 as the metal layers were presented. In 1987, ARALL-3 with Al-7475 layers and ARALL-4 with a new-strong adhesive for military applications were developed. Main feature of ARALL-3 and -4 was that these were cross-ply laminates. However, since the splicing concept (see below) was not invented yet, manufacturing ARALL was found to be expensive. The Netherlands Agency for Aerospace Programs (NIVR) funded research on ARALL and in this respect F-27 wing panels were studied at TU Delft. For some (compressive) load cases, ARALL showed to be not suitable, because the fibre bridging failed. Therefore, the aramid fibres were replaced by glass fibres and a new FML was created: GLARE. In 1987, the patent of GLARE (GLAss REinforced) by AKZO company was filed and in 1991, AKZO and ALCOA worked together to commercialize the material. One of the early applications was a bulkhead, a primary structure of the Learjet 45 designed from GLARE in 1996. A joint venture of the aforementioned companies was founded in 1991 named as “Structural Laminates Company (SLC)”. SLC was engaged in design
Introduction
3
studies at Boeing and was studying to find a successor for Boeing 747 but did not have enough data to apply GLARE on this aircraft. Therefore they investigated the application on the Boeing 777. GLARE found to perform well under impact and as a result, in 1991, GLARE was used in the cargo floor in the Boeing 777, as the first commercial application. Application of GLARE in fuselages of Airbus aircraft were investigated in later years on A320-A330-A340 and A380. In the end, substantial parts of the A-380 fuselage were made from GLARE.
GLARE has better damage tolerance properties compared to aluminum including impact, corrosion, fatigue and residual strength. With these properties GLARE could be used in the following sections:
- Fuselage skin panels under fatigue conditions - Wing skin panels for fatigue and damage tolerance
- Stringers and frames for higher fatigue strength under uni-directional loads
- Floors in passenger and cargo areas: GLARE is better under impact compared to composites and are easier repaired.
- Fire walls made of GLARE present good performance against fire - Bulk heads of the airplane
- Cargo barriers that separate cargo from pilot area are stiffened sheets. Their life under impact can be improved if constructed from GLARE, since glass fibers after material yield can still carry load.
As can be seen, GLARE has applications in different parts of the fuselage and wings of Airbus airplanes and is continuously under development and evaluation in new airplane design projects. The largest application is AIRBUS A-380 fuselage, in which large skin panels are made of GLARE.
Special Features
1.4
In this section, properties of FMLs that made them superior materials for aerospace applications are:
‐ Fatigue behaviour
FML has a good fatigue life due to its low crack growth which is due to the fibers that remain intact. These fibers bypass the load over the cracked aluminum sheet(s), thereby reducing the stress intensity at the crack tip. Limited delamination between the metal and composite layer is a prerequisite for this mechanism. The adhesive under fatigue shear loading separates from metal layer. This helps the fibers that have limited failure strain to stretch so they do not break. On the other hand if delamination is too large, the crack opens and grows faster. Therefore, there is a balance between crack growth and delamination. Another item to mention is that in monolithic aluminium panels, the crack initiation period is more or less the most dominant factor of the fatigue life since crack propagation is so fast and the panel fails in a short time; but in FMLs, crack initiation occurs in the
aluminium but the crack propagates over a much longer period. Therefore, FMLs have better fatigue properties than aluminium but similar to full composites.
‐ Residual Strength (tensile)
Residual strength can be defined as the remaining static strength of a material in the presence of a damage. In GLARE, fatigue cracks seldom lead to through the thickness cracks. A common kind of damage in GLARE is a crack in the aluminium layers while the fibres remain intact. Since the fibres remain intact, the reduction in static strength is limited. The same is true in case of other damages like dents or delamination due to impact. In all these cases, the residual strength is still very high when compared to metal structures. .
‐ Impact
FMLs are superior to both metal sheets (membrane stresses) and full composites. Application in airplane bulk cargo bay floor is because of this property and also leading edges of horizontal and vertical tail planes.
‐ Resistance to environmental influences
The metal sheets “protect” the composites for the moisture absorption, the UV-degradation, etc. and the composite layers limit the effect of corrosion damages to one layer. Therefore, fiber layers in GLARE can inhibit corrosion through the thickness.
‐ Fire resistance
Delamination and carbonization of the composite layers has a significant fire stopping capacity. This was shown to bring additional benefit in GLARE firewall tests. The heat from fire makes the thicker laminates delaminate which creates an insulation. The inside metal layers are not melted so the inner laminates will be in lower temperatures. Therefore, we can say that fire resistance of aluminum is poor, for glass composites, it is better and for GLARE, it outperforms its constituents.
Manufacturing Concepts and Procedures
1.5
In this section, manufacturing procedures of FMLs for airplane fuselage skin panels are briefly discussed:
1.5.1
Layup and curePrepreg and metal layers are laid up together on a layup tool that should result in the desired shape of the panel and the layup is cured in an autoclave process to have a solid laminate.
1.5.2 Splice concept and doublers
The splice concept was introduced in the early 1990s, to reduce the high cost of manufacturing large scale laminates. In this concept, metal layer overlaps plus adhesive films are used to overcome the limited width of the metal sheets. As a result, wider and lighter laminates with lower costs can be
Introduction
5
manufactured by making bigger panels, less joints are required which further reduces costs and weight. Different doubler configurations can be designed. Doublers can be either internal or external. Because of aerodynamic reasons the external ones are not used for the outside of skin panels. In locations where the number of plies in the laminate changes (ply drop-off), single layer (internal or inter-laminar) doublers are utilized.
