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390 Advances in the bonded composite repair of metallic aircraft structure (assumed to apply for this specimen configuration): (13.11) Since, from Figure 13.12, omax is around 160 MPa and a is 33 mm, Kcdt is estimated to be about 56 MPam'/2. Similar results for were obtained from several other unpatched panels. These values for I& are in reasonable agreement with published values for 2024T3 panels of this thickness. For the patched panel, patching theory suggests that K, is approximately 53MPam1/*. Although Ko0 is fairly close to Grit, the former is an upper-bound estimate of stress intensity so it is tentatively concluded that crack propagation in the metal was not the cause of the failure. Strain capacity analysis A direct estimate, using joint theory, of net strain in the patch over the crack indicates a value of 7100 microstrain. However, if the extra load attracted to the patch (as a result of the inclusion effect) is considered, the strain could be as high as 9500 microstrain. Since strain capacity of the boron/epoxy is measured to be about 7300 microstrain, the conclusion is that failure was probably a result of initial failure of the patch. Furthermore, as discussed in reference [ 11 for the patch configuration employed, the ratio (inner-surface strain)/(outer-surface strain) in the patch is significantly greater than unity. In this case it is estimated to be about 2.5. On this basis the inner strain could have exceeded 12 000 microstrain; however, the strain elevation would be very localised. The conclusion is thus reached that failure in the patched panels resulted from initial failure of the patch, possibly associated with the strain concentration at its inner surface. This failure mode may change where significant disbond growth occurs during fatigue cycling for two reasons: 0 Stress intensity K, may exceed Lt allowing the crack to grow catastrophically 0 The strain concentration in the patch over the crack will be reduced if even minor Thus, for a small disbond, say a fewmm, residual strength is likely to increase because of the reduced stress concentration in the patch. Increasing the thickness of the patch, say to nine layers (the current patch is seven layers), should provide some increase in residual strength. However, at higher stress levels, plastic yielding of the metal around the patch (exacerbated by stress concentrations at the ends of the patch) will limit this increase. The failure mode is then expected to change from patch failure to disbonding from at the ends of the patch. under the patch. disbonding occurs. Chapter 13. Boronlepoxy patching efficiency studies 391 450 400 350 I 5 300 250 b m - m ; 200 2 u) 150 100 50 0 I., - :onstant Amplitude a=% FALSTAFF a=39 mm Fllla=38 mm No Fatigue a=30 mm No Fatigue a=33 mm Standard Boron Standard Boron Standard Boron Unpatched I I Standard Boron Fig. 13.13. Histogram showing residual strengths for patched panels with or without prior fatigue testing and for an unpatched panel. The results for the panels with no prior fatigue are plotted in Figure 13.12. Residual strength following fatigue testing Tests were also conducted on panels after fatigue testing under (a) constant amplitude, (b) F-1 11 spectrum loading-representative of the F-1 11 lower wing skin or (c) FALSTAFF spectrum, representative of a standard fighter lower wing skin. Figure 13.13 depicts the results together with those patched after fatigue cracking. Thermographic NDI was used in an attempt to detect disbond damage over the crack region in the fatigue-tested specimen; however, damage could only be detected in the FALSTAFF specimen as a relatively small -2mm ellipse centred on the crack. This does not imply that the other specimens had not suffered damage, only that the disbonds were probably smaller or for some reason less detectable by thermography. The first conclusion is that the residual strength has not been reduced by cyclic loading for cracks in the 30-40mm range. Indeed the strength may have actually increased due to the reduction of stress concentration around the crack caused by any local disbonding. In the case of the 56-mm crack residual strength was clearly reduced compared to the others. Since this crack is approaching the boundary of the patch, it is possible that in this case the critical stress intensity for the crack in the panel was exceeded, rather than the failure stress of the boron/epoxy. In all test panels the strength equalled or exceeded oy - although, with no margin in the case of the panel with the 56-mm crack. As discussed later, there is a case for equating oJ, with DUL. If this case is accepted it can be concluded that the patched panels had adequate residual strength to satisfy most certification requirements. 