Bonded composite repair of fatigue-cracked primary aircraft structure

Bonded composite repair of fatigue-cracked primary aircraft structure

Alan Baker

Airframes and Engines Division, Defence Science and Technology Organisation, FishermenÕs Bend, Aeronautical and Maritime Research Laboratory,

Melbourne, Australia

Abstract

Repairs based on adhesively bonded ®bre-composite patches or reinforcements are more structurally e�cient and much less

damaging to the parent structure than standard repairs based on mechanically fastened metallic patches. As a result of the high

reinforcing e�ciency of bonded patches ``live'' fatigue cracks can be successfully repaired. However, when such repairs are applied

to primary structure a conservative certi®cation approach is often taken in which no credit is given to the patch system for slowing

crack growth or restoring residual strength. Thus, cracks approaching critical size cannot be repaired and inspection intervals must

be based on the predicted growth behaviour of the unpatched crack. To allow credit to be given to the patch the need is to

demonstrate either (a) that the likelihood of patch loss is acceptably low or (b) that its loss can be immediately detected. Two

approaches are discussed: the ®rst approach which addresses (a) is based on a demonstrated ability to predict the patch systemÕsfatigue behaviour and to assure its environmental durability. The second approach that addresses (b) is based on the ``smart patch''

concept in which the patch system monitors its own health. Crown Copyright Ó 2000 Published by Elsevier Science Ltd. All rights

reserved.

Keywords: Aircraft structure; Repair technology; Composites; Adhesive bonding; Fatigue cracking; Smart structures

1. Introduction

As a result of their excellent load transfer character-istics [1] bonded reinforcements or patches provide a sti�alternative load path so they can be used very e�ectivelyto repair ``live'' cracks. In contrast, standard (StructuralRepair Handbook-SRM) repairs, based on mechani-cally fastened metallic patches provide a relativelycompliant alternative load path so they cannot e�ec-tively repair cracks and require prior removal (or someother terminating treatment) of the cracked region.

Mechanical repairs also have several other disad-vantages, compared to bonded repairs as highlighted inFig. 1.

1.1. Background on patching technology

The high-performance ®bre-composites boron/epoxy(b/ep) and graphite/epoxy (gr/ep) are highly suited foruse as a patching or reinforcing material for defective ordegraded metallic structure [2].

Brie¯y the attributes of these composites include:· High Young's modulus and strength, which minimis-

es the required patch thickness (b/ep is around 3� assti� as aluminium).

· Highly resistant to damage by cyclic loads.· Immunity to corrosion, forms excellent protective layer.· High formability, which allows easy formation of

complex shapes.· Low electrical conductivity (b/ep only), which facili-

tates use of eddy current NDI for monitoring thepatched cracks and eliminates concerns with galvaniccorrosion.The main disadvantage of composites as patching

materials results from their relatively low coe�cient ofthermal expansion compared to the parent materialwhich results in residual tensile mean stresses in the re-paired component [2].

Although relatively costly, b/ep is chosen as the patchor reinforcement for most Australian bonded compositerepair applications, mainly because of its excellent me-chanical properties, low conductivity and relatively highcoe�cient of thermal expansion. However, gr/ep be-cause of its better formability is chosen for regions withsmall radii of curvature and sometimes because of itslow cost and much higher availability. Patches aregenerally bonded with an aerospace-grade structuralepoxy-nitrile ®lm adhesives, curing from 80°C to 120°Cto provide a maximum temperature capability of around100°C.

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Composite Structures 47 (1999) 431±443

0263-8223/99/$ - see front matter Crown Copyright Ó 2000 Published by Elsevier Science Ltd. All rights reserved.

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Composite reinforcement [3,4] can be used for a widerange of repairs/reinforcements to metallic aircraftcomponents, many of which have been successfully ex-ploited in Australia, these are categorised as follows:

Reduce stress intensity (crack patching)

± in regions with fatigue cracks,± in regions with stress-corrosion cracks,± increase damage tolerance in safe-life components.Sti�en under-designed regions

± increase static strength,± reduce fatigue strain at stress concentrators,± reduce ¯utter,± reduce de¯ection.Restore residual strength or sti�ness

± after corrosion removal,± after ¯aw/crack removal,± after reshaping to reduce stress concentration,± in regions with widespread cracking.

