Global Design Challenge 2012 Final Report

Team 1: Traveling Forward Fuselage

Michael Cope, Alyssa Doman, North Carolina State University

Vamsi Krishna, S. Soundararajan,Sathyabama University of Chennai

Chelsea Chavez and Cameron MillerWichita State University

December 7, 2012

Table of Contents

1. Executive Summary2. Introduction

2.1. Background2.2. Objectives

2.2.1. General Objectives2.2.2. Forward Fuselage Objectives

2.3. Team Evaluations3. Methodology

3.1. Design Philosophies3.2. Project Breakdown

4. Design Overview4.1. Primary Components4.2. Secondary Components4.3. Tertiary Components4.4 Engines4.5. Mass4.6. Materials

5. Design Justifications5.1. Analytical Stress Analysis

5.1.1 Hoop and Longitudinal Stress5.1.2 Ground Loads5.1.3 Normal Unaccelerated Flight Loads

5.2. Primary Components5.2.1. Skin5.2.2. Frames

5.3. Secondary Components5.3.1. Forward Pressure Bulkhead5.3.2. Landing Gear Attachment5.3.3. Passenger Entry Door5.3.4. Radome5.3.5. Interface5.3.6. Windshield

5.4. Tertiary Components5.4.1. Passenger Seating5.4.2. Fuel Tanks5.4.3. Avionics Bay

5.5. Engine Selection5.6. Mass Properties5.7. Primary Components Materials

5.7.1. Composite vs. Metal5.7.2. Thermoplastic vs. Thermoset5.7.3. PPS and PEI

6. Design Logistics6.1. Primary Components

6.1.1. Skin6.1.2. Frames6.1.3. Floor

6.2. Joining and Repair6.3. Production

7. Conclusions8. Recommendations9. Acknowledgments10. References11. Appendix

1. Executive Summary

The aerospace industry has experienced several major breakthroughs since the birth of the airplane. However, the past few decades have been fairly quiet. The Spirit Aerosystems Global Design Challenge encourages students from universities all over the world to collaborate to create an innovative design with justifications for future implementation. This project deals with the forward fuselage section only of a 170-passenger, commercial airplane. It highlights the use of thermoplastic materials for primary structures and simplified yet effective structural reinforcements through elimination of stringers and crisscrossing frames that are slanted at a 20º angle. It also incorporates innovations for several specialty components of the forward fuselage: the radome, the forward pressure bulkhead and the windshield.

2. Introduction

“Spirit Global Design Challenge is an academic competition sponsored by Spirit AeroSystems [and Infosys Systems] that encourages college engineering students to seek out new thinking in “outside-the-box” fashion while using new technologies to design processes and aviation structure.” 1

The 2012 Spirit Global Design Challenge is comprised of four teams tasked with creatinga 170 passenger commercial airliner. These four teams are split into forward fuselage, center fuselage, empennage and wing teams. Each of the teams assumes the role of a Tier-1 supplier. Spirit and Infosys, acting on behalf of a hypothetical Original Equipment Manufacturer (OEM), created a set of customer requirements for each supplier. These requirements reflect current as well as anticipated needs of the aerospace industry.

The forward fuselage team, authors of this report, includes members from Wichita State University, North Carolina State University, and Sathyabama University Chennai. The forward fuselage team is, of course, responsible for the forward fuselage section of the traveling team’s airplane but also for for engine selection. Propulsion systems were assigned as shared responsibility amongst all four teams. Through inter-team collaboration, specific responsibilities were assigned for the completion of propulsion systems. The forward fuselage was tasked with engine selection.

2.1. Background

According to commercial market outlook research done by a major aerospace company, demand for large commercial jet aircraft is due to significantly increase over the next twenty-year period. Almost 60% of projected new deliveries are expected to be in the area of new growth, with only 40% in fleet replacement2. Single aisle aircraft comprise 63% of the commercial fleet today, and they are expected to increase their share over the next 20 years to 69%. This equates to 23,460 new single aisle aircraft deliveries by 2031.

Figures 1 and 2: Anticipated growth of market and fleet size for commercial airplanes2.

In order to attract prospective buyers of these new airplanes, performance increases must be made in comparison to currently available technology. For example, as of the second quarter 2012, an airline spends 31% of its total operating expenses on fuel. This is the largest single operating expense, and increasing the fuel efficiency of the aircraft could lead to substantial savings to the operators3.