1.5.3
Self‐Forming Technique (SFT)The Self-Forming Technique (SFT) is used in producing laminates under pressure (6-11 bar) in an autoclave cure cycle. Therefore, forming the laminate, adding panels with splices and attachment of small doublers are all performed in a single cycle. Additional adhesive is applied to fill the gaps which arise at the edges of metal layers resulting in higher shear and delamination strength of the laminate so they are not sensitive to shear and delamination under static or fatigue loads.
(Thick) doublers and stringers are bonded to the skin panel in a second cure cycle. In regions with cut-outs like surroundings for windows and doors, doubler packages are used. Fibre direction are aligned with the loads; for example, in doors, also 45-degree fibre orientations are applied.
1.5.4 Production requirements
Some limitations are encountered in manufacturing the designed FML including: number of thickness steps in the laminate, joining method of stringers and thick doublers, number of aluminium sheets, number of splices, etc. FML has some limitations and specific features in production process that the designer should keep in mind:
‐ Machinability
Machining methods are drilling, milling, water-jet cutting and laser-jet cutting. Forces and heat generated during machining may cause delamination of the laminate and also fracture and wear of machining tool. Therefore the feed rate and the temperature (in cool milling) should be controled.
‐ Formability
Formability of FML is limited. Bending as the most common process, can cause failure of constituents. This can occur because of low failure strain of fibres or epoxy resin. Coherence of the laminate may also fail due to inter-laminar shear stresses. Stretch forming is not suitable for GLARE because of the small limit strain and also of its large spring-back. Therefore, double curved shells and stringers are manufactured by laying up separate layers or thin laminates and not by stretch forming.
1.5.5 Manufacturing procedure and phases
Different processes may influence the dimensional accuracy of a fuselage skin panel made of FMLs which are illustrated in Figure 1-1.
Figure 1-1: Effects of different manufacturing phases on residual stress and distortion of FMLs
1.5.6 Effects of making cut‐outs in panels made of FMLs
According to Figure 1-1, after cure and removal from mould, the panel is trimmed and cut-outs are made in the FML panel (see Figure 1-2) using methods like water-jet and milling. This may increase or re-distribute residual stresses and induce additional distortions which needs mathematical simulation with finite element method to predict them. Experimental measurements should be used for validation.
Figure 1-2: A skin panel including cut-outs and reinforcements (doublers and bonded stringers) 1.5.7 Effects of adding non‐symmetry by splices, doublers and stringers
The Splice joints and the reinforcements bring local non-symmetry that produce residual stresses and distortions. In the splicing concept, metal layer overlap plus adhesive films are used to overcome the limited width of the metal sheets to make larger panels (Figure 1-3). Splicing has acceptable tolerances and minimum consequences for the local thickness [8]. Skin panels may be flat, single or
Introduction
7
double curved shells and can be locally reinforced with doublers to reduce stress concentrations. Stringers can also be bonded for increasing the out of plane stiffness and the stability of skin panels against buckling under compressive loads (see Figure 1-2).
Figure 1-3: The overlap splice geometry in a GLARE laminate [8]
Modelling development strategy
1.6
As already mentioned, even properly designed panels do show some distortions and have residual stresses. The ultimate goal of the research, is to have a model to simulate the laminate responses to all manufacturing steps for an integrated fuselage panel made of FML. To achieve this, a modelling strategy is used, as shown in Figure 1-4. As the first step, non-symmetrical FMLs are analysed with
thermo-elastic modelling of the cool-down process, taking into account the panels’ large deflections
through a geometrically nonlinear analysis (model 1). The model is improved by approximation of the
non-thermoelastic effects (model 2). Further improvement of the model accuracy can be obtained by
considering cure-dependent material properties and by simulation of the whole cure process including chemical shrinkage and stiffness increase during curing (model 3). Temperature dependency and the
viscoelastic behaviour of the epoxy will also be added to the model (model 4), for which a complete
material characterization procedure is needed for this purpose. Adding models 3 and 4 leads to a model capable of considering time-cure-temperature dependent properties of GLARE (model 5). Each model is actually an upgrade of the previous model. Furthermore, features like splices, doublers, ply drop-offs and cut-outs are supposed to be investigated that may include some plastic deformation in the aluminium layers. Post-cure processes [1, 8] are also required for the attachment of stringers and large doublers.
Figure 1-4: Modelling development
Experimental strategy
1.7
Experimental investigations and measurements are used either to validate the model or to identify the important factors for modelling. Measurements are also needed to model the material behaviour.
1.7.1 Measurement of distortions and residual stresses
Distortions and/or residual strains can be measured for studying the elastic response of different variants of FMLs during manufacturing and second for model verifications based upon the following approaches:
Approach 1: Direct/indirect measurement of cured samples
The following measurement techniques can be applied on cured FMLs to find the final residual stresses in the laminate:
1) Deflection of a FML strip can be measured using methods like digital image correlation (DIC). Subsequently, residual strain and stresses can be calculated from cylindrical bending equations.