392 Advances in the bonded composite repair of metallic aircraft structure 13.5. An approach to b/ep patch design 13.5.1. Cyclic loading Assuming that environmental degradation of the adhesive is not an issue (through good quality control), the margin of safety, efficiency and durability of a repair to a cracked component can be assessed from estimates of the following: (a) The stress intensity range AK and R in the repaired region. This determines patching efficiency through the crack-growth parameters AR and nR. (b) The tensile strain eR in the b/ep patch which allows estimation of the margin of safety for failure of the patch. It is assumed for a composite patch that fatigue is not an issue; if it were then the range of strain AeR and R ratio would have to be considered. (c) A (validated) damage parameter in the adhesive system (including the composite interface). Possible parameters are the shear strain range Ay or Mode I1 energy release rate AGII. This allows estimation of the fatigue durability of the adhesive system. It is best, if feasible, to design the repair so that the damage threshold of the adhesive system over the crack is not exceeded; however, if it is not feasible the disbond growth rate, db/dN (Section 13.2.3) must be included in the analysis, using Eq. (4). Limited disbond growth over the crack is acceptable, however, and within limits will not dramatically reduce patching efficiency. Another important factor needed for design of the repair system is the length L* available for the patch between obstructions (Figure 13.14), since this can limit the allowable patch thickness. The length LR required for efficient load transfer depends on the patch and adhesive parameters (Figure 13.3) including patch thickness tp and the taper rate at the outer ends of the patch. Assuming largely elastic conditions in the adhesive (as required to avoid patch system fatigue), a conservative estimate of the patch length [l] is given by: 6 LR = - + length of the taper , D (13.12) where /3 is given by Eq. (Id), The taper rate for b/ep we use is around 3 mm per ply. Finally, the residual stress oT, resulting from patch and component thermal expansion mismatch, must be included in the analysis, since this influences Ay, eR and RR. Residual stress CT depends on AT= (Toperating temperature - Tcure temperature), typically 100 "C for a 120 "C curing adhesive and, Aa = (@pat& - acomponent). The length between thermal expansion constraints in the component structure (see Figure 13.13) influences acomponent which for full constraint is 0.5 aP. Based on Rose's analysis described earlier, the author [l] developed a simple algorithm for estimating the minimum thickness patch that could be applied within the installation constraints that would survive the external cyclic loading. It is generally desired to use the thinnest patch feasible for several reasons, including (a) to minimise the residual stress problems, (b) to maintain aerodynamic Chapter 13. Boronlepoxy patching eficiency studies 393 Patch Craack PARAMETERS FIRST CYCLE FOR MIN THICKNESS PATCH Fig. 13.14. Outline of algorithm for designing the minimum thickness patch. acceptability, for example to minimise disturbance to the airflow when repairs are made to an external surface, (c) to minimise balance problems; for example, when repairs are made to a control surface, and (d) to comply with installation restraints, for example, not to exceed available fastener lengths when fasteners must pass through the patch for system requirements, or to maintain clearance between moving surfaces. The logic for the design approach is shown in flow chart form in Figure 13.14, which is based on comparison of the following, as the patch is increased in thickness one ply at a time: 0 The computed patch length LR with the allowable (available) length L* 0 The computed styin in the patch compared with the experimentally determined allowable strain e,; a value of 5000 microstrain was found to be reasonable for b/ eP. 0 The computed shear-strain range compared with experimentally determined allowable A?* = 0.18 was originally used for FM73, but current work suggests that 0.10 may be more appropriate for long life repairs. These patch and adhesive allowables were obtained from tests on representative bonded joints. Increasing patch thickness increases LR but reduces eR and Ay. Assuming constant amplitude fatigue at Bo, and R, Figure 13.15, shows the outcome of a calculation based on the parameters listed. 394 Advances in the bonded composite repair of metallic aircraft structure ~~138, Rz0.1 2024T3 AT=IOO”C FREE EDGES L* = 80 mm 25 mm EXAMPLE A?