40207_39

REPAIR ASPECTS OF COMPOSITE AND 39 ADHESIVELY BONDED AIRCRAFT STRUCTURES Anton L. Seidl

39.1 INTRODUCTION

39.1.1 MANUFACTURABILITY AND MAINTAINABILITY OF COMPOSITE AIRCRAFT STRUCTURE

To the manufacturer, weight reductions, structural requirements, manufacturability and production costs have long been obvious pri- orities. Only recently, however, and only as a consequence of persistent user demands, have maintainability and repairability been added to this list. From the operator’s perspective, nevertheless, composite structures continue to be a mixed blessing. Clearly, and despite state- ments being heard to the contrary, the industry would be loath to give up the many obvious advantages gained through the use of composites and revert to all-metal airplanes. However, the maintenance problems associated with composites cannot be underestimated and may well be regarded as the weak link in the new technology chain.

39.1.2 METAL REPAIRS COMPARED WITH COMPOSITE REPAIRS

Compared to the relative simplicity of conventional metallic structures, composites are replete with complexities that continue to baf- fle and confuse maintenance workers trained only in the traditional, i.e. metalworking, skills. The glossary of terms alone, as used by

Handbook of Composites. Edited by S.T. Peters. Published in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7

the composites industry, is an unfamiliar lan- guage to the uninitiated. Inspectors are often at a loss when attempting to describe a condition they perceive as a defect; the words simply do not exist in their standard lexicon. It is intuitively clear to even the casual observer that repairs using mechanically fastened conventional materials can be effected quickly, under almost any atmospheric conditions, and with minimal investments in tooling, raw materials, and training. In contrast, repairing even relatively minor damage on composite structure requires an array of non-conventional materials, highly skilled and experienced technicians, special tooling and equipment, access to production drawings (to locate and interpret the many hidden features characteristic of composite structures), a controlled environment in terms of temperature and humidity, time-consuming preparatory work, cold storage of shelf-life limited and occasionally hazardous materials, lengthy resin cure cycles, post-repair NDT, and legally mandated record-keeping and follow-up activities.

39.1.3 COMPOSITE REPAIRS: AN AIRLINE PERSPECTIVE

The aim and purpose of this presentation is to highlight the principal aspects of composite structure repairs from an airline perspective. An attempt will be made to:

1. describe some of the more common defects and conditions encountered in service;

858 Repair aspects of composite and adhesively bonded aircraff structures

2. give a brief summary of common mainte-

3. describe a limited number of typical repairs. nance practices;

39.2 DAMAGE ASSESSMENT

39.2.1 IMPACT DAMAGE - NON-METALLIC STRUCTURE

Foreign object impact without skin penetration

Prior to any repair action, it is important to determine the extent of the damage sustained by the structure. One must always assume that the actual damage is more extensive than the visible damage'. This is especially true for carbon fiber-reinforced composites with non-toughened 177°C (350°F) cured matrix resins. After a foreign object impact, there is generally, but not invariably, some visual indi- cation in the form of damaged paint. However, because of the elasticity of high modulus fibers, the laminate often 'springs back', leaving residual subsurface damage in the form of broken fibers, ply separations and in the case of sandwich panels, crushed core and disbonded face sheets. Tap testing is generally sufficient to delineate the extent of the damage and should be conducted before removing any damaged materials. Defects may also propa- gate during the removal process, often as the result of stress relief.

Skin penetrations: holes, cracks, tears, gouges, cuts and abrasions

On the 'wetted' outer surface of the aircraft, even minor penetrations of the face sheet must be regarded as serious because once an opening exists, the part has been rendered permeable to atmospheric moisture and aircraft system fluids. Ingested water can and will degrade the affected part, leading to premature failure. Fluids such as hydraulic oil, when allowed to enter, contaminate both laminates and honeycomb core materials, making

subsequent repairs more difficult to perform. When surface defects are detected, it is imperative to determine the extent of the damage that may already have occurred and if possible, evacuate and decontaminate the panel. As a first line of defense against any further structural deterioration, foil tape should be applied at the earliest opportunity2.

Severe degradation is generally quite obvious, having resulted in visible disbonds and delaminations. If the damage is due to a recent event, and the process of deterioration has only begun, the amount of ingested fluid may still be relatively small and if the precise location of the contaminant can be determined, complete evacuation and purging may be successfully accomplished.

Limitations of moisture detectors3

Commercial moisture detectors are extremely useful devices requiring no special training. Where appropriate, they may be used to determine how far any ingested water has spread into the core cells adjacent to the point of impact. Moisture detectors, however, are effective only on non-metallic (typically glass or aramid-reinforced structures); they cannot be used on panels containing carbon fibers, or in zones reinforced with metals. Nor are moisture detectors effective through surfaces coated with carbon-filled conductive paints, on panels having metallic coatings, metal- coated fibers, or similar lightning protective and EM1 shielding features.

E@cacy of radiography

Radiography (X-ray) is presently the only available practical technique for determining moisture contamination in panels containing electrically conductive elements. Given the limitations of moisture detectors and the limited availability of X-ray equipment in the field, water detection by X-ray and subsequent evacuation are generally carried out only during depot level maintenance opportunities.

Damage assessment 859

Interim repair actions - ‘speed tape’ repairs

When a composite panel is found to have been penetrated, it is important to prevent further deterioration of the panel. When fluid detrimental to adhesion (hydraulic oils, deicing fluid, engine oils, etc.) is present, the affected area must be thoroughly decontaminated before attempting a repair, or the contaminated material removed entirely.

When a permanent repair is to be deferred, fractured material should be trimmed away and the opening covered with foil tape before the aircraft is dispatched to a location where the appropriate repair facilities exist. Foil tapes must be applied with care to prevent their coming loose in flight. Loose foils have been known to create static noises that can interfere with radio communications.

39.2.2 IMPACT DAMAGE ON METAL-SKINNED SANDWICH PANELS

Unlike laminated face sheets, which may show little evidence of an impact having taken place, thin metal face sheets (common on many honeycomb sandwich panels) invariably become dented or gouged by foreign object impact. The resulting surface irregularities are readily seen.

Minor damage - no skin penetration

Shallow dents may be present that do not nec- essarily result in disbonding of the skin, but there will always occur some crushing of the core cells. A tap test will usually, but not always, determine whether the skin is disbonded. Dents that have not resulted in skin disbonds are generally considered negligible damage and may be filled with an appropriate compound to restore aerodynamic cleanness, provided the added weight does not affect the balance of a critical control surface. Flight control surfaces damaged by hailstones frequently exhibit multiple dents that cannot be repaired by dent fillers without creating an out-of-balance condition.

Effects of skin penetration: corrosion, resin plasticization and core dissolution

Any impact damage resulting in skin penetration must be regarded as serious damage. However, unlike non-metallic core materials, which absorb and diffuse water, non-perforated aluminum honeycomb cores tend to keep any ingested water concentrated about the area of the penetration. Left unattended, prolonged exposure will cause the ingested water to migrate to other areas of the panel by gradual, progressive diffusion through the adhesive bondlines and, preferentially, through the core splice adhesives. As the bonding adhesives absorb moisture they become plasticized and their bond strength weakens.

At the same time, unprotected areas of the face sheet, doublers, substructural components, cut edges and fastener holes, i.e. where the anodic and primer protections have been removed during the manufacturing process, and machined edges of the honeycomb core, are rendered vulnerable to corrosion attack. Ingested water, if left unevacuated for long periods, has been known to initiate chemical reactions that lead to complete dissolution of the aluminum honeycomb core.

39.2.3 DAMAGED PROTECTIVE COATINGS AND SEALANTS: LEAK PATHS

Water ingestion and fluid contamination must be presumed to exist whenever the protective coatings or sealants of a panel have been dis- turbed. The cause may be erosion of the protective finish, substrate corrosion, hail damage, minor collisions, or similar foreign object damage episodes. Leak paths, no matter how small, are detrimental to the long-term structural integrity of the panel because they allow atmospheric moisture, aircraft system fluids, or a combination of contaminants, to enter the structure.

