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Repair of mirage III aircraft using the BFRP crack-patching technique
Theoretical and Applied Fracture Mechanics 2 (1984) 1-15 North-Holland
1
REPAIR OF MIRAGE III AIRCRAFT USING THE BFRP CRACK-PATCHING TECHNIQUE
A.A. BAKER, R.J. CALLINAN, M.J. DAVIS, R. JONES and J.G. WILLIAMS Defence Science and Technology Organisation, Aeronautical Research Laboratories, Melbourne, Vic. 3001, Australia
The BFRP crack-patching technique has been applied to the field repair of fatigue cracks in the aluminium alloy wing skins of Mirage III fighter aircraft. Finite-element procedures were used in patch design. The repair was qualified using fatigue-crack propagation studies on panels simulating the cracked and repaired area. A field support unit was designed to allow repairs to be carried out by air force personnel during routine maintenance of the aircraft. To date over 150 patches have been applied and nearly three years of operational history gained. While some crack growth was observed after repair of a few wings, the patch stopped further growth and no wing skin has required further repair.
1. Introduction
Aeronautical Research Laboratories (Australia) have developed methods for repair of cracked metallic aircraft components using an adhesively bonded boron fibre reinforced plastic (BFRP) patch. This report describes experience in developing and applying a BFRP patch to fatigue cracks in the wing skin of Mirage III fighter aircraft. Previously the technique had been applied successfully to stress corrosion cracks in Hercules wing planks [1] and fatigue cracks in Macchi landing wheels [2]. Traditional repair techniques using metallic patches require many additional holes to be drilled for fasteners to provide efficient load transfer into the repair. With adhesively bonded repairs, this is unnecessary. The advantages of using BFRP for the patch material include excellent resistance to corrosion and cyclic loading and ease of production even with complex curvatures. The patch can be produced with the fibres spanning the crack to restrict crack opening under load and hence to prevent crack growth by reducing the stress intensity. The high directional stiffness prevents unnecessary and undesirable increases in component rigidity in directions other than across the crack. Carbon fibre reinforced plastic (CFRP) could also have been used, but BFRP was selected, although more expensive, for its better stiffness and fatigue resistance and its higher thermal expansion coefficient. Table 1 gives values for physical constants for repair materials. Thermal stresses result-
ing from mismatch in thermal coefficient can introduce high residual stresses in systems bonded in place using heat-cured adhesives. As BFRP has negligible electrical conductivity, readily available eddy-current procedures can be used to detect and monitor cracks beneath the patch. Recently, fatigue cracks were discovered in the lower wing skins, close to the main spar of some Australian Mirage aircraft, Figs. 1 and 2. It was decided, in consultation with RAAF, that BFRP crack-patching would be an effective solution for this problem, because the repair would (i) cause no mechanical damage to the skin (i.e. no fastener holes), (ii) cause no strain elevation in the spar, since reinforcement need only occur across the crack, (iii) allow the use of conventional eddy-current procedures to check for crack growth, and (iv) allow implementation in the field during normal servicing, thereby minimising unavailability of the aircraft. Because of the significance of the cracking and the long life desired of the repair, detailed design studies and a considerable amount of further research and development were required before the repair could be implemented. Some aspects of this work are reported here.
2. Patch design and analysis
The fatigue cracks initiated near the fuel decant hole in the lower wing skin, close to the intersection of the main spar and root rib (Fig. 1). The skin consists of aluminium alloy AU4SG, similar
0167-8zlz12/84/$3.00 © 1984, Elsevier Science Publishers B.V. (North-Holland)
2
A.A. Baker et al. / B F R P Crack- patching technique
Table 1 Values of material constants Material
Aiuminium a Alloy (2014-T6) CFRP BFRP AF-126 Adhesive
Elastic Modulus E n (GPa)
Thermal Expansion Coeff. x 1 0 - 6 / o C
Specific Gravity
Strain 'g
72 130 204 0.7 (20 o C) (shear)
23 negligible 5.0 -
2.8 1.5 2.0 -
0.6 (yield) 1.1 (fracture) 0.6 (fracture) -
" The alloys used for experimentation were 2024-T3 and L104. L104 is British Standard equivalent of the French AU4SG, from which the aircraft skin was m~ide. These last two materials are equivalent to 2014-T6 aluminium alloy.
to 2014T6, about 3.5 mm thick and is covered by a fairing. The wing skin forms part of the wing torsion box which is designed to be in shear. This stress state was confirmed by experimental strain studies at ARL. The state of shear results in the fatigue cracks propagating at 45 o to the spar (Fig. 2). For design purposes the maximum size of the
crack (including the fuel decant hole) was taken as 111 ram. Hight loads were estimated from straingauge data and from the manufacturer's stress analysis. The BFRP patch (Fig. 2) is a unidirectional laminate with fibres running perpendicular to the direction of crack propagation. The patch contains
Fig. 1. Silhouette of Mirage III aircraft showing where the fatigue cracks developed in some aircraft. Insert: Fuel decant hole region, showing the nature of the fatigue cracking.
A,A. Baker et al. / BFRP Cr~ck-patching technique Z
C
t -
Tzx
Element of adhesive
=X
Boron patch
Adhesive
Aluminium alloyskin Crack
Fig. 3. Crossectionaldiagram of the seven-layer,internally-steppedBFRP patch.
seven layers of BFRP and is internally stepped, i.e. the largest layer is on the outside to reduce interlaminar shear and peel stresses. The shear stress reduction arises largely as a result of the increased adhesive thickness near the edge of each ply. The transverse stress (peel stress) reduction in the inner plies arises as a result of the opposing stress resulting from the geometry of the plies in the overlap region (see Fig. 3). The beneficial effect of this approach was observed in some preliminary fatigue tests. The proximity of the spar bolts and the root rib bolts to the decant hole necessitated closer spacing of the layer steps in this region. The attachment holes for the decant hole cover were preformed in the patch; the decant hole opening was greatly reduced in size by the patch, so as to obtain maximum reinforcement efficiency. The main objectives of the analytical procedure were to assess the reduction in stress-intensity in the skin, reveal any undue strain elevation in the spar and to estimate the maximum levels of stress in the BFRP patch and adhesive layer. Ideally, a sufficient reduction in stress-intensity should be obtained without exceeding the materials allowables in the patch system, the most critical of which
is the adhesive shear strength under fatigue loading in the operating environment.
Decant ilumbly
Inb'd
SlUr
i
No 1 Bolt
il~
/
A
~.
.-~~_.= = --,lilill @ 0"" = --- O Ollcimthole Ill/ill . / in i . ~Ikin i . , ,~ , Iil~,,il
//0
Hole in petch ~
°k
|
i
I
to .',','.
\
@ ®
o
o
~/ .
.
,/,+," ~'.','%,"--
111 !11 i,iI! I
®
i II 7.'/ #'---S," --III/11
-'111 7I
® Fig. 2. Schematic diagram of fuel decant hole region with outline of BFRP patch superimposed; note the internal stepping of the patch.
4
A.A. Baker et aL / BFRP Crack-patching technique
3. Finite element analysis Initially a detailed, two-dimensional finite element stress analysis of the cracked and unpatched region was undertaken. This model consisted of 252 quadrilateral membrane elements, and 68 constant strain triangular elements representing the wing skin, which is 3.39 m m thick, 16 flange elements representing the integrally-milled stiffeners, and two crack-tip elements which are described in the appendix. The resulting stress-intensity factor, assuming loads typical of 7.5 g flight manoeuvre were K 1 = 72 MPa ¢ m , K 2 = 3.3 MPa ¢ ~ at the tip nearest to the spar, and K 1 --68 MPa V~mm,K 2 = 1.0 MPa ~ at the tip nearest to the root rib. The effect of the patch was estimated by adding a finite-element representation of various BFRP patches to the cracked model (Fig. 4) utilizing a special element described in [3]; this model allows for the separate response of the wing skin, patch and adhesive. In the specific case when the xy axis system coincides with the axis of orthotropy of the patch, the relationship between displacements and shear stresses in the adhesive layer reduce to: ~ x = (Uo - u s ) / ( t J G a
%,. = (% - v s ) / ( t J G
+ 3tJ8Gs + 3to/8G13)' ~ + 3tJ8G,
+ 3to/8G23 )
(1)
which is similar to the relationship developed in [4]. Here %x and ~'~.~are the shear stresses developed in the adhesive; t a, t s and t o are the thicknesses of the adhesive, wing-skin and patch respectively; (u 0, %) and (u s, vs) are the x, y displacements at the mid surfaces of the wing-skin and the patch; Gs and Ga are the shear moduli of the skin and the adhesive respectively, and Gl~ and G23 are the transverse shear moduli of the patch. The through-the-thickness variation of the inplane displacements are given by the following formulae, when the transverse shear moduli of the patch are equal, viz: u = u o + "rsx(Z(t s + t a - t o - z / 2 ) -(t s + ta + to/2)(ts/2
= (u o - 3~s~to/8Go)(Z - ts)/t a + (u s + 3rs~tJ8Gs)(ts
= u~ + (z 2 - t ~ / 4 ) ~ s x / 2 G s t s in the sheet
(2)
and a similar expression for v ( x , y , z ) . As a result of this formulation the value of the
o 8~
Fig. 4. Finite-element mesh for the fuel decant hole region.
