Advances in the Bonded Composite Repair o f Metallic Aircraft Structure phần 9 pptx

Chapter 15. Graphite/epo.xy patching eficiency studies 439 Fig. 15.18. Different damage configurations of "equivalent" width that the combined contribution of Kl, and KIIr to the effective stress intensity factor was less than 8% for the configurations studied. Fatigue testing confirmed that various forms of damage could be repaired effectively with single patches. For example, the fatigue lives of panels containing 40mm diameter holes and either single cracks or four 45 kinked cracks were improved by factors of 5.1 and > 15, respectively, by single-sided repairs with a 1 mm thick patch [80 mm x 80mml. The measured fatigue crack growth rates were Table 15.5 Comparison of stress intensity factor ranges for four 45" kinked cracks and two diametrically opposite cracks, with tips at x, =4Omm and hole radius = 20 mm (R = 0.1, omrx = 65 MPa). Crack configuration Patch thickness, mm AKP AKu AKp/AKL' 4 x 45" kinked 2.0 6.0 22.3 0.30 cracks 1 .o 8.8 22.3 0.39 2 x diam. opposite 2.0 cracks 1 .o 11.3 27.7 0.41 13.7 27.7 0.49 440 Advances in the bonded composite repair of metallic aircraft structure in good general agreement with theoretical predictions. For example, the mean crack growth rate of 9 x 10~9m/cycle measured for four kinked cracks at a half crack length of 35mm during R=0.1, am,,=41.25MPa loading, was in good agreement with predicted values of AKp= 5.7 MPam’I2 and Kmin/Kmax = 0.51, and da/dN-AK data for 2024-T35 1 aluminium alloy sheet. Furthermore, the observed crack paths indicated little effect due to Mode I1 loading. In the case of a 40mm hole and two cracks, double sided patching resulted in crack arrest, in agreement with theoretical predictions. Work is in progress to determine the effectiveness of patch repairs for other damage configurations. 15.10. Future work Although adhesively bonded gr/ep patch repair of cracked metallic structures has been studied extensively and service experience with repairs has been good, it appears that further work is required to address some remaining problems and to assess the full potential of the repair technique. Specific research objectives include the following: (a) To investigate the effect of variable amplitude loading spectra on patch debonding and hence on patch efficiency. There is a clear requirement for a model to predict debonding, and for incorporation of this in a general model, which will enable the effect of patching on fatigue crack growth to be predicted for a wide range of loading spectra. (b) To assess the influence of hot-wet fatigue test environments on patch efficiency, and the effect of long-term pre-exposure to hot-wet environments on such behaviour. (c) To establish the advantages and limitations of repairs carried out by co-cure of prepreg and adhesive. (d) To develop and assess bonded patch repair schemes for applications involving elevated service temperatures. (e) To investigate the effectiveness of bonded patches for the repair of various forms of corrosion damage and battle damage in aluminium alloy structures. (f) To develop and assess patch repairs for applications involving bonding over fasteners. (g) To assess the potential of bonded patches for the repair of SPF/DB titanium alloy structures, and to develop optimum repair schemes. (h) To develop “Smart” patches for monitoring repair performance in service, and improved NDE techniques for (i) inspecting pretreated surfaces prior to bonding, and (ii) assessing the strength and durability bonded patch repairs. 15.11 Acknowledgements 0 British Crown Copyright 2001. Published with the permission of the Defence Evaluation and Research Agency on behalf of the Controller of HMSO. Chapter 15. Graphitelepoxy patching efficiency studies 44 1 References I. Kemp, R.M.J., Murphy, D.J., Butt, R.I., et al. (1983). RAE Technical Report TR 83005. 2. Sutton, G.R., Stone, M.H., Poole, P. et a/. (1984). In: Repair and Reclamation, The Metals Society; 3. Poole, P., Stone, M.H Sutton, G.R., et al. (1986). In: The Repair of Aircraft Structures Involving 4. Sutton, G.R. and Stone, M.H. (1989). RAE Technical Report TR 89034. 5. Dowrick, G., Cartwright, D.J. and Rooke, D.P. (1980). RAE Technical Report TR 80098. 6. Young, A,, Cartwright, D.J. and Rooke, D.P. (1988). Aeronautical J., pp. 41&421. 7. Young, A,, Rooke, D.P. and Cartwright, D.J. (1989). Aeronautical J., pp. 327-332. 8. Ball, AS. (1993). MOD Contract SLS 41B/2093, Final Report BAe-KDD-FCP-0104. 