Subsequent ’ground-air-ground’ and ‘freeze-thaw’ cycling are capable of introduc- ing considerable quantities of water and other

860 Repair aspects of composite and adhesively bonded aircraft structures

fluids into the core of a panel. Sandwich panels with thin face sheets of aramid/epoxy are especially vulnerable to moisture contamination via small cracks in the resin gelcoat and at the resin-fiber interface. This phenomenon has been explained as the result of the thermal expansion behavior of the aramid fiber, which is slightly negative in the longitudinal direction and strongly positive in the transverse direction, leading to excessive strain build-up within the weave itself.

39.3 ENVIRONMENTAL DAMAGE AND DEGRADATION

39.3.1 GENERAL EFFECTS OF AGINGz4

All polymeric materials are subject to degradation over time. For this reason, the importance of maintaining protective coatings and sealants cannot be emphasized too strongly. The 'normal' operating environment of an aircraft exposes composite structures not only to considerable static and dynamic loads, but also to significant temperature gra- dients, extreme variations in humidity conditions, and to a number of chemical agents necessary in aircraft systems, the most detrimental being hydraulic fluid, a powerful solvent.

39.3.2 EXPOSURE OF COMPOSITES TO THE 'NORMAL' FLIGHT ENVIRONMENT

There is abundant evidence that the combined effects of stress, temperature, water and other fluids expose bonded and especially fiber- reinforced composite structures to a far wider range of hazards than their baseline metal analogs. Following is a brief description of the most common environmental hazards composite materials are exposed to.

Moisture

While demonstrably corrosive to metals, moisture is far more pernicious in its effect on composites because it plasticizes resins, degrades their mechanical properties and lowers their glass transition temperature4. The latter effect becomes extremely critical when carrying out hot bonded and/or prepreg repairs that require the heating of the structure.

Atmospheric electricity 22-z

Atmospheric electricity, of negligible con- squence to metal structures having inherent conductivities, can have a crippling effect on non-metallics, which compels the operator to place a high priority on periodic testing and proper maintenance of anti-static and lightning protection schemes, i.e. ground paths, bonding fasteners, bus strips, jumper cables, conductive enamel and/or flame-spray coatings, as well as discharge ports.

Chemical contamination

Aircraft fluids and chemicals that are harmless on metals can effectively destroy a composite, if allowed to penetrate its outer protective layers. Chemical paint strippers routinely used on metal aircraft are occasionally - albeit inadvertently - applied to composite surfaces with destructive consequences, even if the exposure is but of short duration. Presently, composites can only be stripped by abrasive, non-chemical methods5.

Overheat conditions

Heat, except for annealing temperatures, is of minor concern with metals; by contrast, the heat resistance of composites is effectively limited by the maximum use temperature of the polymeric matrix. Heat generated by lightning strikes has been known to vaporize matrix resins and create large areas of delamination and fiber fracturing on composite rudders,

Damage removal techniques 861

ailerons, wing and stabilizer tips, nose domes, and nacelle cowling. When exposed to hot gases over long periods, polymeric resin binders, irrespective of chemistry, can become completely destroyed through a process sometimes described as thermo-oxidation. This condition may be found on all types of composites, including those with inorganic matrices, such as metal matrix composites. Preventive maintenance may consist of the application of heat-resistant, ablative or intumescent coatings. Extensive redesign of the detail may be necessary, using metals or, if a fiber composite is to be used, choosing a polyimide or similar high temperature resistant resin system.

39.4 DAMAGE REMOVAL TECHNIQUES

39.4.1 PLANNING THE REPAIR 'THINK BEFORE CUTTING'

After determining the full extent of the damage, the repair technician must consider a range of possible approaches, based on such considera- tions as damage location, access to the damage, required disassembly to create better access, available tooling and repair materials, as well as the allotted out-of-service time. Because most repairs are 'on-condition', i.e. the result of damage events affecting the structure at unpre- dictable locations in a multiplicity of manners, allowing only limited pre-planning, the technician's experience and intuitive problem-solving abilities are of paramount importance.

39.4.2 AERODYNAMIC SKIN DAMAGE REMOVAL

If the damage affects the outer, aerodynamic or 'tool' side of a panel and the backside (the %ag side') is accessible, it is best to remove material from the backside, and as much of the core material as necessary, to gain access to the damage. Using this method preserves as much of the aerodynamically 'clean' surface as possible. Repair work is more easily accomplished

from the backside, with the additional benefit of causing only minimal disruption to the aerodynamic surface. Figure 39.1 illustrates this principle.

If the backside is inaccessible, the damage must be repaired from the aerodynamic skin side, inevitably enlarging the repair surface and making the repair more difficult to perform. With only one side accessible, the question of how best to apply vacuum pressure is always problematic and requires considerable operator skills. (Applying vacuum pressure for a bonded repair is an art form that must be learned as any other.) As an alternative to field repairs, panels are often removed from the affected structure and routed to a repair facility specially equipped to effect the appropriate restorations. It should be noted that the damages affecting the aerodynamic skin surface normally require 'flush' repairs to preserve the original contour, particularly in zones of the aircraft defined as aerodynamically critical. Except for small damages, the tooling and skill levels required to effect proper repairs do not exist at field stations.

REPAIR PLIES-/ -FILLER

\ CORE PLUG REPAIR PLIES-

-iti--FLUSH SIDE

~ N O N - F L U S H SIDE

Fig. 39.1 Aerodynamic skin side repair.


39.4.3 REMOVAL OF METAL FACE SHEETS AND DOUBLERS

Metal-faced sandwich panels are used on wing spoilers, wing and stabilizer panels, flaps, slats, engine cowling, landing gear strut doors, as well as in a multitude of other applications, including aircraft interiors. Damaged, corroded, or disbonded face sheets are generally peeled away after the application of carbon dioxide pellets ('dry ice'). The dry ice is allowed to dwell on the surface until the thermal shock has weakened the bond strength of the adhesive sufficiently to allow the skin to be removed. If done properly, the face sheets separate, leaving the core cells relatively undamaged.

39.4.4 REMOVAL OF COMPOSITE FACE SHEETS AND DOUBLERS

The outer surface of most non-metallic sandwich panels consists of only a small number of prepreg fabric and/or tape plies co-cured onto non-metallic honeycomb core, although precured laminates, secondarily bonded to the core, are also found. Removing damaged or disbonded face sheet materials requires either a rotary sander (physical abrasion) or the use of a hot air gun combined with peel- ing action. Heating the skin laminate has the effect of weakening the resin fillets sufficiently to allow the technician to peel the face sheet materials with only minimal damage to the core.

39.4.5 REMOVAL OF HONEYCOMB CORE MATERIALS

After the face sheet material has been removed, the condition of the honeycomb core must be determined. Because of the high cost and limited availability of some core materials, repair shops attempt to salvage the original material if at all possible.

Aluminum core removal

Severely damaged aluminum core (crushed, corroded, failed node bonds, etc.) should always be replaced. Removal is generally accomplished by using non-metallic scrapers or chisels mounted in a pneumatic rivet gun. Care must be taken to avoid damaging the intact face sheet on the far side. The remaining adhesive fillets on the far side should be abraded with a rotary sander, provided the adhesive is still firmly attached, thils providing a good base for bonding in the replacement core plug. If the far side adhesive is plasticized or unbonded, or corrosion is found between the adhesive layer and the metal skin, the adhesive must be removed for closer inspection and possible reconditioning of the bonding surface. Corroded skins and doublers are routinely replaced.

Non-metallic core removal

Non-metallic core materials are generally replaced if crushed or split at the nodes, or if irreversibly contaminated by oil, hydraulic fluids, or other contaminants that would inhibit subsequent repair resin adhesion and cure. Repair shops often attempt to decontaminate core by flushing out the cells with solvent, a method not always successful and a potential environmental hazard. If the contaminant is water, dehydration of the core by evaporation, placing the part in an oven at a low temperature, is often possible, allowing the material to be salvaged (see Section 39.5). When core replacement becomes necessary, the affected sections are generally cut out with knives or rotary cutters; the resin fillets remaining on the far side are then removed with rotary sanders, to create a proper surface for bonding in the replacement core Plug.