+ ta - z ) / t a
in the adhesive
Z
---.L / , ' l / l l
+ 3to/4))Got o
in the patch
Spar
/
+ tJ2
A.A. Baker et al.
/ B F R P Crack:patching technique
stress intensity factor varies through the thickness of the wing skin. This variation may be derived from the previous equations and is given by Kl = K , p ( 1 -
X( z2 - t 2 / 4 ) / 2 G s t ~ ) ,
(3)
where KIp is the value of the stress intensity factor at the mid-surface of the wing skin after patching and X is the spring constant X = 1/(tJG
a + 3 t J 8 G s + 3to/8Go).
(4)
The geometry of the patch was constrained by the presence of the spar, root rib, and the fairing attachment. Furthermore the plan area of the patch was limited by the requirement to minimise pressurisation loads during application of the patch. This resulted in a fixed geometric shape with the prime variables being the patch and adhesive thickness and step size between plies. With these restrictions, six B F R P patch configurations were considered each with the plan form
shown in Fig. 2. Each patch was modelled using approximately 380 of the bonded elements, which are triangular in plan form, described above and in greater detail in reference [3]. All of the six patches considered were unidirectional laminates and were internally stepped, i.e. with the longest ply on the outside. The fibre direction was at ninety degrees to the crack. Initially it was uncertain if carrying the fibres over the drain hole was necessary, or how frequently the drain hole was used in service. As a result, in three of the patches considered a hole was left so as not to interfere with the draining of the wing. In the other three patches varying amounts of the hole were covered. In one case one third of the total area of the hole was covered while in the other two cases virtually all of the hole was
covered.
For each of the patches the maximum shear stresses in the adhesive bonding the patch to the wing skin occurred at points A, B, C and D (see
2.0
Oc/O u .9
! !
0')
0
I 3
I ~ i I I 6 9 12 15 Distance from Bolt Hole Number 1 (ram)
I 18
Fig. 5. Stress distribution in the aluminium skin. The curves are for the stress, o, for ratios between three cases denoted by subscripts; p = patched and cracked, u ~ unpatched and uncracked, and c = cracked and unpatched.
6
A.A. Baker et al. / BFRP ('rack-patching technique
Fig. 2). The maximum stresses in the fibres occurred at point D for the patches with a hole in the patch, and in the fibres over the hole in the patches with the hole partially covered. The values for these stresses, along with the percentage reduction of the stress intensity factor K 1 at each tip, achieved by each patch are shown in Table 2. Here Klu and K~p are the values of the stress intensity factors before and after patching respectively, the value of Kip being at the mid surface of the wing skin. All of the six patches achieve a reduction in the stress intensity factor K 1 of at least 91%. Consequently they would all significantly retard crack growth. Similarly for all of the patches the fibre strains are below the m a x i m u m working level of 0.005, which corresponds to a stress of 1000 MPa, although of the six patches numbers 5 and 6 have by far the greatest factors of safety. As a result as shown in [7] the patch design was finally chosen primarily on the basis of the magnitude of the shear stresses developed in the adhesive. On this basis patch numbers 1, 2, 3, and 4 may be rejected. The two remaining patches are patch numbers 5 and 6. Of the two, Table 2 shows that the adhesive shear stresses along the edges of the patch, are substantially higher for patch number 5 than for patch number 6, although both are below the
threshold value for fatigue damage for this adhesive. As a result patch 6 is much less likely to suffer fatigue damage to the adhesive bond. Hence patch number 6 was adopted as the final repair. At locations C and D in patch 6 the shear stress in the adhesive is sufficiently high so as to cause concern over the possibility of fatigue damage occurring in the adhesive. However these high values occur in the interior of the patch at the intersection of the crack with the drain hole, and are very localized. As a result any damage which does occur should not spread and, as shown in [5], should have virtually no effect on the stress intensity factors at the crack tips or on the fibre stresses. The effect of this patch on the maximum principal stress in the skin along the line of attachment to the spar (Fig. 5) showed that the patch effectively restores the stress distribution along the spar to that in a wing with an uncracked panel. Consideration was also given to thermal and residual stress effects in the aluminium skin and to thermal-fatigue effects in the adhesive due to the use of an adhesive curing at elevated temperature ( 120°C) and the mismatch in thermal expansion between the aluminium alloy and BFRP. It was concluded [6,7,8] that these effects posed no serious problems in this particular repair. Table 2 shows the value of the stress intensity
Table 2 Effect of drain hole patch 7.5 g load case Patch Number Maximum path thickness (mm) Adhesive thickness (mm) Thickness of first Layer (mm) Covering of drain hole Adhesive shear stress (MPa) at points: A B C D Maximum fibre stress (MPa) Reduction in stress intensity factor K l, i.e. 1 - Kip/Klu. at: 1) Spar tip 2) Root rib tip
1
2
3
4
5
6
0.762
0.762
0.762
0.889
0.889
0.889
0.102
0.102
0.102
0.203
0.203
0.203
0.127 open
0.254 open
0.127 1/3 covered
0.254 open
29 55 181 153
43 79 179 153
29 55 164 131
31 58 120 98
26 42 63 50
18 30 64 51
953
930
760
911
450
455
91% 99%
91% 99%
91% 99%
91% 99%
0.254 substantially covered
92% 99%
0.127 substantially covered
91% 99%
A.A. Baker et al / BFRP Crad~-patching technique
reduction at each crack tip. After patching, however, the maximum value of the stress intensity factor occurs at the free surface of wing skin (see equation 3 for z = 0) and has the value K_lp(1 + ts/SGs). For this patch Kip = 6,5 MPa ¢ m , and the factor (1 + tJ8G~) is--approximately (1 + tsGa/8Gat,)--= 1.075 so that the maximum value of the stress intensity factor becomes 7 MPa v/~ and occurs on the free surface on the inside of the skin. These values lie in the vicinity of the threshold value for fatigue crack growth. In the above analysis bending of the wing skin was assumed to be negligible, as was consistent with strain gauge readings on the wing skin. However, the presence of bending could result in the threshold value of the stress intensity factor being exceeded on the inner surface of the wing skin.
4. Repair qualification The BFRP repair was qualified mainly by two series of tests (a) a strain survey on a Mirage wing with a patch installed to show that the patch did not significantly elevate the local strain in the spar,
and (b) fatigue tests on aluminium alloy panels configured to simulate the cracked area in the wing (Fig. 6). The aims of the fatigue test were to check that (i) the predicted reduction in stress-intensity could be achieved, and (ii) the patch/adhesive system could endure the fatigue loading. The strain survey showed that the strain levels in the spar after patching increased by less than 3% indicating that no significant strain elevation occurs in the spar after patching. The fatigue results are given below. Fatigue crack propagation studies
Tension-tension fatigue tests were undertaken on the aluminium alloy panels using either constant load cycling or block load cycling. Only the principal tensile stress in the wing could be simulated. The tests were carried out in laboratory environment at a temperature of 23°C. Initially, due to unavailability of 2014T6 materials, tests were carried out on aluminium alloy 2024T3 panel. Although this alloy has significantly greater resistance to crack propagation than the wing skin alloy AU4SG, it was considered that these tests would show whether or not the stress-intensity was
Sealant Decanthole and associated fastenerholes /
Patch pat..chregion
Decant
assembly o.
. . . .
Crack
Sandwichpanel ~
Porousteflon--/
c~ed 91ass
i 1,0,°°l I
7
~I :3.66rnm
Fig. 6. Details of 2024T3 aluminium alloy wing-skin simulation panels used for crack-propagation studies, showing position of patch and vacuum box restraint.
8
A.A. Baker et al. / B F R P Crack-patching technique
reduced by patching and if the fatigue characteristics of the adhesive layer were satisfactory. Cracks were initiated from saw-cuts and propagated to about 100 mm total length prior to BFRP patching, using the procedures described below. During application of the patch, the panels were constrained at their ends to simulate the constraint to thermal expansion that would be experienced in the wing-skin under practical repair conditions. This rate of crack growth was monitored during the initial test by direct observation with a travelling vernier microscope. After patching, eddy-current procedures were used to monitor the crack under the patch. The unit used was a Sperry Products EM-3300 unit. The accuracy with which the crack length could be determined was + 1 mm. In the early tests local bending of patched panels occurred under load, due to local displacement of the neutral axis of the panel by the patch. The effect negated the reinforcing effect of the patch to a large extent. Strain-gauge tests confirmed the presence of substantial strains in the panel on the opposite side to the patch. In the actual wing-skin, local stiffening by the surrounding structure, particularly the spar, root rib and internal stiffeners, would largely prevent bending and a method was therefore devised to minimise its occurrence in the test. This procedure used atmospheric pressure (by evacuating the region between the back of the test panel and the sandwich panel). A thin layer of porous, fluorocarbon-
120
coated fibre-glass between the test panel and the sandwich panel acts to allow air removal in the gap between the test panel and sandwich panel and minimised the frictional load. Strain-gauge studies showed that the bending effects were then reduced by about 50%. Using the vacuum restraint, substantial increases in life prior to crack growth were achieved; Fig. 7 plots crack growth curves for three 2024T3 specimens subjected to constant load cycling at an applied stress of 91 MPa, (equivalent to a nominal 4g manoeuvre load). In all cases initial crack growth rates indicated that failure of the panels (without the repair) would have occurred within a further 104 cycles. Panels were tested under more realistic loading conditions using a block loading sequence. The flight spectrum used was obtained from flight data recordings. This spectrum was truncated at zero and divided into 9 load blocks. The number of cycles and corresponding peak stresses are given in Table 3~ In this table load block numbers refer to the order in which loads are applied. Each program sequence is approximately equivalent to one year of aircraft usage. Table 3 represents 0.6 of a program sequence. Substantial life extensions for the cracked 2024T3 alloy panels, (Fig. 8) indicate that, provided local bending is minimised, the BFRP patch system provides a significant stressintensity reduction. The actual life extension probably depends largely on the degree to which local
Patch Applied
E ~i
Thisspecimen, Load increa~ledby 38%
/
/
,s/
s,~ */
/
= 30MPa ,V'~
No growth afler 105 cycles
80
=0.1 ak Nominal Stress 91MPa
~.O
I ~ 'tO r
I 160
I
I 180
i
/L v
l 370
l
l 390
NO of Cycles (x|0 -3) Fig. 7. Crack length versus cycles for 2024T3 wing-skin simulation panels subjected to constant load cycling; stress intensity range prior to patching is indicated.
