**2.2 Advanced Composites Technology (ACT) program**

The objective of the ACT fuselage program was to develop composite primary structure for commercial airplanes with 20–25% less cost and 30–50% less weight than equivalent metallic structure [3]. The Advanced Technology Composite Aircraft Structure (ATCAS) program was performed by Boeing as the prime

**Figure 9.** *F/A-18E/F fuselage structure.*

contractor under the umbrella of NASA's ACT program and focused on fuselage structures. A large team of industry and university partners also supported the program. The primary objective of the ATCAS program was to develop and demonstrate an integrated technology that enables the cost and weight effective use of composite materials in fuselage structures for future aircraft.

The area selected for study was identified as Section 46 on Boeing wide body aircraft (**Figure 10**). This section contains many of the structural details and manufacturing challenges found throughout the fuselage. This includes variations in design details to address high loads at the forward end and lower portions of the fuselage. The loads decrease toward the aft end and the upper portion of the fuselage, allowing for transitions in the thickness of the structure that are tailored to match the structural loading.

A quadrant panel approach was selected for study as shown in **Figure 11**. The cross section is split into four segments, a crown, keel, and left and right side panels. The circumferential, four quadrant panel approach was selected with the idea of reducing assembly costs by reducing the number of longitudinal splices. This builtup assembly approach is baseline to metallic aircraft manufacturing and is similar to the approach Airbus selected for most of the fuselage of the A350.

Manufacturing process development and design trade studies contributed to the development of Cost Optimization Software for Transport Aircraft Design Evaluation (COSTADE) which allowed for defining and evaluating the costeffectiveness and producibility of various designs. Included in the program were assessments of tooling, materials and process controls needed for future full-barrel fabrication like Boeing selected for the 787.

The structural concepts studied included stiffened skin structures achieved by stand alone or combinations of cocuring, cobonding, bonding, and mechanical attachment of stringers and frames to monolithic or sandwich panel skins (**Table 1**). The crown section study selected fiber placed skins laminated on an IML controlled layup mandrel with the skin subsequently cut into individual panels and transferred to OML cure tools. Hat stiffeners used solid silicone mandrels located longitudinally along the IML of the skin panels for cocuring.

The recommended optimized panel design included cobonding of cured frame elements while cocuring the hat stiffeners and the skin. The cured frames were demonstrated using braided textile preforms and resin transfer molding (RTM). One of the main challenges of the crown panel concept was the bond integrity between the precured frames cobonded to a skin panel that is stiffened with cocured hat stringers. Alternative concepts the team considered during the review process included mechanically attached Z-section frames instead of cobonded J's.

**103**

**Table 1.**

*ACT structural concepts [3].*

**Figure 11.**

*ACT quadrant panels [3].*

**Details Process**

Skins AFP (tow, hybrid AS4/S2)

Frames Braiding/resin transfer molding (triaxial 2-D braid)

Panel assembly Cocured/cobonded stringers, cobonded frames

Compression molding

Pultrusion/pull forming

Stringers Hat—ATLM/drape forming (cocured, thickness variation) "J"—pultrusion

*The Evolution of the Composite Fuselage: A Manufacturing Perspective*

The mechanically fastened frame approach greatly reduces the complexity of IML tooling needed to cocure the hat stiffeners and cobond the frames. This is especially true at the intersections of the frame and hat. Flexible caul plates and custom fit reusable bags became part of the tooling system needed to accomplish the fully integrated skin/stringer/frame structure. Producibility issues are complicated by the blind nature of the IML of the skin being completely covered by flexible cauls and the reusable bagging system. The structural arrangement shown in **Figure 12** is very similar to the configurations that ended up on both the 787 and A350 programs. The program studied the pultrusion process for producing skin stringers. Continuous resin transfer molding (CRTM) developed by Ciba-Geigy was one of the more promising technologies studied. Improved process control and reduced waste are among the perceived advantages; process maturity, constant cross-section stringers and costs associated with secondary bonding or cobonding are among the disadvantages.