2) Direct measurement of residual strains of cured laminates can be done using different methods previously applied to full composites and/or metals including:
‐ Hole drilling method: mainly for metals that can be used for outer aluminium sheet of FMLs
‐ Incremental slitting method: can be used for both metal and prepreg layers of FMLs
‐ Layer removal (Peel-ply) by chemical etching: a convenient method for measuring strain through the thickness of FMLs
Approach 2: Cure monitoring/measurement during cure:
1) Simultaneous measurement of temperature (using embedded thermocouples) and curvature of panels during cure using DIC or a displacement probe
Model 1: Thermo‐elastic (cooling) (Small/large Displacement) Model 2: Thermo‐elastic with Curing Effects Model 3: Cure‐dependent thermo‐elastic Model 4: Cure‐independent thermo‐viscoelastic Model 5: Cure‐dependent thermo‐viscoelastic
Introduction
9
2) Simultaneous measurement of temperature and residual strains using embedded strain gages or fibre optic sensors
1.7.2 Material Characterization
For the modelling of the complete cure process as the current research aims for, the following measurements are needed:
1.7.2.1 Thermal‐mechanical properties of the cured material
For improved modelling of the cool-down, change of material properties with respect to temperature should be considered, since the epoxy resin properties are dependent on temperature. Change of the coefficient of thermal expansion (CTE) of the epoxy or prepreg layer with temperature can be measured using Thermo-Mechanical Analysis (TMA). Glass-transition temperature (Tg) can
also be obtained from TMA results. The temperature dependent stiffness of the prepreg layer can be measured using Dynamic Mechanical Analysis (DMA). If the measurements are also dependent on time, viscoelastic properties of cured epoxy can be derived from DMA and used for modelling the behavior of the material’s dependency on temperature and time.
The phenomena occurring in the epoxy cure cycle are measured on the cured epoxy:
- Change of thermal expansion coefficient (CTE) with temperature using Thermo-Mechanical Analysis (TMA)
- Measuring the time-temperature dependent stiffness using Dynamic-Mechanical Analysis (DMA) on the cured epoxy adhesive
1.7.2.2 Cure kinetics and propertied during cure
Cure kinetics is modelling the completion of cure of a thermosetting resin with respect to applied temperature and time. Differential Scanning Calorimetry (DSC) is used to measure the heat flow into and from the sample during the cure cycle. Onset and completion of cure, degree of cure and glass-transition temperature (Tg), all can be obtained from DSC. Furthermore, cure dependent material
properties and viscoelastic effects during cure can be derived using DMA during cure of pure epoxy or prepreg.
Following measurements would be carried out on the curing epoxy:
- Cure kinetics of the epoxy adhesive using Differential Scanning Calorimetry (DSC) - Cure shrinkage of epoxy/prepreg
- Cure-dependent material properties of prepreg/epoxy using DMA for determination of E-modulus dependent on temperature, time and degree of cure
Thesis Objectives
1.8
The output of this research is aimed on two different aspects. First, for the scientific content, the impact of the manufacturing processes on the elastic responses of FML materials will be investigated.
Second, the output from the scientific part of the research will be used to build a model capable of predicting the effects and final state of an integrated FML skin panel in the design phase. In other words, final stress state and dimensions with a proper tolerance can be predicted by the designer and this makes production of the aircraft skin panels more accurate.
To follow the mentioned strategy, software simulations, analytic modelling, finite element solutions and experiments are needed to characterize and model each process.
Outline of the thesis
1.9
With the aforementioned procedure for the research objectives, the following chapters of this thesis are presented:
After this introductory (chapter 1), chapter 2 is dedicated to the study of the state of the art of the subject. Related research areas are discussed and previous works on full composites, metals and FMLs are reviewed.
In chapter 3, in order to have primary understanding of the distortions and residual stresses after cure, a simple thermo-elastic model is developed and implemented on simple (non-featured) FMLs.
The contribution of different parts of the cure cycle on the residual stress build-up, is investigated in chapter 4.
Material characterization is performed on the adhesive used in GLARE. Cure kinetics, thermal and viscoelastic properties of epoxy FM-94 are measured and the results are explained in chapter 5.
In chapter 6, an analytical cure model is developed including the cure development and the chemical shrinkage of the adhesive for any orthotropic material capable of predicting the distortions and residual stresses after curing of the laminate.
Chapter 7 describes the developed model for thermo-viscoelastic analysis of orthotropic composites with no limitation on orthotropic relaxations and geometry.
The latter model together with the curing stresses obtained using the cure model, is used to get the final deflection and residual stresses of GLARE panels after the cure cycle (chapter 8). The developed models can be easily extended to any other orthotropic material like full composites.
Finally, in chapter 9, concluding remarks and the possible future work are presented.
Introduction
11
References
1. Ad Vlot WG. Fibre metal laminates: An introduction. The Netherlands: Kluwer Academic Publishers.
2. Vermeeren C. An historic overview of the development of fibre metal laminates. Applied Composite Materials. 2003; 10:189-205.
3. Asundi A, Choi AYN. Fiber metal laminates: An advanced material for future aircraft. Journal of Materials Processing Technology. 1997; 63:384-94.
4. Alderliesten RC. Fatigue crack propagation and delamination growth in GLARE. The netherlands: Delft university of technology, 2005.
5. Homan JJ. Fatigue initiation in fibre metal laminates. International Journal of Fatigue. 2006; 28:366-74.
6. Sinke J. Development of fibre metal laminates: Concurrent multi-scale modeling and testing. Journal of Materials Science. 2006; 41:6777-88.
7. Alderliesten RC, Benedictus R. Fiber/metal composite technology for future primary aircraft structures. Journal of Aircraft. 2008; 45:1182-9.
8. Sinke J. Manufacturing of GLARE parts and structures. Applied Composite Materials. 2003; 10:293-305.
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HAPTER
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2 Literature review
Abstract
This chapter reviews relevant literature concerning the development of residual stresses and shape distortions during manufacturing of laminated materials including full composite and FMLs. Different causes that influence the shape and dimensions of laminated panels are recognized and the previous works done on the modelling and predicting the effects are reviewed.
