This study investigates the static and fatigue performance of bonded, mechanically fastened, and hybrid step-lap metallic joints in primary thick aircraft structures, focusing on both baseline configurations, representing optimally assembled joints with designed inherent strength, and those assembled with predefined bondline defects to simulate practical imperfections, that can go undetected by current non-destructive inspection techniques. The results demonstrate that hybrid joints, which integrate fasteners with adhesive bonding, significantly enhance static and fatigue failure resistance compared to traditional methods. While purely bonded joints nearly restore original stiffness, they remain prone to abrupt failure, particularly in the presence of undetectable bondline defects. In contrast, the hybrid joints tested in this study extended the fatigue life of the structure to more than nine times that of mechanically fastened joints, surpassing the aircraft’s service life by over twofold. The inclusion of fasteners effectively arrested crack propagation, preventing catastrophic failure and improving overall durability. Visual inspections, strain gauges, and optical monitoring confirmed the bolts’ role in reducing Mode I opening and peeling stresses. These findings underscore the potential of hybrid joints to enhance the durability and safety of thick aircraft structures, leading to significant cost savings by reducing the frequency of repairs and downtime.

This study investigates the static and fatigue performance of bonded, mechanically fastened, and hybrid step-lap metallic joints in primary thick aircraft structures, focusing on both baseline configurations, representing optimally assembled joints with designed inherent strength, and those assembled with predefined bondline defects to simulate practical imperfections, that can go undetected by current non-destructive inspection techniques. The results demonstrate that hybrid joints, which integrate fasteners with adhesive bonding, significantly enhance static and fatigue failure resistance compared to traditional methods. While purely bonded joints nearly restore original stiffness, they remain prone to abrupt failure, particularly in the presence of undetectable bondline defects. In contrast, the hybrid joints tested in this study extended the fatigue life of the structure to more than nine times that of mechanically fastened joints, surpassing the aircraft’s service life by over twofold. The inclusion of fasteners effectively arrested crack propagation, preventing catastrophic failure and improving overall durability. Visual inspections, strain gauges, and optical monitoring confirmed the bolts’ role in reducing Mode I opening and peeling stresses. These findings underscore the potential of hybrid joints to enhance the durability and safety of thick aircraft structures, leading to significant cost savings by reducing the frequency of repairs and downtime.

The applicability of a damage slow growth management strategy to bonded joints/patch repairs of primary aircraft structures was evaluated through an experimental and computational study. Fatigue tests were conducted to investigate the entire process of disbond growth from initiation up to joint ultimate failure. The residual static strength of the joint as a function of disbond length was established using finite element modelling, in which the mesh size was calibrated using the static strength of the specimens measured in room temperature and dry (RD) and hot-wet (HW) conditions, based on the characteristic distance approach. A virtual crack close technique (VCCT) approach was utilised to assess the strain energy release rates (SERRs) as a function of disbond crack length. The measured disbond growth rates were correlated with the SERRs using a modified Paris law that enabled prediction of joint fatigue life. The fatigue test results indicated that for a joint having a sufficient static strength safety margin under a typical fatigue loading that would propagate disbond, the disbond growth would be stable in a particular length range. Thus, the slow growth approach would be feasible for a bonded joint/patch repairs if the patch is designed to be sufficiently large to allow extended damage propagation (whilst in the case when patch size must be limited, safe-life design for the patch termination region in critical repairs must be considered. Should disbond growth occur in this case, the joint must be repaired or replaced). The work presented in this paper validated the framework/procedure proposed previously by the authors (Tanulia et al., 2020) for managing damage slow growth in bonded joints/patch repairs. In the last part of this paper the planned follow-on research is briefly described.

