29th ICAF Symposium – Nagoya, 7–9 June 2017

Fatigue behavior and damage tolerant design of bonded joints for aerospace application on Fiber Metal Laminates and composites

Kruse, Thomas 1, Thomas Körwien², Roman Ruzek³, Robert Hangx1, Calvin Rans1

1 Technical University of Delft, Netherlands 2Airbus Defence and Space, Germany ³VZLU, Czech Republic

Abstract: Today, the application of adhesive bonding technology for primary aerospace structures is limited due to the certification regulations. State of the art is the widely used “chicken rivet” as crack arrestor which is limiting the benefits of bonding technology, particularly in composite bonded joints. In this paper results from testing of novel design approaches for damage tolerant high load transfer (HLT) joints as e.g. Panel Joints or large bonded repairs will be discussed for CFRP and for Fiber Matel Laminates´(FML) adherents. Results from fatigue testing with Wide Single Lapshear (WSLS) specimen will be presented for different configurations to proof the crack stopping behavior of sate of the art fasteners as reference crack arrestor concept.

STATE OF THE ART BONDING TECHNOLOGY

With the entry into service of the A350XWB a consequent evolution of the usage of CFRP for primary structures within Airbus Group has reached the next milestone. After a long and excellent experience with CFRP in civil and military applications, first applied on secondary structures and since 1983 for the vertical as first major primary structural component for civil aircrafts, Airbus Group has now reached the next step in the transition from a metallic to a composite with the first CFRP of an Airbus aircraft on A350 XWB. One key technology for the future development of composite aircraft structures is a suitable joining technology. Mechanical fastening remains the state of the art joining method for primary structures, whether it is made from metallic or composite materials. Bonding is one of the most promising alternative joining technologies especially for composite structures. At the same time bonding is enabling new disruptive structural concepts based on new integration sequences, structure mechanic principles and joint geometries which are also beneficial for future aircraft concepts. Beside CFRP as base material Fibre Metal Laminates (FML) can be considered as potential material for future large aerospace components due to its excellent Fatigue performance. Bonding for metallic structures including FML is considered as State of the Art technology. Nevertheless an application for assembly bonding of large parts (as e.g. fuselage panel or barrels) has not been implemented yet. One major obstacle is the missing damage tolerance behavior of bonded joints as well as missing trust in bonding as technology for critical parts.

Classification of Bonding Technologies Figure 1 shows the three main classifications of bonded joints representing the different stages of integration. The principle has been derived from CFRP as adherent material. 2 Thomas Kruse, Thomas Körwien, Roman Ruzek, Robert Hangx, Calvin Rans

Figure 1. Classification of composite bonded joints

Co-Curing represents the highest stage of part integration, resulting in a fully integrated component. The joining mechanism is chemical cross-linking Principle only valid for FRP / Polymer adherents

Co-Bonding represents an intermediate stage of product integration. An uncured part is joined with one or more cured parts to a component, typically with an additional layer of adhesive. The joining mechanism between the adhesive and the cured part is adhesion. Between the un-cured part and the adhesive chemical cross-linking is taking place. Principle valid for FRP / polymer adherents and hybrid bonding with one uncured polymer adherent

Secondary Bonding represents the lowest stage of integration. Previously cured parts are joined by a film or paste adhesive to an assembly as e.g. a component. The joining mechanism between adhesive and adherend is adhesion. Principle valid for all adherent materials

Definition of potential failure initiation modes

The following three failure initiation modes are describing the most important origins of potential failures of bonded joints. There are different root causes for these initial failure modes and only major effects will be discussed in this paper. The basic principles are identical for FRP, FML, metallic and hybrid joints. The specific characteristic of the damage and criticality for the joint is different for each material configuration and will not be discussed in this paper.

Disbond is an initial area within a bonded joint without connection between adherend and adhesive. A potential cause is a missed cover ply of a prepreg or film adhesive or failures within the adhesive application process (e.g. gaps within the adhesive layer). A disbond is detectable by means of nondestructive inspection technologies (NDI) as e.g. ultrasonic inspection within the individual limits of the detection threshold per technology and application case.