Decontamination 863

39.5 DECONTAMINATION

39.5.1 EVACUATION AND DECONTAMINATION OF POLYMER MATRIX COMPOSITE STRUCTURES

Vulnerability of polymers to fluids

Organic matrix composites typically absorb between one and two percent of their dry weight in moisture under normal service conditions. There exists a certain risk when such assemblies are subjected to the elevated temperatures routinely applied during bonding and laminating repairs. During hot bond repairs, the absorbed moisture volatilizes. The effect on the repaired structure may manifest itself in the form of porosities in the bondline or in the laminate. In severe cases, such as when water is present in the core cells, the pressure resulting from the entrapped steam often results in uncontrollable skin-to-core disbonding. For these reasons, it is always advisable to pre-dry composite panels when moisture contamination exists in detectable quantities, or may be presumed to exist, given the general condition of the part. One should keep in mind that fibers, with the exception of aramid, do not absorb moisture. Moisture absorption is a phenomenon that affects primarily the resin matrix and, secondarily, non-metallic core made from aramid fibers. As a general rule, resin systems cured at 170°C (350°F) or above are more resistant to moisture pick-up than resin systems cured at lower temperatures, which includes all room temperature-cured repair resins and cold- bond adhesives-.

Effect of contaminants on weight and balance

Fluids absorbed by, or otherwise introduced into a structure, induce weight gains and may cause out-of-balance conditions in flight control surfaces. Contamination detected should always be evacuated, the leakage paths identi- fied, repaired and the structure resealed.

Effects of contaminants on structural integrity

Dimensional swelling and plasticization of the resin matrix generally result from exposure to high humidity at high temperatures, exposure to many aircraft fluids, to chemical paint strippers, and to a variety of common solvents. Absorbed moisture lowers the glass transition temperature of the laminate* and may be con- ducive to additional microcracking within the matrix which, in turn, increases the potential for additional moisture absorption. Micro- cracks are considered irreversible, since they remain after the laminate has been completely dehydrated. Absorbed chemicals may or may not affect the structural or mechanical properties of the composite, but generally render the affected part unrepairable because they inhibit repair resin adhesion and cure.

39.5.2 GENERAL PRECAUTIONS

Water evacuation under vacuum pressure at elevated temperatures

Removing water is mandatory in all cases, but the process becomes especially critical if the repair requires the application of elevated cure temperatures under vacuum pressure, which is typical or many in situ heating blanket type repairs. The operator must be aware that, under a vacuum bag, lowering the vapor pressure also lowers the boiling point of the water; at the same time, increasing the temperature increases the steam pressure inside the sandwich (Fig. 39.2). The result is often a failed repair: blown core and disbonded face sheets.

Removing moisture barriers (coatings and films)

Evacuation of composite laminates is best accomplished by first removing any protective coatings and moisture barrier films that may still be present and intact. Barrier materials may be various enamel finishes, sealer

864 Repair aspects of composife and adhesively bonded aircraft structures

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Fig. 39.2 Pressure and evacuation guidelines for honeycomb core repair.

coats, or bondable plastic films. Their removal is essential to create a path for volatiles to escape rapidly.

Removing face sheet materials

Evacuation of honeycomb sandwich panels is most effectively done by removing one of the face sheets. This method exposes the core and allows the thorough flushing of any contaminants with an appropriate solvent. Complete drying should be performed under vacuum pressure at moderate heat.

Flushing contaminants with solvents

Evacuation of chemical contaminants may be accomplished by flushing the core cells with

solvents. Often, considerable quantities may be necessary to purge the contaminated core, creating potential environmental hazards. In many cases, complete core replacement may be the only appropriate action.

39.5.3 SPECIFIC EVACUATION TECHNIQUES

Evacuation of fluids from core with face sheet removed ('open core' evacuation)

Visible liquids should be evacuated by blow- ing filtered, compressed air across the surface. This should be followed up by flushing the core with an oil-free solvent and then allowing the solvent to evaporate completely. Next, several layers of breather fabric are stacked over the panel, the assembly is envelope-bagged

Typical repairs 865

and a vacuum of approximately 67 kPa (20 inHg) is applied. The panel is then heated slowly to approximately 74°C (165°F) and allowed to remain at temperature for a minimum of one hour.

Evacuation of fluids from core with face sheets intact (‘through-the-facesheet’) evacuation

First, all protective coatings and moisture barrier plies must be removed from the areas to be evacuated. Then the gelcoat of the outermost ply should be abraded to expose the fibers. (Fibers inadvertently damaged during this process require subsequent repair.) Next, several layers of breather fabric are applied and the assembly envelope-bagged. Then a vacuum of 34-40 kPa (10-12 in Hg) is applied and the panel heated very slowly (5°C per minute maximum heating rate) to approximately 75°C (165°F) and maintained at that temperature and vacuum pressure for a minimum of 24 h. After this initial drying cycle, the temperature should be increased to 107°C (225°F) and maintained for an additional four hours.

Handling of dried details - inspection and storage

After drying, details should be re-examined and, if satisfactory, stored in a clean, dry environment until the appropriate repair actions can be taken.

39.6 TYPICAL REPAIRS

39.6.1 WET LAY-UP REPAIRS AT AMBIENT OR ELEVATED TEMPERATURES

So-called ’wet lay-up’ repairs are the most frequently recommended because they require only the most basic in terms of equipment, tooling, and repair materials. On the other hand, they are also the most limited in terms of size and applicability because such repairs do not restore the full, pre-damage strength of

the structure. Wet lay-ups normally involve the use of the same type of fabric used in the original construction, in conjunction with a laminating resin capable of room temperature cure under vacuum pressure. The quality of the repair is generally enhanced by applying moderate heat (100°C max) by means of heating blankets, heat lamps or hot air.

Heating blankets

Heating blankets used in conjunction with vacuum pressure repairs should have an out- put (watt density) of no less than 7750 W/m2 (5 W/in’). To facilitate draping over curved surfaces, heating blankets with silicone rubber-embedded elements are preferred over mineral fiber-insulated pads, because of their inherent flexibility. Stiffer pads should only be used on flat surfaces.

Heat lamps

Heat lamps that are used either as the primary heat source, or as a means of augmenting other heat sources, should be 250-300 W tungsten or quartz tube, explosion-proof types. When using heat lamps as the primary source, the effective heat input is controlled by the stand- off distance, as shown in Fig. 39.3. To avoid overheating any portion of the assembly being repaired, thermocouples should be placed at several locations to monitor the temperature throughout the cure cycle. The stand-off distance or the positioning of the lamp should be adjusted as necessary to maintain the cure temperature within specified limits.

Hot air blowers

Hot air blowers similar to hair dryers are often used to accelerate resin cure; they may also be used for reticulation of unsupported film adhesives. Such devices are typically designed with 1000-2000 W heater elements and fan drive motors.

866 Repair aspects of composife and adhesively bonded aircraft structures

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39.6.2 USE OF ADDITIONAL PLIES OVER WET LAY-UPS

Recognizing that wet lay-ups are inherently inferior to autoclave-cured laminates, many repair specifications9 call for the addition of two or more plies of the type of material used in the original construction, as a means of compensating for the loss of stiffness implicit in wet lay-ups. The added plies do, however, result in weight gain and some loss in aerodynamic cleanness.

39.6.3 TYPICAL WET LAY-UP REPAIR PROCEDURES

Damage assessment and removal

1. Determine perimeter of damaged area by

2. Clean area with solvent. NDTlO.

3. Sand off any protective or decorative finishes and coatings; scrape off sealants, especially silicone sealants.

4 Inspect detail for presence of water or other fluid contamination.

5. Evacuate panel using one of the methods described in Section 39.5.

6. Remove damaged materials - face sheets, doublers, and core, using the appropriate techniques described in Section 39.4.

Note: Steps (5) and (6) may be inverted, depending on the condition of the part.

Core plug repair

1. Obtain and prefit replacement core plug, using same as original material, cell size, and density (or an approved substitute).