This paper evaluates the applicability of a damage slow growth management strategy to patch repairs or bonded joints of primary aircraft structures established earlier by the authors utilising wide bonded metal joints through a computational study using MSC Marc software. The adhesive element failure criteria was applied to establish the residual static strength of the joint as a function of disbond length. A cohesive zone element model implemented in a Ucohesive subroutine was used to evaluate the strain energy release rates (SERRs) as a function of disbond crack extent and predict the disbond growth in the joint. Similar to the 2D analysis conducted in the past, the results showed that for a wide joint with sufficient static strength safety margin under a typical fatigue loading that would propagate disbond, the disbond growth would be stable within a particular length range. Therefore, the slow growth approach would be viable when the patch is modelled to be large enough to allow expanded damage growth. Furthermore, numerical results on the load redistribution effect indicate an overall significantly slower disbond growth and longer fatigue life of the joint with part width disbond than that with full-width disbond.

Variable amplitude fatigue testing is a time consuming process and a necessary part of certification for aircraft structures. While the frequency of loading can be increased for testing metals, for composites this approach is limited due to heating effects. Hence it is more pertinent to develop effective methods for reducing/compressing fatigue test spectra for composite structures. Effective spectrum reduction techniques mainly include a spectrum truncation method which removes load cycles that negligibly effect fatigue damage initiation/growth, and a cycle merging method that combines cycles to form a cycle with the equivalent effect on fatigue damage initiation/growth. Application of these methods requires understanding of the fatigue mechanisms. These techniques have seen wide use for metallic structures, with comparatively small amounts of literature existing for composite materials. For composites this is more complex due to their non-isotropic, non-homogeneous, multi-failure mode characters. Due to the lack of research concerning spectrum reduction for composites, there are no widely accepted simplification approaches with the exception of omitting low stress levels. Research is therefore required to develop and test methods for accelerating variable amplitude fatigue tests for composite aircraft structures. On the basis of review of literature, new research directions in this area are identified.

This chapter describes the application of adhesively bonded fibre composites to repair or reinforce highly loaded structures. Although the focus is mainly on patch repair of primary metallic or composite airframe structures suffering, respectively, from fatigue cracking or mechanical damage, it is equally applicable to demanding nonaircraft repairs and to other types of damage. While the use of adhesive bonding to attach repair patches is highly advantageous when compared with mechanical fastening, it is challenging for several reasons. Due to the current inability of nondestructive inspection procedures to detect weak bonds in practical situations, the greatest concern is in obtaining airworthiness certification, since for repairs to primary structure confidence must be provided that the risk of patch disbonding is extremely low. To meet this requirement, alternative approaches for assessing the initial and ongoing structural integrity of the adhesive bond are outlined, based on proof testing and/or structural health monitoring. The use of composite reinforcement to extend the fatigue life of primary metallic components by reducing strain in areas of local strain concentration is generally less critical than repairs so potentially easier to certify. Adhesively bonded reinforcements can be applied, for example, to rectify fatigue-prone design features ('hot spots'), reduce stress on welded repairs, and restore strength or stiffness that is lost by corrosion damage.

This paper assesses the suitability of a damage slow growth management strategy to bonded joints or patch repairs of primary aircraft structures. The entire process of a disbond crack growth from initiation to the ultimate failure of a typical double lap metallic joint is investigated. The residual static strength of the joint as function of disbond crack length is established using the finite element method with adhesive element failure criteria and a progressive failure analysis. For a joint having sufficient static strength safety margin under a typical fatigue loading that would propagate disbond crack, a finite element fracture mechanics analysis indicates that the disbond growth would be stable in a significant length range of the disbond crack, initiated from either "disbond tolerant zone" or "safe-life zone". The analysis further suggests that for local (part width) disbond, due to load redistribution effect the disbond growth rate would be reduced as the disbond propagates. These results suggest the slow growth approach would indeed be feasible. With the above analyses, the framework to implement the slow growth approach, predict allowable fatigue life and determine inspection interval, in accordance with the guidance provided in FAA AC 20-107B [1], is established. Furthermore, this study assesses the effect of rigidity imbalance between inner adherend and outer adherend and shows how varying the adherend thickness could affect the adhesive bond strength and disbond growth rate. This information would be useful in design of validation experiment.

The objective of this chapter is to highlight some key topics for bonded composite repairs and to suggest approaches to repairs and reinforcement that contribute to extending the lifespan and/or the inspection interval when applied to primary airframe structures.