1.2. Current limitations of crack patching

The economic bene®ts of patching live cracks arefully realised only when full credit can be given to thepatch for the restoration of residual strength and re-duction in fatigue crack growth rate. However, whileapplication to non-¯ight critical structure is becomingwidely accepted, application of bonded composite re-pairs to cracked primary structure is acceptable in mostcountries only on the basis that a margin on design limitload (DLL) capability is retained in the loss (total ab-sence) of the patch [5,6]. This is the Fail-Safe approach,which can be applied:· When a cracked single load path component main-

tains su�cient residual strength to withstand DLLtimes a safety factor (often 1.2). In this case inspec-

tion is required to ensure that the critical crack lengthfor failure at this stress is not exceeded, based on thepredicted growth rate for the unpatched crack.

· When a redundant (multiple) load path componentcan withstand DLL times the safety factor if thecracked path has failed. This is a standard procedurenow for such structures so should not pose any par-ticular problems. In this case inspection is requiredto ensure that the alternate load paths are notcracked to the extent that they could not withstandthe elevated stress caused by the failure.The single load path and redundant load path

structural designs are, respectively, typical of ®ghter andlarge transport aircraft.

The inspection requirement based on growth of theunpatched crack in the single load path case may beacceptable (for example, if it coincides with the currentinspection interval) where the initial crack size is small(or non-existent, for example, in the case of a preven-tative doubler rather than a patch). However, whilst theFail-Safe approach is highly desirable, there will besome repairs, for example, those with crack approachingcritical size or with small cracks having a high predictedgrowth rate, where this inspection requirement will notbe cost-e�ective. The repair can then be certi®ed only bygiving full credit to the reinforcement in restoring re-sidual strength and reducing the rate of crack growth.

The need to certify such a repair arose a few years agoin Australia, where a bonded composite repair [7] wasdeveloped for a fatigue crack in the lower wing skin of anF111 aircraft. This crack was very near to the critical sizeat DLL. Fig. 2 provides details on the repair, which hasprevented crack growth in three years of further service.

Recently this cracking problem has become a ¯eet-wide issue with F111 since several similar cracks have

Fig. 1. (a) Some disadvantages of standard mechanically fastened repairs and (b) advantages of bonded composite repairs.

432 A. Baker / Composite Structures 47 (1999) 431±443

been detected. The aim is to patch these or, in the ab-sence of cracking, to apply preventative reinforcementson all Australian F111 aircraft. The Fail-Safe approachcan be adopted initially since the new cracks are belowthe critical size.There are essentially two ways of ob-taining this justi®cation. These are by showing:

(a) that the risk of signi®cant loss in patching e�cien-cy (de®ned later) is acceptably low, or(b) that signi®cant loss can be detected before it be-comes a safety issue.

Fatigue strength of the patch system is considered not tobe a concern where the static strength of the repair joint(crack spanned by the patch) is designed to exceedgreatly the ultimate strength of the parent material [8,9].However, fatigue is a concern with highly loaded pri-mary structure where this margin cannot be achieved. Inall repairs, thin or thick skin, environmental durabilityof the patch system is a concern.

Thus two approaches are discussed here: the ®rstwhich addresses (a) is based on a demonstrated ability topredict the patch systemÕs fatigue behaviour and to as-sure its environmental durability. The second, whichaddresses (b), is based on the ``smart patch'' concept inwhich the patch system monitors its own health.

2. Justifying credit for patching e�ciency ± fatigue

concerns

This section addresses the situation where reductionin patching e�ciency due to fatigue damage is the main

concern. This includes consideration of degradation infatigue properties of the patch system due to environ-mental agents such as temperature, moisture, fuel, hy-draulic oil, etc.

Essentially the requirement is to prove by test oranalysis that the patch can restore the residual strengthof cracked structure to the required level (>1:2�DLL�for the remaining service life. Thus, it is important toshow that unacceptable loss in patch e�ciency cannotresult from the service cyclic loading (or from mechan-ical damage) and that the fatigue crack in the parentstructure will grow at an acceptable rate which can bepredicted for the purpose of setting inspection intervals.

Patching e�ciency is de®ned here as (a) the recoveryin residual strength following patching and (b) the re-duction in crack growth rate of the fatigue crack pro-vided by the patch. The patching e�ciency dependslargely on the reduction of stress intensity followingpatching and possibly on the level of mean stress [10]resulting from thermal mismatch between the patch andparent structure.