2. Apply resin compound to edges of core plug to provide a shear tie and insert the plug into the cavity.

3. Apply release film, breather fabric, thermocouples, and vacuum bagging materials. Apply vacuum and check bag for leaks.

4. Cure core splice (shear tie) resin, observing the appropriate time/ temperature relationship specified for the core splicing resin. Maintain vacuum pressure throughout the cure cycle.

5. Remove bagging materials and thermocouples.

Face sheet repair

1. Taper and splice joint area. 2. Sand core plug flush with innermost ply. 3. Vacuum up sanding dust, solvent clean

repair surfaces and allow solvent to dry completely.

4. Using same as original fiber type and weave style, and observing proper yarn orientation, prepare and impregnate each repair ply of fabric with an appropriate laminating resin mixture.

5. Apply repair plies, observing ply stacking sequence and fiber orientation.

Typical repairs 867

6. Apply perforated release film, breather/ bleeder fabric, thermocouples, and vacuum bag. Apply vacuum and check bag for leaks.

7. Cure laminate under 67-81 kPa (20-24 in Hg) vacuum pressure, while observing the appropriate time/ temperature relationship specified for the repair resin. Maintain vacuum pressure throughout the cure cycle.

8. Remove bagging materials.

Restoration of coatings, finishes and sealants

1. Ensure resin is fully cured (must be hard when tapped and resistant to solvents when wiped with solvent-soaked cheesecloth).

2. Reactivate surface by mild abrasion. 3. Clean surface and allow to dry. 4. Reapply finishes, including any anti-static

and lightning-protective coatings that may be required.

5 . Reapply any sealants or other coatings removed for repair.

Materials, tooling, equipment and repair environment 25

1. A two-part epoxy laminating resin of the required chemistry.

2. A compatible core splice resin or compound.

3. Same as original type and style of fabric (unidirectional tape may be replaced with two plies of fabric of equivalent thickness, if allowed by the local structural repair manual).

4. Same as original type, cell size, and density core material, or an approved substitute.

5 . Release film materials, both solid and perforated.

6. Breather and bleeder fabrics. 7. Vacuum bagging film, vacuum gage, and

8. Abrasive discs, hand-held pneumatic vacuum sealer tape.

motor.

9.Oil-free solvents and clean cheesecloth wipers.

10. Heat lamps and/or blankets, hot air gun. 11. Thermocouples and temperature monitor-

ing equipment. 12. Compressed air and vacuum source capa-

ble of being regulated. 13. Environmental conditions: Work should be

done indoors, under conditions of moderate temperatures (ambient) and low relative humidity (40-65%).

39.6.4 WET LAY-UPS USING PRECURED PATCHES

Instead of repairing damaged face sheets 'ply- for-ply', using dry fabrics and laminating resins, prepreg materials may be precured, between layers of peel ply fabric, under autoclave conditions, and stored for later use as patching materials. Precured patches should be perforated to facilitate resin flow and to provide vacuum contact. (Perforations should be of sufficient diameter and spacing to provide a vacuum path and resin bleed, without causing resin starvation at the bondline.) Perforated precured materials may be used with laminating resins, film adhesives, or adhesive paste. When using laminating resins, 3-5 wt% fumed silica ('CAB-O-SIL'TM, made by Cabot Corporation, is generally specified) or an equivalent thickener should be used to improve resin filleting on the honeycomb core.

Limitations of precured materials

Precured carbon/epoxy patches are normally applied only over flat surfaces. Precured glass/epoxy patches (or similar low modulus fiber) may be applied over mild curves. Since precured patches are basically 'scab' patches, they should not be used if the surface requires a high degree of aerodynamic cleanness. If precured patches are needed for repairing surfaces having compound shapes, special contour molds must be fabricated and used as a strongback to precure the material, so as to


produce a precise contour match. The precured patch may be regarded as the composite equivalent of a metal stamping.

Necessity of reducing vacuum pressure for bonding precured details and for co-curing prepregs and film adhesives

Available test data indicate that precured patches, as well as prepregs and film adhesives being co-cured (unless cured in an autoclave under positive pressure conditions) should be processed under 3340 kPa (10-12 in Hg) vacuum pressure only. Significant reductions in bond strength have been observed when such repairs are cured under heating blankets and at full vacuum pressure. Many combination repair techniques, utilizing both precured and resin-impregnated dry fabrics, prepregs and film adhesives have been developed for specific damage conditions and damage locations. All such repairs should be cured under reduced vacuum pressure.

Wet lay-ups cured at elevated temperatures

Unlike low temperature (65°C maximum) curing resins, laminating resins capable of being cured under vacuum pressure and up to 150°C (300"F), so-called 'room temperature set/elevated temperature post-cure' resins, produce high quality repairs and are therefore considered desirable alternatives to prepreg repairs. Because of the hazards inherent in all elevated temperature repairs, especially non-autoclave repairs performed under vacuum pressure only, a cautious approach is necessary. It is almost universally recommended to cure the repair under the lowest possible cure temperature at the expense of elapsed time. A given resin may be curable in two hours at 150°C (302°F) and may require six hours at 85°C (185°F). To reduce the risk of part failure during the cure, it is generally advisable to opt for the longer cure cycle at the lower temperature.

39.6.5 PREPREG REPAIRS

Autoclave repairs

Restoring damaged laminates by utilizing the same as original preimpregnated fabric or tape, at the same as original cure temperature and pressure, is normally recommended when full restoration of the original design properties is a requirement. In practical terms, however, full restoration should only be attempted by depot level facilities, since any such action necessitates:

1. removal of the affected part from its parent assembly;

2. availability of strongback tooling to maintain contours;

3. an autoclave capable of meeting the original cure parameters;

4. availability of same as original materials of construction and requisite facilities, equipment and NDT capabilities.

Clearly, only major operators have the necessary capabilities to conduct what can only be described as a remanufacturing operation. At the present time, only a limited number of the major airlines have the requisite equipment to perform rebuilds to OEM specifications. To satisfy market demands, a number of repair facilities have been granted remanufacturing authority under Part 145 of the Federal Aviation Regulations.

Non-autoclave repair methods

For limited damage requiring only partial restoration, there are approved alternative repair methods. Nearly all these utilize prepregs and film adhesives that are normally cured by means of heating blankets under 'vacuum pressure only' conditions. Such repairs can be carried out with minimal capital investments. Because such repairs yield lower than original design strengths, size limitations usually apply. These limitations are contingent upon the specific location of the damage as

Typical repairs 869

defined in the Structural Repair Manual for the aircraft in question. As a rule, repairs in the vicinity of a load path, as defined by finite ele-

part.) It is generally accepted that, before con- templating a prepreg repair, the following factors be given serious consideration.

ment analysis, are severely restricted. The allowable repairs in so-called 'field areas', i.e. at some predetermined distance away from spars, ribs, hinge and latch points, etc., are more generous in terms of size as well as repair method.

-

1. The part must be completely dry (see Section 39.5).

2. If at all possible, the part should be envelope-bagged to prevent backskin disbonding during the cure.

3. The cure should always be effected at the

Vacuum pressureheating blanket repairs (using prepregs and film adhesives)

In situ prepreg repairs are often preferred over wet lay-up/elevated temperature repairs because the resin content of the repair is more easily controlled by using a prepreg. One of the risks associated with the use of production prepregs and adhesive films is that these products were formulated for production and normally require high cure temperatures which, when applied to damaged parts likely to contain residual moisture, may cause severe disbonding of the remaining, thus far undamaged, structure. (The repair action thus

4.

lowest permissible temperature specified for the product. If at all possible, a repair prepreg and/or film adhesive should be selected that is curable at a temperature 40-60"C (104156°F) lower than the original cure temperature. This is of particular importance when repairing structures originally cured in the 170-180°C (338-356°F) temperature range. (Several such products are becoming available as a result of persistent industry demands. Representative products are listed in Table 39.1. This listing is given for reference only and does not imply endorsement of any given product.)

severely damages or effectively destroys the

Table 39.1 Repair adhesives and resins curable at reduced temperatures

Product

FM300-2 FM250 FM73

EA9680 EA9394 PL795 CYCOM 919 SP377 Epon9410 DER 329 Epocast35A/927

FM123-5

Class

Film Film Film Film Film Paste Film Resin Resin Resin Resin Resin

Manufacturer

Am Cy Am Cy Am Cy Am Cy Hysol Hysol BFG Am Cy 3M Shell Dow Furane

Min. cure temperature Max. use temperature

"C "F "C O F

120 250 112 230 105 225 95 200

120 250 95 200

120 250 112 235 95 200 80 175

Room temperature R.T. + postcure

150 80

120 120 150 150 177 70

105

300 180 250 250 300 300 350 160 220

See Note See Note See Note

Note: Postcure raises upper use temperature.