The inability of current non-destructive inspection (NDI) procedures to confirm bond integrity has greatly limited the application of adhesively bonded repairs to primary aircraft structure, especially in applications where failure of the repair would lead to safety in flight concerns. Given these concerns, applications to primary structure are generally limited to situations where the residual strength of the parent structure in the absence of the repairs can exceed the design limit load by an acceptable factor, most conservatively as high as 1.5, which is the design ultimate strength.This paper proposes technologies and associated strategies that could lead to some relaxation of the current residual strength requirements. This makes possible the wider application of adhesively bonded repairs to primary metallic structure suffering fatigue cracks and composite airframe structure suffering visible impact damage.The detection of weak, and even absent (“Kissing”), adhesive bonds generally requires the application of a significant stress to the region of the adhesive bond which cannot be achieved by conventional NDI – ultrasonic techniques for example. Since it is generally not feasible to stress the actual repair patch, a “Proof Test” has been developed. This test requires the application of shear stress to a bonded repair coupon (BRC) made of the patch material. The BRC is bonded to the parent structure concurrently with the patch. This test can confirm both the initial and through-life structural integrity of the repair bond. However, it must be agreed by the appropriate authority that the BRC is fully representative of the patch system.In addition to the proof test more critical repairs also require through-life structural health monitoring (SHM) with a focus on detecting patch disbonding and a secondary focus in the case of repairs to metals of monitoring crack growth. It is concluded that to detect patch disbonding, a simple approach using resistance strain gauges, or more robustly optical fibre sensors, holds the most promise, at least in the short term. Disbonding is detected from measurements of reductions in strain transfer from the parent structure into the patch, an approach previously demonstrated during full-scale fatigue testing of a repaired F111 wing.Based on use of the Proof Test together with SHM, as appropriate, decision charts are presented for the management of adhesively bonded repairs to fatigue cracks in metallic structure in which the crack is not removed and to composite structure following removal of visual impact damage. A simple approach to the management of stress-reducing reinforcements for metallic components is also presented.

Design and experimetal validation were carried out for a bonded repair to a composite helicopter main rotor spar subject to ballistic damage. The research addressed a number of challenges for bonded repairs for primary structure applications. Building block design and testing of 3D and 2D specimens were combined in an optimum way to progress the repair design-validation process efficiently and cost-effectively. FEM analyses and experimental work were carried out interactively to achieve a reliable repair design. Novel repair concepts were applied that facilitated meeting certification/authorisation requirements and other special requirements for battle damage repairs. These include adopting shape optimisation for damage removal to increase the residual strength prior to application of bonded repairs, and special bonded patch design that eliminated adhesive failure as a critical failure mode in the repair system. The experimental results agreed well with the FEM model prediction and confirmed the effectiveness of the repair design.

The availability of an efficient, cost-effective repair technology is an important maintenance requirement to restore structural integrity to metallic and composite airframe structures damaged in service. Generally repair involves attachment of a reinforcing structural element or patch to replace the damaged load path. Traditionally, the reinforcements are attached to the structure with rivets or bolts; however, attachment by adhesive bonding offers many structural and cosmetic advantages.However, bonded repairs of primary structure are very difficult to certify this is because available non-destructive procedures, such as ultrasonics or thermography are unable to detect weak adhesive bonds. In view of the limitation of non-destructive inspection an alternative approach is to directly apply stress to the actual repair bond region or to a very close simulation of the region.In this paper, further work is documented on a proof test of bonded repair coupons (BRCs) that are bonded to the parent structure at the same time as bonding of the repair patch. Therefore, the BRCs are close representation of the actual repair bond strength. To assess the bond strength, immediately after patch application and also possibly through the life of the repair, the BRCs are subject to a previously determined proof load in torsion.The aim of the study is to improve the Technical Readiness Level of the test when applied to various parent-structure/patch-repair systems, including carbon-epoxy/carbon-epoxy; aluminium/boron-epoxy and aluminium/aluminium. Improved BRC application methods were developed to increase the reliability and consistency of the results, and sensitivity to cure condition, surface treatment, contamination, and fatigue damage were evaluated.A detailed finite element (FE) study was undertaken to: a) simulate stresses in the BRC, adhesive and parent structure during the proof test, b) compare the stresses in the patch and BRC when the parent material is under stress and c) investigate the influence of BRC proximity to the patch tip when the parent material is under stress.A conclusion from the FE analysis and fatigue study was that a BRC with the appropriate ply configuration could represent the bondline stresses experienced at the patch tip, and hence could also be used to monitor fatigue damage.