Introduction
2.1
Few works are performed mainly on the topic of manufacturing-induced distortions and residual stresses in FMLs. From the study of the literature, however, one can argue that some research performed on metals and especially on full composites can be extended to FMLs. In this way, same modelling approaches can be assumed for common responses between full composite and FMLs. Although, specific considerations should be added for FMLs.
In this chapter, causes of residual stress and dimensional deviations in production of composite laminated panels including full composites and fibre metal laminates are discussed and reviewed. Based on this review study, one can find the lacking knowledge on FMLs and the needed solution procedure to improve knowledge on this.
Manufacturing processes on metals
2.2
Since metal layers exist in FML, first a review on compensation of dimensional changes of metal sheets due to manufacturing is performed.
2.2.1 Sheet metal forming processes on metals
Since FMLs have metal layers, the metal sheet forming processes should be considered as they react the same in the cold forming of FMLs. It should be noted that these forming techniques induce both elastic and plastic deformations and that the elastic energy restored in the material is partly released as spring-back and partly remains in the material as residual stress. In the curing of simple curved shells from FML, no plastic strain is encountered. However, in forming FML stringers as reinforcement of skins, bending is required that takes the aluminum layers of FML into the plastic
region. The conventional processes in manufacturing of metallic shapes include bending, stretching and deep drawing.
2.2.2 Dimensional instabilities (spring‐back)
In forming metal shapes, elastic and plastic deformations are present. The product springs back due to the elastic recovery when the loads are removed. The shape deviation may be large and a compensation may be needed.
Research has already been performed to produce metallic parts with less shape deviations, some of which are presented here to exemplify the used concepts.
When the metallic part is bent, the part angle increases and deviates from the bending angle during the process. Without predicting the distortions, corrections are made usually based on experiments that takes a long time and are also costly. The shape after bending and spring-back of the sheet metal is taken into consideration in the bending die design [1]. As a result, the spring-back angles of different sheet metals with different bending angles were obtained. The tool geometry is optimised in [2] to compensate spring-back within the deep drawing process using FE analyses. For more detail on the calculation and compensation of spring-back in metal forming processes, one can refer to other research results available [1-6].
Manufacturing of precise full composites
2.3
2.3.1 Introduction
Composite structures are used in aerospace and automotive applications due to their high strength and stiffness to weight ratios. The manufacturing processes consist of impregnation of fibres and resin in the laminae, layup of the laminae into laminates and finally curing of the resin.
The exact process of manufacturing of composite materials depends on the type of its constituents and more specifically the type of polymer used as the matrix. Matrix polymers can be generally divided to thermosets and thermoplastics. Thermoset polymers are cured at elevated temperatures with crosslinking of the polymer. They are more brittle than thermoplastics and not able to be thermally recycled, so they are cured and shaped to their final geometry. On the other hand, thermoplastics can be reheated, melted and reshaped and have higher viscosity. Thermoplastics can be welded and their toughness, storage life, and chemical resistance is rather high. As far as thermoset polymers are used in prepreg layers of FML’s, this type of composite will be considered in more detail in the literature review of the curing-induced phenomena.
In the following sections, different aspects of the manufacturing process of composite laminates are discussed and the research results in the literature are reviewed.
Literature review
15
2.3.2 Manufacturing‐induced phenomena in composites
From so many years ago (1970’s), efforts are made to study the process of curing and manufacturing composite materials [7,8].
Autoclave processing, using pre-impregnated uni-axial prepreg, is a common manufacturing method to produce high performance fibre reinforced composite laminates. Generally speaking, the cure cycle is divided into three parts: heating to the cure temperature (above the gelation temperature), curing isothermally (at constant cure temperature) and finally cool-down to the ambient temperature. Note that the isothermal cure part can be done in one or two stages (White & Hahn, 1992 [9]). In a two-step cure cycle, the material is held at the dwell temperature for a dwell period of about 1 hour. The temperature is increased to the second temperature level and held constant for the cure period (2-8 hours). The purpose of the dwell period is to allow gases (entrapped air, water, or volatiles) to escape the matrix material and to allow the matrix to flow, facilitating compaction of the part. Thus, the viscosity must be low during the dwell. Typical viscosity versus temperature profiles of polymer matrices show that as the temperature is increased, the viscosity of the polymer decreases until a minimum viscosity is reached. As the temperature is increased further, the polymer begins to cure rapidly and the viscosity increases rapidly too. The first dwell temperature must be chosen judiciously to allow the viscosity of the resin to be low while keeping the cure to a minimum. Isothermal viscosity versus time profiles are useful in determining the pot life of the polymer: the maximum length of time at a specific temperature for the polymer to maintain a specified viscosity for handling of the resin. A certain minimum temperature must be reached before the crosslinking reaction begins. It is here that the strength and related mechanical properties of the composite are developed. Demands for increased performance have recently led to the development of several temperature resins. These high-temperature resins retain good mechanical properties at elevated service high-temperatures. One of the problems encountered when processing at higher temperatures is the increased residual stresses.
Some portions of the produced residual stresses are released and create shape deviations (distortions) and some remain as residual stresses. In order to get the part within the preset dimensional tolerances, geometrical compensation of the tool is necessary. Curing-induced stresses reduce the load capacity and the fatigue life of the product. The shape deviations may result in extra assembly forces producing internal forces and stresses.
With the prediction of these responses in the design phase, automating the manufacturing process would be possible and this will be a great opportunity to the industry. In this section, different mechanisms and factors leading to residual stress and dimensional changes in composite panels are discussed and the related literature is reviewed.
