Chapter 3 Crack Repair Techniques
3.1 Basic Principles of Repair
There are some basic principles regarding the patch of repairing aircraft structure with the hope of restoring the original strength and properties.
3.1.1 Shape of Patch
It is better to use a circular patch instead of rectangular because circular patches reduce the risk of getting any further cracks that could form at the corners as what rectangular shape would have. If the patch that must be used is a rectangular shape, then the corners have to be shaped like a curve to prevent a sharp corner.
3.1.2 Material of Patch
The patches must also be the same or similar material as the original material of the aircraft structure. According to Army Institute for Professional Development (1994), if we cannot use a similar material, we can use material of a gage heavy enough to give an equivalent cross-sectional strength because it is not safe if we use a lighter gage but stronger material for the patch. The danger that the Army Institute for Professional Development state is that one material can have greater tensile strength than another but less compressive strength.
3.1.3 Size of Rivet
Next is rivet for joining the patch. A way of determining the size the rivets is to use the rivets in the next parallel row inboard on the wing or forward on the fuselage according to Army Institute for Professional Development (1994). One more method to determine the rivets’ size is to multiply skin thickness by three and use the closest larger size rivet based on the result. For example, if the skin thickness is 0.041 inch, multiply 0.041 by three, the result then is 0.123, based on this result, use the next larger size rivet which is 4/32 inch (0.125 inch).
3.1.4 Original Contour
All repairs are done with a reminder that it must be done in a way that contour is as similar as possible with the original contour. We have to be more careful with the patches that are on the external skin of the structure because it changes the external contour and the change with the air flow of the aircraft. These changes, if they are not done carefully, could lead to more damages that can occur to the aircraft.
3.1.5 Minimum Weight
Because an aircraft needs to be as light as possible, the repairs must not have a huge weight and it is better if the weight is the most minimum weight it can be. This is achieved by making the patch as small as it can be and also limit the amount of rivets that are gonna be used. Sometimes the aircraft structure’s balance is disturbed due to the repairs. It is because the adding of the weight when it gets repaired.
3.2 Bolted Repair
Bolted repair or mechanically fastened repair is the type of repair which used mechanical joints to repair cracks on aircraft structure. This technique commonly has two main elements which are the fastener and the plate. The simple explanation of this repair technique is making a circle hole around the cracked surface and remove the surface along the with the crack, drilling small holes of the surface of the structure around the hole, place a plate that are the same material and thickness of the surface skin and already has similar holes as with the surface of aircraft structure on top of the surface that has hole, fasten the plate with the aircraft surface using bolt or rivet.
Bolted repair has stress concentrations because of the drilling and machining of the holes and is more damaging to the material of the aircraft structure than bonded repair. Ratwani (2000). Bolts cause localized stress from clamping pressure and the patch itself causes stress due to differential expansion and contraction. The differential expansion of the patch and the base material pushes or pulls against the bolts which focuses all the stress into the bolt holes.
As the author have mentioned before in the literature review, there are three concepts of bolted repair according to Ratwani (2000) which are :External Patch with Backup Plate
This concept uses an external chamfered metal patch bolted to the panel being. The bolts thread into nut plates mounted on metal backup plates that are on the side of the repaired panel. The backup plate can be split into two or more pieces and slipped through the opening.
External Patch with Blind Fasteners
This concept is similar to the previous one, except that the backup plates are not used. Blind fasteners are not as strong as bolts and nut plates, but if acceptable strength can be restored, this concept is easier to use.
Bolted Internal DoublerThis concept has been used as a standard repair for metal structures. Access to the backside is required to install the doubler. The doubler cannot be installed through the hole as a separate piece because the doubler has to be continuous to carry loads in all directions. Filler is used to provide a flush outer surface, and is not designed to carry loads.
For more understanding Ratwani provide Figure 2 below.
Figure 2. Bolted Repair concepts
3.3 Bonded Repair
Bonded repair techniques is a more modern technique compared to mechanically fastened repair and is also preferred than mechanically fastened repair for repairing the damaged metallic structures. This is because according to Ratwani (2000) mechanically fastened repair require drilling of holes for additional fasteners that affect the structural integrity of the structure. Bonded repairs have no stress concentrations that are caused by holes, which is happening for the bolted repair. Bonded repair are less damaging to the material of the aircraft structure because bonded repair technique doesn’t need drilling or machining holes. Bonded repair are also more aerodynamically and structurally efficient than bolted repair. Due to the development of repair technology, bonded repair technique is now possible to repair damaged structures which previously could not be repaired with the traditionally mechanically fastened repair. Bonded repair or adhesive bonding offer more life service over the mechanically fastened repair which is also another reason why bonded repair is more desirable.