A.A. Baker et al. / B F R P Crack-patching technique ?
Growth detected
120
/ Test terminated tch applied L104 Specimen
~0
0
I
I
20
~0
I
A //,
I
V
100
I
I
120
No. of programmes Fig. 8. Crack length versus n u m b e r of programs of block loading for 2024T3 and 2014T6 wing-skin simulation panels.
bending is reduced in practice. The long life of the repair indicates that the patch/adhesive system is capable of withstanding sustained fatigue loading, at least in a laboratory environment. Currently tests are being conducted using panels of 2014T6 alloy, the material used in the aircraft wing skin. These panels are prepared with the rolling direction oriented similarly to that in the wing skin. Tests to date show that crack propagate in these panels very much faster than in 2024T3; the results for one such panel are included in Fig. 8. It is worth noting that the notched 2014T6 panels could not withstand one complete Table 3 Load spectrum Load block number
N u m b e r of cycles
Peak stress (MPa)
1 2 3 4 5 6 7 8 9
9999 1027 870 341 579 381 31 3 0.32
57.9 69.2 79.0 88.2 103.0 121.0 139.3 151.0 156.1
program of the loading sequence; crack growth to the 'repair size' occurred at load level 5 ( - 5 g ) , whereas the loading sequence goes to load level 9 ( - 9g). This observation suggests that the stressing level employed in these tests may be higher than that encountered by this region in the aircraft, since, in practice, cracks are not found to grow this rapidly.
5. Selection of materials and processes Adhesive
An epoxy-nitrile structural adhesive, 3 M's AF126, was chosen on the basis of previous background studies into suitable adhesives for crack patching, which included work on fatigue and stress relaxation [5], and practical experience on repairs to stress corrosion cracks in Hercules aircraft over a period of more than four years [2]. AF126 cures at about 120°C and requires a pressure of 70-350 kPa. Originally, an adhesive curing at a lower temperature and pressure was sought to simplify patch application and minimise thermal and residual stress problems. Unfortunately, none of those studied was suitable; future development
10
A.A. Baker et al. / B F R P Crack-patching technique
of a suitable adhesive curing at lower temperatures and pressures would greatly aid patching technology.
Cr~k
C--
Adher~ve layer
Metal adherend WqN:~etest specimen
In previous repairs, non-chemical surface-treatments of the aluminium alloy component were used to simplify patch application procedures and to avoid the danger of stress-corrosion cracking. This treatment was degreasing, followed by gritblasting with 50 #m alumina powder; the same procedure was used for preparing the BFRP patch. However, it was considered that the present repair required a higher level of environmental durability than would be provided by this procedure, because of the critical nature of the cracking and the long life required of the repair. It was decided therefore, to evaluate the phosphoric acid gel anodising procedure recently developed by the Boeing Aircraft Co., for field repairs to adhesively bonded aluminium components [9]. Table 4 lists details of the solution used and the anodising conditions. Boeing wedge tests were used to compare the effectiveness of various surface treatments, including alumina grit-blasting, gel anodising, and standard tank anodising (factory procedure). The results (Fig. 9) show that gel anodising gives adequate durability. Although the use of primer gave further improvement, its application was considered too difficult to control under field conditions. Prior to final acceptance of the gel anodising procedure, further wedge tests showed the bond to be resistant to fuel immersion, the main environmental contaminant in the present repair. Other tests showed that the gel anodise procedure did not cause stress-corrosion cracking in the 2014T6 material.
Table 4 Materials and conditions for surface anodising (Boeing process)
Time at 6 volts Total anodising time
~ , J
Wedge I
Surface treatment
Acid Gel: Orthophosphoric acid A R grade 889[, Distilled water Aerosil 200 Voltage required Rate of change of voltage
Metal ~ l h e r e n d
0.15 1 to 2.0 I 0.240 kg 0 - 6 volts 1 volt steps at 1 min. intervals 10 min. 15 n-fin.
80 rit b l i ~
Field anodise Field a r m d i ~
and ptin~r
~0
Factory anodi~ |
I
I
Time lhours)
Fig. 9. Crack length versus time in hours (log scale) for Boeing wedge test specimens bonded with adhesive AF126, after various pre-bonding surface treatments; schematic diagram of a wedge specimen is shown above.
Patch manufacture Patches were manufactured from pre-preg tape obtained from Composite Technology Inc as type 4.0/296 impregnated with 3 M's PR-286 epoxy. The tape was initially cut into plies by a numerically controlled laser which also made the cut-outs for the bolts of the decant assembly and the fuel drain opening. The plies were then assembled on a jig, using the bolt holes as datum points, and were consolidated in a heated platten press. Vacuum was applied during hear up prior to gellation; the vacuum was vented to atmosphere and pressure applied via a silicone rubber pad. Patches were produced in batches of 8. Fig. 9 shows a typical batch of patches, encapsulated in peel ply, together with quality control specimens. It is of interest to note the wash of boron particles around the lower quality control specimen, Fig. 10 which was prepared by mechanical cutting of the BFRP pre-preg and compare this with the clean edges produced by laser cutting. The quality control specimens were used to determine interlaminar shear strength and void content.
A.A. Baker et aL / B F R P Crack-patching technique
11
Fig. 10. Batch of finished BFRP patches and quality control specimens encased in peel ply material.
Environmental protection
The completed repair is illustrated in crossection in Fig. 11.
In addition to the use of the phosphoric acid gel anodise procedure a number of other measures were taken to improve the durability of the patch system. Contact of the bond surface of the patch with fuel was minimised by bonding an aluminium plug into the decant hole with polysulphide rubber (PR1422) prior to application of the patch. The edges of the patch and adhesive bond were then sealed by bonding an aluminium tube through the hole in the aluminium plug and patch. Finally, to prevent contact of the outer surface of the patch with moisture, aluminium foil was bonded over the BFRP patch using polysulphide rubber.
PR-1422 Seal over ~ crack zone andant around blankingplug.
Patch removal procedure A method for removing installed patches was developed to allow replacement of incorrectly applied or out of life patches. Patch removal is accomplished by use of a proprietary epoxy disintegration solvent ("Decap" produced by Dynaloy Inc) which is heated to a temperature of about 120°C and sprayed against the surface of the patch, using either a paddle wheel or a pump to provide the spray. Using this procedure a patch can be removed in less than 24 hours with no chemical or mechanical damage to the wing.
T Originalholeblankedoff
Aircraftskin 7
Crack~
Adhesive layer--J
/
Fueldraintube
-- ~
Boronpatch~
"
Crack
~
£ Decanthousing
PR-1422Sealant 0.1mmAluminiumprotective foil
Fig. 11.Schematicrepresentationof theenvironmentalprotoctionmeasures.
12
A.A. Baker et al. / B F R P Crack-patching technique
6. Repair implementation Patch application Application of BFRP patches to the Mirage wings under field conditions during routine aircraft servicing required the development of several techniques, the design and construction of a 'CrackPatching Unit' (CPU) and the development of specialised training procedures for RAAF technical personnel. The CPU (Figs. 12 and 13) contains all the equipment necessary for making the repair, including a grit-blaster and anodising unit for surface-treating the skin, and temperature controllers and hydraulic pressure indicators for controlling the skin heating and patch pressurisation systems during the adhesive cure. Detailed operating procedures for the CPU and the patch application were prepared; Fig. 14 indicates the various steps employed in the patching procedure. The turn around time per aircraft is about 3 days or about 4.5 man days per wing. A finished patch, complete with aluminium foil environmental protection coating is shown in Fig. 15. The main problems in developing the system were to obtain uniform heating of the wing-skin and pressurisation of the patch system. The phosphoric acid gel anodising of the underside of the wing posed no particular problems, provided gel of sufficient consistency was used; however rigor-
Fig. 13. Hydraulic system in place supporting the heater system during cure of the adhesive.
Fig. 12. Crack Patching Unit (fight) and equipment chest (left) under Mirage wing; grit-blast container box can be seen in the background.