2.3.3 Types and sources of distortion and residual stress
The distortions can be from different mechanisms that can be thermoelastic (reversible) or non-thermoelastic (non-reversible). Radford & Rennick in 2000 [10] and Wisnom in 2006 [11] have well defined these mechanisms which can be explained and separated as follows.
In this research, the deviations from the designed shape are called distortions. Distortions can be of several types: spring-in, spring-back and warpage. Spring-in and spring-back occur in curved composite laminates as a change in the enclosed angle of the part due to elastic response of the laminate. When the enclosed angle decreases, it is called spring-in and when it is increased, it is called spring-back. Another type of distortion occurs in flat and balanced laminates. Warpage is mainly because of the mismatch in thermal properties of the mould (tool) and part (laminate).
2.3.3.1 Anisotropy
Anisotropy or different directional properties in orthotropic materials plays a role in the following phenomena:
Chemical shrinkage (matrix cure shrinkage during polymerisation)
Shape distortions and residual stresses during manufacturing of composite parts are closely related to the development of cure in the resin. During processing of thermoset composites, the resin transforms from a viscous fluid of monomers first to a rubbery and then to a cross linked network. During this network formation, the free space occupied by the polymer molecules reduces and this causes a chemical shrinkage which is usually referred to as cure shrinkage. The chemical reactions involved in the cure of a resin result in the generation of heat.
The extent of the cure reactions is described by the degree of cure, which is quantified as the fraction of heat generated to that point relative to the total heat generated through the complete cure. There are some researches that have assumed and observed a linear relation between the volume change due to chemical shrinkage and degree of cure (q)[12-15]. The two latter papers model the cure process and measure the evolution of shape and volumetric changes caused by chemical and thermal shrinkage. In the foregoing sections different works relating to the cure modelling of composite are reviewed.
In a paper from G. Kelly et Al. in 1996 [16], the effect of chemical shrinkage on the residual stress and warpage of moulding compounds used in plastic encapsulated integrated circuit packages are determined to be as high as 70%. In this way, they have concluded that convenient predictive models that account only for the thermal source in cool-down, truly underestimate the results.
Thermal contraction (during cool-down)
During cool down process in composite laminates, contraction or shrinkage occurs in all the constituents of the material. Since they (fibres and matrices) have different CTE and stiffness E, a
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mismatch causes distortions in the laminate from the initial shape and residual stresses. This is the most dominant mechanism in production of laminated structures [7,8].
2.3.3.2 Fibre Volume Fraction (Material Property) Gradient
Volume fraction gradient arises as a result of bleeding the resin from one side of the part, i.e. the vacuum bag side. Radford & Rennick in 2000 [10] and Darrow and Smith in 2002[17] have considered this phenomenon. Therefore this effect is important when resin bleed is present in the manufacturing of a composite laminate. Since in using prepreg materials, no bleed process is present [18], distortion due to volume fraction gradients through the thickness of prepreg layers and also in FMLs can be ignored.
2.3.3.3 Tool‐Part Interaction (Stress Gradient)
Even in thin, flat and balanced (symmetric) laminates that are manufactured on a flat tool (mould), a convex warpage would occur after cure if the mould material has a different CTE from the composite (Ersoy et Al., 2005 [19]). During processing, as the prepreg is heated under pressure, the fibres closest to the mould (which are held against the mould surface by the processing pressure) can be stretched by the mould. The mould stretched layer develops residual tensile stress when the part cools down after cure. When the part is removed from the mould, the strain is equilibrated across the laminate thickness and the part warps concave down. Besides CTE mismatch, the degree of mould stretching can be affected by the mould surface roughness, cure temperature and the applied pressure during cure (Kappet et Al.,2011[20]).
Under the vacuum pressure, shear stresses are generated at the mould-laminate interface during various phases of the cure cycle. Cho et Al.[21] in 1998 have studied different factors like thermo-mechanical properties of tool-laminate, influencing the distortion shapes. Twigg et Al. in 2003[22] investigated the shear stress development at the interface between the tool and laminate. Complete cure cycle is evaluated in this respect with eight strain gauges placed on the aluminum tool, six oriented longitudinally and two transversely. Both sticking and sliding conditions occur in the cure cycle and the shear stress at which the parts slides (τsliding) is dependent on the degree of cure and also
the pressure. So the cure cycle parameters can influence the amount of residual stress and the final shape. The same authors in 2004 [23], carried out an experimental survey on the effects of part aspect ratio and processing conditions. Warpage was observed to be affected mostly by cure pressure and the length and thickness of the laminate. In this regard, tool surface condition had neither significant effect nor was it predictable. In a companion paper [24], they studied the effects of different parameters in the interaction, numerically. They concluded that both part-tool shear stress and in-plane stress distribution in the part (laminate) are important to consider. A similar work has been done by Ersoy et Al. in 2005[19], in which the frictional shear stresses between prepreg layers and also prepreg-tool interface is measured experimentally as a function of degree of cure. As a numerical finite element
work, Zeng & Raghavan in 2010[25] used ABAQUS subroutines to develop a process model to numerically study this mechanism of distortion. Kappel et Al. in 2011 [20] have presented a finite element simulation and measured the warpage of some test specimens, made from two common prepreg systems, as a function of thickness.