Airworthiness certification is required when bonded repairs are made to primary composite structure in situations where damage has reduced or has the potential to reduce residual strength to below the design ultimate strength. Generally, certification of bonded primary structure poses many difficulties. As most repairs are one-off events meeting these certification requirements is especially challenging since demonstration by testing will generally not be possible or cost-effective. This paper discusses options for addressing the two key issues relating to certification: (a) how to validate initial and enduring bond strength of adhesive bonds, mainly given the inability of conventional non-destructive inspection to provide this assurance and (b) how to develop acceptable generic design allowables for bonded repairs which represent actual failure modes – especially for cyclic loading, since validation by testing of simulated repairs will generally be infeasible. It is concluded that proof testing of bonded repair coupons is a promising approach for validating bond strength and fatigue testing of representative bond joint specimen can provide generic allowables for patch design. For hidden structure or very high value repairs structural health monitoring of repairs based on a strain-transfer approach offers considerable promise.

This paper describes the development of a proof test to evaluate the through-life integrity of structural adhesive bonds. The test is focussed on adhesively bonded patch repairs for aircraft structure; especially those where flight safety depends on the integrity of the repair patch. The test could be used either as an alternative or as an addition to structural health monitoring of bonded repairs to increase confidence to the extent that they could be certified for application to flight-critical structure. The implementation of the test is as follows: thin coupons of the patch material are bonded to the surface of the parent structure simultaneously with, and therefore under very similar conditions to the repair patch. These coupons are proof tested periodically in shear using a torque wrench. Failure of the coupon below a predetermined proof load provides an indication that the adhesive bond to the patch or (possibly) the patch itself has degraded and should be replaced. It was concluded that this test is a very promising cost-effective approach for detecting defective or deteriorated adhesive bonds; however, to raise the technology readiness level to the extent that it could be considered for aircraft applications more work is required to improve the test database and increase practicality.

This paper focuses on the difficult issue of the certification of adhesively bonded repairs in applications where credit has to be given to the patch for restoring residual strength in flight-critical structure. The scope of the paper includes both adhesively bonded composite repairs to composite components and composite repairs to metallic components. After discussing typical bonded repairs and, as a baseline, procedures currently used to certify new structure, a proposal is made which may constitute an acceptable basis for the structural certification of repairs. The key requirement is to demonstrate an acceptably low probability of patch disbonding during the remaining life of the structure. The focus is on one-off repairs where development of a comprehensive certification procedure based even on limited testing will be infeasible: Firstly, a decision process is undertaken to establish if there is indeed a certification issue. That is situations where flight safety depends on the structural integrity of the repair patch.Secondly, especially if there is a certification issue, a rigorous repair design approach is undertaken, based on a generic data base previously obtained from the testing of bonded joints representing the critical regions in the repair.Thirdly, the actual repair system is validated to ensure the required strength and potential durability has been achieved. This includes Boeing wedge testing, a proven quality control procedure for the adhesive bonding process.Finally, if this approach is considered inadequate then the option is to include proof testing and/or structural health monitoring for assessment of the actual through life structural integrity of the repair system; however, this would only be feasible for special cases where the considerable extra cost and complication can be justified.

Carbon/bismaleimide (BMI) composite is increasingly employed in critical load carrying aircraft structures designed to operate at temperatures approaching 180°C. The high post-cure temperature (above 220°C) required to fully react the BMI resin, however, renders existing on-aircraft prepreg or wet layup repair methods invalid. This paper presents a new on-aircraft repair technique for carbon/BMI composites. The composite prepregs are first warm-staged to improve the ability to evacuate entrapped air. Then the patch is cured in the scarf cavity using the vacuum bag technique, followed by off-aircraft post-cure. The fully cured patch then can be bonded using a structural adhesive.