CTLM (contoured tape lamination machine, 12″ tape)

Stretch forming (thermoplastic, discontinuous fibers)

Cocured/cobonded stringers, fastened frames Sandwich panels, cobonded frames

*DOI: http://dx.doi.org/10.5772/intechopen.82353*

**Figure 10.** *ACT fuselage section [3].*

*The Evolution of the Composite Fuselage: A Manufacturing Perspective DOI: http://dx.doi.org/10.5772/intechopen.82353*

**Figure 11.** *ACT quadrant panels [3].*

*Aerospace Engineering*

to match the structural loading.

fabrication like Boeing selected for the 787.

along the IML of the skin panels for cocuring.

contractor under the umbrella of NASA's ACT program and focused on fuselage structures. A large team of industry and university partners also supported the program. The primary objective of the ATCAS program was to develop and demonstrate an integrated technology that enables the cost and weight effective use of

The area selected for study was identified as Section 46 on Boeing wide body aircraft (**Figure 10**). This section contains many of the structural details and manufacturing challenges found throughout the fuselage. This includes variations in design details to address high loads at the forward end and lower portions of the fuselage. The loads decrease toward the aft end and the upper portion of the fuselage, allowing for transitions in the thickness of the structure that are tailored

A quadrant panel approach was selected for study as shown in **Figure 11**. The cross section is split into four segments, a crown, keel, and left and right side panels. The circumferential, four quadrant panel approach was selected with the idea of reducing assembly costs by reducing the number of longitudinal splices. This builtup assembly approach is baseline to metallic aircraft manufacturing and is similar to

Manufacturing process development and design trade studies contributed to the development of Cost Optimization Software for Transport Aircraft Design Evaluation (COSTADE) which allowed for defining and evaluating the costeffectiveness and producibility of various designs. Included in the program were assessments of tooling, materials and process controls needed for future full-barrel

The structural concepts studied included stiffened skin structures achieved by stand alone or combinations of cocuring, cobonding, bonding, and mechanical attachment of stringers and frames to monolithic or sandwich panel skins (**Table 1**). The crown section study selected fiber placed skins laminated on an IML controlled layup mandrel with the skin subsequently cut into individual panels and transferred to OML cure tools. Hat stiffeners used solid silicone mandrels located longitudinally

The recommended optimized panel design included cobonding of cured frame elements while cocuring the hat stiffeners and the skin. The cured frames were demonstrated using braided textile preforms and resin transfer molding (RTM). One of the main challenges of the crown panel concept was the bond integrity between the precured frames cobonded to a skin panel that is stiffened with cocured hat stringers. Alternative concepts the team considered during the review process included mechanically attached Z-section frames instead of cobonded J's.

composite materials in fuselage structures for future aircraft.

the approach Airbus selected for most of the fuselage of the A350.

**102**

**Figure 10.**

*ACT fuselage section [3].*

The mechanically fastened frame approach greatly reduces the complexity of IML tooling needed to cocure the hat stiffeners and cobond the frames. This is especially true at the intersections of the frame and hat. Flexible caul plates and custom fit reusable bags became part of the tooling system needed to accomplish the fully integrated skin/stringer/frame structure. Producibility issues are complicated by the blind nature of the IML of the skin being completely covered by flexible cauls and the reusable bagging system. The structural arrangement shown in **Figure 12** is very similar to the configurations that ended up on both the 787 and A350 programs.

The program studied the pultrusion process for producing skin stringers. Continuous resin transfer molding (CRTM) developed by Ciba-Geigy was one of the more promising technologies studied. Improved process control and reduced waste are among the perceived advantages; process maturity, constant cross-section stringers and costs associated with secondary bonding or cobonding are among the disadvantages.


