Chapter 2. Materials selection and engineering 23 If the problem causing the need for the repair was fatigue or corrosion, it may be more appropriate to use a composite for the repair as these materials are effectively immune to these problems (composite repair layups generally have fibre dominated properties which are immune to fatigue whereas layups with matrix dominated properties may be susceptible to fatigue). The repair material chosen can also be important where subsequent inspections are required and in many cases the use of boron/epoxy composites is advantageous as eddy current methods can be used to readily detect the crack underneath the repair. This is usually more difficult if a metallic or graphite fibre patch is used due to the fact that these materials are electrically conducting. Metallic materials will require the use of stringent surface preparation and surface treatment processes to obtain a durable bond, however, if a corrosion inhibiting primer is used, these processes could be conducted elsewhere and the patch stored prior to use. Composite repairs using thermosetting matrices such as epoxies are comparatively easier to prepare for bonding, although the processes required are still important [2]. Thermoplastic composites are in general harder to bond to than the more commonly used thermoset composites. Finally metals lend themselves best to relatively flat repair locations due to the difficulty in accurately forming a metallic sheet to a curved profile. This is one of the strengths of composites where the desired shape can be formed into the repair during cure. Further considerations for the selection of a metallic material may include corrosion and patch thickness. To avoid galvanic corrosion problems between dissimilar metals, a sensible choice would be to use the original material for the repair material as well. Where this is not possible, a check should be made to ensure that different repair materials would not be susceptible to corrosion. For example, repairs to a graphite/epoxy component will often be performed with a graphite/ epoxy material as well. Use of an aluminium material in this situation would be unusual as the aluminium will readily corrode if in galvanic contact with the graphite fibres. The adhesive should serve as an electrically insulating layer, however, the more usual alternative to a graphite patch in this situation would be titanium which will not corrode should the insulation break down. In situations where the thickness of the repair is critical (on an aerodynamic surface for example) consideration may be given to either steel or titanium to repair aluminium. The greater stiffness of these materials should permit the design of a thinner patch than would be possible with aluminium. Again consideration should be given to possible galvanic coupling and potential corrosion problems in this situation and it is possible that the choice of a composite may be preferable. Laminated metallic materials have been developed in the Netherlands which consist of layers of composite sandwiched between thin aluminium alloy sheets [3]. Where the composite used is kevlar (or aramid) the laminate is referred to as ARALL (aramid reinforced aluminium laminate) and if the composite used is glass fibre, the laminate is referred to as GLARE (Chapter 14). The fundamental idea behind the development of these materials is to combine the traditional advantages of both metals and composites. The composite component confers increased fatigue strength and damage tolerance to the structure, while the aluminium allows the use 24 Advances in the bonded composite repair of metallic aircraft structure of conventional metallic forming, fastening and manufacturing processes for reduced cost. GLARE has been proposed as a possible material for use in bonded repairs and in particular has been used as a material for the repair of damage to the fuselages of transport aircraft. The principal advantage of GLARE in this situation is the high coefficient of thermal expansion. Work by Fredell et al. [4] and Chapter 14, has shown that for repairs to thin fuselage skins which will mostly see pressurisation loads at cruising altitudes (-55 "C), the higher coefficient of thermal expansion of GLARE provides structural advantages compared with composite alternatives (see Section 2.6 for further discussion). On the other hand the low specific stiffness of GLARE results in a much thicker patch than for a high modulus composite material, and this needs to be carefully considered in the design to ensure that bending effects due to neutral axis offset are not excessive and that high stresses at the ends of the patch are alleviated by tapering for example. Finally, it may be possible to use nickel as a repair material in some specific circumstances for example where geometry is complex. The repair of a crack in the comer of a bulkhead pocket is a good example. Nickel can be electroformed to replicate the surface of a mould with very high precision, and therefore it should be possible to produce an electroformed nickel patch which will fit precisely into the pocket. As mentioned above, the isotropic nature of the nickel would be an advantage in this situation, although care needs to be made to ensure that the electroforming process does not produce planes of weakness within the electro- form. Work is underway to evaluate this method as a repair option for a damaged army gun support structure [5]. In situations such as this where a certain degree of rough handling can be expected, the hard, damage resistant surface of the nickel provides another important advantage over a fibre composite repair. 