A.A. Baker et al. / B F R P Crack-patching technique L;
13
AI RCRAFT PREPARATION
NDI
,]
HEAT DRY
I
SOLVENT CLEAN
]
...,q...-
GRIT BLAST
Elapsed T
:'=.,±
ANODISE
I Water
AIR DRY
criticall
[
~ reak t~t
POLARISER TEST
I
BOND PATCH
I
INSPECT
[
I I
I "' CRA TSERV,CE I Fig. 14. Sequence of steps involved in the patch application procedure.
Fig. 15. Finished BFRP patch, complete with external environmental protection in position on a Mirage wing.
ous quality control was essential to ensure that the surface was uniformly and correctly anodised. A considerable amount of effort was required to devise methods which would achieve uniform heating of the local area of the wing; the main problems were the proximity of the massive spar which acted as a heat sink and which was not allowed to exceed 100 o C. Other workers [10] had shown that exposure to temperatures above 100°C could result in loss of residual stress (cold work) around fastener holes with a consequent reduction in cycles to fatigue crack initiation. The maximum allowable temperature in the skin was 135°C to minimise overaging effects, and the minimum l l 0 ° C to ensure cure of the adhesive in the required time. During heating trials, a dummy patch with an array of thermocouples bonded into the pre-cured adhesive layer was used to estimate the temperature distribution. The method finally adopted was to heat the area with three independently controlled heaters situated along the spar and root-fib and over the skin region; the last also acted as a
caul plate for the hydraulic ram (Fig. 13). The danger of damaging the skin during pressurisation of the patch at temperature was reduced by limiting the pressure to 140 kPa. Uniform pressure over the patch region was obtained by using a silicon rubber pad, moulded to accept the patch, under the caul plate; a thin metal frame was incorporated into the rubber to prevent pressure loss at the edges of the patch. The efficiency of the pressurisation and heating procedure was assessed by examining the adhesive for voids, flow and cure in tests where the adhesive layer was separated from the surface of the wing and patch by a fluorocarbon film during cure.
Quality control and NDI Periodic checks on adhesive flow and cure are carried out (as above) by the repair team to ensure correct operation of the heating and pressurisation system. The surface-treatment procedure is checked by manufacture of wedge-test specimens.
A.A. Baker et al. / BFRP Crack-patching technique
14
7. Discussion and conclusions
This repair has been applied to over 150 wings in service. Nearly three years flight experience has been gained on the earliest repairs. In a few cases (five) some crack growth was reported, but the crack immediately arrested and no further growth has been reported. No repaired wing has required removal from service for unacceptable crack growth. Short-term laboratory tests cannot yet resolve uncertainties about long-term environmental degradation of adhesive bonds. Further, the precise nature of the stress in this region of the wing, particularly the degree of bending, is not fully known and no allowance was made for biaxial effects, due to the transverse compressive stress, which could promote local displacement of the crack faces. However, it has been shown that BFRP crack-patching is highly cost-effective and that it can lessen aircraft unavailability because it can be successfully applied under quite difficult field conditions by specially trained service personnel with no previous experience of structural
adhesive bonding. The advantages of crack-patching over conventional repair procedures are: (a) no mechanical damage to surrounding structure, no fastener holes, (b) crack can be inspected through the patch by conventional eddy-current NDI, (c) the cracked area is protected from further external corrosion, and internal fuel leakage, by the patch system, (d) reinforcement is only in the direction required, no undesirable stiffening in other directions, and (e) patches can be removed and replaced with no damage to the surrounding structure. Although development of the CPU and repair procedure involved considerable effort at ARL, most of this work will not be repeated in further repairs, although changes may be required to the heaters and pressurisation systems. With the addition of a vacuum pump (for which provision has been made), the CPU could be used for general crack-patching, repairs to adhesively bonded metallic components, repairs to fibre composite components and for general repair of battle damage.
8. Appendix
The crack tip element used in this study was octagonal in shape with the crack tip located at the centroid of the octahedron. The u and v displacements used within this region were obtained from the well known Williams' solution, viz:
U(r,0)=-~-~
E (-1)"
r
az,_,[-----~cos[~}O+(3.5-n +C2n_ 1[ - ~
+(-1)"+'( r kro
s i n ( ~ - - ~ ) 0 + (3.5 -
).
n
40) 4o) s i n ( ~ ' - ~ )
{ d 2 . [ ( n + 1) cos(n+ 1 ) 0 + ( 3 - 4 o - n )
0] } c o s ( n - 1)01 (A1)
+ c 2 . [ ( n - 1) sin(n + 1)0 + ( 3 - n - 40) s i n ( n - 1)O] }. and N
VCr,
n- 1/2
E (-1)" nil
l--T-cos t
-(2.5 + n - a a ) c o s
+d2n_l[(2.5+n-4o)sin(~---)O
2n-3
[2n-3) 1 ~
0
]
15
A.A. Baker et al. / BFRP Craek:patching technique +(
1)n+t(rl"{d2n[(n+ -
3-4o)sin(n-1)O-(n+
l)sin(n+
l)O]
~rol
+ c 2 n [ ( n - - 1) c o s ( n + 1 ) 0 - - ( n + 3 - - 4 o ) c o s ( n - - 1 ) 0 ] },
(A2)
where o = v for p l a n e strain. v / 1 + v for p l a n e stress.
(A3)
H e r e r0 is a c h a r a c t e r i s t i c length for the element, while U, a n d V are c o m p o n e n t s of the d i s p l a c e m e n t field in the p o l a r c o o r d i n a t e system r, 0. This c o o r d i n a t e s y s t e m has r = 0 at the c r a c k tip while the faces of the crack lie at 0 = +,n. T h e u a n d v d i s p l a c e m e n t s are r e l a t e d to U a n d V b y u= V(r, 0) cos0-
V(r, 0) sin0+u
0-~0y,
v = U ( r , 0 ) sin 0 + V ( r , 0 ) cos 0 + v0 + ~ x
(A4) (A5)
a n d where u0, v0 a n d ~0 are rigid b o d y degrees of freedom. H a v i n g specified the n a t u r e of the d i s p l a c e m e n t field it is a s i m p l e m a t t e r to o b t a i n the stiffness m a t r i x using s t a n d a r d finite e l e m e n t procedures. T h e details of this m a y b e f o u n d in reference [11]. H a v i n g o b t a i n e d the stiffness m a t r i x a n d s u b s e q u e n t l y o b t a i n e d the d i s p l a c e m e n t s at the n o d a l p o i n t s the stress i n t e n s i t y factors are o b t a i n e d using a m e t h o d a n a l o g o u s to the well k n o w n J integral a p p r o a c h . This p r o c e d u r e is d e s c r i b e d in detail in [11] a n d is also a p p l i c a b l e for elastic analysis [12].
Acknowledgements M a n y p e o p l e at A R L a n d in R A A F were conc e r n e d in the d e v e l o p m e n t of this repair, n o t all c a n be a c k n o w l e d g e d individually. However, we p a r t i c u l a r l y t h a n k E n g i n e e r i n g Facilities D i v i s i o n A R L , for d e v e l o p m e n t of the C r a c k P a t c h i n g U n i t , Mr. B.C. Hoskin, Sq.Ld. M. R e y m o n d , Mr. J.D. R o b e r t s , Mr. A . W . Rachinger, Mr. T.G. Hill, Mr. B.C. Bishop a n d Mr. T. van Blaricum for their significant c o n t r i b u t i o n s to various p a r t s of the project.
References [1] A.A. Baker, Composites, Vol. 9, 1978, pp. 11-16. [2] A.A. Baker and M.M. Hutchison, "Fibre composite reinforcement of cracked aircraft structures" Proceedings of 33rd Annual Technical Conference of the Plastics Industry, 1978, Section I7-E. [3] R. Jones and R.J. Cailinan, J. Struct. Mech. Vol. 7, 1979, pp. 107-130.
[4] R.A. Mitchell, R.M. Wooley and D.J. Chivirut, A.I.A.A., Vol. 13, 1975, pp. 744-749. [5] A.A. Baker, S.A.M.P.E. Journal, Vol. 15, 1979, pp. 10-17. [6] A.A. Baker, MJ. Davis and G.A. Hawkes, "Repair of fatigue cracked aircraft structures with advanced fibre composites: Residual stress and thermal fatigue studies", to be published in Proceedings of the lOth International Committee on Aeronautical Fatigue Symposium, May 1979. [7] R. Jones and R.J. Callinan, J. Fibre Science and Technology, Vol. 14, 1981, pp. 99-111. [8] R. Jones and R.J. Callinan, J. Struct. Mech., Vol. 8, 1980, pp. 143-149. [9] M.C. Locke, W.M. Scardino and H. Croop, "Non-tank phosphoric anodize method of surface preparation for repair bonding", 7th S.A.M.P.E. Tech. Conf., 1975. [10] J.R. Hart-Smith and F.E. Kiddie, "Effects of heat on fatigue in aircraft structure', R.A.E. Reports and Memorand~ No. 379.8, 1975. [11] R. Jones and R.J. Callinan, Int. J.:l=racture, Vol. 13, 1977, pp. 51-64. [12] R. Jones, K.C. Waiters and R.J. Callinan, "Hybrid elements and elastic plastic fracture mechanics" Proc. 4th Int. Conf. in Australia on Finite Element Methods, Melbourne, Aug. 1982.