Voids and porosities in vacuum-pressure cured laminates and bondlines

Major disadvantages of ’vacuum-pressure- only’ cures are a reduction in the compaction of the laminate and the inevitable formation of porosities in the laminate and/or adhesive bondline. The finished repair yields, as a rule of thumb, approximately only 80% of the strength of an autoclave-cured part in terms of shear and flexural properties. The problem of compacting thick laminates may be overcome to some extent by hot debulking each ply, or a stack of several plies of a laminate, under vacuum pressure before the final cure. This method is labor-intensive but useful; it draws off entrapped gasses, improves resin flow, fiber wet-out and therefore overall laminate quality.

Prepregs co-cured with film adhesives

Repair technicians often use a compatible film adhesive together with a prepreg when making a repair. A layer of film adhesive is especially desirable as a bond ply over honeycomb core because it enhances the honeycomb peel strength by providing a deeper glue fillet than would be achieved with prepreg alone. There is, however, beside the added cost, a slight weight gain that must be considered when repairing a weight and/or balance critical part.

Prepregs applied over metal substrates

Prepregs applied over metallic substrates always require the use of a layer of film adhesive between the metal and the non-metal. The metallic substrate also requires the normal surface preparations applicable for metal bonding, by one of the methods described in Section 39.6.8. Prepregs are often used to provide debris protection in damage-prone areas of thin-skinned sandwich panels, notably wing flaps and other panelling in line with the landing gear. Occasionally, prepregs are used

as panel edge close-out in preference over metal stampings.

39.6.6 SURFACE PREPARATION FOR NON- METALLIC SUBSTRATES

Abrasion and cutting of plies

Taper-sanding is the preferred method of creating a scarf joint at the substrate/repair interface, especially if the substrate material is made from a woven fabric. Repairs in unidirectional tape laminates often use step joints, with each repair ply butted against the original ply. Instead of sanding, the splice joint is then prepared by cutting each ply carefully with a sharp instrument such as an ’Exacto’ knife. It is common to use a lap of 13-19 mm (0.5-0.75 in) per ply, although there is lack of agreement with respect to the optimum lap distance or the stacking sequence of fabric plies, i.e. whether the smallest or the largest ply should be placed first. Some authorities calculate the overlap as a function of materials thickness (e.g. L = 187‘) whereas others recommend a straightforward 13 mm (0.5 in) overlap per fabric ply and a 25 mm (1 in) overlap per tape ply in the zero degree orientation3, 11-13.

Use of peel plies

Multi-stage processes using precured laminates often use peel ply fabrics which, upon removal, yield a surface that requires no further cleaning or abrading. Chapter 29 contains some important observations about peel piles.

Grit blasting

Grit blasting followed by solvent wiping is sometimes used to prepare non-metallic substrates for subsequent bonding and laminating operations. Plastic media with a Mohs hardness of 3.0-3.5 (US Plastic and Chemical Corporation’s Polyextra and Polyplus granulated plastics, sieve size 30/40,

Typical repairs 871

propelled at a low incident angle (15-30") and at moderate nozzle pressure (25-30 psig) have been demonstrated to remove coatings effectively without damage to fibers, and to leave surface conditions of high quality5.

39.6.7 BASIC REPAIR JOINT PREPARATION

Whatever the specific surface preparation method, the focus must be on producing a smooth, contamination-free, activated bond surface capable of promoting adhesion and, after the cure, capable of transferring the structural loads across the joint with minimal disruption of the load path and minimal stress build-up. Stress risers of any kind, abrupt changes in thickness, brittle adhesives, the wrong scarf angle, poor detail fit-up, preloads, etc. should be avoided.

Cleaning, deoxidizing, anodizing, bonding primer application /cure

For optimum joint strength and bond durability, all metal surfaces that are to be adhesively joined require the following essential steps: (1) degreasing; (2) alkaline cleaning; (3) deoxidizing; (4) low voltage anodizing in chromic or phosphoric acid; (5) application, and (6) prebaking of a bonding primer. For other than complete rebuilds, which imply complete tear- down of the bonded elements, stripping of all adhesive residues, and full reprocessing of details through solution tanks, tank etching/anodizing and primer prebaking are often omitted at the expense of repair quality and longevity. Comparable values of various pre- bond surface treatments are shown in Fig. 39.4.

Non-tank anodizing

A process known as PANTA (Phosphoric Acid Non-Tank Anodizing) exists but requires extensive preparatory work, equipment and special skills, and has therefore not been fully accepted by the industry at the present time. Parts processed in this manner have been demonstrated to be almost equivalent to tank- processed parts in terms of bond strength and d~rability'~.

Surface preparation for in situ non-autoclave repairs

39.6.8 REPAIR OF METAL BONDMENTS

Honeycomb panels with metal face sheets

Because thin-skinned honeycomb sandwich panels are the most easily damaged, structures of this type are most often in need of repair. Several kinds of repair activity are considered typical by the industry:

1. minor repairs consisting of the application of cold or hot bonded metal patches;

2. partial skin and/or core replacement with or without the benefit of autoclave pressure;

3. rebuild or remanufacture (considered depot level repair).

Aluminum surface preparations

The quality of the repair is directly related to the quality of the surface to which the adhesive is applied. Poorly or inadequately prepared bonding surfaces are the primary reason why bonded repairs fail.

Typically, the repair patches or partial replacement skins are cleaned, acid etched, anodized, primed and prebaked. Structure not amenable to tank solution processing, i.e. the lap joint areas of the structure being repaired, is typically prepared with an acid paste, followed by a deionized water rinse, air drying, and spray application of a bonding primer without, however, the benefit of prebaking. Elevated temperature prebaking is generally impossible without exposing the structure to heat damage and is therefore


LOSS IN SHEAR STRENGTH OF 2024 -T3 SAMPLES BONDED WITH FM 123 - 5 ADHESIVE, AFTER 30 DAYS

AT 120OF AND CONDENSING HUMIDITY

UNEXPOSED I EXPOSED 1 TANK ETCH

+ PHOS. ANODRE + CIP

UNEXPOSED I , EXPOSED I TANK ETCH + CIP

OSFD I EXPOSED I TANKETCH

UNEXPOSED EXPOSED

I I SCOTCHBRITE. MEK + PASA - JEL

I I I I I I I

0 1000 2000 3000 4000 5000 8ooo

SHEAR STRENGTH, (PSb

Fig. 39.4 Effect of various aluminium surface treatments on repair bond strength and durability.

omitted, at some sacrifice in terms of bond strength and durabilityI5.

Abrasive cleaning of lap joints

For reasons of expediency, many repairs are effected under conditions considered marginal. One common practice is to abrade the joint area with aluminum oxide paper, followed by solvent wiping and the application of the adhesive. Repairs of t h s type, whether the adhesive selected be a paste or a film, are rarely of long duration and should be considered ’interim’ repairs only. On the other hand surface preparations using three-dimensional abrasives such as Scotchbrite@, a product of the 3M Company, in conjunction with high quality bonding

primers, have been demonstrated to produce joints of considerable durability and should be encouraged in preference over abrasion with aluminum oxide paper only.