2.2.2. Non-metallic materials The two main non-metallic materials used are boron/epoxy and graphite/epoxy composites. Glass fibre composites are not used due to their low stiffness and kevlar composites while strong and stiff in tension have relatively poor compression performance. Boron fibres were first reported in 1959 and were the original high modulus fibre before the development of graphite fibres in the 1960s. Boron composites were used to produce aircraft components such as the skins of the horizontal stabilisers on the F-14 and the horizontal and vertical stabilisers and rudders on the F-15. The use of boron composites in large-scale aircraft manufacturing has largely stopped now due to the development of more cost-effective graphite fibres. The production process for boron fibres is time consuming and does not lend itself to mass production in the same way as modem methods for producing graphite fibres. For this reason the price of boron fibres has not dropped as significantly as that of graphite fibres which are now at around I/lOth the cost. Boron fibres are manufactured individually by chemically vapour depositing boron onto a heated tungsten wire substrate from boron trichloride gas in a reactor. The fibres are Chapter 2. Materials selection and engineering 25 available from Textron Speciality Materials in 100 and 140 micron diameters and commercial pre-pregs are available with either 120°C or 175°C curing epoxies. The fibre diameter is significantly larger than normal graphite fibres due to the presence of the tungsten core. Attempts have been made in the past to use a carbon filament precursor to reduce the production costs, however, these boron-carbon filaments have generally not had the high level of strength that can be produced with the tungsten filament precursor. Boron fibre is an extremely hard material with a Knoop value of 3200 which is harder than tungsten carbide and titanium nitride (1800 -1880) and second only to diamond (7000). Cured boron composites can be cut, drilled and machined with diamond tipped tools and the pre-pregs are readily cut with conventional steel knives. In practice the knives cannot actually cut the hard fibres, however, gentle pressure fractures the fibres with one or two passes. “Snap-off’ knife blades are commonly used as the cutting edge is rapidly worn by the hard fibres. Although it is possible to cut complex shapes with the use of templates, laser cutting has been shown to be the most efficient way to cut a large amount of non-rectangular boron plies. Circular patches, for example, are readily cut using a laser cutter with the pre- preg supported on a backing material such as Masonite. The combination of very high compressive stiffness, large fibre diameter and high hardness means that boron fibres can readily penetrate skin and care must be exercised in handling boron pre-preg to reduce the chance of splinter-type injuries. If a fibre does enter the skin, it should be removed very carefully with he tweezers. Trying to squeeze the fibre out must be avoided as the fibre may fracture into smaller segments. The stiffness and diameter of boron fibres also restricts their use in small radius corners. The 100 micron diameter fibre can be formed into a radius of 30 mm, but this is about the limit than can be comfortably achieved. The smaller diameter of graphite fibres makes it the choice for smaller radii situations. In most other aspects, boron pre-pregs handle and process in a similar fashion to the more common graphite pre-preg materials. As a repair material, boron/epoxy composites have a number of advantages [ 1,6] including; 0 an intermediate coefficient of thermal expansion which helps to minimise the level of thermally induced residual stress which results from an elevated temperature cure. This contrasts with graphite fibres mentioned below. 0 relatively simple NDI is possible using eddy currents through the repair patch to detect the extent of the defect. This is possible due to the non-conducting nature of the fibres. 0 no galvanic corrosion problems when bonded to common airframe materials. 0 a good combination of high compressive and tensile strength and stiffness (the compressive strength of a unidirectional B/EP composite is 2930 MPa compared with 1020MPa for HMS GR/EP) Graphite fibres are now available in a very wide range of properties and forms and improvements in manufacturing processes has seen the cost of the fibres reduce over the past 25 years. Although the fibres are not as hard as boron, the cured 26 Advances in the bonded composite repair of metallic aircraft structure composites are very abrasive and diamond tipped tools are normally used for cutting or machining. The fine graphite laden dust from such operations is believed to be a health hazard and so measures to control this hazard must be taken. This electrically conducting dust can also cause problems with electrical equipment if it is not removed and filtered