2.1. In¯uence of fatigue on patching e�ciency

To satisfy these requirements there is a need to pro-vide assurance that service loads cannot cause thepatching e�ciency to fall below an acceptable level dueto fatigue damage in the patch system.

The patch e�ciency could fall due to transverse fa-tigue failure (an unlikely event because of the relativefatigue immunity of these composites when loaded in the®bre direction, ®bre-dominated mode) or to disbondingcaused by fatigue crack growth through one or more ofthe following matrix/interface dominated modes:· The adhesive.· The adhesive (or primer) to metal interface.· The adhesive composite interface.· The surface matrix resin of the composite.· The near-surface plies of the composite.

Fig. 3 illustrates some of these failure modes [2] inpatched test specimen subjected to constant amplitudefatigue, R � ÿ1. In each case the fatigue crack initiatesin the high shear and peel region at the ends of the patch(strip).

In (a) the disbond propagates ®rstly through the ad-hesive very close to the metal interface and then into themiddle of the adhesive layer; in (b) it propagates mainlythrough the adhesive, at approximately �45° as a Mode1 crack (normal to the nominal principal stresses); in (c)in which the b/ep patch is bonded with a epoxy-®lmadhesive (similar to that used in most Australian re-pairs) the crack has a preferred propagation paththrough the Dacron carrier ®bres (presumably due tothe weak ®bre/resin interface) and through the surfacematrix layer of the b/ep ± above the glass ®bre mat layer.

Fig. 2. Details of the location of a 45-mm fatigue crack in the F 111

lower wing-skin, at FASS 281 with an outline of the repair patch,

which is made of 14 ply b/ep �02;�45; 03�S bonded with adhesive.

Patch dimensions are approximately 500 mm spanwise � 350 mm

chordwise.

A. Baker / Composite Structures 47 (1999) 431±443 433

The predominantly matrix path for disbonding is oftenfavoured for b/ep or gr/ep composites.

To discuss the in¯uence of fatigue disbonding of thepatch system the patch can be considered as two zones,Fig. 4(a):· Safe-life zone at the tapered ends of the patch where

damaging strains in the adhesive system are shear andpeel.

· Damage-tolerant zone in the middle where damagingstrains are mainly shear.To prevent rapid loss of the patch due to fatigue

disbonding, strains in the safe-life zone at the patch endsmust not exceed the fatigue damage initiation thresholdfor a signi®cant number of cycles. Even slow initialdisbond growth in this region is unacceptable since oncethe disbond reaches the inner end of the taper region,

Fig. 4. (a) Schematic of an external bonded patch showing safe-life zone (no cracking allowed in patch system) and damage-tolerance zone (slow

crack growth allowed in patch system). (b) Plot of crack growth versus cycles for three panels tested at 138 MPa, showing inset disbond size over

crack. Taken from Ref. [4].

Fig. 3. Micrographs of taper sections through specimen with bonded b/ep or boron/aluminium (b/al) strips (patches) with the following patch-

adhesive combinations (a) b/al-epoxy amine paste, (b) b/al-epoxy polyamide paste and (c) b/ep-epoxy nitrile ®lm. The thickness of the adhesive is

around 0.1 mm. Taken from Ref. [2].


strains dramatically increase and growth will becomerapid. Thus the patch system in this region must bedesigned on a safe-life basis. This is achieved [11] bylimiting strains in the critical regions by design (e.g.,appropriate end tapering, increased adhesive thickness,etc.) and use of appropriate strain allowables. Withoptimum patch design the strain levels in the taperedregion can be designed to be quite low [12] so theyshould not be a concern in most repairs.

The forgoing considerations apply mainly to thickhighly loaded structure. In thin-skin structure disbond-ing in the tapered region can be tolerated [13] as the loadtransfer stresses are acceptably low, even in the absenceof the taper.

In the damage tolerance region, Fig. 4(a), limiteddisbond growth in the patch system is quite acceptable.Damage growth in this region is generally stable, and, ifnot large, has only a fairly small e�ect on reinforcinge�ciency, Fig. 4(b) [4]. Since the patch/metal bond lineis highly stressed in this region signi®cant damagegrowth can be expected, particularly in demanding re-pairs involving thick skins and high loads.

