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15:30   Fatigue Life Enhancement Methods and Repair Solutions Chair: Aslak Siljander 15:30 30 mins BONDED REPAIR OF A F-16 BULKHEAD FLANGE David Wieland, David Stargel, Joseph Wahlquist Abstract: The Fuselage Station (FS) 341 Bulkhead of the F-16 is experiencing cracking in the keel beam attachment flange. The main landing gear attaches to this bulkhead and it forms the aft enclosure boundary of the main landing gear wheel wells. The FS 341 bulkhead is shown in Figure 1 after it has been removed from the aircraft. The lugs where the main landing gear struts are attached can be seen clearly in this figure. Cracks are developing in the radius of the bulkhead flange that attaches to the centerline keel beam. The typical cracking at this location is shown in Figure 2. Development of a bonded repair for this location is difficult due to irregular loads, varying thickness, fasteners and complex geometry. These factors impose severe limitations on locations for bonded repair doublers. Of significant technical interest to the United States Air Force (USAF) was the capability of single sided bonded repair doublers to sufficiently control crack growth. If doublers had to be applied on both sides of the bulkhead, the amount of disassembly of the aircraft required to gain access to the back side of the bulkhead would significantly reduce the appeal of a bonded repair doubler approach. This single sided repair strategy for the bulkhead structure provides the major appeal for a bonded repair and became a major focus of the development of a bonded repair for this area. A practical solution for the bulkhead cracking has been developed and permits aircraft repairs without removing the bulkhead from the airframe. Thus, significant return on investment (ROI) has been achieved for this repair. These ROI benefits include both reduced aircraft repair costs and a reduction in aircraft downtime. This paper will discuss the successful design development, analysis and testing of a bonded repair for the F-16 FS 341 bulkhead flange. This presentation will also discuss the successful prototype bonded repair installed on an F-16 at Hill AFB (Figure 3). The bonded repair has been approved for installation on USAF F-16s. The Air Force Research Laboratory not only funded the effort but was also instrumental in taking the prototype repair procedure and developing a depot level repair procedure. The repair has been installed on various aircraft and is currently allowing the aircraft to return to service without any flight restrictions. No anomalies have been observed in fleet use to date. 16:00 30 mins BONDED REPAIR FOR FUSELAGE DAMAGE: AN OVERALL BENEFIT TO COMMERCIAL AVIATION Domenico Furfari, H.J.M. Woerden, Rinze Benedictus, Adrie Kwakernaak Abstract: Adhesively bonded repair of aircraft structures is frequently used in military applications, especially with the Australian and United States Air Forces. Although bonded repairs have the capacity to provide a maintenance-free repair solution to a variety of structural damages when compared to conventional mechanically fastened repairs, commercial operators still only use them on a very limited scale. In this paper the fatigue behaviour of bonded repairs to aluminium fuselage skins structures are discussed. A fatigue test campaign with increasing complexity of the test specimen (“pyramidal approach”) was carried out. Fatigue test results of cold and hot bonded repairs were compared to those of conventional riveted repairs. Two types of environmental conditions during application of bonded repairs have been considered: the “production” environment meaning a bonded doubler with surface pre-treatments and bond line curing conditions as will be used in production environment (Grit-blast Silane, autoclave); and the “repair” environment as the one typical when a repair is applied to an aircraft in-field (SolGel® surface pre-treatment, vacuum bag and heat blanket). Tests ranged from simple coupon specimens (butt-joints) to aircraft structural components (curved stiffened fuselage panels), in which both a conventional riveted repair and the hot bonded repair were embodied. Flat unstiffened panels with increasing cutout (i.e. damage) size were also tested using the different bonded repair solutions. All specimens were subjected to constant amplitude fatigue loading at stress ratio R=0.1. The applied loading conditions of the curved stiffened panels were internal pressure and longitudinal load at R=0.1, simulating in-service loading conditions of an aircraft fuselage. The fatigue life of the conventional riveted repair was used as reference to estimate the fatigue life enhancement for the bonded repairs. The influence on fatigue life of surface pre-treatment, environment (“production” or “repair”) as well as adhesive type (cold or hot bonding) was studied. The results of the coupon specimens were used as input for design of the next specimen type in the “pyramidal approach” by selecting the combination of the parameters showing the most promising and most interesting results. The bonded repairs on the curved stiffened fuselage panels were applied using SolGel® as surface pre-treatment, and a vacuum bag and heat blanket cure setup to simulate in-field repair conditions as will most usually be met with for a real aircraft. It was found that the amount of cycles to initiate a fatigue crack in bonded repair solutions on the coupon specimens strongly depends on the adhesive type used. The fatigue lives of cold bonded coupon specimens were not longer than that of the reference riveted repair. The hot bonded coupon specimens showed a significant enhancement in fatigue performance with fatigue lives up to 6 times longer than the conventional riveted repair. In general the hot bonded repair solution, regardless of the surface pre-treatment or the application environment showed a fatigue life at least 2 times longer than the reference riveted repair solution. These results were confirmed for hot bonded repairs applied to the most realistic (top of the “pyramid”) curved stiffened fuselage panel specimens. 16:30 30 mins FATIGUE & DAMAGE TOLERANCE OF HYBRID MATERIALS & STRUCTURES - SOME MYTHS, FACTS AND FAIRYTALES Rene Alderliesten, Rinze Benedictus Abstract: In the past decades, the Fibre Metal Laminate (FML) material concept has been developed with success for aeronautical applications. Although originally aiming for lower wing skin panels with FMLs containing aramid fibres (Arall), the concept has been brought to a technology readiness level for fuselage skin applications with FMLs containing glass fibres (Glare). Inspired by this successful concept, many researchers all over the world have pursued research on various FML types to investigate the behaviour of these materials under mechanical-, thermal- and environmental loading. As a consequence, the list of internationally refereed journal papers concerning FMLs has become impressive. However, there seems to be an enormous gap between the work reported in these journal papers and the work being done by institutes and companies on actual FML applications. The group of Aerospace Materials & Structures at Delft University of Technology has been working in both fields; scientific research often performed by PhD researchers, and applied research together with among others OEMs. From the cross-fertilisation between scientific and applied research on FMLs within this group, some insight and understanding has been gained on the significance of each field towards the other. Additionally, it has been observed that the lack of interaction between science and application seems to contribute to the misperceptions on the hybrid material technology at both sides. A lot of academic research reported in the literature describes a selection of experiments or (numerical) analyses on a particular combination of metallic and fibre reinforced polymer constituents. The research topic is approached with a specific scientific question or objective in mind. Often this leads to detailed understanding of a particular problem. However, despite claims and conclusions, the investigated combination of materials never leads to actual applications, because of the lack of a thorough development process towards applications. Applied research on the other hand, often approaches the material concept from the perspective of a particular design criteria or problem. In the attempt to obtain a specific material performance, the nature of the hybrid combinations seem to be neglected and the scope of research limited. As a consequence, the possible design options remain limited. This paper provides a brief overview of the reported applied and academic research performed over the last two decades on fatigue and damage tolerance of hybrid materials. The most important aspects observed in both fields will be highlighted. Subsequently, the lessons learned from both the academic and applied research fields are being addressed. These lessons are then used to discuss the (recent) developments on hybrid material concept technology in the context of optimizing and designing to increase the fatigue performance and damage tolerance of structures.