from the room air. Graphite pre-pregs are commonly available as 120°C and 175°C curing systems and lower temperature cure resins are also available now for use in repair situations. Graphite fibre is an unusual material in that it has a slightly negative coefficient of thermal expansion, which means that the fibres contract slightly in the axial direction when heated. This results in relatively high levels of thermally induced residual stress if the cured composite is bonded to the structure with an elevated temperature curing adhesive. As well, the fibres are electrically conducting and will cause galvanic corrosion of aluminium if the two are in electrical contact. Due to the electrical conductivity it is more difficult to use eddy-current NDI methods with these materials to check the position of a crack under the patch for example. Graphite composites are significantly cheaper than boron composites and are available from a very wide range of suppliers. They offer a wide range of properties for design and with epoxy resin matrices are readily processed and can be cured to complex shapes to suit the damaged structure. If a repair is required to a tight comer with a small radius, graphite fibres would be preferred to boron as mentioned above. Repairs to aircraft are usually weight critical and so the specific properties of the various repair materials are therefore of interest. Table 2.2 compares the mechanical and thermal properties of some candidate patch or reinforcing materials. This comparison includes boron/epoxy (b/ep) and graphite/epoxy (gr/ep), the metal/ composite laminates GLARE and ARALL and typical high-strength aluminium and titanium alloys - which also represent the metals to be repaired. 2.2.3. Patch material selection Many of the criteria for selection of a successful repair material have been discussed in the above two sections. The reader is referred to Sections 2.1 and 2.2 for a complete discussion of the issues and in this section a summary of the main points is given referring to the four main repair materials and some of the main design issues that are commonly faced. 0 Patching efficiency: High tensile stiffness is required to minimise the crack opening displacement after repair and therefore keep the stress intensity and crack growth down. The fibre composite materials are naturally more efficient than either the conventional or laminated metallic materials (refer Table 2.2 for specific stiffness i.e. modulus divided by destiny). 0 Operating temperature: For sustained high temperature operation over 1 50"C, a titanium patch may prove to be the best solution. Conventional aluminium alloys and the laminated metals would need to be carefully investigated as there are a range of upper temperature limits depending on the alloy and heat treatment involved. In general, most aluminium alloys could withstand extended Chapter 2. Materials selection and engineering 21 Table 2.2 Relevant materials mechanical and physical properties for component and patch materials. Thermal Shear Critical Fatigue expansion Modulus modulus strain Strain Density coefficient Material GPa GPa x 10-~ x 10-~ (g/cm3) oc x IOP Aluminium alloy Aluminium alloy Titanium alloy 6 Boron/epoxy b/ep (unidirectional) Graphite/epoxy gr/ ep (unidirectional) Aluminium laminate GLARE 2 Aluminium laminate ARALL 3 Electroformed Nickel 1015 T6 2025 T3 A1/4V 12 12 110 208 max 20 min 148 max 12 min 65 68 201 21 21 41 1 5 na na 16 6.5 3.3 4.5 3.3 8.8 6.8 7.3 1.0 13 12.0 5.2 3.3 8.9 3.3 1.7-3.4 na 2.8 2.8 4.5 2.0 1.6 2.5 2.3 -9 23 23 9 4.5 min 23 max - 0.3 min 28 max - 15 - 16 13 Notes: (a) Maximum modulus and minimum expansion coefficient are in the fibre direction, other values are for the transverse direction, (b) shear modulus values for the composite are for through-thickness deformation, (c) critical strain refers to failing strain for the composites and yield strain for the metals, (d) fatigue strain refers to approximate strain for crack initiation at lo6 cycles, R - 0. periods at 120°C, which is slightly higher than the normal operating temperature of 105°C for a 175°C curing composite pre-preg. Higher temperature curing resins are available for composites, although the availability is not as high and depending on the system involved, processability may be reduced. 0 Residual stress: If a repair (cured at elevated temperature) is likely to see extended service at low temperatures (for example a fuselage repair to a transport aircraft - [4]), the best choice may be either a conventional or laminated metallic material where the coefficient of thermal expansion is more nearly matched to the structure. In this situation, graphite/epoxy repairs and to a lesser extent boron/epoxy repairs will result in higher levels of thermally induced residual stress [7]. 0 Cost: Although not usually a major driver, conventional metallic materials would offer the lowest material costs, followed by the laminated metals, graphite composites and the boron fibre composites are the most expensive. Analysis of repair costs need to be done carefully as often a composite repair may prove to be cheaper than a metallic repair despite greater material costs. This is largely due to the excellent formability of composites and the reduced time required to form the repair patch to the desired shape. 