Thus, assuming that the initial residual strength re-quirements are satis®ed, full credit for the patch inslowing crack growth could be justi®ed as follows:· Test/design certi®cation approach: Demonstration

that: (a) strains are below threshold for fatigue dam-age initiation in the safe-life zone; (b) disbond growthis slow and predictable in the damage-tolerant zone,as described in the next section.

· Continuous safety by inspection approach: Based onself-assessment of the patch system integrity usingthe smart patch approach, described later.

2.2. Obtaining patch system fatigue allowables

To minimise certi®cation costs suitably designedbonded joints [6] can be tested to provide a generic database on patch system fatigue allowables. The proposedtests are outlined in Table 1. The joints, shown as insetin Table 1, have been used as the generic joints [2,14] to

assess disbond initiation and growth rate in b/ep patchrepairs to aluminium alloy specimen.

To characterise the allowable strains in the double-overlap joint the approach evaluated is to use the the-oretical or the measured strain range in the adhesive Dcas the damage severity criteria. This is plotted againstdisbond growth rate db/dN, as shown in Fig. 5, where bis the disbond width and N is the number of cycles. Thedotted line is arbitrarily de®ned as the initiationthreshold in this case.

The patch-termination specimen shown in Table 1 isthe more applicable generic joint specimen for cha-racterising the threshold for damage growth in the safe-life zone, since this is representative of the geometry ofthis region and also represents the combination of peeland shear stresses that develops in this region. The totalenergy release rate GT appears to be a suitable param-eter to use as a damage severity criterion.

Since the patch system is expected to be insensitive tolow loads and load sequences the allowables can prob-ably be based directly on elevated stress constant-am-plitude loading rather than using the constant-amplitudedata to predict life under spectrum loading. A conser-vative approach could be to eliminate cycles at low loadsfrom the spectrum (shown in prior testing) not to bedamaging, then elevated all other loads to DLL andbase allowable life on the number of cycles to damageinitiation or acceptable damage growth.

A major issue is the relation of the generic joint to theactual loading situation in a particular repair. Providingthe generic test is similar to the practical situation,limited testing using representative joints could be un-dertaken to ®nd the knockdown factor for the di�er-ence. For example, the F111 repair was subjected tosigni®cant peel stresses [7] which could be simulatedwith an appropriate structural detail.

Table 1

Generic joint test program to obtain repair system allowables

Fig. 5. Plot of disbond growth rate in the adhesive versus shear strain

range obtained from tests on a double overlap joint, shown inset. The

adhesive is FM73, around 0.3-mm thick, the inner adherend is 2024 T6

aluminium, around 3-mm thick, and the outer adherends are b/ep,

around 1-mm thick.


2.3. Validation of patching analysis

The aim is to validate patch design approaches basedon generic repair con®gurations. The approach is basedon tests on a range of generic structural detail speci-mens, as indicated in Table 2. Again limited testing maybe required on representative structural detail specimento validate actual repair design, if this di�ers signi®-cantly from the generic structural detail [6].

Fig. 6(a) depicts a structural detail specimen testedunder constant amplitude loading and Fig. 6(b) plots logda/dN versus log DK predicted using RoseÕs patchingmodel [15]; where a is the crack length and K the stress

intensity. These and other results [10] help to validateuse of this model as a practical approach for predictingK in simple cracking geometries. The growth of dis-bonds in the damage-tolerant zone can be incorporatedin the model [10] based on data from the generic jointtests [14].

Some success has been achieved in using the crackgrowth parameters obtained in the constant amplitudestudies to predict growth of patched cracks underFALSTAF and F111 spectrum loading [16] using amodi®ed crack closure model based on that developedby Newman [17]. The results for FALSTAF loading areshown in Fig. 7 using results taken from Ref. [18]. It canbe seen that the predictions based on the equivalentcrack method are in good agreement with the experi-mental results, noting that the e�ect of the thermal re-sidual stress has been included. It is also interesting tonote that, due to the increased mean stress resultingfrom the thermal residual stress, simple predictionsbased on simply integrating the crack growth equationon a cycle-by-cycle basis, using the steady-state crack-closure stress, are slightly conservative.

In addition to the fatigue studies, for direct com-parison with model predictions [19] the stress intensityK was measured using K gauges bonded as the tip of apatched crack. The specimen used in this study wassimilar to that used in the fatigue experiments. Asshown in Fig. 8 agreement between measured andpredicted K is quite good up to a temperature ofaround 70°C.

