1 Curing of composite materials
Curing cycles are cycles of time, temperature, and pressure used to cure thermoset resin systems or prepregs. The curing of the repair is as important as the curing of the original part material. Unlike metal repairs where the material is prefabricated, composite repairs require technicians to fabricate the material. This includes all storage, processing, and quality control functions. The restoration cycle for aircraft maintenance begins with material storage. Improperly stored materials will begin to cure before they can be used for restoration. All times, temperatures, and requirements must be met and documented. Refer to the Aircraft Structural Repair Manual to determine the correct repair cycle for the part to be repaired.
1.1 Room Temperature Curing
Room temperature curing is most beneficial in terms of energy savings and portability. Room temperature cured wet lay-up repairs do not restore the strength or durability of the original 250℉ (121℃) or 350℉ (176.67℃) cured component and are typically used for wet fibreglass lay-up repairs on non-critical components. Room temperature cured repairs can be accelerated by heating. Maximum performance is achieved at 150℉ (65.56℃). A vacuum bag can be used to consolidate the laminate and provide a path for air and volatiles to escape.
1.2 High temperature curing
All prepregs are cured using high temperature cycles. Some wet ply repairs use elevated curing cycles to increase the strength of the repair and speed up the curing process. Curing ovens and hot bonders use vacuum bags to consolidate the laminate and provide a path for air and volatiles to escape. Autoclaves use vacuum and positive pressure to consolidate the laminate and provide a path for air and volatiles to escape. Most heaters use programmable computer controls to run the cure cycle. The operator can select from a menu of available cure cycles or write his or her own programme.
Thermocouples are placed near the repair and they provide temperature feedback to the heating unit. Typical cure temperatures for composites are 250℉ (121℃) or 350℉ (176.67℃) The temperature of a large part cured in an oven or autoclave may be different from the temperature of the oven or autoclave during the cure cycle because they act like a heat sink. The temperature of the part is most important for proper curing, so thermocouples are placed on the part to monitor and control the temperature of the part. An oven or autoclave air temperature probe used to measure oven or autoclave temperature is not always a reliable device for determining part cure temperature. If the part or tool is acting as a heat sink, the oven temperature and the part temperature can be very different.
The high temperature curing cycle consists of at least three parts.
-Warming up: the heating unit warms up at a set temperature, usually between 3℉ (-16.1℃) and 5℉ (-15℃) per minute.
-Holding: the heating unit maintains temperature for a predetermined period of time.
-Cooling: the heating unit cools at a set temperature. The cooling temperature is usually below 5℉ per minute. When the heating unit falls below 125℉, the part can be disassembled. When curing parts with an autoclave, be sure to relieve the pressure in the autoclave before opening the door. As shown in Figure 53.
Figure 53: Autoclave curing process
The curing process is accomplished by applying heat and pressure to the laminate. As the temperature increases, the resin begins to soften and flow. At lower temperatures, very little reaction occurs. Any volatile contaminants, such as air or water, are extracted from the laminate with a vacuum during this time. The laminate is compacted by applying pressure, usually a vacuum (atmospheric pressure); the autoclave applies additional pressure, usually 50-100 psi As the temperature approaches the final cure temperature, the rate of reaction increases considerably and the resin begins to gel-harden. Holding at final cure allows the resin to finish curing and obtain the desired structural properties.
2 Repair of composite honeycomb sandwich
A large proportion of current aerospace composite components are damage-prone lightweight sandwich structures. Since sandwich structures are bonded structures with thin panels, damage to sandwich structures is usually repaired by bonding.
Repair of sandwich honeycomb structures uses similar techniques to the most common panel materials such as glass fibre, carbon fibre and Kevlar®. Kevlar is usually repaired with glass fibre. As shown in Figure 54.
Figure 54: Typical repair of honeycomb sandwich structures
2.1 Classification of damage
Short-term repairs can meet strength requirements but are limited by time or flight cycles. At the end of the life of the repair, the repair must be removed and replaced. Temporary repairs may restore the required strength of the component. However, this repair will not restore the required durability of the component. Therefore, it has different inspection intervals or methods. A permanent repair is a repair that restores the required strength and durability of the component. The repair has the same inspection methods and intervals as the original component.
2.2 Microdamage to sandwich structure stack cores (packing and potting repair)
Encapsulation repair can be used to repair damage to sandwich honeycomb structures less than 0.5 inches. The honeycomb material can be left in place or removed and filled with a potting compound to restore some strength. Encapsulated repairs do not restore the full strength of the part.
The potting compound is usually an epoxy filled insulating glass, phenolic or plastic microspheres, cotton, a mixer or other material. The potting composite can also be used as a filler for decorative repair edges and skin panels. The potting compound is also used in laminated honeycomb panels as a hard point for bolts and screws. The potting composite is heavier than the original core, which may affect flight control balance. The weight of the repair must be calculated and compared to the flight control weight and balance limits specified by the SRM.