0 Inspections: If full use is made of the benefits of bonded repair technology and the defect is left in the structure under the repair, it is likely that future non- 28 Advances in the bonded composite repair of metallic aircraft siructure destructive inspections will be required to confirm that the defect has not grown significantly in size. Boron composites are well suited to such circumstances, as the routine use of eddy currents will detect the presence of fatigue cracks for example under the patch. The detection of defects with eddy currents under highly curved boron repairs is more difficult as is the detection of defects under any sort of graphite repair due to the conductivity of the fibres. Detection of defects under bonded metallic repairs can be difficult and may involve the use of X-rays or ultrasonics. 0 Weight: If the repair is to be made to a weight critical component such as a flight control surface, materials with the highest specific properties are desirable. The composite materials will enable repairs with greatly reduced weight compared with the metallic materials. This same point is also of relevance where aerodynamic smoothness is important. Composite repairs will typically be one- third the thickness of an aluminium repair and so will provide significantly less drag. 2.3. Adhesive systems Adhesive technology has undergone rapid growth over the past 50 years and adhesives are now widely used in markets such as automotive, aerospace, construction, packaging and consumer appliances. Most common adhesives can be usefully categorised as belonging to one or more of the following classes; structural, hot melt, water-based or pressure sensitive. Of these only the structural class is of interest in this book. Structural adhesives are defined as those adhesives capable of withstanding significant loads and capable of bonding together adherends also capable of carrying significant loads. For the purposes of this book, shear strengths of 10MPa would be seen as the minimum requirement. 2.3.1. Adhesive types Within the structural adhesive class are a number of adhesive types based on chemistry. The most important are epoxies, modified acrylics, polyurethanes, cyanoacrylates, anaerobics, phenolics and polyimides. Anaerobics cure in the absence of oxygen by free radical polymerisation and are widely used in threaded assemblies to prevent loosening of nuts. They can develop high shear strengths but generally have limited temperature capability and are not used for Bonded Repairs. Cyanoacrylates cure due to the presence of water molecules on the adherends which act as initiation sites for polymerisation. They have excellent shear strength but are comparatively brittle with poor peel strength, are not suitable for filling gaps and are degraded by moisture. Relatively high shrinkage stresses on cure also mitigate against their use in Bonded Repairs. Polyurethanes have good toughness and flexibility, but tend not to have the high shear strength and temperature capabilities that are required for bonded repairs. Phenolic adhesives were the original structural adhesives used in aircraft construction but tended to be very brittle until the Chapter 2. Materials selection and engineering 29 introduction of modified phenolics (the “Redux” adhesives) which had higher peel strength. Phenolic adhesives exhibit excellent bond durability and the modern nitrile modified phenolics are widely used in a range of demanding applications. In general, however, they require high cure temperatures and pressures which may be difficult to accommodate in a repair situation. The other main structural adhesives are those capable of very high temperature operation such as the polyimide (PI) or bismaleimide (BMI) adhesives. These could be considered in specialised repair applications, however, compared with epoxies or acrylics they tend to be difficult to cure. The two adhesive types used most successfully for Bonded Repairs are the epoxies and modified acrylics. The properties of these adhesives are discussed in greater detail in the next section. Acrylics are normally produced in paste form, however, epoxies are commonly available in both paste and film versions. Film adhesives have the resin and curing agents pre-mixed at the factory and are then coated onto a thin carrier cloth or scrim in the form of a thin film. The advantages of this are that mistakes can’t be made in mixing the correct ratio of hardener, the film makes it easy to achieve uniform thickness bondlines and film adhesives are much easier to apply and handle than pastes. Disadvantages are increased cost and the resin is effectively curing as soon as the hardener is mixed and therefore film adhesives must be refrigerated to provide a reasonable shelf life. 2.3.2. Adhesive properties Epoxies come in a very wide range of formulations and types but are generally characterised by high levels of strength, good temperature capability, low shrinkage stresses on cure and the ability to form durable bonds. Epoxies are normally considered to be the most expensive of the common adhesive types (although are not as expensive as the high temperature polyimides). The ability to form durable bonds is highly dependent on the level of surface treatment that is applied to metallic adherends in contrast to the