Application of bonding pressure

Vacuum bagging and bondline thickness control

For non-autoclave repairs, the most common method of applying bonding pressure is by means of a vacuum bag. Film adhesives used for repair are normally scrim-supported and thus provide bondline thickness control. (A listing of representative film adhesives available with supporting scrims is provided in Table 39.2. This listing is for reference only and

Typical repairs 873

Table 39.2 Scrim-supported film adhesives

Manufacturer Product designation

Curing at 120°C (250°F) Am Cy FM73

FM123-2 FM123-5, FM137

Hysol EA9628 3M AF126-2, AF163-3

AF3109-2 Narmco a Metlbond 1113

Metlbond 1133 B.F. Goodrich Plastilock 7178

Cure temperature range Max. use temperature -

- "C

107-150 107-120 95-120 113-120 113-120 107-177 95-143 95-135 107-120

Curing at 177°C (350°F) Am Cy

Hysol

3M

Narmco a

B.F. Goodrich

FM61, FM150-2 FM96 FM300 FM400

FM350 EA9689 EA9649R AF191

FM300-1

AF131-2 AF143-2 AF147 Metlbond 328 Metlbond 329 Metlbond 1515 Plastilock 729-3

O F "C O F

225-300 225-250 200-250 235-250 235-250 225-350 200-290 200-275 225-250

120 250 120 250 120 250 95 200

120 250 See Note See Note See Note

82 180

163-177 160-177 163-177

150-177 163-1 77

171-182 177-182 175-180 175-1 80 175-180 175-180 175-180 163-19 1 135-185 163-177 171-182

325-350 320-350 325-350 325-350 300-350 340-360 350-360 345-355 345-355 345-355 345-355 345-355 325-375 350-365 325-350 340-360

150 300 177 350 177 350 204 400 150 300 177 350 177 350 177 350 177 350 204 400 177 350 150 300 150 300 204 400 150 300 177 350

Note: Use temperature increases as a function of cure temperature. a Now marketed by American Cynamid

not an endorsement for any given product.) When using paste adhesives, scrim cloth is normally inserted between the adherends to prevent adhesive squeeze-out and resin starvation in the bondline.

Bondline porosities resultingfiom vacuum pressure

The repair technician must be aware that not all film adhesives are equally suitable for bonding under vacuum pressure; indeed, most products are formulated for positive (i.e. autoclave) pressure applications. After curing under vacuum, some adhesives exhibit bond-

line porosities that inevitably result in lowered bond strength, which must be taken into account during the repair design. In an effort to overcome these negative effects, a unique bagging method called 'double-bagging' was developed a number of years ago. This method provides for an inner, 'low vacuum' bag 34 kPa (under 10 in Hg) for expelling volatiles, and an outer, 'high vacuum' bag 81-98 kPa (24-29 in Hg) to provide the equivalent of 4147KPa (12-14 psig) bonding pressure on the assembly. The intent of this method is to minimize the effect of full vacuum pressure on the resin during cure by isolating the laminate within a separate diaphragm.


Application of mechanical pressure

Mechanical pressure applications are sometimes used when both sides of the part are accessible for clamping. Anacoustical (sound suppression) panels having perforated or otherwise permeable skins make the application of vacuum impossible unless the panel is envelope-bagged. To do so generally requires extensive tear-down of the assembly and removal of the affected panel. Permeable face sheet materials (perforated metal, feltmetal, permeable glass fiber /polyimide laminates, or various fine mesh wire cloth acoustic sheet materials) may be successfully bonded under mechanical pressure, using liquid adhesives, pastes, or unsupported reticulating film adhesives.

Sand bags

Pressure application methods employing sand bags, shot bags, etc., exist but are cumbersome and yield repairs of marginal quality and questionable durability. Such repairs should always be rendered 'fail-safe' by the addition of mechanical fasteners to provide a secondary load path should an adhesive failure occur during subsequent flight service.

lnj7atable bladders

A novel pressure application method using an inflatable rubber bladder system has report- edly been successful16. The loads generated by the inflating bladder must be reacted out against one or several hard points on the aircraft structure, which requires equipment of model-specific design geometry.

Specific risks associated with pressure

Parts may be damaged during repair through improper pressure application. Most damages occur in the autoclave and are the result of poor fixturing and of insufficient attention being paid to proper vacuum bagging techniques.

Pressure damage affects primarily details made from light-weight honeycomb core, which is easily crushed if not adequately protected by well-anchored support blocks. Bondments are especially vulnerable to damage when the integrity of the vacuum bag is breached and compressed gasses enter the assembly. Well- designed fixtures, proper padding of potential puncture sites, bagging films of high quality, pressure levels appropriate for the materials and part configuration, as well as constant monitoring of the pressure cycle are imperative to prevent damaging parts during the cure. Table 39.3 shows the bonding pressures considered typical. It should be noted that for assemblies incorporating honeycomb (sandwich structures) the recommended pressures are predicated on the compressive strength of a core being simultaneously subjected to both autoclave pressure and elevated temperatures. Panel edges are occasionally collapsed during cure, if not properly supported against side loads. Bevelled edges should have angles between 15 and 20°, as shown in Fig. 39.5. Steeper angles require that the core be stabi- lized with additional resins or core fillers to prevent collapse under pressure.

Application of heat

How to introduce the proper amount of heat for curing the repair adhesive or resin has long been considered problematic. Unlike manufacturing processes, which can be optimized through cure cycle verification by destructive testing, most repairs are rather unique and influenced by a multitude of factors not easily controlled. Heating blankets of a constant watt density tend to overheat thin sections, e.g. the trailing edges of a bondment, while undercur- ing the bondlines located over a heat sink, e.g. a heavy metal fitting or a spar. It has often been found necessary to protect thin sections against overheating by inserting silicone rubber pads between the heating blanket and the part, thus reducing the effective cure temperature in selected areas, while allowing the

Typical repairs 875

Table 39.3 Recommended bonding pressures

Core material Thickness/cell size Density Max pressure cm in kg/m3 p.c& kPa Psig

Aluminum over 12.7 over 0.5 48 or higher 3 or higher 118 35 Honeycomb Sandwich and MetaYMetal Panels (Aluminum Bondments)

less than 3 84 25 Aluminum under 12.7 under 0.5 48 or higher 3 or higher 35 40

less than 3 101 30 No core n/a n/a 169-338 50-100

Laminates and Panels Containing Non-metallic Honeycomb Cores Aramid* 3.1 1/8 48-64 3.04.0 118 35

4.7 3/16 48 or lower 3.0 or lower 84 25 56-88 3.5-5.5 101 30

635 1 /4 all 84 25 9.5 3/8 all 84 25

No core n/a n/a 152 or 45 or higher higher

* Nomex HRH, HRP, Hh4X or similar core materials.

colder portions to reach the appropriate cure temperature. Occasionally, heat lamps or other auxiliary means must be employed in conjunction with heating blankets to provide additional heat inputs at critical locations to make sure the resins are fully cured. It is imperative that thermocouples be used at as many locations as necessary to monitor the cure cycle and to ensure the repair meets spec- ification requirements when completed.

Specific risks associated with heat

The principal risks associated with repair activities on structures that require the use of thermosetting resins and adhesives are:

1. Water or residual moisture in any portion of the assembly may vaporize and cause additional damage such as ply separation, core node bond separation, or skin-to-core bond failure (see Section 39.5).

2. Overheat conditions may develop under a heating blanket, causing irreversible damage, occasionally a fire. Constant monitoring or the incorporation of overheat

alarms may prevent part damage during the cure cycle.

3. Heat sinks may drain away heat energy required for resin cure, leaving residual uncured materials of unacceptable structural value. Hot bonding should not be carried out during adverse atmospheric conditions or while the aircraft is cold- soaked.

4. Improper heat-up rate control may cause resin flow and gel anomalies resulting in a product of marginal quality. Heat-up rates must be monitored or appropriate control devices used.

5. Foaming adhesives may generate exother- mic reactions resulting in irreversible damage. This hazard can be avoided by minimizing the width of splice gaps to be filled through careful sizing and fit-up of details prior to and during lay-up. Non- metallic core details should be joined by crush splicing rather than by adhesive foams to reduce the amount of reactive polymers present in the panel during the cure cycle.