2.3 Damage to one or both side panels requiring replacement and repair
NOTE: The following steps are for reference only and should not be considered directly applicable to all repair methods.
Step 1: Inspect for damage
Thin laminates can be visually inspected and percussion tested to determine damage. Thicker laminates require more in-depth NDI methods such as ultrasonic inspection as shown in Figure 55. Check for water, oil, fuel, dirt, or other foreign matter entering in the vicinity of the damage. Water can be detected with an x-ray, backlight, or moisture detector.
Figure 55: Knockout test technique
Step 2: Remove Water from Damaged Area
Water needs to be removed from the honeycomb core before the part can be repaired. As shown in Figure 56, if the water is not removed, it will boil during the high temperature cure cycle and the panel will expand the core, causing more damage. Water in the honeycomb core can also freeze at low temperatures at high altitudes, which can cause panels to peel off.
Figure 56: Vacuum bagging method for drying parts
Step 3: Remove Damage
Trim the damaged area on the part to a smooth shape with rounded corners, or a round or oval shape. Do not damage undamaged layers, cores, or surrounding material. If the core is also damaged, trim the core to the same contour as the core skin. As shown in Figure 57.
Figure 57: Removing Core Damage
Step 4: Presetting the Damaged Area
Use a soft disc sander or rotary pad sander to sand around the cleaned damage with an even taper. Some manufacturers give a taper ratio, such as 1:40, while others specify a taper distance, such as a 1-inch overlap of the existing taper distance for each layer. Remove the exterior finish, including the conductive coating in an area at least 1 inch larger than the taper border. Remove all sand and dust with dry compressed air and a hoover. Clean the damaged area with a clean cloth moistened with permitted solvent. As shown in Figure 58.
Figure 58: Sanding the repair area
Step 5: Install Honeycomb Core (Wet Layer)
Cut the replacement core with a knife. The core plugs must be of the same type, grade, and quality as the original core. The core cells should be orientated in the same direction as the honeycomb of the surrounding material. Plugs must be trimmed to the proper length and solvent cleaned with an approved cleaner.
For wet lay-up repairs, cut two layers of woven fabric suitable for the inner surface of the undamaged skin. Impregnate the fabric layers with resin and place in the hole. Use infusion compound around the core and place it in the hole. For prepreg repairs, cut a piece of adhesive film to fit the hole and use foam adhesive around the blockage. The plug should touch the sides of the hole. Align the core core of the blockage with the original primed material. Repair the area with a vacuum bag and cure the replacement core using an oven, autoclave or hot blanket. Wet ply repairs may be cured at room temperature to 150℉ (65.56℃). Prepreg repairs must be cured at 250℉ (121℃) or 350℉ (176.67℃). Typically, replacement cores are cured in a separate cycle rather than co-cured with the patch. After curing, the blockage must be sanded flush with the surrounding area. This is shown in Figure 59.
Figure 59: Core Replacement
Step 6: Prepare and Install Repair Layers
Refer to the repair manual for the correct repair material and the number of layers required for the repair. Typically, install one more layer than originally installed. Cut the layer thickness to the correct size and orientation. The repair layer thickness must be installed in the same direction as the original layer being repaired. Impregnate the layer with resin for wet lay-up repairs or remove the backing material from the prepreg. The layers are usually laid using a minimum layup first taper layup sequence. As shown in Figure 60.
Figure 60: Patch Installation
Step 7: Vacuum Bagging the Repair
Once the paving material is in place, use vacuum bag installation to remove air and pressurise the repair to cure. See Figure 61 for vacuum bag installation instructions.
Figure 61: Vacuum Treatment
Step 8: Curing the repair
Repair during the desired repair cycle. Wet pavement repairs can be cured at room temperature. The temperature may be raised to 150℉ (65.56℃) to accelerate curing. Prepreg repairs need to be cured during the elevated cure cycle. Components removed from the aircraft to be repaired can be cured in a hot room, oven, or autoclave as shown in Figure 62. Heated blankets are used for repairs on aircraft.
After curing remove the encapsulated material and inspect the repair. The repair should be free of pits, blisters, and areas of resin gain or loss. Lightly sand the repair patch with sandpaper to give a smooth surface that will not damage the fibres. Apply surface treatment and conductive coating (light wave resistant).
Figure 62: Cured Repair
Step 9: Post-Repair Inspection
A visual, tap or ultrasonic inspection is usually used to do an inspection of the thickness of the repair. If defects are found, the repaired patch is removed. This is shown in Figure 63.
Figure 63: Check after repair
If the flight control panel was repaired, perform a balance check and ensure that the repaired flight control panel is within the range of the SRM. Failure to do so may result in flight control chatter and compromise flight safety.
To be continued
Source "Composites Frontier" Public Website