Table 2

Validation of design approaches based on generic structural detail

program

Fig. 6. (a) Generic structural detail test specimen for patched cracks in metallic structures used to validate the patching model and obtain crack

growth parameters. (b) Plot of experimental log da/dN versus log DK, predicted from RoseÕs [15] analytical patching model.


3. Justifying credit for patching e�ciency ± environmental

durability concerns

A key issue with the use of adhesive bonded repairs(as with structural adhesive bonding generally) is theneed to assure environmental durability of the adhesivebond.

Environmental durability in bonded repairs impactsdirectly on structural integrity of the repair. Unlike fa-tigue it is not feasible to de®ne either a strain thresholdfor environmental disbonding or a disbond growth rate.With inadequate environmental durability the thresholdfor disbonding can fall to zero so the disbond growthrate can be catastrophic. Thus it is not possible to set asafe inspection interval for this type of degradation.

It is important to note that environmental durabilityis generally a signi®cant problem only with metals not

with polymer±matrix composites such as b/ep or gr/ep.This is because of the moisture sensitivity of the metalinterface if pre-bonding surface preparation is inade-quate. Hydration of an unstable surface metal oxideresults in a weak, easily disrupted interface.

The key point is that NDI techniques can detect onlyobvious disbonds, voids, etc. Bonds with potentiallypoor environmental durability cannot be detected aslong as intimate contact at the bond interface is main-tained. Developing an NDI capability to detect weakbonds is a very long-term prospect and may well proveinfeasible. However, pre-bond NDI, that is, the devel-opment of techniques to check that the surfaces arecorrectly treated prior to bonding is feasible, as dis-cussed later.

3.1. Assurance of patch system environmental durability

The requirement is to provide assurance at an ac-ceptable level of probability that patch bond failure willnot occur in the service environment during the requiredlifetime of the repaired component ± or alternativelythat the risk of bond failure is acceptable for the re-quired service life.

This assurance revolves about quality control of thepre-bonding process because:· This is the most critical process in the repair and is of-

ten applied under di�cult conditions.· The danger of incorrect application or surface con-

tamination, for example, by organic species fromfuel, etc, in the repair environment is signi®cant.Thus a simple-to-apply (non-hazardous) surface

treatment, such as the silane process [2,20], increasesapplication reliability, even though other more complexrepair surface treatments, for example, those based onphosphoric acid anodising, may provide bonds withmarginally superior environmental durability.

To provide the required assurance a suitable test isrequired to: (a) assess the process evaluation, (b) assessthat it has been correctly applied, (c) qualify technicians.It is proposed that the Boeing wedge test (BWT), de-picted in Fig. 9, ful®ls this requirement for the followingreasons:· The BWT has been a US industry standard for adhe-

sively bonded airframe components for many yearsand is an ASTM test.

· The BWT is a simple low-cost test which requires aminimum of equipment.

· Australian experience based on assurance of bonddurability through BWT testing has been very posi-tive, as described in the next section. Assurance isnow based on RAAF Engineering Standard C5033Composite Materials and Adhesive Bonded Repairs[9].However, there are questions concerning whether the

severity of the BWT is appropriate and how to quantify

Fig. 7. Comparison between experimental and predicted growth

behaviour of patched cracks under FALSTAFF spectrum.

Fig. 8. Comparison of experimental and theoretical results for K1 for

a � 30 mm, patched specimen. The models used for the theoretical

predictions are RoseÕs model and a related model developed more

recently by Rose and Wang. Taken from Ref. [19]


it in terms of a failure probability. Neither of thesequestions can satisfactorily be answered at present.

E�orts are being made in Australia to establish if therequired statistical data are available in the literature.There should also be a suitable statistical database heldby the major aircraft manufacturers, but this is likely tobe of a proprietary nature. At present the only infor-mation on which a decision can be based is on our ex-tensive service experience discussed in Section 3.2.

Australia uses a severe version of the standard BWT.For example, an acceptable test result is generally lessthan 19 mm growth of the disbond in 1 h; whilst theAustralian requirement is less than 5 mm growth in 24 hor 7 mm in 48 h.