behaviour of acrylic adhesives. The temperature capability of the adhesive is dependent on the cure temperature and so for repairs to structure that sees high temperatures, an elevated temperature cure is required. Room temperature curing epoxies are commonly available in paste form (usually two components) and these adhesives can often provide moderate temperature capability with a post cure to above the operating temperature. Some pastes can also provide higher temperature capability, however, for service at 100°C or higher, film adhesives are commonly used. Unmodified epoxies are inherently brittle materials like phenolics and so most commercial systems are modified with the addition of the toughening agent which is commonly an elastomer. Modified acrylics or second generation acrylics were developed during the 1960s from the original acrylics which were too brittle to be of practical use in structural joints. The rubber toughened acrylics have good shear and peel strengths although the shear strengths are generally not as high as those of the epoxies. They usually cure rapidly at room temperature, in some cases within 1 to 2min, and they have 30 Advances in the bonded composite repair of metallic aircraft structure the ability to readily bond a range of different adherend materials. The ability of these adhesives to develop good adhesion strengths with limited surface treatment is due to the acrylic monomer which is a free flowing liquid of low surface tension. Modern acrylics are able to produce strong, durable bonds to unprepared aluminium and steel surfaces; epoxy adhesives are unable to achieve this. Commercially available systems now do not require mixing of two components but instead can use an activator applied to one adherend and the adhesive to the other which simplifies the use compared with two-part epoxies. Disadvantages include an odour that some people find objectionable, limited temperature capability and limited pot life which can be a problem for larger repairs. Acrylics are widely used in industrial applications where the ability to rapidly bond poorly prepared steel sheet is an important advantage and is able to replace the use of spot welding or riveting. 2.3.3. Adhesive selection The designer of a bonded repair has a very wide range of adhesives to choose from, although in practice the selection is usually made from those adhesives that are readily available to the company. The two most important selection criteria are temperature and load carrying capability. A conservative approach is to use an adhesive for the repair of equal temperature capability to the original structure. This is typically 120 "C cure for commercial (subsonic) aircraft and 175 "C cure for military (supersonic) aircraft. However, the use of a 175 "C curing adhesive during manufacture does not necessarily mean the structure will be exposed to such high temperatures. Often a 175 "C adhesive is used in manufacture to be compatible with the 175°C curing pre-preg so that the part can be cured and bonded in one autoclave cycle. If the actual operating temperature of the component can be shown to be 60°C for example, it is possible to produce a sound repair with a 120 "C curing adhesive. It should be noted that the use of 175 "C curing adhesives for repair has in itself caused significant problems when the structure to be repaired contains honeycomb core and water is present within the core. At around 140 "C, the pressure generated inside the core by the air and water exceeds the flat-wise tension strength of the skin to core adhesive and the skin can be disbonded by the pressure. The risk of such damage occurring is greatly reduced at 120°C and at least one adhesive manufacturer has developed a 120°C version of the standard 175°C adhesive system for use during repair to honeycomb structure. Bond durability (particularly for epoxies) is generally related to cure temperature and it is common to find excellent bond durability for 175°C systems, good durability at 120 "C but only fair to good durability for room temperature curing adhesives. The improvement in durability for the 175 "C cure, however, needs to be weighed up against the other problems which can develop such as blown skin to honeycomb core bonds and increased thermally-induced residual stresses. The required load carrying capacity of the adhesive needs to be carefully considered. Some manufacturers of structural adhesives are now beginning to Chapter 2. Maierials selection and engineering 31 provide design data in the form of shear stress/shear strain data. The more common lap shear strength is not suitable for use in a bonded repair and is generally only useful in comparing one adhesive to another. Details of the data that is required for design based on adhesive properties is given in Chapter 4, and if it is necessary to generate this data, appropriate test methods are described in Section 2.5 and Chapter 4. Two key parameters are the shear strength and plastic strain to failure. The adhesive needs to have sufficient shear strength so as not to yield excessively under the design loads, and care should be taken in designing with relatively brittle adhesives which cannot provide a soft, yielding type of failure under high loads. Less well understood is the ability of the adhesive to