Figure 2. Failure initiation mode disbond

2 Fatigue behavior and damage tolerant design of bonded joints for aerospace application on FMLs and composites 3

Weak bond is characterized by an adhesive failure mode between adherend and adhesive. Its strength ranges from close to zero to almost full strength. The root cause is an insufficient adhesion of the adherend interfaces e.g. due to contaminations of the surface or unfavorable process conditions.

Figure 3. Failure initiation mode weak bond

A weak bond is not detectable by means of today’s NDI methods due to the absence of a detectable interface layer. Research addresses the problem but results are not expected short- to mid-term for industrial usage.

Impact events within manufacturing and in service can lead to initial damages of the adherend and/or the adhesive. Damages resulting from impact are detectable by NDI within the individual limits of the detection threshold.

Figure 4. Failure initiation mode impact

Certification compliance

The certification authorities differentiate between bonding of FRP and Bonding of metallic surfaces. Metal bonding is a certified process applied for large components as demonstrated in the past for several commercial and military aircrafts. Nevertheless a usage for critical high load transfer joints in major load path as e.g. fuselage panel joints have not been implemented due to leak of confidence and missing damage tolerance behavior compared to a multi-point connection by individual fasteners.

For Composite structures today‘s certification guidelines according AC 20-107B [5] are limiting the certification of composite bonded joints within the standard means of compliance to the following possible approaches for civil aircraft applications:

“For any bonded joint, the failure of which would result in catastrophic loss of the airplane, the limit load capacity must be substantiated by one of the following methods: (i) The maximum disbonds of each bonded joint consistent with the capability to withstand the loads in paragraph (a)(3) of this section must be determined by analysis, tests, or both. Disbonds of each bonded joint greater than this must be prevented by design features, or (ii) Proof testing must be conducted on each production article that will apply the critical limit design load to each critical bonded joint, or (iii) Repeatable and reliable non-destructive inspection techniques must be established that ensure the strength of each joint." [5]

Today, no suitable NDI method to fulfill the requirement [5]; (iii) of a secured measurement of the failure strength of a joint is in place. Moreover, it is not affordable to establish a full single part testing of each bonded joint within an industrial environment of a commercial aircraft manufacturing according to requirement [5]; (ii). Therefore, the only requirement [5]; (i) is practically taken into account for the sizing and certification of bonded joints. The state of the art to certify a structural composite joint is to follow approach [5]; (i) by the usage of additional fasteners which have to be capable to carry the relevant loads taking into account a global failure of the bondline. This boundary condition and the corresponding technical concept of additional fasteners are limiting the benefits of the application of composite bonded joints in terms of weight, cost and performance.

CRACK ARRESTING APPROACH

Context for certification

4 Thomas Kruse, Thomas Körwien, Roman Ruzek, Robert Hangx, Calvin Rans

To follow directly the directive provided by AC20-107B [5]; (i) the limitation of the maximum disbond to a non- critical size for each structural application is one feasible way within todays certification boundaries. The European funded Project BOPACS has focused on the development of crack stopping concepts for CFRP adherents to improve today’s state of the art additional fasteners concept by two means: First, by understanding of the crack stopping mechanism in composite bonded joints for primary structures with high and low load transfer configuration. Second, by development of novel crack stopping features as alternative to state of the art used fasteners. Complementary to the work in BOPACS the transferability of the principle towards FML application is under investigation at TU-Delft.

Validation approach

The target of the current work is to demonstrate a secured crack stopping under fatigue loads in case of the presence of a local defect as e.g. a weak bond. For the validation of the crack arresting principle a two-step approach has been established. For comparison of the individual crack arresting capability of different design features the Cracked Lap Shear (CLS) test hast been selected. The CLS specimen consists of a lap adherent and a partially bonded strap with an artificial disbond. Under tension it features a mixed mode load (in plane shear & peeling) at the bondline interface. The significance has been demonstrated within the coupon test campaign [1].

To demonstrate a more realistic application scenario the wide single lap shear (WSLS) specimen has been developed within BOPACS. It represents a typical high load transfer (HLT) configuration as e.g. a fuselage longitudinal joint. The implementation of artificial disbonds and different disbond stopping features is part of the validation concept.