9. Poole, P., Brown, K. and Young, A. (1990). RAE Technical Report TR 90055. pp. 17.1-17.6. Composite Materials, AGARD-CP-402, pp. 9.1-9.21. 10. Poole, P., Lock, D.S. and Young, A. (1991). In: Aircraft Damage Assessment and Repair. The 11. Poole, P. and Young, A. (1992). In: Theoretical concepts and Numerical Analysis of Fatigue [A.F. 12. Baker, A.A. (1988). In: Bonded Repair of Aircraft Structures (A.A. Baker and R. Jones, eds.), 13. Poole, P., Young, A. and Ball, AS. (1994). In: Composite Repair of Military Aircraft, AGARD- 14. Poole, P., Lock. D.S. and Young, A. (1997). In: Proc. of 1997 USAF Aircrufi Structural Infqrit.)’ 15. Poole, P., Brown, K., Lock, D.S et a/. (1999). In: Proc. of IW9 USAF Aircruft Structural Infegrit?, 16. Poole, P., Stone, M.H., Sutton, G.R., el al. (2000). In: Proc. of 2000 USAF Aircraft Structural Institution of Engineers, Australia, pp. 85-91. Blom and C.J. Beevers, eds.], EMAS, pp. 421438. Martinus Nijhoff, pp. 107-173. CP-550, pp. 3.1-3.12. Conf., USAF. ConA, USAF. Integrity Conf., USAF. Chapter 16 REPAIR OF MULTI-SITE DAMAGE R. JONES and L. MOLENT" Defence Science and Technology Organisation, Air Vehicles Division, Monash University, Wellington Rd, CEayton, Victoria 3168, Australia 16.1. Introduction The phenomenon of multi-site damage (MSD) in aircraft has been under examination in recent years by many in the aviation industry. This section investigates the feasibility of applying advanced bonding technology to commercial aviation structures containing MSD. The validity of this technology has already been proven in its application to fatigue and stress corrosion in military aircraft, as described in other chapters of this book. The consequence of the undetected presence of MSD was dramatically illustrated through the in-flight failure of a fuselage lap joint on an Aloha Airlines B-737 aircraft on April 28, 1988. Essentially this failure occurred due to numerous small cracks along a fastener line linking together, causing the residual strength of the fuselage to be exceeded under pressurization. A test programme was conducted to reproduce this type of failure, and an adhesively bonded boron/epoxy doubler for use as a repair or preventative measure has been developed. This chapter presents the results of a fatigue test programme, which also considers environmental and damage tolerance aspects, conducted using specimens representative of wide-bodied commercial aircraft fuselage lap joints. This work was reported in detail in [l-lo]. Two separate generic specimens were considered, one representative of Boeing Commercial Aircraft Company (Boeing) and the other of Deutsche Airbus GmbH (Airbus) aircraft fuselage lap joints. * Air Frames and Engines Divbion, Aeronautical and Maritime Research Laboratory, Fishermum Bend, Virtoriu 3207. Australia. Baker, A.A., Rose, L.R.F. and Jones, R. (ea's.), Advances in the Bonded Composite Repairs of Metallic Aircraft Structure Crown Copyright 0 2002 Published by Elsevier Science Ltd. All rights reserved. 443 444 Advances in the bonded composite repair of metallic uircraft structure Following the development of a bonded-composite repair for MSD in the fuselage lap-joint of wide bodied transport aircraft a number of full scale demonstration repair/reinforcements were undertaken. 16.2. Specimen and loading 16.2.1. Boeing lap joints Figure 16.1 details a typical configuration of a modern Boeing wide bodied aircraft pressurized fuselage construction. For the purpose of this work attention is focused on the lap joint area. Local details of this location vary depending on the age of the aircraft and specific manufacturers’ details. It should he noted that, in many cases, in addition to fasteners, the lap joints are bonded together, either using hot or cold setting adhesives. This is done for the purpose of increasing the fatigue life of the joint. In service, environmental degradation may cause this bond to become ineffective, and corrosion of the mating skins could accelerate the onset of MSD (as was the case in the Aloha incident). For these reasons bonded lap joints are not considered in this work. In the present investigation a worst-case scenario was assumed, viz: a non- bonded, full depth, upper plate, counter-sunk configuration as shown in Figure 16.2(a). Here the counter-sunk rivet hole, if accompanied by improper rivet head seating, leads to a phenomenon known as “knife-edging’’ (i.e. the tip of the counter-sunk in the top plate “cutting” into the lower plate). This in turn leads to a reduction in the fatigue life of