1
REPAIR OF MIRAGE III AIRCRAFT USING THE BFRP CRACK-PATCHING TECHNIQUE
A.A. BAKER, R.J. CALLINAN, M.J. DAVIS, R. JONES and J.G. WILLIAMS Defence Science and Technology Organisation, Aeronautical Research Laboratories, Melbourne, Vic. 3001, Australia
The BFRP crack-patching technique has been applied to the field repair of fatigue cracks in the aluminium alloy wing skins of Mirage III fighter aircraft. Finite-element procedures were used in patch design. The repair was qualified using fatigue-crack propagation studies on panels simulating the cracked and repaired area. A field support unit was designed to allow repairs to be carried out by air force personnel during routine maintenance of the aircraft. To date over 150 patches have been applied and nearly three years of operational history gained. While some crack growth was observed after repair of a few wings, the patch stopped further growth and no wing skin has required further repair.
1. Introduction
Aeronautical Research Laboratories (Australia) have developed methods for repair of cracked metallic aircraft components using an adhesively bonded boron fibre reinforced plastic (BFRP) patch. This report describes experience in developing and applying a BFRP patch to fatigue cracks in the wing skin of Mirage III fighter aircraft. Previously the technique had been applied successfully to stress corrosion cracks in Hercules wing planks [1] and fatigue cracks in Macchi landing wheels [2]. Traditional repair techniques using metallic patches require many additional holes to be drilled for fasteners to provide efficient load transfer into the repair. With adhesively bonded repairs, this is unnecessary. The advantages of using BFRP for the patch material include excellent resistance to corrosion and cyclic loading and ease of production even with complex curvatures. The patch can be produced with the fibres spanning the crack to restrict crack opening under load and hence to prevent crack growth by reducing the stress intensity. The high directional stiffness prevents unnecessary and undesirable increases in component rigidity in directions other than across the crack. Carbon fibre reinforced plastic (CFRP) could also have been used, but BFRP was selected, although more expensive, for its better stiffness and fatigue resistance and its higher thermal expansion coefficient. Table 1 gives values for physical constants for repair materials. Thermal stresses result-
ing from mismatch in thermal coefficient can introduce high residual stresses in systems bonded in place using heat-cured adhesives. As BFRP has negligible electrical conductivity, readily available eddy-current procedures can be used to detect and monitor cracks beneath the patch. Recently, fatigue cracks were discovered in the lower wing skins, close to the main spar of some Australian Mirage aircraft, Figs. 1 and 2. It was decided, in consultation with RAAF, that BFRP crack-patching would be an effective solution for this problem, because the repair would (i) cause no mechanical damage to the skin (i.e. no fastener holes), (ii) cause no strain elevation in the spar, since reinforcement need only occur across the crack, (iii) allow the use of conventional eddy-current procedures to check for crack growth, and (iv) allow implementation in the field during normal servicing, thereby minimising unavailability of the aircraft. Because of the significance of the cracking and the long life desired of the repair, detailed design studies and a considerable amount of further research and development were required before the repair could be implemented. Some aspects of this work are reported here.
2. Patch design and analysis
The fatigue cracks initiated near the fuel decant hole in the lower wing skin, close to the intersection of the main spar and root rib (Fig. 1). The skin consists of aluminium alloy AU4SG, similar
0167-8zlz12/84/$3.00 © 1984, Elsevier Science Publishers B.V. (North-Holland)
2
A.A. Baker et al. / B F R P Crack- patching technique
Table 1 Values of material constants Material
Aiuminium a Alloy (2014-T6) CFRP BFRP AF-126 Adhesive
Elastic Modulus E n (GPa)
Thermal Expansion Coeff. x 1 0 - 6 / o C
Specific Gravity
Strain 'g
72 130 204 0.7 (20 o C) (shear)
23 negligible 5.0 -
2.8 1.5 2.0 -
0.6 (yield) 1.1 (fracture) 0.6 (fracture) -
" The alloys used for experimentation were 2024-T3 and L104. L104 is British Standard equivalent of the French AU4SG, from which the aircraft skin was m~ide. These last two materials are equivalent to 2014-T6 aluminium alloy.
to 2014T6, about 3.5 mm thick and is covered by a fairing. The wing skin forms part of the wing torsion box which is designed to be in shear. This stress state was confirmed by experimental strain studies at ARL. The state of shear results in the fatigue cracks propagating at 45 o to the spar (Fig. 2). For design purposes the maximum size of the
crack (including the fuel decant hole) was taken as 111 ram. Hight loads were estimated from straingauge data and from the manufacturer's stress analysis. The BFRP patch (Fig. 2) is a unidirectional laminate with fibres running perpendicular to the direction of crack propagation. The patch contains
Fig. 1. Silhouette of Mirage III aircraft showing where the fatigue cracks developed in some aircraft. Insert: Fuel decant hole region, showing the nature of the fatigue cracking.
A,A. Baker et al. / BFRP Cr~ck-patching technique Z
C
t -
Tzx
Element of adhesive
=X
Boron patch
Adhesive
Aluminium alloyskin Crack
Fig. 3. Crossectionaldiagram of the seven-layer,internally-steppedBFRP patch.
seven layers of BFRP and is internally stepped, i.e. the largest layer is on the outside to reduce interlaminar shear and peel stresses. The shear stress reduction arises largely as a result of the increased adhesive thickness near the edge of each ply. The transverse stress (peel stress) reduction in the inner plies arises as a result of the opposing stress resulting from the geometry of the plies in the overlap region (see Fig. 3). The beneficial effect of this approach was observed in some preliminary fatigue tests. The proximity of the spar bolts and the root rib bolts to the decant hole necessitated closer spacing of the layer steps in this region. The attachment holes for the decant hole cover were preformed in the patch; the decant hole opening was greatly reduced in size by the patch, so as to obtain maximum reinforcement efficiency. The main objectives of the analytical procedure were to assess the reduction in stress-intensity in the skin, reveal any undue strain elevation in the spar and to estimate the maximum levels of stress in the BFRP patch and adhesive layer. Ideally, a sufficient reduction in stress-intensity should be obtained without exceeding the materials allowables in the patch system, the most critical of which
is the adhesive shear strength under fatigue loading in the operating environment.
Decant ilumbly
Inb'd
SlUr
i
No 1 Bolt
il~
/
A
~.
.-~~_.= = --,lilill @ 0"" = --- O Ollcimthole Ill/ill . / in i . ~Ikin i . , ,~ , Iil~,,il
//0
Hole in petch ~
°k
|
i
I
to .',','.
\
@ ®
o
o
~/ .
.
,/,+," ~'.','%,"--
111 !11 i,iI! I
®
i II 7.'/ #'---S," --III/11
-'111 7I
® Fig. 2. Schematic diagram of fuel decant hole region with outline of BFRP patch superimposed; note the internal stepping of the patch.
4
A.A. Baker et aL / BFRP Crack-patching technique
3. Finite element analysis Initially a detailed, two-dimensional finite element stress analysis of the cracked and unpatched region was undertaken. This model consisted of 252 quadrilateral membrane elements, and 68 constant strain triangular elements representing the wing skin, which is 3.39 m m thick, 16 flange elements representing the integrally-milled stiffeners, and two crack-tip elements which are described in the appendix. The resulting stress-intensity factor, assuming loads typical of 7.5 g flight manoeuvre were K 1 = 72 MPa ¢ m , K 2 = 3.3 MPa ¢ ~ at the tip nearest to the spar, and K 1 --68 MPa V~mm,K 2 = 1.0 MPa ~ at the tip nearest to the root rib. The effect of the patch was estimated by adding a finite-element representation of various BFRP patches to the cracked model (Fig. 4) utilizing a special element described in [3]; this model allows for the separate response of the wing skin, patch and adhesive. In the specific case when the xy axis system coincides with the axis of orthotropy of the patch, the relationship between displacements and shear stresses in the adhesive layer reduce to: ~ x = (Uo - u s ) / ( t J G a
%,. = (% - v s ) / ( t J G
+ 3tJ8Gs + 3to/8G13)' ~ + 3tJ8G,
+ 3to/8G23 )
(1)
which is similar to the relationship developed in [4]. Here %x and ~'~.~are the shear stresses developed in the adhesive; t a, t s and t o are the thicknesses of the adhesive, wing-skin and patch respectively; (u 0, %) and (u s, vs) are the x, y displacements at the mid surfaces of the wing-skin and the patch; Gs and Ga are the shear moduli of the skin and the adhesive respectively, and Gl~ and G23 are the transverse shear moduli of the patch. The through-the-thickness variation of the inplane displacements are given by the following formulae, when the transverse shear moduli of the patch are equal, viz: u = u o + "rsx(Z(t s + t a - t o - z / 2 ) -(t s + ta + to/2)(ts/2
= (u o - 3~s~to/8Go)(Z - ts)/t a + (u s + 3rs~tJ8Gs)(ts
= u~ + (z 2 - t ~ / 4 ) ~ s x / 2 G s t s in the sheet
(2)
and a similar expression for v ( x , y , z ) . As a result of this formulation the value of the
o 8~
Fig. 4. Finite-element mesh for the fuel decant hole region.