withstand through-thickness stresses, i.e. those perpendicular to the plane of the joint. Conventional design wisdom with adhesive joints is to eliminate such stresses by the use of different design techniques. In many cases it is possible to eliminate or greatly reduce the magnitude of these stresses simply by the use of sensible design features such as tapering of the end of the repair. In some circumstances, however, it is not possible to reduce these stresses and some examples are given in Chapters 30 and 33. In repairs to structure involving a high degree of curvature, the question then becomes one of determining the capacity of the adhesive to withstand the through-thickness or peel stresses that are present. There is currently no generally agreed test method to generate design data for this situation, although a novel test specimen has been proposed which may be suitable for this purpose [8,9]. Any repair design where high levels of peel stress are likely to be present needs to be very carefully considered and would be expected to require extensive analysis and experimental validation for certification. The work described in [8] is aimed at increasing the understanding of the performance of adhesives under peel stresses, however, while this may lead to some easing of certification requirements, the sound engineering practice will continue to be to design peel stresses out of an adhesive joint where ever possible. Other criteria which may be important in the selection of a repair adhesive could be availability and the ability to cure at low temperatures. Availability and the requirement for refrigerated storage could be important at some forward Air Force bases for example, where only a very limited range of adhesives may be available at short notice. When rapid repairs have to be made in primitive conditions, for example to battle damage, it may not be possible to provide refrigerated storage and therefore only two-part adhesives would be available. As described in Section 2.6, thermally-induced residual stresses are produced when the repair material has a different coefficient of thermal expansion to the substrate and an elevated temperature cure is necessary. The obvious way of reducing the level of such stresses is by reducing the cure temperature of the adhesive as much as possible. Some adhesives are able to cure at temperatures lower than their advertised cure temperature although this is not always the case [lo]. Film adhesives are often sold as either 120 "C or 175 "C curing systems (partly for compatibility with other pre- pregs etc.), however, a careful examination of the thermodynamics of cure can indicate that the optimum cure temperature is different from these advertised temperatures. Considerable care must be taken if a decision is made to cure at 32 Advances in the bonded composite repair of metallic aircraft strucmre temperatures other than those advertised to ensure that other properties are not compromised. The ability of the adhesive to remain durable in the operating environment is normally of critical importance and consideration may need to be given to the influence of solvents or chemicals which the adhesive may be exposed to. For example some repairs have been applied inside aircraft fuel tanks or in regions where the adhesive is exposed to hydraulic oil. Most epoxies and acrylics have very good resistance to solvents and chemicals and so these types of exposures have not been of major concern to date, but do need to be checked on an individual basis Where possible it is recommended that repairs are cured under positive (as compared to vacuum) pressure and further details are given in Chapter 25. When the use of vacuum bag pressure is the only alternative, consideration may need to be given to the void content in the cured adhesive bondline (Section 6.2). Some adhesives do not cure well under vacuum and heavily voided bondlines can result. There is some evidence to suggest that moderate amounts of voids do not adversely affect fatigue strength, however, in general significant void contents in structural adhesive bondlines are to be avoided. v11- 2.4. Primers and coupling agents A range of different chemicals may be required for effective surface preparation and a detailed scientific discussion of these is given in Chapter 3. This section will look at some of these chemicals from a materials engineering perspective and consider some of the common factors that may be need to be considered in the overall design of the repair. From Section 2.1 it is clear that significant attention must be paid to the surface treatment of metallic adherends prior to bonding if a strong, durable adhesive bond is to be produced. There are two major types of treatments for aluminium alloys that for historical reasons have developed in Europe and North America. In Europe, the preferred treatment is the use of a chromic acid etch to produce a hydration resistant oxide, whereas in North America the use of phosphoric acid is preferred. Both treatments have been used successfully in aircraft manufacturing and are capable of producing highly durable bonds. Components are dipped into tanks of acids and other chemicals in the factory to produce the required oxide structure for bonding. The difficulty comes in transferring this technology to a repair situation. For example when acids are used on an assembled aircraft structure, care must be taken to completely remove the acids or corrosion may result. Boeing in particular have developed procedures whereby the same technology as used in manufacturing can be applied to some repairs. The phosphoric acid containment system (PACS) uses vacuum bags over the repair site to transport the acid across the surface. This contains the acids to minimise health and safety concerns and permits a final flush with water to remove the acid from the aircraft surface. The anodisation is carried out under the bag as well. This [...]