This analysis of the forecasted market of the aerospace industry is what establishes the driving forces behind this project. Like any good business model, the product must respond to customer’s needs. In this report, various performance factors are considered for the design of an innovative aircraft that incorporates those concerns which customers will have in the next 20 years.

2.2. Objectives

2.2.1. General Objectives

The general objective of Spirit Global Design Challenge is for the four teams that make up the traveling team to collaborate together in order to create a unified aircraft design. This design should try to meet the emerging needs of the forecasted market, as mentioned in the previous section. The project’s primary sponsor, Spirit Aerosystems, has identified various anticipated needs as requirements for the aircraft. Please refer to Appendix A for a listing of these requirements.

2.2.2. Forward Fuselage Objectives

Each team also has a specific set of requirements that correspond to their section. Please refer to Appendix B for a listing of these requirements. Team 1 also derived the following additional requirements based on priorities emphasized in the team’s design philosophy (section 3.1) and the abilities of the forward fuselage:

● Bulkhead, radome and windshield improvements● Cargo capacity● Maximized component lifetimes● Lean manufacturing

2.3. Team Evaluations

This challenge takes on a competitive nature by the division and evaluation of individual teams from both the traveling and stationary aircraft teams. Evaluation of the each team’s work is done partly through periodic progress reports but primarily through a final report. A panel of industrial professionals will assign a grade to each team to determine the winner of the 2012 Global Design Challenge. The judging rubric and the weighting system used for the various progress reports and final paper are listed in Appendix C.

3. Methodology

3.1. Design Philosophies

A design philosophy is paramount in the creation of a product. The design philosophy gives guiding direction to this creation process and guards against scope creep. In this project, the product can be described as a forward fuselage design complete with production logistics. Two design philosophies were created: one for the entire aircraft (encompassing all four teams) and one specific to Team 1, the traveling forward fuselage team.

Aircraft design philosophy:To utilize current and emerging technologies to design an economical, energy-

efficient commercial airliner that extends the lifetime of modern aircraft.This design philosophy places more emphasis on cost effective use of technologies over the pursuit of undeveloped technology readiness level (TRL) technologies. Generally, the newer a technology, the higher the associated costs with implementing it. Creating an economical aircraftthat is energy-efficient and has an extended lifetime demonstrates that the team is sensitive to thecustomer’s needs.

Forward fuselage design philosophy:To optimize the roles of individual components and improve production

techniques to produce a low cost forward fuselage.This design philosophy places emphasis on achieving target requirements by utilizing weight reduction and placement strategies while maintaining structural integrity for individual components as well as improving the manufacturing process by reducing part complexity and part count.

3.2. Weighted Criteria Matrix

Team 1 identified seven criteria for a well-functioning aircraft: shape, weight, material, aerodynamic efficiency, structural strength, propulsion system and fuel efficiency.These criteria were set against the general requirements of the project to identify where weight should be placed in designing the aircraft. This resulted in a weight criteria matrix which can be found in Appendix D.

4. Design Overview

Figures with labeled components and dimensions can be found in Appendix E.

4.1 Primary Components: Skin, Frames and Floor

The cross section of the skin begins at the nose with a horizontal ellipse shape and tapers into a large horizontal ellipse at the interface between forward- and center-fuselage sections. All surfaces are smooth and the skin is a single unit. Cut-outs for the windshield, passenger entry door, flight deck exit door, cargo door and landing gear attachments are identified.

The frames are made of composite and T-shaped in cross section. They are placed at positive and negative 20º angles that overlap and create a crisscrossing pattern. Each frame flushly follows the inside radius of the skin.

The passenger floor is a composite sandwich supported by vertical beams that connect it to the cargo floor. It is horizontally supported by I-beams that run from the left to the right side of the fuselage. It also has seat tracks that run fore to aft for passenger seating.

4.2. Secondary Components: Bulkhead, Landing Gear Attachment, Interface, Radome, and Windshield

The forward pressure bulkhead is a deformable dome construction as opposed to a traditional flat bulkhead design. The dome is canted forward at an angle of approximately 5 degrees. A high impact smart composite is used in the construction to absorb high impacts from birdstrike and foreign objects in the nose of the aircraft.

The landing gear attachment is composed of a box section whose top is parallel to the passenger floor and whose bottom follows the contour of the inner skin surface. Horizontal beams, located at the top of the box, and vertical beams, located on the sides, distribute the load from the landing gear to the frames.

The interface is the splicing section between the forward and center fuselages. Because the center fuselage has a triple-bubble cross section and the forward fuselage chose a horizontal elliptical one, a part that makes full contact with the perimeters of both is necessary.