2.3.4 Researches on cure‐cycle induced residual stresses
In this section, the works investigating the stress build-up during cure and cool-down are reviewed. As an early and highly cited work in the literature, Hahn & Pagano in 1975 [7] have simulated the cure process. They have assumed that at the start of the cool-down stage, the material is in the stress-free state. They claim that the residual stresses due to chemical shrinkage of the matrix will disappear due to the viscoelastic relaxation. Therefore, they state that after this stage, no considerable residual stress remains in the matrix so the stage of cure is not needed to be considered. They have used incremental constitutive equations in conjunction with temperature dependent (nonlinear) material properties. A method of curing stress analysis based on total thermal strains was formulated for resin matrix composites. The method decomposes the total strain into mechanical and thermal parts. Temperature dependent elastic behavior was assumed, i.e., elastic compliances and thermal strains were allowed to depend on temperature.
H.T. Hahn continued his work in 1976 [8] using a linear elastic approach for the prediction of residual strains from fabrication and how moisture contributes to the residual stresses and compared these with the experimental data. He stated, in contrast to their claim in their paper in 1975 [7], that the stress free temperature is lower than the cure one.
In 1979, Weitsman [26] calculated the thermal residual stresses within cooling from the cure temperature. He accounted for temperature dependence of properties and the viscoelastic response. Comparisons with linear elasticity indicated that viscoelastic relaxation may reduce the residual stresses by about 20 percent. So only the cool-down phase was modeled and the solution was linear viscoelastic.
In 1988, Favre [27] presented the first review of works done before that on residual stress development in composites for both thermoplastic and thermoset resin types.
In 1989, Kim & Hahn[28] studied the process in which residual stresses develop during processing of a thermoset (graphite/epoxy) composite. The laminate deflection was monitored during intermittent curing of a non-symmetric cross-ply laminate. Mechanical properties were also measured as functions of the cure time. They state that for residual stresses after complete cure, linear elastic predictions are adequate provided that the change in the matrix modulus is accounted for in the analysis. They have investigated the complete cure cycle and observed that residual stress and warpage are important after the gel point. Therefore, we can see that the stress free temperature (which in their case was near the cure temperature) has to be known to calculate the warpage and residual stresses either when an elastic
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(linear) or viscoelastic solution is considered. Cure was done within 4 hours in a vacuum oven at 177°C. The temperature required to make the panel flat was assumed to be the stress-free temperature. The same approach for determination of stress-free temperature is used by Crasto and Kim in 1993[29].
In a proceeding work in 1990 by White & Hahn[30], mechanical property changes and development of residual stresses during cure were investigated. Thus, they have considered only the cool-down stage and the strains due to chemical shrinkage were small and negligible, regarding their specific processing conditions.
Bogetti & Gillepsie simulated the cure process of thick-section composites for the distribution of temperature and degree of cure as a function of the autoclave temperature history with two-dimensional modelling in 1991[31] and with a one two-dimensional model in 1992 [12]. They compared their results with the similar approach from Kim & Hahn, 1989 [28]. They investigated the effect of cure parameters. Cure shrinkage, resin modulus, and composite mechanical properties are assumed as cure dependent. Effects including thickness, resin modulus development, cure cycle, cure shrinkage, non-symmetric curing (about the mid-plane of the laminate), stacking sequence, amount of the stress-free temperature and the influence of resin shrinkage on the assumed stress-stress-free temperature were all investigated in their detailed study. Indeed, since the evolution of stress throughout the curing process is predicted, no stress-free temperature assumption is required. The results demonstrate the significant influence that the resin shrinkage can have on the assumed stress-free temperature, and thus the magnitude of the resulting residual stress distributions.
White & Hahn in 1992 [32], modelled the complete cure process with linear viscoelastic formulation and concluded that the material dependence on temperature should be considered and chemical shrinkage has small contribution to the residual stresses if relaxation time in the second phase of cure is enough. In their cure cycle, the contribution of chemical shrinkage was lower than 4% but they declare that this part cannot always be neglected. If a fast cool-down occurs during or shortly after chemical shrinkage, a considerable part of residual stress due to chemical shrinkage cannot be relaxed and would remain in the matrix. They state as a result that a residual stress process model should incorporate viscoelastic material response, chemical and thermal shrinkage effects, and mechanical property development during cure. They investigated the mathematical modeling with experiments in the same year[32]. Thermal and chemical strains were also measured by strain gauges and used as required inputs for the viscoelastic model.
Some researchers have also discussed internal stresses and corresponding distortions developed during curing and investigated the effects of pre-stressing on the increase of load capacity of laminates [33].
In 1995, Wang et Al.[34] measured residual stresses, including a curing shrinkage stress and a cooling shrinkage stress, automatically and continuously during curing and cooling. One important
note we can extract from their work is that if the cure temperature is low and the test time is shorter than the apparent gelation time, no residual stress will remain in the matrix. They state that in the first stages of cure, from ambient temperature to cure temperature and until the middle of the cure stage, the degree of cure in the epoxy resin is low and therefore the residual stresses due to cure shrinkage don’t remain and will be relaxed. However, after gelation, the stresses are locked (due to the low rate of stress relaxation) and the gelation point is the start of creation of residual stresses. From tgel up to
the start of cool-down, some cure shrinkage stress will be created which are much smaller than the cooling shrinkage stresses. The ratio to the total residual stresses is about 5%. In their work, a finite strip is used and the stresses are calculated from Timoshenko beam theory based on linear elastic behaviour.
Stress build up process for thermosets is also dependent on the cure temperature. In 1997, J. Lang et Al. [35,36] studied the residual stress development in thermoset resins in two cases: cure below and above the resin’s glass transition temperature (Tg). They first investigated the cure process below Tg.