+ ta - z ) / t a
in the adhesive
Z
---.L / , ' l / l l
+ 3to/4))Got o
in the patch
Spar
/
+ tJ2
A.A. Baker et al.
/ B F R P Crack:patching technique
stress intensity factor varies through the thickness of the wing skin. This variation may be derived from the previous equations and is given by Kl = K , p ( 1 -
X( z2 - t 2 / 4 ) / 2 G s t ~ ) ,
(3)
where KIp is the value of the stress intensity factor at the mid-surface of the wing skin after patching and X is the spring constant X = 1/(tJG
a + 3 t J 8 G s + 3to/8Go).
(4)
The geometry of the patch was constrained by the presence of the spar, root rib, and the fairing attachment. Furthermore the plan area of the patch was limited by the requirement to minimise pressurisation loads during application of the patch. This resulted in a fixed geometric shape with the prime variables being the patch and adhesive thickness and step size between plies. With these restrictions, six B F R P patch configurations were considered each with the plan form
shown in Fig. 2. Each patch was modelled using approximately 380 of the bonded elements, which are triangular in plan form, described above and in greater detail in reference [3]. All of the six patches considered were unidirectional laminates and were internally stepped, i.e. with the longest ply on the outside. The fibre direction was at ninety degrees to the crack. Initially it was uncertain if carrying the fibres over the drain hole was necessary, or how frequently the drain hole was used in service. As a result, in three of the patches considered a hole was left so as not to interfere with the draining of the wing. In the other three patches varying amounts of the hole were covered. In one case one third of the total area of the hole was covered while in the other two cases virtually all of the hole was
covered.
For each of the patches the maximum shear stresses in the adhesive bonding the patch to the wing skin occurred at points A, B, C and D (see
2.0
Oc/O u .9
! !
0')
0
I 3
I ~ i I I 6 9 12 15 Distance from Bolt Hole Number 1 (ram)
I 18
Fig. 5. Stress distribution in the aluminium skin. The curves are for the stress, o, for ratios between three cases denoted by subscripts; p = patched and cracked, u ~ unpatched and uncracked, and c = cracked and unpatched.
6
A.A. Baker et al. / BFRP ('rack-patching technique
Fig. 2). The maximum stresses in the fibres occurred at point D for the patches with a hole in the patch, and in the fibres over the hole in the patches with the hole partially covered. The values for these stresses, along with the percentage reduction of the stress intensity factor K 1 at each tip, achieved by each patch are shown in Table 2. Here Klu and K~p are the values of the stress intensity factors before and after patching respectively, the value of Kip being at the mid surface of the wing skin. All of the six patches achieve a reduction in the stress intensity factor K 1 of at least 91%. Consequently they would all significantly retard crack growth. Similarly for all of the patches the fibre strains are below the m a x i m u m working level of 0.005, which corresponds to a stress of 1000 MPa, although of the six patches numbers 5 and 6 have by far the greatest factors of safety. As a result as shown in [7] the patch design was finally chosen primarily on the basis of the magnitude of the shear stresses developed in the adhesive. On this basis patch numbers 1, 2, 3, and 4 may be rejected. The two remaining patches are patch numbers 5 and 6. Of the two, Table 2 shows that the adhesive shear stresses along the edges of the patch, are substantially higher for patch number 5 than for patch number 6, although both are below the
threshold value for fatigue damage for this adhesive. As a result patch 6 is much less likely to suffer fatigue damage to the adhesive bond. Hence patch number 6 was adopted as the final repair. At locations C and D in patch 6 the shear stress in the adhesive is sufficiently high so as to cause concern over the possibility of fatigue damage occurring in the adhesive. However these high values occur in the interior of the patch at the intersection of the crack with the drain hole, and are very localized. As a result any damage which does occur should not spread and, as shown in [5], should have virtually no effect on the stress intensity factors at the crack tips or on the fibre stresses. The effect of this patch on the maximum principal stress in the skin along the line of attachment to the spar (Fig. 5) showed that the patch effectively restores the stress distribution along the spar to that in a wing with an uncracked panel. Consideration was also given to thermal and residual stress effects in the aluminium skin and to thermal-fatigue effects in the adhesive due to the use of an adhesive curing at elevated temperature ( 120°C) and the mismatch in thermal expansion between the aluminium alloy and BFRP. It was concluded [6,7,8] that these effects posed no serious problems in this particular repair. Table 2 shows the value of the stress intensity
Table 2 Effect of drain hole patch 7.5 g load case Patch Number Maximum path thickness (mm) Adhesive thickness (mm) Thickness of first Layer (mm) Covering of drain hole Adhesive shear stress (MPa) at points: A B C D Maximum fibre stress (MPa) Reduction in stress intensity factor K l, i.e. 1 - Kip/Klu. at: 1) Spar tip 2) Root rib tip
1
2
3
4
5
6
0.762
0.762
0.762
0.889
0.889
0.889
0.102
0.102
0.102
0.203
0.203
0.203
0.127 open
0.254 open
0.127 1/3 covered
0.254 open
29 55 181 153
43 79 179 153
29 55 164 131
31 58 120 98
26 42 63 50
18 30 64 51
953
930
760
911
450
455
91% 99%
91% 99%
91% 99%
91% 99%
0.254 substantially covered
92% 99%
0.127 substantially covered
91% 99%
A.A. Baker et al / BFRP Crad~-patching technique
reduction at each crack tip. After patching, however, the maximum value of the stress intensity factor occurs at the free surface of wing skin (see equation 3 for z = 0) and has the value K_lp(1 + ts/SGs). For this patch Kip = 6,5 MPa ¢ m , and the factor (1 + tJ8G~) is--approximately (1 + tsGa/8Gat,)--= 1.075 so that the maximum value of the stress intensity factor becomes 7 MPa v/~ and occurs on the free surface on the inside of the skin. These values lie in the vicinity of the threshold value for fatigue crack growth. In the above analysis bending of the wing skin was assumed to be negligible, as was consistent with strain gauge readings on the wing skin. However, the presence of bending could result in the threshold value of the stress intensity factor being exceeded on the inner surface of the wing skin.
4. Repair qualification The BFRP repair was qualified mainly by two series of tests (a) a strain survey on a Mirage wing with a patch installed to show that the patch did not significantly elevate the local strain in the spar,
and (b) fatigue tests on aluminium alloy panels configured to simulate the cracked area in the wing (Fig. 6). The aims of the fatigue test were to check that (i) the predicted reduction in stress-intensity could be achieved, and (ii) the patch/adhesive system could endure the fatigue loading. The strain survey showed that the strain levels in the spar after patching increased by less than 3% indicating that no significant strain elevation occurs in the spar after patching. The fatigue results are given below. Fatigue crack propagation studies
Tension-tension fatigue tests were undertaken on the aluminium alloy panels using either constant load cycling or block load cycling. Only the principal tensile stress in the wing could be simulated. The tests were carried out in laboratory environment at a temperature of 23°C. Initially, due to unavailability of 2014T6 materials, tests were carried out on aluminium alloy 2024T3 panel. Although this alloy has significantly greater resistance to crack propagation than the wing skin alloy AU4SG, it was considered that these tests would show whether or not the stress-intensity was
Sealant Decanthole and associated fastenerholes /
Patch pat..chregion
Decant
assembly o.
. . . .
Crack
Sandwichpanel ~
Porousteflon--/
c~ed 91ass
i 1,0,°°l I
7
~I :3.66rnm
Fig. 6. Details of 2024T3 aluminium alloy wing-skin simulation panels used for crack-propagation studies, showing position of patch and vacuum box restraint.
8
A.A. Baker et al. / B F R P Crack-patching technique
reduced by patching and if the fatigue characteristics of the adhesive layer were satisfactory. Cracks were initiated from saw-cuts and propagated to about 100 mm total length prior to BFRP patching, using the procedures described below. During application of the patch, the panels were constrained at their ends to simulate the constraint to thermal expansion that would be experienced in the wing-skin under practical repair conditions. This rate of crack growth was monitored during the initial test by direct observation with a travelling vernier microscope. After patching, eddy-current procedures were used to monitor the crack under the patch. The unit used was a Sperry Products EM-3300 unit. The accuracy with which the crack length could be determined was + 1 mm. In the early tests local bending of patched panels occurred under load, due to local displacement of the neutral axis of the panel by the patch. The effect negated the reinforcing effect of the patch to a large extent. Strain-gauge tests confirmed the presence of substantial strains in the panel on the opposite side to the patch. In the actual wing-skin, local stiffening by the surrounding structure, particularly the spar, root rib and internal stiffeners, would largely prevent bending and a method was therefore devised to minimise its occurrence in the test. This procedure used atmospheric pressure (by evacuating the region between the back of the test panel and the sandwich panel). A thin layer of porous, fluorocarbon-
120
coated fibre-glass between the test panel and the sandwich panel acts to allow air removal in the gap between the test panel and sandwich panel and minimised the frictional load. Strain-gauge studies showed that the bending effects were then reduced by about 50%. Using the vacuum restraint, substantial increases in life prior to crack growth were achieved; Fig. 7 plots crack growth curves for three 2024T3 specimens subjected to constant load cycling at an applied stress of 91 MPa, (equivalent to a nominal 4g manoeuvre load). In all cases initial crack growth rates indicated that failure of the panels (without the repair) would have occurred within a further 104 cycles. Panels were tested under more realistic loading conditions using a block loading sequence. The flight spectrum used was obtained from flight data recordings. This spectrum was truncated at zero and divided into 9 load blocks. The number of cycles and corresponding peak stresses are given in Table 3~ In this table load block numbers refer to the order in which loads are applied. Each program sequence is approximately equivalent to one year of aircraft usage. Table 3 represents 0.6 of a program sequence. Substantial life extensions for the cracked 2024T3 alloy panels, (Fig. 8) indicate that, provided local bending is minimised, the BFRP patch system provides a significant stressintensity reduction. The actual life extension probably depends largely on the degree to which local
Patch Applied
E ~i
Thisspecimen, Load increa~ledby 38%
/
/
,s/
s,~ */
/
= 30MPa ,V'~
No growth afler 105 cycles
80
=0.1 ak Nominal Stress 91MPa
~.O
I ~ 'tO r
I 160
I
I 180
i
/L v
l 370
l
l 390
NO of Cycles (x|0 -3) Fig. 7. Crack length versus cycles for 2024T3 wing-skin simulation panels subjected to constant load cycling; stress intensity range prior to patching is indicated.