876 Repair aspects of composite and adhesively bonded aircraff structures

5052

HRH

0' 100%

100%

a

30 ' 45 goo 15' 80% 76% 1 9% 2%

86% 70% 1 5% 2%

*O f \ 0 8'.

I 1

I I - - _ - - O0 1 5 O 30° 450 900

OFF-AXIS BARE COMPRESSIVE STRENGTH OFHONEYCOMB CORE

b

C

e Effect of core edge bevel on core stability under

bonding pressure

Fig. 39.5 (a) Bare compressive strength of honeycomb at various angles of loading; (b) Off-axis bare compressive strength of honeycomb core; (c) Effect of core edge bevel on core stability under bonding pressure.

Typical repairs 877

Lower cure temperatures enhance repair safety be the choice for repairs on aluminum, because ~"

Problems associated with hot bonding increase exponentially as a function of cure temperature. Cure temperatures in excess of 180°C (360°F) are several times more likely to result in a failed part than repairs performed at lower temperatures. Given the option, repairs should always be conducted at the lowest practical cure temperature, using a suitable adhesive or resin system. See Table 39.1 for typical products.

39.6.9 COMPOSITE REPAIRS APPLIED TO METAL STRUCTURES

Resin-impregnated fiberglass cloth repairs on aluminum

Wet lay-up, epoxy-impregnated fiberglass cloth repair patches have been approved repair methods via OEM Structural Repair Manuals and Military T.0.s for many years1. Utilizing room temperature curing resins in conjunction with fabrics, wet lay-up repairs can be applied over flat as well as curved surfaces with a minimum in equipment and under almost any conditions. Experience has shown, however, that in terms of overall quality and durability, these repairs are the least desirable and should be applied only when more advanced methods are unavailable. All resin-impregnated cloth repairs over metal require the use of primers (typically nitrile rubber based liquids) to pro- mote resin-to-metal adhesion.

Repairs utilizing advanced fibers

Repairs utilizing boron/epoxy and graphite/ epoxy prepregs over aluminum substrates have been under active consideration for use on military aircraft for some time17-21. Only recently, a large freight carrier made the decision to apply boron/epoxy patches on some of its large trans- port category aircraft, malung this the first time that boron is being used for this purpose on a commercial fleet in the USA. Boron appears to

of its galvanic compatibility and its favorable CTE with respect to aluminum. Graphite is inherently incompatible for the same reasons and may be suitable only for repairing titanium substrates. Composite repairs can be made by applying multiple layers of prepregged fabrics or unidirectional tapes, which may be cured by means of heating blankets under vacuum or mechanical pressure. Overlays of this type have been reported to enhance the fatigue life of conventional metal structures by several orders of magnitudez1. One major disadvantage is the need for a chemically prepared surface involving the use of acids which, if entrapped under the repair, could cause corrosion and premature structural failure, making periodic NDT of the repair mandatory for the remainder of the air- frame life. For maximum effectiveness, phosphoric acid non-tank anodizing (PANTA) and the use of bonding primers are essential. When using graphite as the backbone fiber, a barrier ply of fiberglass is necessary to prevent galvanic coupling between the repair material and the substrate. The risks involved have thus far inhibited the use of graphite on aluminum.

39.6.10 MECHANICALLY FASTENED REPAIRS

With increased use of composite materials in primary and principal structure not readily removable from the aircraft after a damage incident, bolted repair concepts are being vali- dated for major skin/stringer and skin/chord damage repairs. Utilizing mainly precured composite elements together with metal doublers and splice angles, such repairs can be effected where access is limited to one side of the structure only. Essentially, such repairs are a logical extension of, and quite similar to, conventional mechanically joined metal repairs, except that both metallic and precured composite elements are utilized. The Boeing B-777 is the first major program to approve this type of repair on its primary structures, chiefly its all-graphite composite empennage.


39.7 TECHNICIAN TRAINING AND SKILL REQUIREMENTS

Personnel engaged in designing and carrying out repairs to bonded and composite aircraft structure should be familiar with the funda- mental concepts listed in Table 39.4.

39.8 CONCLUSION AND SUMMARY

39.8.1 TECHNOLOGICAL EVOLUTION OUTPACING TRADITIONAL AIRCRAFT MAINTENANCE SKILLS AT ALL LEVELS

It has been observed that with every new gen- eration of commercial aircraft, there is an increase in the utilization of composite materials and a corresponding increase in the complexity of its design. The transition from simple hand layed-up bonded aluminum honeycomb sandwich and fiberglass-skinned Nomex panels to monolithic carbon fiber structures produced largely by means of auto- mated equipment has been a long, inexorable, and not altogether painless process from the operators’ point of view.

The end users, principally the world’s commercial airlines, are finding it increas- ingly difficult to keep up with the rapid technological changes thrust upon them by the manufacturers. Despite the large volume of technical literature available on the subject of composites, there is a dearth of practical information. One is tempted to say that the industry is encumbered by a surfeit of highly specialized data that is impenetrable to all but the experts. There is no denying that this research is both necessary and beneficial; there is, however, a dire need to make this data amenable to all through thoughtful dis- tillation. The worker in the field must know the practical effects of this research on the daily exercise of his craft. He must be kept abreast of technological advances and given a chance to upgrade his skills in order to meet ever-changing demands.

39.8.2 INSUFFICIENCY OF TECHNICAL TRAINING

Technical school curricula as well as regulatory guidance materials dealing with composite aircraft repairs are suffering from a technology lag that can, ultimately, only be bridged by greater emphasis on education and training at all levels. Community colleges and vocational schools should encourage the active participation of people experienced in the field (even though they may not possess the requisite academic credentials) and seek to enlist the help of subject matter experts.

39.8.3 DESIGN AIRCRAFT FOR MAINTAINABILITY

Design criteria focused on manufacturability without regard to maintainability may ultimately result in compromising flight safety, especially now that composites are finding increasing use in primary and principal aircraft structure. New regulations mandating ’damage-tolerant’ designs should be of great value in the determining the design criteria of future aircraft. Airline customer involvement in the design of new aircraft must go beyond payload, range, and other marketing concerns; the time has come for the designer to solicit the comments and suggestions of the maintenance engineer, the inspector and the mechanic.

39.8.4 NEED FOR STANDARDIZATION

Standardization of repair methods, practices and especially, repair materials is long over- due. The emphasis must be placed on the typical and generic, rather than the peculiar and proprietary. Cooperative efforts involving manufacturers, materials suppliers, airlines, repair facilities, regulatory agencies, profes- sional societies as well as academia will be needed to ensure the long-term viability of composites in aircraft structures.

Conclusion and summary 879

Table 39.4 Composite repair training topics

Components of Composite Materials Fibers and filaments Product definitions Fiber mechanical properties Finishes and sizings Specialty fabrics Bonding adhesives, resins, prepregs

Procedures of Fabrication and Processing

Glass and quartz; carbon and aramid; boron Fiber; roving; strand; yam; woven fabrics, unidirectional tape; milled fibers Density; strength; modulus; coefficient of thermal expansion

Chrome; silanes; plasma treatments; resin solutions Scrims; peel plies; bleeders and breathers; ceramics Epoxies; polyesters; phenolics; polyimides; bismaleimides; catalysts and hardeners; cyanate esters; acrylics; anaerobics; liquid adhesives; primers; coupling agents; film adhesives; prepreg fabrics and tape

Laminating

Other structures Curing methods

Vacuum bagging Adhesive bonding Sandwich panel construction Core materials

Surface preparation for metals Surface preparation for non-steps Joining and fastening

Machining of composites

The 'Laminate Code'; isotropic, anisotropic, quasi-isotropic laminates; cross-plied laminates; hybrid laminates; anacoustic laminates Filament/tape winding; RTM; braiding; pultrusions Autoclave; non-autoclave; single and multi-stage cures; postcuring; cure monitoring - flow/gel/set Bagging films; sealant tapes; breathers and bleeders; bagging techniques Pastes; liquids; films; cements; pressure application Face sheets, doublers and close-outs/pans; core; properties of sandwich construction: static strength and rigidity; adhesive filleting; shear ties Metal/non-metal honeycomb; cell sizes and shapes: hexagonal, overexpanded, flexcore; core density/weight; directional properties; compressive and shear strengths Cleaning, etching, anodizing, primer application; aluminum, steel, titanium, other metallic adherends Abrasion; grit blast; taper sanding lap joints; step joints; use of peel plies