More severe tests can be performed. These include thedouble-cantilever beam test, the constant-compliancecantilever-beam test [21] and the slow strain-rate test[22]. The advantage of these is that they are quantitativein that a Mode 1 fracture energy can be measured andcorrelated with exposure conditions and time. This isgenerally not possible with the BWT as with tougheradhesives some plastic bending of the metal adherendarms occurs when the wedge is driven in.

The disadvantages of the alternative tests are thatthey are much more costly to make and test comparedwith the BWT and they (probably) cannot be correlatedwith service behaviour, since a suitable database forthese tests is unavailable. The ability to quantify a du-rability parameter, e.g., Mode 1 strain energy releaserate G1, is of limited use since it is not a useable designparameter. Thus these tests are more suitable for re-search purposes, for example, in the evaluation of newsurface treatments.

Another issue is the value of a BWT made during therepair process, rather than simply to qualify techniciansand processes. Separate test specimens made up inparallel and, as far as possible, under the conditions of

the repair cannot capture all the variables associatedwith the repair process; for example, in the exact parentmaterial, in the type and degree of contamination and inthe cure cycle. Nevertheless, it is reasonable to assumethat a good BWT result from the parallel specimen willprovide considerable extra con®dence in the repairprocedure.

Other than the BWT for quality control of surfacetreatment there is also the possibility of pre-bond NDI,including simple wetting measurements. However, moresophisticated measurements for assessing, for example,contamination, degree of surface blasting [23], andproduction of the correct surface condition are alsofeasible and can be adapted for ®eld use. These includesurface potential measurement and infrared (IR) surfaceanalysis.

3.2. Australian experience on service durability

Table 3 provides details on Australian bonded com-posite repair applications. Many repairs have survivedover 20 years of service, without evidence of environ-mental or fatigue durability problems.

However, a few environmental durability failureswere found in the Mirage III wing repairs. In these re-pairs the surface treatment was based on the phosphoricacid gel-anodising process (PAA). The failures were at-tributed to (a) the extreme porosity, which developed inthe bond line of some of these repairs caused by mois-ture absorption under the tropical conditions in whichthese repairs were applied and (b) the absence of cor-rosion inhibiting primer. Use of staged adhesive andmore careful drying of the surfaces prior to bondingwould have reduced the moisture problems.

Fatigue durability of the patch system has been aproblem in only one case: the F111 wing-pivot-®tting b/ep doublers which were primarily designed to reducestrains in the ®tting during a cold ()40°C) proof loadtest ± this they did successfully. A secondary aim for thedoublers was to reduce fatigue strains in the ®tting inservice to increase the inspection interval (by a factor ofaround 3); however, fatigue failures are being experi-enced. Failure occurs mainly in the surface resin of theb/ep, as seen in part of Fig. 3(c). Whilst analysis of thecause of failure is not complete, the failures are thoughtto result from use of an excessive taper angle at the endsof the doubler resulting in adhesive strains above thefatigue threshold.

4. Justifying credit for patching e�ciency ± the smart

patch approach

An alternative to certi®cation procedures based ontesting is the smart patch approach [6,24] illustrated inFig. 10. The aim here is to develop a capability for

Fig. 9. The BWT to assess bond durability. The metallic adherends are

made the aluminium alloy under investigation, generally about 3-mm

thick bonded with the appropriate adhesive. The wedge is driven in

under standard conditions.


continuous autonomous (self) monitoring of patch per-formance to ensure that the patch is performing in ser-vice as required. Although the need to follow the correctpatch design and process procedures is unchanged, the``smart'' approach should allow considerable relaxationin the certi®cation requirements. However, failure of thepatch system will still be a costly and therefore a highlyundesirable event.

The most direct approach to assess the ``health'' ofthe patch system is to measure the level of load transferin the safe-life zone, Fig. 4. This approach was suc-cessfully demonstrated on test specimen during the de-velopment of the b/ep doubler for the F111 wing-pivot

®tting [25], using resistance strain gauges bonded at theends of the tapered region, but was premature for in-¯ight application. The monitoring could also includeautomatic assessment of cracking in the parent structureand disbond growth in the patch system in the damagetolerance zone.