Experimental investigation – Wide Single Lap Shear (WSLS) specimen

Principle definition: The evaluation of the damage tolerance and crack arresting behavior for HLT configurations of bonded joints is based on the expected principle behavior as sketched in Figure 5.

Figure 5. Structure mechanic principle on high load transfer (HLT) joints

4 Fatigue behavior and damage tolerant design of bonded joints for aerospace application on FMLs and composites 5

The sizing of a bonded joint has to fulfil the no growth criteria according the certification requirements. In case of a local manufacturing defect (weak bond) or in-service damage (impact) of the bondline the load transfer is interrupted. This leads to an increase of the stress peak next to the damage (Figure 5, position ). Depending of the initial disbond size, the stress peak will exceed the no growth load level. Therefore, it is assumed that cyclic loading will lead to an accelerated crack growth with increasing stress concentrations at the edges of the disbond (Figure 5, position  to ). Within BOPACS the target was to demonstrate for this configuration a safe, controlled disbond growth at loads below limit load (typical fatigue spectrum) similar to the slow growth criteria for metallic structures. A secured crack arresting has to be demonstrated before the critical disbond length is reached to ensure limit load capability with arrested disbond growth. In this arrested condition criteria AC-20- 107B [5]; (i) “The maximum disbonds of each bonded joint consistent with the capability to withstand the loads in paragraph (a)(3) of this section must be determined by analysis, tests, or both” is again the baseline for certification. For FML adherents also the fatigue effects of the metallic layers have to be considered.

Following this principle a suitable test setup has to enable a crack growth perpendicular to the load direction. Also the width has to be defined in accordance with the desired maximum disbond growth to be investigated to limit the risk of a static rupture by increased mean stress in the far field of the undamaged bondline. Also a relevant pitch for placement of crack arrestors has to be ensured.

WSLS TEST SETUP AND RESULTS CFRP

The WSLS test configuration has been defined with a width of 500 mm and has been equipped with a bolted load introduction with metallic doublers as displayed in Figure 6. Adherent material is made of Hexcel 8552 / IM7 prepreg tape with quasiisotropic stacking, the used adhesive is EA9695. The initial disbond for the conducted tests is located in the center of the specimen by means of a Teflon strip. The inspection of the crack growth has been performed by means of ultrasonic at damage growth adapted intervals.

Figure 6. WSLS geometrical definition and load introduction

The proof the stress distribution resulting from the load introduction has been performed by direct image correlation (DIC) technique under static loading using a DANTEC Q-400 DIC system. A homogeneous strain/stress distribution over the panel width has been observed, correlating with previous simulation. The results of the DIC measurement is shown in figure 7.

6 Thomas Kruse, Thomas Körwien, Roman Ruzek, Robert Hangx, Calvin Rans

2

1,75

1,5

1,25 Line 1 1 Line 2

Strain [milistrain] Line 3 0,75

0,5

0,25

0 0 100 200 300 400 500 panel width [mm]

Figure 7. DIC measurement of stress distribution of WSLS panel without initial damage

Fatigue testing has been conducted to determine the no growth strain level first. With the given configuration this strain level was found at 30% of the static strength.

Figure 8. WSLS test comparison

6 Fatigue behavior and damage tolerant design of bonded joints for aerospace application on FMLs and composites 7

Figure 8 shows the disbond propagation evaluated by ultrasonic inspection in k-cycles for three representative configurations with a similar size of the initial disbond. Beside the two reference configurations without disbond stopping feature (figure 8, WSLS panel 1 & 2) a configuration with lockbolts as crack arrestor (figure 8, WSLS panel 3) have been tested.

Figure 9 gives the comparison of the disbond area growth of the selected configurations. A difference between WSLS 1 & 2 (reference panel without crack stopping feature) in terms of crack growth rate and final failure size before rupture was observed. Therefore, further repetitions have to be conducted to validate the results for all configurations. For specimen WSLS 3 (incl. disbond stopping features) no significant progress of the damage has been observed after more than 3.000.000 cycles. After increasing the Loadlevel by 33% a continued crack growth has been observed without significant acceleration for additional one million cycles without rupture.