the joint, relative to the case where the counter- sunk does not fully penetrate the plate, due to the sharp corners accelerating the initiation of cracks. The basic specimen used in this investigation (referred to as the “Boeing” type) consisted of two 2024-T3 clad aluminium alloy sheets 1.016mm (0.04 inch) thick, fastened with three rows of BACR15CE-5, 100” shear head counter-sunk rivets, 3.968 mm (5/32 inch) diameter, as shown in Figure 16.3. The width of the specimen was chosen to coincide with the typical distance between tear straps of a B-737 aircraft. The upper row of rivet holes contained crack initiation sites, induced by means of an electrical spark erosion technique, on either side, nominally 1.2 mm long. This length was chosen so that the defect was obscured by the fastener head, which is representative of possible undetectable flaws. These flaws were achieved by drilling the rivet holes undersize (3.85mm diameter), spark eroding the initiation sites to 1.225 mm, and then machining the counter-sunk (4.039 mm) to achieve the required hole diameter. The accuracy to which the latter was performed determined the final configuration of the defects. In some cases the defects only remained on one side of the hole, the other being removed by the tool. Following this the fasteners were inserted. The specimens were manufactured by the then Australian Airlines (now QANTAS), from material supplied by them, to aircraft standards. End tabs were Chapter 16. Repair of multi-site damage Frame station 445 Fig. 16. I, Typical wide body fuselage construction (from Boeing). bonded to the base specimen to ensure failure would not initiate from the specimen ends (see Figure 16.3). Since the amount of out-of-plane bending due to fuselage curvature in a typical fuselage joint was unknown, the local bending was prevented by testing the specimens bonded back-to-back and separated by a 12.5 mm thick honeycomb 446 Advances in the bonded composite repair of metallic aircraft structure Joint description configuration (a) Base line 3 rows - 5/32’ Csk. Rivets 1.13” space .w to .w (b) 2 rows - 5/32 Csk. Rivets 1 row -3/16 Universal .04 to .04 (c) 3 rows - 3/16 Universal 0.04 to 0.04 (d) 3 rows - 3/16 Csk. Rivets 0.02 Bonded Doublers 0.04 to 0.04 (e) 3 rows - 3/16 Csk. Rivets 1.30 space 0.056 to 0.056 (9 3 rows - 3/16 Csk. Rivets External Doubler 0.04 to 0.04 (9) 3 rows - 5/32 Csk. Rivets Hot Bonded 0.02 Doubler 0.04 to 0.04 Bond Fig. 16.2. Various fuselage lap joint configurations (from Boeing). Chapter 16. Repair of multi-site damage 447 2024-T3 plate thickness = 0.040in End tab 3.0 5.0L I I Dimensions in inches End tabs to be bonded Fig. 16.3. Uniaxial "Boeing" type lap joint specimen. Note rivet numbering used in this chapter. core. Details of the procedure used to construct the test specimens can be found in [2]. In this configuration strain gauge results indicated no global bending or parasitic stiffening due to the honeycomb. One drawback of this method of testing is that the failure of one face (i.e. the base specimen) terminates further testing of the other. The over-riding advantage of this technique is that due to symmetry, a heat cured repair can be applied to the base specimens without inducing extensive bending due to the thermal expansion mismatch of the parent material and that of the repair. A view of the assembled test specimen is given in Figure 16.4. The specimens were tested in various capacity servo-hydraulic test machines. The specimens were loaded in tension to give a remote plate stress of 92 MPa (13.4 ksi). This figure was determined from operational data obtained for the US DOT MSD Committee Review Board, see Table 16.1 (from [ll]), for the B-737 aircraft. 448 Advances in the bonded composite repair of nietallic aircraft structure Fig. 16.4. Back to back bonded lap joint specimens. Table 16.1 MSD committee review results [l 11. Typical maximum normal operating stresses for Boeing 727 and 737 fuselage splices Primary skin stress is pressure hoop stress Actual PR/T At frames highest Comment Aircraft psi B727 15900 10000 13200 Midway between frames B737 16100 9800 13400 Middle of waffle strap area 10400 Midway between frames Maximum applied shear stresses are less than 25% of the 13000 psi hoop stresses Maximum principal stresses are: Tension ~ less than 110% of hoop stress Shear - less than 60% of hoop stress [...]