**Table 1.** *ACT structural concepts [3].*

**Figure 12.** *ACT crown panel structural arrangement [3].*

Airbus has studied automating stringer fabrication using both pultrusion and RTM but felt limited by aspects of both processes. As an answer, Airbus developed their version of pultrusion RTM. **Figure 13** shows equipment completed in 2011 that is being used to develop and qualify the process [4]. This hybrid fabrication approach allows the use of preform laminates instead of being limited to unidirectional reinforcements like traditional pultrusion and supports continuous production instead of batch processing associated with the traditional RTM. Instead of dipping the preform stack through a resin bath, it is pulled into an RTM tool that is open on both ends. To overcome resin being pushed out at both ends of the open tool, Airbus worked with resin suppliers to develop an epoxy resin with a parabolic temperature/viscosity curve. At 120°C resin viscosity is very low with high flow characteristics, but at both room temperature and at 180°C and higher, it is very viscous. The tool entry is cooled so the resin is too viscous to flow out; the middle is heated to obtain resin flow and cure; more heat is added at the end to increase resin viscosity to make sure it does not flow out and reduce cure pressure.

**105**

*The Evolution of the Composite Fuselage: A Manufacturing Perspective*

Even in the early days of development, industry leaders believed in the possibility of higher layup rates using AFP than was possible with hand layup, but the capabilities and the scale that the industry has achieved today is astounding. Almost as astounding as how the industry reinvented itself from a raw material cost saving technology to an enabling technology for large aircraft structural components.

In the late 1980s and early 1990s Northrop and ATK/Hercules worked on several joint projects sponsored by the Air Force which included fiber placement development and application. The technology was in its infancy as ATK was developing tow placement (as it was more commonly referred to originally) from its roots in filament winding technology. The main prize in the early days was \$5 per lb. high modulus carbon fiber and \$15 per pound high temperature/high performance resin instead of the \$60+ per pound price of prepreg. A wet process of running fiber through a resin bath prior to placement onto the layup mandrel was never able to realize the quality and consistency required by the design. This same process has been used in the large wind blade manufacturing process and it reminds us of how challenging (and messy!) that approach can be. In addition, the wind blade manufacturing industry has learned some valuable lessons from those early days of "build it as cheap as you can" using the lowest cost material you can deal with. While those early blades were built with lower manufacturing costs, the argument can be made that many of those blades failed very early in their lifecycle and required costly repairs or replacement to generate electricity. If the blade cannot turn because it has delaminated, it is not generating any electricity in addition to

Not only did the technology not realize the cost savings of dry fiber and wet resin, it was forced to adopt prepreg technology into the process—namely dealing with backing paper and ADDING to the cost of unidirectional prepreg tape by requiring it to be slit into prepreg tows. At the time of the ATCAS program, the AFP process was still evolving from what was originally envisioned as a much lower raw material cost build up starting with a dry fiber/wet resin process instead of a costly unidirectional fiber prepreg. The baseline process the ATCAS program selected for fabricating fuselage skins was AFP using prepreg tow. The dry fiber/wet resin tow had evolved to prepreg tow in an attempt to improve process consistency. The process was selected based on several factors including the potential for reduced material cost (compared to prepreg tape), the potential to achieve high lay-up rates over contoured surfaces, and the potential to efficiently support a significant amount of ply tailoring. In addition, the fact that tow material does not require backing paper eliminated a perceived risk of greater machine downtime.

When compared with the quality and consistency of parts made with prepreg tape, tow preg and subsequent prepreg tow, was not acceptable. The variability seen in the quality of the resultant panels would require compensation in the design of the part, resulting in weight penalties. But this did not prove fatal to the technology,

There have been many studies of the AFP process that have helped to shape and refine the characteristics and capabilities that exist in today's equipment offerings. But the ACT program allowed Boeing to better understand, study, define and refine the process to guide the technology development based on the needs of the user community. Everything from tack of the initial plies to the tool surface, to overlaps and gaps in the laminate; the most efficient ways to handle window/door cutouts, laminate thickness transitions, lay-up rates for flat, curved, cylindrical and duct shaped parts, etc., etc. What has ended up on production on the 787 is not the direct result of that ACT program, but the ACT program created the path for subsequent

instead tow placement reinvented itself (**Figure 14**).

AFP development to follow and improve upon.

*DOI: http://dx.doi.org/10.5772/intechopen.82353*

**2.3 Automated fiber placement**

the cost of repair or replacement.

**Figure 13.** *Airbus continuous pultrusion equipment [4]. Source: CTC Stade.*