*. ~0.18 e*R= 5x1 O3 t. = 0.1 9 mm 3 mm - 7 plies blep eR = 3x1 o3 ATA ~0.16 A K, = 12.5 MNm’” A K, = 40 MNm’” uT=67 MPa LR= 57 mm Fig. 13.15. Outcome of an analysis for the minimum patch thickness, AKa is the stress intensity for the unpatched case. Once AK, is estimated the inspection interval N can be determined from Eq. (13.2) and (la) or (if disbonding is a consideration) from Eq. (13.4) as: (13.13) where ai is the initial crack size and ax is the size chosen for inspection. Typically ax would be less than one third patch width to provide at least three chances of finding the crack before it grows out from under the patch. As shown in Figure 13.14, if the inspection interval is too short, (the AK reduction is inadequate) there is an option to increase the thickness of the patch providing it can still fit within the allowable length. 13.5.2. Spectrum loading Crack-growth analysis is significantly more complex under spectrum loading. It is feasible to assess crack growth for the cracked component and damage growth in the adhesive system on a cycle-by-cycle basis for the various values of effective AKo, and R. Chapter 13. Boronlepoxy patching efJiciency studies 395 If the spectrum is unknown, design can be based on a standard spectrum: FALSTAFF or TWIST for fighter or large transport aircraft respectively. If the peak stress in the spectrum (the design limit stress, CTDLL) is unknown, an estimate can be made based on the material yield stress o,, as described in the next section. The patch length LR can then be estimated for the estimated patch thickness t,, to obtain the required K reduction. However, this may be over-conservative since by definition DDLL is expected to occur only once (although in fighter aircraft it can occur many times) in the life of the aircraft. Thus LR could be based on say 0.5 or 0.6 CJDLL - and still provide acceptable residual strength at say 1.2 x ODLL (see final section). A simplified estimate of patching efficiency could be obtained by increasing stresses in all cycles in the spectrum above the threshold for crack growth to the peak stress CJDLL. As this is a severe assumption for both the cracked component and patch system, it provides an over-conservative estimate. A complication with using this approach is that the threshold stress will reduce with disbond growth. 13.5.2.1. Estimating the design limit stress (a) The most conservative is to equate it with material yield o,,. Thus the nominal stress at the design ultimate load DUL is 1.50,,, which marginally exceeds the material ultimate strength ou. For example for 2024T3 and 7075T6, respectively, ou/o,, = 1.4 and 1.3. (b) A less conservative but (in the author’s opinion) more reasonable assumption [2] is to equate the stress at DUL with o,,. Thus in accord with the requirements for DLL, where limited yielding is allowed at stress concentrations but no large-scale yielding leading to permanent deformation. As an example, Table 13.1 provides cDLL and o,, values for the F-1 1 1 lower wing skin, which is made of aluminium alloy 2024 T581. This shows that the ratio CT,,/ODLL exceeds 1.5, as required. Use of approach A would result in a 3040% overdesign. (c) By direct strain measurement, either from a static calibration or in flight. (d) From a knowledge of the external aerodynamic loads and the availability of a full F-E model of the aircraft and local region to estimate internal loads. There are several options to estimate the CDLL: Table 13.1 Data on design limit stress UDLL for F-111 for several (DADTA) data points in the lower wing made of aluminium alloy 2024 T581, compared with the yield stress uy 67 202.9 400.2 462.3 1.2 266.8 1.97 2.28 70 167.0 400.2 462.3 1.2 266.8 2.40 2.77 70a 204.2 400.2 462.3 1.2 266.8 1.96 2.26 78 149.7 400.2 462.3 1.2 266.8 2.67 3.09 154 171.8 400.2 462.3 1.2 266.8 2.33 2.69 194 165.6 400.2 462.3 1.2 266.8 2.42 2.79 396 Advances in the bonded composite repair of metallic aircraft structure Approach (c) is very time consuming and likely to be prohibitively expensive in most repairs. Approach (d) depends on having the loads and F-E model available, and even then will be costly and time consuming. However, this is the preferred approach for critical repairs and was the procedure adopted in a bonded composite repair developed for the F-1 1 1 lower wing skin [lo]. Of the two simple approaches the result of assuming approach (a) is that a thick repair would be designed resulting, in the case of composite patches, in large residual stresses and in large parasitic stress concentrations. This is not a major concern for thin-skin components (skin thickness <2 mm) where approach (a) is probably quite acceptable. 