They concluded that in general the stress build-up generally depended on the crosslink density. Stress induced in the cure phase ranged from less than 1% of the total residual stress (for cure and also cool-down) in a lightly linked epoxy to more than 30% of the total residual stress in densely cross-linked epoxies and acrylates. So they conclude that the contribution of the isothermal curing part should be generally considered. When the cure temperature is above the Tg, the epoxy, as available
from the literature, exhibited no detectable stress during the curing reaction, nor during cooling down to the glass transition temperature, but develops stress below Tg. However, the acrylate generated
considerable stress with the major part above Tg , throughout the reaction and cooling,.
In the research done by Theriault & Osswald in 1999 [37], chemical shrinkage is not considered. They stated that the stress built up in cure of the matrix is released when cooled from Tcure to Tg and
the residual stresses should be considered from Tg and during the cool-down process. In this respect,
Madhukar et Al. in their papers in 2000[38-40] have noted that in different resins, the contribution of matrix volume change in the residual stress is different and therefore chemical shrinkage contribution may be large and may not be neglected. By modeling the cure cycle, they present that changing the cycle that also changes the resulting residual stresses. Reduction of chemical (cross-linking) shrinkage stresses is done by a combination of stress relaxation and thermal expansion and completing the cure in a short time.
As can be expected, process-induced residual stresses in a laminate have a considerable influence on the sensitivity to damage loads during the life-time performance. Some research was done to minimize the amount of residual stresses by varying parameters of the cure cycle. In 1993, after simulation of the cure process, White & Hahn [41] investigated the control and reduction of process-induced residual stresses by modifying processing conditions for a graphite/BMI composite material. Residual stresses are decreased by changing parameters including: cure time, dwell time, dwell
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temperature, cool-down rate and pressure. Also, Olivier & Cottu in 1998 [42] have optimised the cure cycle in order to minimize residual curing stresses while the laminate mechanical characteristics remain constant. A similar thing is done by Madhukar et Al. in 2000 [38-40], in order to reduce the residual stress in fibres.
It should be noted that in some researches that only cool-down process is considered for estimation of residual stresses and the resulting distortions like in [43], just the Tcure is chosen in the temperature
difference. In a relevant work for FMLs [44], a similar approach is chosen to account for the residual stresses.
Park & Lee in 2001[45], have used finite element analysis to model the entire cure process and compared this to the experimental results available in the literature. The procedure is useful for computing the residual stresses during cure. Again modelling the entire cure process, Zhu et Al. in 2001[46], have used a finite element method to solve the three-dimensional thermo-chemo-viscoelastic formulation, considering the heat transfer in the laminate to find the distribution of temperature. Through this effort, they state that a major part of the residual stress develops before cool-down. Since, they have simulated the dimensional change of L-shaped graphite-epoxy profiles and using this formulation, larger spring-in is measured compared with the case that only the cool-down stage is modelled using either elastic or viscoelastic modelling. Johnston et Al. in 2001[47], have also modeled the complete cure cycle. All the sources like heat transfer, cure kinetics (thermal expansion and cure-shrinkage), tool-part interfaces and post-processing tool removal are included in their model. The model is applied to calculate the distortion of some L-shaped parts and had good accuracy for prediction of both the spring-in angle and the warped shape.
Oota & Saka in 2001[48], measured both cure and thermal shrinkages and presented a special method for cure shrinkage measurement. They verified that both sources contribute to the warpage.
In 2004, Svanberg & Holmberg[49,50] investigated the cure process by a viscoelastic model with strain, degree of cure and temperature as the state variables. The variables were path dependent instead of rate dependent in conventional viscoelasticity. Thermal and chemical parts were included and used in an ABAQUS subroutine. The model was validated with some experiments.
Shokrieh & Kamali in 2005[51], have also considered only cool-down stage of cure process and calculated the stresses and the curvature.
The most recent review paper on the development of residual stresses in composites is for thermoplastic ones only (Parlevliet et Al. in 2006 and 2007[52-54]). They investigated the effective factors in three levels of the material consisting micromechanical (constituents) level, macro mechanical level (ply to ply) due to lamina anisotropy and global level (annealing and tool-part interaction).
In 2010, Abou-Msallem et Al.[55] evaluated the development of cure residual stresses of an epoxy matrix composite, considering the cure-dependent chemical shrinkage. Indeed, peel-ply method was
utilized to measure the residual stresses. In the same year, they solved the thermal, chemical and mechanical equations for the cure cycle using a FE code and by experimental verification, stated that stresses after gelation and before cool-down are considerable[15]. In 2012, Mergheim et Al. [56] stated that chemical shrinkage is important to be considered in the cure modelling of thermoset materials.
2.3.5 Researches focusing on the distortions
According to the above descriptions of the mechanisms governing the distortion of composites, different research topics are performed in this area.
In 1996, L. Peeters et Al. [57] revised some models that were previously developed to predict the shape of laminates after manufacturing. In their paper, they performed some extended experimental studies on cross-ply and angle-ply laminates on a flat mold.
It is desirable to have a mathematical model to predict the distortion of the structure parametrically. However, solving the equations governing the manufacturing processes is not possible analytically. Therefore, investigating the problem or solving the equations using finite element method would be the best choice. In 2001, Oota & Saka [48] measured the total shrinkage of the epoxy matrix of a laminate used in an electronic device and included the shrinkage in their finite element analysis.
Spring-in and warpage in angled parts made from laminated composites were investigated in 2002 by Albert & Fernlund [58].
In 2002, G. Fernlund et Al. [59] carried out some experiments to study the factors influencing the dimensional changes of laminated structures. They stated that different factors other than thermal expansion and resin cure shrinkage are important like: cure cycle, tool surface (interaction), part geometry and lay-up. They showed that if proper material models are used to represent the stress transfer between tool-part, together with large deformation solution, experimental results can be obtained using finite element models.