Baker 1987 This technique is using a patch that is bonded to the original metallic structure to strengthen the cracked zone and to restore the structure to its original design properties. According to Ratwani (2000), bonded composite repair has many advantages over conventional mechanically fastened repair, namely:
More efficient load transfer from a cracked part to the composite patch due to the load transfer through the entire bonded area instead of discrete points as in the case of mechanically fastened repairs,
No additional stress concentrations and crack initiation sites due to drilling of holes as in the case of mechanically fastened repairs,
High durability under cyclic loading,
High directional stiffness in loading direction resulting in thinner patches, and
Curved surfaces and complex geometries easily repairable by curing patches in place or prestaging patches. The crosssection of a typical 16-ply graphite/epoxy patch bonded to an aluminum sheet is shown in Figure 3.
Figure 3. Cross-section of a Typical Composite Repair Patch
According to Ratwani (2000), the steps of progress for this type of repair are:
Adhesive material selection,
Composite repair material selection, and
3.3.1 Surface Preparation
Surface preparation holds an important consideration in bonded structures. The process to prepare the surface consists of paint removal, anodizing and priming. Liquid chemical paint strippers are not recommended, as they may become entrapped in cracked areas and faying surfaces of adjoining structures, thereby causing a corrosion problem.
Aluminum oxide abrasive cloth has been found to be suitable for small repair areas. Both silane and phosphoric acid non-tank anodize (PANTA) are also suitable. The silane process has the advantage over PANTA for being non-acid process. PANTA process might be desirable when look at the long term durability of repairs, as sufficient test data is available on this process.
To prevent contamination and improve long-term durability, primer is applied to the aluminum surface after anodizing with PANTA. BR-127 primer has been found to be suitable for FM-73 adhesive.
3.3.2 Adhesive Material Selection
The service temperature requirements is 180F (82C) in the majority of aircraft repair applications which make room temperature cure adhesives are not considered suitable. Room temperature cure adhesives are paste adhesives and generally do not result in uniform bond line thickness in the repair. Thus, affecting the load transfer to composite patch. Hence, high temperature film adhesives are preferred. Also, long term durability of room temperature adhesives is not well characterized. A 350F (177C) cure film adhesive is not considered desirable, as the curing at such a high temperature is likely to cause undesirable high thermal stresses. Also, an aluminum structure exposed to a 350F (177C) temperature will undergo degradation in mechanical properties. A 250F (121C) cure adhesive system is considered suitable for the composite patch repair of aluminum structure. Ductile adhesives such as FM-73 are preferred over brittle adhesives such as FM-400 due to the tendency of the brittle adhesives to disbond around the damage area, thereby reducing the load transfer to the repair patch.
3.3.3 Composite Repair Material Selection
Both boron/epoxy and graphite/epoxy composites are suitable for the repairs. The choice between boron or graphite fibers should be based on availability, handling, processing and the thickness of the material to be repaired. Boron has higher modulus than graphite and would result in thin repair patches. Thin patches are more efficient in taking load from damaged parts as compared to thick patches. For repairing relatively thick parts, boron may be preferred over graphite. It is considered desirable to use highly orthotropic patches, having high stiffness in the direction normal to the crack, but with some fibers in directions at 45 and 90 degrees to the primary direction to prevent matrix cracking under biaxial loading and inplane shear loads which exist for typical applications. This patch configuration can be best obtained with unidirectional tape. Woven material has greater formability and could also be used, although it would not make a very efficient patch.
The composite patches may be precured, prestaged or cured in place. For locations where vacuum bagging represents a problem, a precured patch may be prepared in an autoclave and then secondary bonded to the repair area. For relatively minor contours, a prestaged patch may be used. For curved surfaces the patch may be cured in place during the bonding operation.
3.3.4 Bonding Operation
Bonding of repair patches requires a proper temperature control within + 10F and -5F in the repair area. Thermal blankets are available to provide temperature in excess of 1000F (538C). A proper temperature control within tolerances is necessary for bondline to achieve desirable strength. A large aircraft structure compared to a small repair area may act as a heat sink and jeopardize maintaining desired temperature control for the required duration. Proper heat blankets for surrounding areas may be required for such cases.