... Unreinforced Unreinforced Reinforced Reinforced Conditionede Conditioned Unreinforced Unreinforced Unreinforced Reinforced Reinforced Reinforced Reinforced Reinforced Conditioned Conditioned Unreinforced Unreinforced Reinforced' Reinforced' Unreinforced Unreinforced Unreinforced Unreinforced Precrdcking cycles at 17.8 ksi Cycles to failure at 13.4ksi - 97 7600" 97 7600a >2100000@ >2100000@ 321 795 0@ 321 795 09. .. surface is particularly localised Had the upper and lower skins been bonded, a substantial proportion of the load would have been transferred prior to the first row of fasteners The boron/epoxy doubler, described in Section 16.3.2.1, uses this concept to increase the damage tolerance of the joint In this approach a boron/epoxy laminate is bonded to the 456 Advances in the honiied composite repair o metallic. .. be in recognition of leaving the initial crack in the metal unchanged but also may cover the presence of an un -bonded region in the joint." The bonded repairs/reinforcements described below were designed and tested to fulfil the above requirements 16.3.2 Repair details A boron fibrelepoxy composite was chosen for the repair applications because of its high stiffness, relatively high coefficient of thermal... 16.16 Fleet crack growth data I00000 lMWO 464 Advunces in the bonded composite repair of metallic aircraft structure It was observed that before initial failure (Le two cracks linking) of a specimen, the remaining ligament length was consistently of the order of 2mm In each case the number of cycles required to achieve initial failure did not differ significantly from that to final failure of the specimen... held constant at 60 "C Although tap water was used, on removal the specimens showed signs of extensive corrosion of the aluminium One week was considered sufficient time to thoroughly soak the boron doubler and the adhesive Water temperature was continuously monitored throughout this time 466 Advances in the bonded composite repair o metallic aircraft structure f Fig 16.17 Reinforced lap joint specimen... fatigue performance of a representative fuselage lap joint containing MSD The experimental program showed that after at least 100000 cycles the bonded- composite doubler was capable of withstanding the imposed fatigue loading Also, the cracks (cut) in the lower (hidden) row had not grown 16.5.3 Environmental evaluation o repairs f The initial testing of the reinforced and unreinforced specimens demonstrated... evaluate a possible bonded repair or life enhancement for mechanically fastened fuselage lap joints I6.3.1 Repair philosophy In 199 0, with the support of the then Australian Civil Aviation Authority (CAA, now Civil Aviation Safety Authority), a world wide study into the commercial application of bonded repair technology was performed [12] Thirty four organisations in eight countries, including ten manufacturers... two most prominent cracks for a number of specimen was plotted, from the same common initial crack length as used in data given by Boeing (see Figure 16.16) In these figures only data for the most significant cracks occurring on a specimen were plotted In each figure the two curves corresponding to the least number of cycles, were the cracks which first linked Comparison of the experimentally obtained... undergoes another decrease at the second line of rivets At this scale the stress concentration, in the bulk stress, at the first row of rivets, was not apparent A more detailed view of the bulk stress field around the first row of rivets was taken for region 2 [6] This confirmed the rapid decay in the load at the first row of rivets, due to load transfer to the lower skin Again a t this scale, the stress... monitored using eddy current techniques The results of periodic monitoring of the reinforced specimens was presented in [2] None of the reinforced or repaired specimens experienced failure in the test section under fatigue loading Given the large number of cycles experienced, this implies that the reinforcement also sufficiently suppresses the failure mode in the lower skin This investigation also . 444 Advances in the bonded composite repair of metallic uircraft structure Following the development of a bonded- composite repair for MSD in the fuselage lap-joint of wide bodied. the final configuration of the defects. In some cases the defects only remained on one side of the hole, the other being removed by the tool. Following this the fasteners were inserted. The. this concept to increase the damage tolerance of the joint. In this approach a boron/epoxy laminate is bonded to the 456 Advances in the honiied composite repair of metallic aircraft structure