Considering the influence of cooling rate, in 2002 by Sun & Pang [60], curvatures of AS4/8552 non-symmetric laminates were compared when quenched (cooled quickly) and again cooled slowly, and there was not much difference. Therefore, it was resulted that the cooling rate does not have any effect on the final spring-in angles.
A FE model is developed in 2002 by Nawab et Al. [14] to account for different contributions in the spring-in of a thermoset laminate. Shrinkage, volume fraction gradient and tool-part interaction were included in the their model which was up to 80% accurate. A similar approach is chosen by Darrow & Smith in 2002 [17] and Bapanapalli & Smith in 2005 [61], accounting for thickness shrinkage, mould stretching (tool-part interaction) and fibre volume fraction gradient in the material. They concluded, from experimental investigations and linear solutions with the FE model, that spring-in was dominated
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by thickness shrinkage, which contributed approximately 75% of the measured distortion. For measurement of spring-in, they have processed digital images from the test samples.
Cure kinetics of the epoxy resin was successfully modeled in 2010 by Ersoy & Tugutlu [62]. They proposed a method that measures the through-the-thickness cure shrinkage strains during curing of the composite in a conventional DMA equipment. Spring-in of a thermoset composite with C-shape was investigated in 2010 by Ersoy et Al. [63] using a FE model. Similar experiments are performed as explained in [64] by the same authors to measure the angles in curved laminates in 2005.
In 2011, K. Magniez et Al. [65] measured the shrinkage during curing using density measurements and it was less than 3%, so the main part of the shrinkage was due to the cool-down process.
2.3.6 Tool compensation design to have precise composite laminates
Conventionally, the expected shape deviation in manufacturing composite laminates is compensated for in the tool in the design phase using past experience such that the warped part has the desired dimensions and shape. Despite the simplicity of the concept, the process is unique for each part with its own features like layup, etc., resulting in a costly and time-consuming effort to do this for each laminate type. Therefore, researches are followed in order to compensate the distortion occurred during manufacturing composite laminates. As an instance, one can refer to the work done by Jung et Al. in 2006 [66], in which spring-back of two types of open laminated shells were estimated with experiment and FE analysis. Using FE analysis, the deviation angle of the mould was changed in each run of ANSYS, and the difference between the calculated and real product angles becomes zero with different modeling properties, etc.
In 2007 [67], a FE model was developed by Capehart et Al. utilizing a simpler elastic constitutive relation for the equilibrium mechanical response of the laminate layers and uses semi-quantitative models for estimating bounds on the lateral stress produced by chemical shrinkage during thermoset cure. The FE analysis is evaluated by test but 20% deviation (error) existed. The corrected moulds are designed and solved for manufacturing distortions in a trial and error process to converge to an optimal solution. The procedure for the corrective mould yielded 45% reduction in the final distortions.
In the same year (2007) by Jung et Al. [68], spring back in a composite beam was measured using experimental measurements, modelling with classical laminate theory and and FE analysis using ANSYS. For compensation, CLT predictions are used and the web based strategy, corporates the online modifications in the CNC code for machining the updated mould.
As can be seen from above articles, a similar procedure can be followed for FML’s to have an accurate predictive model in order to design a revised mould to manufacture skin panels from FMLs.
Design for manufacturing accurate skin panels made of FMLs
2.4
2.4.1 Introduction
As mentioned earlier, different shape deviations may occur when manufacturing composite laminates (and FML’s). Here, a review of the literature is done that gives the basic knowledge to build the predictive model for accurate (and with least residual stress) fuselage or wing panels.
2.4.2 Manufacturing‐induced residual stresses and distortions in FMLs
Different types of distortions may be present after manufacturing a FML skin panel. There are two manufacturing stages in which residual stresses are created: first layup-cure and second post-cure processes.
Spring-back can occur within the layup process of FMLs, due to forming of the metal part. Kim et Al. in 2007 [69], considered the brake forming process for producing GLARE stringers. The spring-back measurements were performed by a 3D laser scan. Effect of design and process parameters were studied on the value of the spring-back angle, including punch radius, punch speed, forming load, and forming temperature. Krimbalis et Al. in 2008[70] calculated the residual stresses in rectangular symmetric FMLs using simple force-equilibrium equations. In 2003, Hofslagare [71] measured the residual stress in the aluminium layer of a FML with three methods: X-ray diffraction, neutron diffraction and strain measurement during stress release induced by delamination which showed good agreement with each other.
2.4.3 Effective parameters making distortion in FMLs
In the cure cycle of FMLs, distortions are actually due to the mechanisms already described in detail for full composites where both chemical shrinkage of the resin and the thermal shrinkage due to different properties of the constituents play a role. Different CTE between prepreg layers and metal layers is the main source. Further investigation should be made on the distortions due to metal forming of stringers.
According to the review made on full composites and FMLs, influential factors present in different development mechanisms of residual stresses and distortions in FMLs, can be listed as:
- Stacking Sequence (laminate layup)
Difference in CTE of constituents in different directions
Difference in shrinkage of the prepreg in different directions during cure - Cure Cycle parameters like temperature, pressure, heating/cooling rates, etc. - Material properties of the ingredients (resin type, fibre material, metal type) - Thickness of layers in the laminate
- Tool parameters like material properties and the friction between the tool and the laminate After cure of FMLs, due to special applications and designs, some other processes may be required that may change the stress distribution and dimensions of the laminate. These are listed as below:
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- Creating cut-outs for doors and windows of fuselage panel
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