Figure 9. Disbond area evolution comparison CFRP

Conclusions WSLS for CFRP Joints

The developed wide single lap shear test setup has proven to be suitable for evaluating the crack arresting capability for high load transfer joints. Also the general assumed and theoretically predicted crack growth behavior transverse to the load direction of a WSLS configuration has been validated. The selected crack arresting concept by Hi-Lock Fasteners readily available provides a significant arresting of the crack and slow growth behavior of the damage. Further crack arresting features will be tested in future campaigns and compared. Currently, the development of suitable numerical fatigue crack growth and arresting prediction methods for bonded joints is ongoing within BOPACS [4]. To verify these developments additional tests have to be performed. Nevertheless is has been demonstrated that a significant increase of fastener pitch beyond the “second loadpath at limit load via bolted joints” is feasible without negative impact on the damage tolerance of the bonded joint.

WSLS TEST SETUP AND RESULTS FML

As a complementary study to the work on CFRP joints performed in BOPACS, an initial study of the transferability towards bonded Fiber Metal Laminate (FML) joints has been carried out at.

Looking at FML joints introduces several new considerations in terms of the effectives of the crack stopping feature; namely, the presence of a metal-to-metal bond, the increased bearing performance and ductility, and the potential for metal fatigue cracks in FMLs could alter the effectiveness of the crack stopping feature. Furthermore, the superior damage tolerance of FMLs increases the relative criticality of the overall bonded joint area in a bonded FML structure compared to bonded monolithic metal structures.

8 Thomas Kruse, Thomas Körwien, Roman Ruzek, Robert Hangx, Calvin Rans

In order to investigate these potential differences, FML specimens conforming to the basic geometry of the WSLS test according the definition of the previous chapter of this paper has been manufactured and tested. The artificial initial disbond represents 16% of the total bonded area, with and without Hi-Lock Fasteners acting as disbond stopping features. The adherent material is GLARE 2A-4/3-0.3 joint by an FM94 epoxy film adhesive. The loading has been set to 128% of initial load level applied at the CFRP Specimens based on initial trials. Fatigue testing was performed at the Delft Aerospace Structures and Materials Laboratory in a 500kN capacity MTS test frame. FML WSLS 1 is without and FML WSLS 2 with disbond stopping features applied.

Figure 10. Disbond area evolution comparison FML.

The comparison of the disbond growth behaviour between the WSLS test with and without disbond stopping features is shown in Figure 10. A significant reduction of the crack propagation could be observed. The bolts were able to increase the total growth life of the disbond by a factor of 8.5. Monitoring of the disbond growth in the FML specimens could not be performed using the same in situ ultrasonic inspection methods used in BOPACS due to reflections in material transition interfaces and reflections between layers. As a result, a combination of Digital Image Correlation and periodic removal of the test specimen from the test frame for ultrasonic C-scan inspection has been used. Figure 11 shows the processed C-scan image results for the WSLS specimen containing crack stopping bolts showing the disbond before and after it envelops the crack stopping bolts. Correlating the cycles in Figure 11 to Figure 10, it is clear that the rapid increase in disbond growth correlates well with this point of envelopment.

Comparing disbond shapes between CFRP and FML configurations, one can observe an inward growing tendency behind the disbond arresting features in the FML specimen. This shape has also been observed for the CFRP configuration with similar load level of 133%.

332000 cycles (a)

372000 cycles (b)

8 Fatigue behavior and damage tolerant design of bonded joints for aerospace application on FMLs and composites 9

Figure 11. Ultrasonic C-scan images of disbond before (a) and after (b) is envelops the bolts.

As observed in the composite specimens in the BOPACS project, joint failure was found to consist of a combination of disbonding along the overlap and interlaminar failure within the adherend (see Figure 12). For an FML, this comprises of delamination between metallic and composite plies in the adherend and cracking of the metallic layers. However, the base fatigue behaviour and mechanisms in CFRP composites and FMLs are not entirely the same. The metal layers represent fatigue sensitive layers that do not contain internal crack stopping features (ie: fibres) and thus rely on load transfer across the interlaminar interface between the composite and metal layers; the so-called bridging load. This bridging load can contribute to delamination growth within the laminate that couples with the metal crack growth behaviour. Thus, it is currently uncertain if the observed delaminated and cracked aluminium regions are the result of final static failure of the joint or could be (at least partially) due to fatigue crack growth within the FML laminates. The possibility of fatigue crack initiation, particularly at the fastener holes) could contribute to a reduction in the crack arresting behaviour of the bolts and may offer an explanation as to why the bolts failed to completely arrest the disbond in the FML specimens. Further study to validate this hypothesis is ongoing.