A.A. Baker et al. / B F R P Crack-patching technique ?
Growth detected
120
/ Test terminated tch applied L104 Specimen
~0
0
I
I
20
~0
I
A //,
I
V
100
I
I
120
No. of programmes Fig. 8. Crack length versus n u m b e r of programs of block loading for 2024T3 and 2014T6 wing-skin simulation panels.
bending is reduced in practice. The long life of the repair indicates that the patch/adhesive system is capable of withstanding sustained fatigue loading, at least in a laboratory environment. Currently tests are being conducted using panels of 2014T6 alloy, the material used in the aircraft wing skin. These panels are prepared with the rolling direction oriented similarly to that in the wing skin. Tests to date show that crack propagate in these panels very much faster than in 2024T3; the results for one such panel are included in Fig. 8. It is worth noting that the notched 2014T6 panels could not withstand one complete Table 3 Load spectrum Load block number
N u m b e r of cycles
Peak stress (MPa)
1 2 3 4 5 6 7 8 9
9999 1027 870 341 579 381 31 3 0.32
57.9 69.2 79.0 88.2 103.0 121.0 139.3 151.0 156.1
program of the loading sequence; crack growth to the 'repair size' occurred at load level 5 ( - 5 g ) , whereas the loading sequence goes to load level 9 ( - 9g). This observation suggests that the stressing level employed in these tests may be higher than that encountered by this region in the aircraft, since, in practice, cracks are not found to grow this rapidly.
5. Selection of materials and processes Adhesive
An epoxy-nitrile structural adhesive, 3 M's AF126, was chosen on the basis of previous background studies into suitable adhesives for crack patching, which included work on fatigue and stress relaxation [5], and practical experience on repairs to stress corrosion cracks in Hercules aircraft over a period of more than four years [2]. AF126 cures at about 120°C and requires a pressure of 70-350 kPa. Originally, an adhesive curing at a lower temperature and pressure was sought to simplify patch application and minimise thermal and residual stress problems. Unfortunately, none of those studied was suitable; future development
10
A.A. Baker et al. / B F R P Crack-patching technique
of a suitable adhesive curing at lower temperatures and pressures would greatly aid patching technology.
Cr~k
C--
Adher~ve layer
Metal adherend WqN:~etest specimen
In previous repairs, non-chemical surface-treatments of the aluminium alloy component were used to simplify patch application procedures and to avoid the danger of stress-corrosion cracking. This treatment was degreasing, followed by gritblasting with 50 #m alumina powder; the same procedure was used for preparing the BFRP patch. However, it was considered that the present repair required a higher level of environmental durability than would be provided by this procedure, because of the critical nature of the cracking and the long life required of the repair. It was decided therefore, to evaluate the phosphoric acid gel anodising procedure recently developed by the Boeing Aircraft Co., for field repairs to adhesively bonded aluminium components [9]. Table 4 lists details of the solution used and the anodising conditions. Boeing wedge tests were used to compare the effectiveness of various surface treatments, including alumina grit-blasting, gel anodising, and standard tank anodising (factory procedure). The results (Fig. 9) show that gel anodising gives adequate durability. Although the use of primer gave further improvement, its application was considered too difficult to control under field conditions. Prior to final acceptance of the gel anodising procedure, further wedge tests showed the bond to be resistant to fuel immersion, the main environmental contaminant in the present repair. Other tests showed that the gel anodise procedure did not cause stress-corrosion cracking in the 2014T6 material.
Table 4 Materials and conditions for surface anodising (Boeing process)
Time at 6 volts Total anodising time
~ , J
Wedge I
Surface treatment
Acid Gel: Orthophosphoric acid A R grade 889[, Distilled water Aerosil 200 Voltage required Rate of change of voltage
Metal ~ l h e r e n d
0.15 1 to 2.0 I 0.240 kg 0 - 6 volts 1 volt steps at 1 min. intervals 10 min. 15 n-fin.
80 rit b l i ~
Field anodise Field a r m d i ~
and ptin~r
~0
Factory anodi~ |
I
I
Time lhours)
Fig. 9. Crack length versus time in hours (log scale) for Boeing wedge test specimens bonded with adhesive AF126, after various pre-bonding surface treatments; schematic diagram of a wedge specimen is shown above.
Patch manufacture Patches were manufactured from pre-preg tape obtained from Composite Technology Inc as type 4.0/296 impregnated with 3 M's PR-286 epoxy. The tape was initially cut into plies by a numerically controlled laser which also made the cut-outs for the bolts of the decant assembly and the fuel drain opening. The plies were then assembled on a jig, using the bolt holes as datum points, and were consolidated in a heated platten press. Vacuum was applied during hear up prior to gellation; the vacuum was vented to atmosphere and pressure applied via a silicone rubber pad. Patches were produced in batches of 8. Fig. 9 shows a typical batch of patches, encapsulated in peel ply, together with quality control specimens. It is of interest to note the wash of boron particles around the lower quality control specimen, Fig. 10 which was prepared by mechanical cutting of the BFRP pre-preg and compare this with the clean edges produced by laser cutting. The quality control specimens were used to determine interlaminar shear strength and void content.
A.A. Baker et aL / B F R P Crack-patching technique
11
Fig. 10. Batch of finished BFRP patches and quality control specimens encased in peel ply material.
Environmental protection
The completed repair is illustrated in crossection in Fig. 11.
In addition to the use of the phosphoric acid gel anodise procedure a number of other measures were taken to improve the durability of the patch system. Contact of the bond surface of the patch with fuel was minimised by bonding an aluminium plug into the decant hole with polysulphide rubber (PR1422) prior to application of the patch. The edges of the patch and adhesive bond were then sealed by bonding an aluminium tube through the hole in the aluminium plug and patch. Finally, to prevent contact of the outer surface of the patch with moisture, aluminium foil was bonded over the BFRP patch using polysulphide rubber.
PR-1422 Seal over ~ crack zone andant around blankingplug.
Patch removal procedure A method for removing installed patches was developed to allow replacement of incorrectly applied or out of life patches. Patch removal is accomplished by use of a proprietary epoxy disintegration solvent ("Decap" produced by Dynaloy Inc) which is heated to a temperature of about 120°C and sprayed against the surface of the patch, using either a paddle wheel or a pump to provide the spray. Using this procedure a patch can be removed in less than 24 hours with no chemical or mechanical damage to the wing.
T Originalholeblankedoff
Aircraftskin 7
Crack~
Adhesive layer--J
/
Fueldraintube
-- ~
Boronpatch~
"
Crack
~
£ Decanthousing
PR-1422Sealant 0.1mmAluminiumprotective foil
Fig. 11.Schematicrepresentationof theenvironmentalprotoctionmeasures.
12
A.A. Baker et al. / B F R P Crack-patching technique
6. Repair implementation Patch application Application of BFRP patches to the Mirage wings under field conditions during routine aircraft servicing required the development of several techniques, the design and construction of a 'CrackPatching Unit' (CPU) and the development of specialised training procedures for RAAF technical personnel. The CPU (Figs. 12 and 13) contains all the equipment necessary for making the repair, including a grit-blaster and anodising unit for surface-treating the skin, and temperature controllers and hydraulic pressure indicators for controlling the skin heating and patch pressurisation systems during the adhesive cure. Detailed operating procedures for the CPU and the patch application were prepared; Fig. 14 indicates the various steps employed in the patching procedure. The turn around time per aircraft is about 3 days or about 4.5 man days per wing. A finished patch, complete with aluminium foil environmental protection coating is shown in Fig. 15. The main problems in developing the system were to obtain uniform heating of the wing-skin and pressurisation of the patch system. The phosphoric acid gel anodising of the underside of the wing posed no particular problems, provided gel of sufficient consistency was used; however rigor-
Fig. 13. Hydraulic system in place supporting the heater system during cure of the adhesive.
Fig. 12. Crack Patching Unit (fight) and equipment chest (left) under Mirage wing; grit-blast container box can be seen in the background.