13.5.3. Check on residual strength It is most important to check that residual strength of the repaired region will exceed oDLL by an acceptable factor F generally between 1.2 and 1.5 x (the latter being the design ultimate). If this is not the case the thickness of the patch will need to be increased beyond that required for the fatigue stress level. The residual strength of the patched cracked component appears to be dependent on the strain capability of the reinforcement (including strain concentration) and the adhesive rather than on the stress intensity in the patched crack. However, a first test should be made to check that at F x ODLL, KR < Ke the effective critical stress intensity for the cracked material. If this is not the case then the patch thickness must be increased. The main test check is to ensure that the patch static-strength allowables, obtained from tests on representative bonded joints, are not exceeded. For the adhesive the allowable shear strain will be greatly increased (for FM73, Ay* = 0.5); however, the allowable patch strain eR* is unchanged, since for b/ep the static strength allowable is the about same as the fatigue allowable. At the ultimateload the adhesive yield shear stress will be greatly exceeded so, in principle, a much longer length than predicted by Eq. (13.12) would be required. However, since the ultimate load case is a check load (where large-scale yielding in both the metallic structure and adhesive is acceptable, as long as failure does not occur) the length given by Eq. (13.12) for the fatigue case should still provide an adequate strength margin. References 1. Baker. A.A. (1988). Crack patching: Experimental studies, practical applications. Chapter 6 in Bonded Repair of Aircraft Structures, (A.A. Baker and R. Jones, eds.) Martinus Nijhoff, pp. 107- 173. 2. Baker, A.A. (1994). Bonded composite repair of metallic aircraft components, Paper 1 in AGARD- CP-550 Composite Repair of Military Aircraft Structures. 3. Rose, L.R.F. (1988). Theoretical analysis of crack patching. Chapter 5 in Bonded Repair ofAircraft Structures, (A.A. Baker and R. Jones, eds.), Martinus Nijhoff, pp. 107-173. Chapter 13. Boronlepoxy patching efficiency studies 397 4. Wang, C.H. and Rose. L.R.F. (1998). Bonded repair of cracks under mixed mode loading. Int. J. of 5. Baker, A.A. (1 993). Repair efficiency in fatigue-cracked panels reinforced with boron/epoxy patches. 6. Chalkley, P.D. and Baker, A.A. (1999). Development of a generic repair joint for certification of 7. Baker, A.A. and Chester, R.J. (1992). Minimum surface treatments for adhesively bonded repairs. 8. Baker, A.A. and Beninati, 0. (1997) Repair efficiency in composite patched panels after removal of 9. Baker, A.A. (1997). On the certification of bonded composite repairs to primary aircraft structure. 10. Baker, A.A., Rose, L.R. and Walker, K.F. (1999). Repair substantiation for a bonded composite Solids, 35, pp. 2148-2113. Fatigue and Fracture of Engineering Materials and Structures, 16, pp. 753-765. bonded composite repairs. Int. J. Adhesion and Adhesives 19, pp. 121-132. Int. J. of’ Adhesives and Adhesion, 12, pp. 13-18. corrosion damage. Proc. of Int. Aerospace Conf. 1997, Sydney, Australia, pp. 53-60. Proc. of ICCM II, Gold Coast Australia, July, Volume 1, pp. 1-24. repair to F-I11 lower wing skin, Applied Composite Materials, 6, pp. 251-267. Chapter 14 GLARE PATCHING EFFICIENCY STUDIES R. FREDELL and C. GUIJT Department of Engineering Mechanics, Center for Aircraft Structural Life Extension, US Air Force Academy 14.1 Introduction Most bonded composite crack patching has been accomplished on small areas of thick structures using high-modulus boron/epoxy composites. Extending the lives of aging transport fuselage structures, however, may involve repairs to large areas of thin fuselage skins and lap joints. These structures often see their highest mechanical stresses (due to pressurization) at the low temperatures encountered at cruise altitude. Hence, more attention to the thermal properties of composite materials may be needed when fuselage structures are being repaired. This chapter presents the results of detailed parametric studies of thermal effects on bonded repairs to cracked pressurized transport fuselage structures. The hybrid glass/epoxy/ aluminum materials known as GLARE are offered as an alternative to boron/epoxy for this special crack patching application. Experiments performed at room temperature, and at the low temperatures encountered at high altitudes, show that bonded GLARE 2 patches can out-perform boron-epoxy in selected repairs to thin skins. These results are discussed with the conclusion that, under certain circumstances, thermal compatibility can be the driving factor in repair material selection in pressurized fuselage skin repairs. 14.1.1. Overview and background of jibre metal laminates The Fiber metal laminate (FML) GLARE 2 is a hybrid material of moderate modulus, combining 2024-T3 aluminum with high-strength unidirectional S-glass/ epoxy composite in a sheet like laminate [2-31. It is known for its excellent fatigue resistance due to the “crack-bridging” effect of the fibers and its high residual 399 Baker, A.A Rose, L.R.F. and Jones, R. (eds.), Advances in the Bonded Composite Repairs of Metallic Aircraft Structure Published by Elsevier Science Ltd. [...]