Bonded joints; mechanical joints; fastener types and alloys; hole spacing; edge distances; hole sizes for composite joints Sawing; routing; drilling; sanding and grinding; water jet cutting; laser cutting

Protective Coatings and Sealants Organic polymers

Anti-static and lightning protection Heat and fire protection

Pinhole fillers; sanding sealers; primers; surfacers; enamels: epoxy, polyurethane; polysulfide coatings and sealants Carbon-filled enamels; flame spray coatings; EM1 shielding materials

Silicone coatings; ablative and intumescent coatings; heat-resistant enamels

Environmental Effects Moisture

Temperature

Effect of moisture on uncured resins Effect of moisture on cured systems Effect of temperature on uncured resins Effect of temperature on cured systems; glass transition temperature; heat deflection temperature

Corrosion prevention Effect of rain and particulates

Corrosion Galvanic corrosion: carbon/metal couples

Erosion Radiation Ultraviolet; thermal; nuclear Continued on next page


Table 39.4 (Continued)

Atmospheric Static charges; lightning strikes electricity Chemicals

Aging Standard Tests for Adhesives and Prepregs Metal adhesion Prepregs Cured laminates

Aircraft system fluids: oils, hydraulic, deicing Accidental exposures: paint stripper, solvent spills Fatigue and embrittlement effects on composites

Tensile shear; T-peel; honeycomb peel; crack extension (wedge) test Volatile and resin content; resin flow; gel time; tack Interlaminar shear; short beam shear; flexure; 45" in-plane shear; tensile and compressive strength and modulus; sandwich beam; hot/wet strength

Inspection and Quality Controls Non-instrumented Instrumented

Process quality controls

Post-repair NDI

Visual inspection; tap test; penetrants X-ray; moisture detector; pulse-echo ultrasonic; through-transmission ultrasonic; resonance ultrasonic; eddy current Raw materials handling and storage; environmental controls; processing materials controls; facilities and equipment controls; tool design and alteration; detail preparation; in-process sampling inspections and witness coupons Verification of compliance with specifications: cure cycle chart review; physical tests and checks

Damage Assessment, Failure Analysis, Preventive Maintenance In-service damage Failed repairs 'Lesson learned' Periodic inspections Coatings and sealants

Foreign object impact; environmental degradation

Electrical continuity check

Specific Repair Methods Repair categories Repair preparation Repair materials selection Surface preparations Metals; non-metals Adhesive/resin cure Applying heat and pressure; cure and postcure; cure monitoring Assembly completion Reassembly; finishes; weight and balance; final inspection

Construction type; original materials; original cure temperature Stripping; damage assessment and removal; decontamination Adherends and adhesives; auxiliary/ processing materials

Health and Safety Aspects of Composites Chemical exposure routes Hazard levels Material safety data Basic industrial hygiene Toxicology of composite materials coatings

Dermal; ocular; inhalation; ingestion

Acute vs. chronic toxicity The MSDS and how to interpret Engineered control systems; personnel protection

Resins and catalysts; solvent and diluents; fibers and fiber dust; sealants and

References 881

REFERENCES

1. Anon., Advanced composite repair guide, Contract No. F33615-79-3217, Air Force Wright Aeronautical Laboratories, Wright-Patterson AFB, OH 45433,1982.

2. Anon., Environmental Durability of Speed Tape. Summary Report by Boeing Materials Technology, Renton, WA, 1982.

3. Anon., Guidance Material for Design, Maintenance, Inspection and Repair of Thermosetting Epoxy Matrix Composite Aircraft Structures. Montreal/Geneva: IATA, Doc. Gen/3043 1991.

4. McKague, Lee, et al., Test of graphite-fiber sizing effects upon laminate properties. SAMPE I., Nov/Dec, 1979.

5. Egan, William, Composite Paint Stripping Development. Manufacturing Development Report No. 6-35081. Boeing Commercial Airplane Company, Seattle, WA 92124,1985.

6. Hertz, Julius, Moisture effects on the high temperature strength of fiber-reinforced resin composites. Convair Aerospace Division of General Dynamics/Hercules Inc. Joint Study

7. Dexter, H. Benson and Donald J. Baker, Flight Service Environmental Effects on Composite Materials and Structures. NASA Langley Research Center, Hampton, VA 23556. 73rd AGARD Structures and Materials Panel Workhop, San Diego, CA October 7-8,1991.

8. Crossman, F.W. and Flaggs, D.L., Dimensional stability of composite laminates during environmental exposure. Lockheed Palo Alto Research Laboratory, Palo Alto, CA. SAMPE J. July/August, 1979.

9. Anon., Structural Repair Manuals (All Models), Boeing Commercial Airplane Company, Seattle WA, 98124.

10. Seidl, A.L., Inspection of composite structure. Report prepared for ATA/IATA/SAE Commercial Aircraft Composite Repair Committee (CACRC), Washington, D.C. Meeting, December 3-5,1991.

11. Anon., Bonded Component Repair Manual (BCRM), Boeing Commercial Airplane Company, Seattle, WA 98124. Document D6- 51169,1983.

12. Anon., Repair Procedures for 250/350 deg. F Cured Aramid Fabric/Epoxy and Aramid/Graphite Fabric/Epoxy Hybrid Composite Structures. Boeing Commercial Airplane Company, Seattle, WA 98124. Document D6-48908,1982.

NAS 8-27435’1972.

13. Kuperman, M.H., Graphite/Epoxy Repair Program Test Results, Internal Report for United Airlines, San Francisco, CA 94128,1983.

14. Locke, Melvin C., Non-Tank Phosphoric Acid Anodize Method of Surface Preparation of Aluminum for Repair Bonding. Report prepared for inclusion in Adhesive Bonded Aerospace Structures Standardized Repair Handbook, Technical Report AFML-TR-77-206, Air Force Wright Aeronautical Laboratories, Wright- Patterson AFB, OH 45433,1978.

15. Kuperman, M.H., Bond Strength and Bond Durability Study for the Evaluation of Various Surface Treatments for Aluminum Adherends. Internal Report for United Airlines, San Francisco, CA 94128, 1975.

16. Molent, L. et al., Design of an All-Boron/Epoxy Doubler Reinforcement for the F-111C Wing Pivot Fitting: Structural Aspects. In Composite Structures Oxford: Elsevier Science Publishers, 1989.

17. Anon., Adhesive Bonded Aerospace Structures Standardized Repair Handbook Air Force Materials Laboratory (AFSC) , Wright-Patterson AFB, OH 45433. Technical Report AFML-TR-77- 206,1978.

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19. Baker, A.A., Boron Fibre-Reinforced Plastic Patching for Cracked Aircraft Structures. Lecture delivered to the Melbourne, Australia, Branch of the Royal Aeronautical Society. Aircraft, September 1981.

20. Sandow, Forrest A. and Raymond K. Cannon, Composite Repair of Cracked Aluminum Alloy Aircraft Structure. Final Report AD-A190-514, Flight Dynamics Laboratory, Wright-Patterson AFB, OH 45433, September 1987.

21. Kelly, Larry G., Composite Repair of Cracked Aluminum Structure (Fatigue Life Extension Study). Air Force Wright Aeronautical Laboratories, Wright-Patterson AFB, OH 45433, Undated Report

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23. Fisher, Franklin, and Plumer, J . Anderson, Lightning Protection of Aircraft. NASA Reference Publication 1008, National Aeronautics and Space Administration, Washington, DC 20456, 1977.

882 Repair aspecfs of composite and adhesively bonded aircraft structures

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25. Anon., DoD/NASA Structural Composites Fabrication Guide, prepared under Contract No. F33615-79-C-5125 by Lockheed-Georgia Company for Air Force Wright Aeronautical Laboratories, Wright-Patterson AFB, OH 45433, 1982.

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