The demonstrator smart approach being developedfor ¯ight trials [24] is also based on strain gauges,bonded to the ends of the taper region, Fig. 11, and onthe surface of the component away from the patch andmonitoring the ratio (patch strains)/(strain in the com-ponent) during service life. Any decrease in the ``patchhealth'' ratio is an indication of disbonding of the patch

Table 3

Australian bonded patch repairs to military aircraft, experience on environmental durabilitya

Aircraft Problem Surface treatment (patch

system)

Remarks on bond durability

C130 Stress corrosion cracked sti�eners in

wing, aluminium alloy 7075

GB initially GB + S later

(b/ep + FM73)

Over 20 years of service. No bond

durability problems where bonding

carried out as speci®ed

Mirage III Fatigue cracking in lower wing skin,

aluminium alloy AU4SG

PANTA (b/ep + FM73) 180 wings repaired or reinforced. Eight

bond durability problems over around

eight years. Failures were associated with

adhesive voiding caused by extreme

humidity in the tropical repair station

F111-C Secondary bending in wing pivot ®ttings

leading to a fatigue problem. Steel D6ac

fastened to aluminium alloy wing skin

GB + S (b/ep + FM73) No bond durability failures to steel or

aluminium surface over 10 years

F111-C Stress corrosion cracking in weapon bay

longeron ¯ange, aluminium alloy 7075T6

GB + S (gr/ep cloth + ep) Over 10 aircraft repaired. No bond

durability problems over around eight

years

F-111-C Stress corrosion cracking in longeron

adjacent to refuel receptacle, 7049-T6

GB + S (b/ep + EA 9321) Over 10 aircraft repaired. No bond

durability problems in eight years

F-111-C Metal-to-metal and sandwich structure

repairs. RAAF adopted GB + S and

changed to FM 300 adhesive in 1992

GB + S FM300 FM 73,

EA 9321

No bond durability failures in over seven

years

F-111-C Pork-chop panel (lower fuselage). Panels

rebuilt after repeated in-service failures

GB + S FM 300 Repeat rebuild rate reduced from 95% to

0. No bond durability failures in seven

years

C-141 (USAF) Fatigue cracking in wing riser weep

holes, 7075T6

GB + S + P (b/ep +

FM73)

No bond durability problems around ®ve

years

P-3C Full depth corrosion damage in

horizontal tail, aluminium alloy 7075 T6

GB + S (al alloy

+ FM73)

No bond durability problems over

around 10 years

Boeing 747 Simulated repairs to several regions

including fuselage lap-joint, wing leading

edge, trailing edge ¯ap and engine thrust

reverser cowl

GB + S + P (b/

ep + FM73 or acrylic)

Demonstrator repairs; 37 000 ¯ying

hours, 7020 landings over a nine-year

period with no signi®cant bond

durability problems

Sea King helicopter Fatigue crack in frame GB + S + P (b/ep +

FM73)

Operated in am o�shore ship-borne

environment for four years with no

problems

F-111-C Fatigue cracking in lower wing skin at

fuel ¯ow hole under forward auxiliary

spar

GB + SB/ep No bond durability problems in over 2.5

years of service

a GB � alumina grit blasting; S � Silane; P � Primer; PANTA � Phosphoric Acid Non-Tank Anodising; b/ep � boron/epoxy; gr/ep � graphite/

epoxy; AF 126, FM73, FM300 epoxy-nitrile ®lm adhesives; ep � epoxy-paste adhesive; acrylic � toughened acrylic adhesive Flexon 241 EA 9321

epoxy paste adhesive.


in this critical end region. In this approach there is norequirement for measurement of the actual loading,disbonding is indicated by the reduction in relativestrain. The gauges are automatically monitored by an``on-board'' miniaturised system that provides processed

information. Data are transmitted to an external com-puter by an IR link.

The prototype device, shown in Fig. 12, was notoptimised for size and weight; however, it easily ®ts inthe palm of an average hand and has a mass of about25 g and is powered by a single 3-V battery. Futureapproaches will be based on micro electro mechanicalsystem (MEMS) [26] and will be very much smaller andmore robust with much lower power requirements. Thelong-term aim is for a fully embedded system. How-ever, in the near future the system (still fully self-contained) would have to be located in a nearby smallcavity.

Experiments were undertaken to demonstrate thisconcept on the skin-doubler specimen depicted inFig. 11. Resistance-foil strain gauges are located on thefar ®eld and the fourth step of the boron doubler oneach side of this specimen. Two 28-lm thick piezoelec-tric ®lm (PVDF) sensors, one located in the far-®eld onthe aluminium adherend (parent structure) and the otherlocated on the ®rst two steps of the doubler. Fig. 13shows the patch health ratio as a function of loadingcycles. It is seen that the device satisfactorily monitorsthe damage growth at the end of the doubler.