... model only when six boron plies are used At this point, the extensional stiffness of the boron patch approaches that of the plate These results allowed sufficient confidence in Rose’s basic approach to proceed with the parametric analysis, outlined in the following section The study assesses the thermal considerations to be accounted for in the selection of patch materials Crack patching of aircraft. .. during FALSTAFF loading Debond FALSTAFF cycles oo x I O 6 25 x 50 x 75 x 88 x IO6 IO6 IO6 IO6 Fig 15.4 Development of debond during FALSTAFF loading (specimen 16/17P) Chapter 15 Graphitelepoxy patching efficiency studies 42 1 fatigue lives and reductions in crack growth rates due to patching were much more pronounced for constant amplitude loading than for FALSTAFF loading, and debonding was observed for... damage, effect of service temperature, effect of exposure to hot-wet environments, repair of battle damage and future work 415 Baker, A.A., Rose, L.R .F and Jones, R ( e d s ) , Advances in the Bonded Composite Repairs of Metallic Aircraft Structure Q 2002 Elsevier Science Ltd All rights reserved Advances in the bonded composite repair of metallic aircruft structure 416 15.2 Repair of thin skin components... stiffened fuselage structure is approximately equal to that of boron/epoxy during the cure cycle of the adhesive (local heating) Thus, in fuselage skin repairs, boron/epoxy is not substantially affected by a change in the cure temperature The large thermal effects with boron/epoxy occur in the cooling from room to cruise temperature of the complete structure, now the CTE of the structure AI 2024-T3 105 Altitude... [b/ep] patch repairs have been compared and the effects of bondline defects and in- service variables on patch efficiency have been studied This chapter reviews DERA research on bonded composite patch repair of aluminium alloy structures; it contains nine main sections covering repair of thin skin components, repair of thick sections, gr/ep versus b/ep patches, effect of bondline defects, effect of impact... (crack-opening) load will exist at the crack tip However, the change in cure cycle could (adversely) affect the adhesive properties Figure 14.7 shows the effect of various cure temperatures on the patching effectiveness of GLARE 2 and boron/epoxy patches in the Rose model at the cruise altitude situation described before In the analyzed case, the effective expansion coefficient of the stiffened fuselage structure. .. usually involves local heating of the repair area During the curing process, the unheated structure surrounding the repair area constrains the thermal expansion of the heated area But the patch, which is entirely inside the heated region, expands freely In stiffened structures, the “effective” coefficient of thermal expansion (CTE) of the constrained structure is much less than the material CTE Figure... was monitored for patched and unpatched specimens, and Figure 15.3 shows crack growth rate data obtained for FALSTAFF loading For some specimens, ultrasonic and SPATE techniques were used to monitor debonding of patches during fatigue testing Extensive debonding was observed in the case of FALSTAFF loading, as illustrated in Figure 15.4, but no significant debonding was detected in the case of the patched... excellent fatigue resistance of advanced composite materials, while retaining the machinability and cold-formability of aluminum alloys In addition, the GLARE laminates approach the performance of titanium alloys as fire barrier materials Fiber metal laminates are in service on the C-17 Globemaster I11 (aft cargo door), Boeing 777 (cargo floors and liners) aircraft, and as bonded repairs on a USAF C-5A... and in these cases crack arrest occurred Some of the important conclusions from this work are summarised below: (a) The rate of crack growth after patching depended on the loading conditions used for precracking prior to patching Thus, load shedding during precracking was recommended in order to ensure that the plastic zones associated with the final stages of precracking do not cause crack growth . Ob 0@ 0 8 8 0 0 0 8 oo Patch strength 8QQ Patch durability 8 8 0 0 @ 1 creep anchor 80 088 001 00 eooooo 1 oe0ooo @@ 00 88 Fig. 14.2. Failure modes for bonded. situation described before. In the analyzed case, the effective expansion coefficient of the stiffened fuselage structure is approximately equal to that of boron/epoxy during the cure cycle of the. stiffness of the boron patch approaches that of the plate. These results allowed sufficient confidence in Rose’s basic approach to proceed with the parametric analysis, outlined in the following