Figure 12. Fracture surface along the bonded overlap showing evidence of disbond and delamination failure.

Conclusions WSLS for FML Joints

Overall, although failing to generate complete arrest of the fatigue disbond, the FML WSLS tests demonstrated that the crack stopping fasteners could be an effective means of at least delaying the growth of disbonds arising from potential weak bonds or damages in a HLT bonded FML structure.

CONCLUSION & COMPARISON CFRP VS. FML

The principle concept of crack arresting under fatigue loading has been demonstrated for both investigated material combinations. Although a similar initial reduction in the growth behaviour of the disbond was observed for the FML specimens relative to the composite specimens, the bolts did not succeed in completely arresting the disbond. This effect is based on the different absolute load levels for the CFRP (100%) and FML (128%). Comparing the crack shape and also the disbond growth rate as displayed in Figure 13 shows a similar crack growth rate for similar load levels. This implied that the influence of the base material is quite similar in the range of the investigated test configurations.

The total disbonded area of the fastened WSLS specimens Figure 13. comparison disbond growth rate differs between the BOPACS specimen and the FML specimen. CFRP (load level 133%) vs. FML (load level This deviation can be explained by additional load uptake of the 128%) eight fasteners used in the FML specimen which were used at the four outer edges of the total bonded surface to decrease the risk of inward disbond growth from the bonded edges.

10 Thomas Kruse, Thomas Körwien, Roman Ruzek, Robert Hangx, Calvin Rans

Furthermore, similarities between the composite and FML specimens were observed in terms of the contribution of first-ply failure and interlaminar failure of the adherends to the overall damage growth. These findings can be exploited to study further how to optimize the joint configuration and the crack stopping bolt configuration in order to maximize their combines disbond arresting potential as a means for certification of bonded HLT joints.

Further work is needed to quantify the basic observed effects and to understand especially the failure mechanism observed within FML configurations. Nevertheless the test campaign has shown the principle applicability of crack arresting as promising approach to implement assembly by secondary bonding in high load transfer joints on a short- to medium-term basis.

ACKNOWLEDGMENTS

The research leading to these results has received funding from the European Union Seventh Framework Programme (FP7/2007-2013) under grant agreement n° 314180 (Project BOPACS: Boltless Assembling of Primary Aerospace Composite Structures).

Examples:

[1] T. Kruse, T. Körwien, S. Heckner, and M. Geistbeck, Bonding of primary aerospace structures – crack stopping in composite bonded joints under fatigue, Proceedings of the 20th International Conference on Composite Materials ICCM-20, Copenhagen, Denmark, July 19-24 2015

[2] T. Kruse, T.-A. Schmid Fuerstes, Bonding of CFRP primary structures – Boundary conditions for certification in relation with new design and technology developments, proceedings of SAMPE Seattle 2014,June, 2-5 2014

[3] Boltless assembling Of Primary Composite Structures (BOPACS), "Description of Work," EU FP7 Research Project, 2012.

[4] R. Sachse, A.K. Pickett, M. Gnädinger, P. Middendorf, Mechanisms to arrest a crack in the adhesive bond line of fatigue loaded CFRP-Joints using a rivetless nutplate joint, Proceedings of the 20th International Conference on Composite Materials ICCM-20, Copenhagen, Denmark, July 19-24 2015

[5] U.S. Department of Transportation - Federal Aviation Administration (FAA), Advisory Circular - Composite Aircraft Structure, AC-107B, FAA, 2009.

[6] T. Kruse, T. Körwien, R. Ruzek, Fatigue behaviour and damage tolerance design of composite bonded joints for aerospace application, Proceedings of the 17th European Conference on Composite Materials ECCM- 17, Munich, Germany, June 26-30 2016

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