IR (rather than radio waves) is being used for thedata transmission. IR has the signi®cant advantage ofnot interfering with (or being interfered by) any of theaircraft electrical functions, but may be di�cult to usewhere the repair is deep in the airframe structure.However, it may prove viable to ``pipe'' the IR to anaccessible region using optical ®bres.

The viability of this smart approach revolves largelyon how to prove its reliability under the aircraft tem-perature and environmental operating conditions. Thisis the reason why ¯ight trials will be so important. Thesetrials are planned within the next year to monitor the

Fig. 12. Photographs and ¯ow chart of the demonstrator experimental set-up.

Fig. 11. Schematic of specimen showing strain gauge and piezoelectric

®lm sensor locations on the steps at the ends of the doubler.

Fig. 10. Schematic diagram of the smart patch concept.


performance of a b/ep doubler system recently devel-oped [27] to extend the life of F/A-18 aileron hinges.

An alternative smart approach, being developed byWilson [28], is based on the use of special chemicalsensors embedded in the adhesive layer. These sensorsdetect chemical species produced at the metal surfaceduring the process of bond degradation. This approach,complementary to the strain approach previouslydiscussed, should provide warning of impending bonddeterioration.

5. Discussion

A decision chart for management by slow-crackgrowth is provided in Fig. 14, based on the previous

Fig. 14. Decision chart for certi®cation of bonded patch repairs to cracked structure. The actions inside the dotted ellipse represent the conservative

Fail-Safe approach.

Fig. 13. Plot of variation of patch state of health (ratio of critical

region strain/far-®eld strain) with increasing number of applied load-

ing cycles.


discussion. This shows that certi®cation requirementscan be made much less stringent by the smart patchapproach, as immanent loss or degradation of patchreinforcing ability will be immediately detected. How-ever, where certi®cation is based on this approach thekey requirement becomes proving the smart system re-liability in the aircraft operating environment. Gener-ally, the smart patch approach appears technicallyfeasible but will be economically viable mainly for re-pairs to costly primary structure

Where the smart patch approach is not used the keyissue is how to provide assurance that the probability ofunacceptable loss in patching e�ciency is acceptablylow. If bond environmental durability is shown not to bea concern then management can be based on patchsystem fatigue allowables, giving full credit for the patchin slowing crack growth and recovering residualstrength. However, bond degradation is generally themajor concern so the main issue here is of providingcon®dence that environmental degradation of the patchsystem will not occur. If this is not possible in the ab-sence of the smart patch approach the repair must bebased on the Fail-Safe approach, giving no credit to thepatch.

To provide assurance of bond durability a validatedaccelerated test of bond quality is required. It is pro-posed that the BWT or some derivative of it is a suitabletest. The BWT can be used to qualify surface treatmentprocesses and technicians, as in the C141 repair program[29] and can be used for in-line quality control purposes.The question of what BWT result is and what is notacceptable is the current challenge. Su�cient informa-tion on service experience may exist in the literature onbonded aircraft components and in the Australian ex-perience on bonded repairs to help provide the answer.

6. Conclusions

The certi®cation requirements to permit managementof bonded composite repairs by allowing full credit forthe presence of the patch in slowing crack growth andrecovering residual strength have been addressed.

The major structural integrity issue is to assure thatdisbonding of the patch due to environmental degra-dation will not occur during the required service life.

Methods based on a combined safe-life/damage tol-erance approach can be applied to the patch systemwhere bond environment durability is not a concern,based on patch system fatigue allowable and knock-down factors.

It is proposed that the BWT should be adopted as theprincipal accelerated test for quality control for bondingsurface treatment. The main challenge is to correlate theBWT with a probability of failure, possibly based onservice experience.

Use of the smart patch approach, based on self-monitoring would considerably alleviate the certi®cationrequirements, since loss of patching e�ciency would bedetected automatically. However, this approach bringsits own problem of reliability assurance.

Acknowledgements

The author would like to acknowledge many stimu-lating discussions with his colleagues on this topic in-cluding: Max Davis, Loris Molent, Kevin Walker andDrs. Richard Chester, Steve Galea, Chun Wang, FrancisRose and David Bond.

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