A.A. Baker et al. / B F R P Crack-patching technique L;
13
AI RCRAFT PREPARATION
NDI
,]
HEAT DRY
I
SOLVENT CLEAN
]
...,q...-
GRIT BLAST
Elapsed T
:'=.,±
ANODISE
I Water
AIR DRY
criticall
[
~ reak t~t
POLARISER TEST
I
BOND PATCH
I
INSPECT
[
I I
I "' CRA TSERV,CE I Fig. 14. Sequence of steps involved in the patch application procedure.
Fig. 15. Finished BFRP patch, complete with external environmental protection in position on a Mirage wing.
ous quality control was essential to ensure that the surface was uniformly and correctly anodised. A considerable amount of effort was required to devise methods which would achieve uniform heating of the local area of the wing; the main problems were the proximity of the massive spar which acted as a heat sink and which was not allowed to exceed 100 o C. Other workers [10] had shown that exposure to temperatures above 100°C could result in loss of residual stress (cold work) around fastener holes with a consequent reduction in cycles to fatigue crack initiation. The maximum allowable temperature in the skin was 135°C to minimise overaging effects, and the minimum l l 0 ° C to ensure cure of the adhesive in the required time. During heating trials, a dummy patch with an array of thermocouples bonded into the pre-cured adhesive layer was used to estimate the temperature distribution. The method finally adopted was to heat the area with three independently controlled heaters situated along the spar and root-fib and over the skin region; the last also acted as a
caul plate for the hydraulic ram (Fig. 13). The danger of damaging the skin during pressurisation of the patch at temperature was reduced by limiting the pressure to 140 kPa. Uniform pressure over the patch region was obtained by using a silicon rubber pad, moulded to accept the patch, under the caul plate; a thin metal frame was incorporated into the rubber to prevent pressure loss at the edges of the patch. The efficiency of the pressurisation and heating procedure was assessed by examining the adhesive for voids, flow and cure in tests where the adhesive layer was separated from the surface of the wing and patch by a fluorocarbon film during cure.
Quality control and NDI Periodic checks on adhesive flow and cure are carried out (as above) by the repair team to ensure correct operation of the heating and pressurisation system. The surface-treatment procedure is checked by manufacture of wedge-test specimens.
A.A. Baker et al. / BFRP Crack-patching technique
14
7. Discussion and conclusions
This repair has been applied to over 150 wings in service. Nearly three years flight experience has been gained on the earliest repairs. In a few cases (five) some crack growth was reported, but the crack immediately arrested and no further growth has been reported. No repaired wing has required removal from service for unacceptable crack growth. Short-term laboratory tests cannot yet resolve uncertainties about long-term environmental degradation of adhesive bonds. Further, the precise nature of the stress in this region of the wing, particularly the degree of bending, is not fully known and no allowance was made for biaxial effects, due to the transverse compressive stress, which could promote local displacement of the crack faces. However, it has been shown that BFRP crack-patching is highly cost-effective and that it can lessen aircraft unavailability because it can be successfully applied under quite difficult field conditions by specially trained service personnel with no previous experience of structural
adhesive bonding. The advantages of crack-patching over conventional repair procedures are: (a) no mechanical damage to surrounding structure, no fastener holes, (b) crack can be inspected through the patch by conventional eddy-current NDI, (c) the cracked area is protected from further external corrosion, and internal fuel leakage, by the patch system, (d) reinforcement is only in the direction required, no undesirable stiffening in other directions, and (e) patches can be removed and replaced with no damage to the surrounding structure. Although development of the CPU and repair procedure involved considerable effort at ARL, most of this work will not be repeated in further repairs, although changes may be required to the heaters and pressurisation systems. With the addition of a vacuum pump (for which provision has been made), the CPU could be used for general crack-patching, repairs to adhesively bonded metallic components, repairs to fibre composite components and for general repair of battle damage.
8. Appendix
The crack tip element used in this study was octagonal in shape with the crack tip located at the centroid of the octahedron. The u and v displacements used within this region were obtained from the well known Williams' solution, viz:
U(r,0)=-~-~
E (-1)"
r
az,_,[-----~cos[~}O+(3.5-n +C2n_ 1[ - ~
+(-1)"+'( r kro
s i n ( ~ - - ~ ) 0 + (3.5 -
).
n
40) 4o) s i n ( ~ ' - ~ )
{ d 2 . [ ( n + 1) cos(n+ 1 ) 0 + ( 3 - 4 o - n )
0] } c o s ( n - 1)01 (A1)
+ c 2 . [ ( n - 1) sin(n + 1)0 + ( 3 - n - 40) s i n ( n - 1)O] }. and N
VCr,
n- 1/2
E (-1)" nil
l--T-cos t
-(2.5 + n - a a ) c o s
+d2n_l[(2.5+n-4o)sin(~---)O
2n-3
[2n-3) 1 ~
0
]
15
A.A. Baker et al. / BFRP Craek:patching technique +(
1)n+t(rl"{d2n[(n+ -
3-4o)sin(n-1)O-(n+
l)sin(n+
l)O]
~rol
+ c 2 n [ ( n - - 1) c o s ( n + 1 ) 0 - - ( n + 3 - - 4 o ) c o s ( n - - 1 ) 0 ] },
(A2)
where o = v for p l a n e strain. v / 1 + v for p l a n e stress.
(A3)
H e r e r0 is a c h a r a c t e r i s t i c length for the element, while U, a n d V are c o m p o n e n t s of the d i s p l a c e m e n t field in the p o l a r c o o r d i n a t e system r, 0. This c o o r d i n a t e s y s t e m has r = 0 at the c r a c k tip while the faces of the crack lie at 0 = +,n. T h e u a n d v d i s p l a c e m e n t s are r e l a t e d to U a n d V b y u= V(r, 0) cos0-
V(r, 0) sin0+u
0-~0y,
v = U ( r , 0 ) sin 0 + V ( r , 0 ) cos 0 + v0 + ~ x
(A4) (A5)
a n d where u0, v0 a n d ~0 are rigid b o d y degrees of freedom. H a v i n g specified the n a t u r e of the d i s p l a c e m e n t field it is a s i m p l e m a t t e r to o b t a i n the stiffness m a t r i x using s t a n d a r d finite e l e m e n t procedures. T h e details of this m a y b e f o u n d in reference [11]. H a v i n g o b t a i n e d the stiffness m a t r i x a n d s u b s e q u e n t l y o b t a i n e d the d i s p l a c e m e n t s at the n o d a l p o i n t s the stress i n t e n s i t y factors are o b t a i n e d using a m e t h o d a n a l o g o u s to the well k n o w n J integral a p p r o a c h . This p r o c e d u r e is d e s c r i b e d in detail in [11] a n d is also a p p l i c a b l e for elastic analysis [12].
Acknowledgements M a n y p e o p l e at A R L a n d in R A A F were conc e r n e d in the d e v e l o p m e n t of this repair, n o t all c a n be a c k n o w l e d g e d individually. However, we p a r t i c u l a r l y t h a n k E n g i n e e r i n g Facilities D i v i s i o n A R L , for d e v e l o p m e n t of the C r a c k P a t c h i n g U n i t , Mr. B.C. Hoskin, Sq.Ld. M. R e y m o n d , Mr. J.D. R o b e r t s , Mr. A . W . Rachinger, Mr. T.G. Hill, Mr. B.C. Bishop a n d Mr. T. van Blaricum for their significant c o n t r i b u t i o n s to various p a r t s of the project.
References [1] A.A. Baker, Composites, Vol. 9, 1978, pp. 11-16. [2] A.A. Baker and M.M. Hutchison, "Fibre composite reinforcement of cracked aircraft structures" Proceedings of 33rd Annual Technical Conference of the Plastics Industry, 1978, Section I7-E. [3] R. Jones and R.J. Cailinan, J. Struct. Mech. Vol. 7, 1979, pp. 107-130.
[4] R.A. Mitchell, R.M. Wooley and D.J. Chivirut, A.I.A.A., Vol. 13, 1975, pp. 744-749. [5] A.A. Baker, S.A.M.P.E. Journal, Vol. 15, 1979, pp. 10-17. [6] A.A. Baker, MJ. Davis and G.A. Hawkes, "Repair of fatigue cracked aircraft structures with advanced fibre composites: Residual stress and thermal fatigue studies", to be published in Proceedings of the lOth International Committee on Aeronautical Fatigue Symposium, May 1979. [7] R. Jones and R.J. Callinan, J. Fibre Science and Technology, Vol. 14, 1981, pp. 99-111. [8] R. Jones and R.J. Callinan, J. Struct. Mech., Vol. 8, 1980, pp. 143-149. [9] M.C. Locke, W.M. Scardino and H. Croop, "Non-tank phosphoric anodize method of surface preparation for repair bonding", 7th S.A.M.P.E. Tech. Conf., 1975. [10] J.R. Hart-Smith and F.E. Kiddie, "Effects of heat on fatigue in aircraft structure', R.A.E. Reports and Memorand~ No. 379.8, 1975. [11] R. Jones and R.J. Callinan, Int. J.:l=racture, Vol. 13, 1977, pp. 51-64. [12] R. Jones, K.C. Waiters and R.J. Callinan, "Hybrid elements and elastic plastic fracture mechanics" Proc. 4th Int. Conf. in Australia on Finite Element Methods, Melbourne, Aug. 1982.