... the predominant forces involve hydrogen bonds in which the hydroxyl 41 Baker, A.A Rose, L.R .F and Jones, R (e&.) Advances in the Bonded Composite Repairs of Metallic Aircraft Structure Crown Copyright 020 02 Published by Elsevier Science Ltd All rights reserved  42 Advances in the bonded composite repair of metallic aircraft structure groups on the metal oxide interact with hydroxyl groups in the polymer... Pre-existing debris is consolidated into the surface during impact deformation of the metal surface The grit-blast treatment improves the hydrophilic wetting of the surface (Figure 3.15) [36] and the durability of the bond (Figure 3.9) over that of the abraded surface Some of the improvement in wetting can be attributed to a decrease in the concentration of hydrophobic contaminant and some to the roughening... organic contaminants on the adherend surface prior to subsequent preparation steps Factory facilities often use vapour degreasing to reduce the concentration of organic contaminant on components to be bonded A solvent such as trichloroethylene is evaporated in a closed space then allowed to condense and drip from the soiled components Organic contaminants are slowly dissolved in the 58 Advances in the bonded. .. microscopy will provide macroscopic information concerning the locus of fracture and the presence of voids or defects The term cohesional failure describes fracture totally within the adhesive, leaving adhesive on both separated adherends The term adhesional failure describes a fracture at one interface with the adherend, resulting in one face 46 Advances in the bonded composite repair o metallic aircraft. .. developed In this test, the critical factor is the correct ratio of the support span to specimen thickness For the 5 521 /4 B/Ep composite, an ILS value of 97 MPa or above, indicates the material is in good condition An advantage of the use of metallic materials for repair patches is their infinite shelf life No testing is required before use, other than to confirm that the alloy and heat treatment are correct... from factory or field experience since most laboratories are held to close environmental tolerances and do not resemble the workshop environment 3.1.4.3 Constraints for on-uircruft repairs On -aircraft repairs impose additional constraints on processes and procedures The considerations include: accessibility of the area, limitations in the use of corrosive chemicals, adequacy of environmental controls... interface The adhesive bond durability is very sensitive to the presence of hydrophobic contaminant on the adherend, but the dependence involves a complex combination of the nature, the concentration and the distribution of the contaminant Studies of bond durability with one epoxy film adhesive following deliberate contamination of prepared aluminium adherends showed sensitivity to the nature of the. .. model that describes moisture ingress to the adhesive metal interface and three degradation reactions The magnitude of microcavities ahead of the crack tip would control the rate of moisture diffusion and the dominant degradation reaction path would determine the position of bond weakening and the dominant locus of failure moisture ingress, the adhesive bond can degrade by any one of three reaction... and the engineering tools available for the through-life management of adhesively bonded structure are primitive The books by Kinloch [13] and Minford [ 141 are, respectively, an excellent introduction to adhesion and adhesives and a compendium for adhesion with aluminium alloys It is not the intent of the authors to reproduce a summary of these works here The focus will be on surface treatments for repair. .. repair bonding, giving consideration to the atomic nature of the bond interface and the relationship between microscopic behaviour and macroscopic mechanical properties It cannot be over emphasised that a strong adhesive 44 Advances in the bonded composite repair o metallic aircraft structure f bond does not imply a durable bond The influence of adherend surface treatment on bond durability is therefore . tolerance to the structure, while the aluminium allows the use 24 Advances in the bonded composite repair of metallic aircraft structure of conventional metallic forming, fastening and manufacturing. referred to Sections 2. 1 and 2. 2 for a complete discussion of the issues and in this section a summary of the main points is given referring to the four main repair materials and some of the. for the composite are for through-thickness deformation, (c) critical strain refers to failing strain for the composites and yield strain for the metals, (d) fatigue strain refers to approximate