The radome consists of a sandwich composite that allows both radar transparency and impact resistance as well as lightning diverter strips to allow electric current to travel safely fromthe radome to the skin.

The windshield consists of two panes, one of which is a structural composite layer that transfers loads between the windshield and skin, and the other which acts as a fail-safe and an abrasion-resistant layer. Heating of the panes is done through an electroconductive ply.

4.3. Tertiary Components: Passenger Entry Door, Fuel Tanks, Avionics Bay, and Passenger Seating

The passenger entry door opening is 34 inches wide by 72 inches tall. The door itself is only a plug. The loads from the flames are carried by the door using a skin doubler and door frame.

There are two fuel tanks located in the forward fuselage, underneath the passenger floor and one on each side of the cargo bay. Each tank has a volume of 930 gallons. Fuel lines are routed beneath the cargo floor to the engines, which are mounted on the empennage section.

Avionics components are accessible via doors in the cabin wall that open into the crawlspace below the cockpit floor. Electrical wiring routing travels directly from the avionics bay into the cockpit for flight deck controls. Other wire bundles, running aft, are routed as necessary either underneath the cargo floor, underneath the passenger floor, or behind interior trim panels.

Passenger seating for 21 business customers is comprised of three rows of 2x3x2 arrangement, where “x” represents an aisle. Each seat is about 24 inches wide and 38 inch seat pitch, fairly standards for a business class chair. The seats have the ability to lay nearly flat--that is to say, not completely flat like a bed but much more so than an economy customer’s seat.

4.4. Engine

Two Pratt and Whitney PW1127G geared turbofans are mounted on the empennage. Each engine delivers 27,000 lbs of thrust. They have a fan diameter of 81 inches, and feature a 12:1 bypass ratio.

4.5. Mass

Using an average structural density of 1.51 g/cm3, and a fuel density of 0.773 g/cm3, the calculated mass of the forward fuselage comes to approximately 20,000 kg for the entire section.

4.6. Primary Component Materials

The material selections for primary components only are mentioned in Table 1. Althoughmaterial selections are specified for secondary components in their respective justification sections (see 5.3), they are outsourced and for this reason are not discussed in materials or manufacturing sections. Only primary components are manufactured in-house.

Table 1: Material selections for all primary components

Component MaterialSkin Polyetherimide Composite

Frames Polyetherimide Composite

FloorPolyetherimide Composite Sandwich with

Honeycomb Core

5. Design Justifications

5.1. Analytical Stress Analysis

5.1.1. Hoop and Longitudinal Stresses

Due to the internal cabin pressure necessary for the passenger transport airliner design, the forward fuselage skin had to be tested to withstand the hoop (circumferential) and longitudinal stresses imparted on the structure. The elliptical cross sectional area prevents a simple hoop stress analysis because the internal pressure not only creates a hoop stress within the

skin membrane but it also imparts a bending moment on the cross section. This bending moment is produced due to the non-uniform curvature of the ellipse wall. Due to the complexity of analyzing this additional bending moment induced on the ellipse, a geometric approximation technique was used in order to estimate the normal stress in the skin of the fuselage9. An obroundshape, shown in Appendix F, was used to approximate the shape of the elliptical fuselage cross section. Equations for the normal stress in the skin at critical locations are also given in Appendix F. Using the given geometry approximation for the forward fuselage, a skin thickness of 12mm, and the given ultimate internal pressure of 18psi (124 kPa), locations A and D are found to be in compression with magnitudes of 9800 and 8534 MPa respectively, while locationsB and C were found to be in tension with magnitudes of 9844 and 8604 MPa respectively. Even without frames, which provide additional strength in the hoop direction, the fuselage skin can withstand the ultimate internal pressure.

Unlike hoop stress, longitudinal stress can be analyzed using traditional methods9. Equation 2 in Appendix G shows the relationship between the longitudinal stress and the given cross sectional area shape. Equation Set 3 was used as an acceptable approximation for the circumference of the ellipse. Using the given geometries, the longitudinal stress was determined to be 13 MPa, a value much less than the hoop stress in the skin of the forward fuselage. If the longitudinal stress were the only normal stress present in this direction in the fuselage, then the skin would be sufficiently analyzed for normal stresses; however, bending of the fuselage in flight and on the ground add to the normal stress experienced by the frames and skin of the fuselage. Analysis of the stresses produced from ground and normal unaccelerated flight loads are analyzed in Sections 5.1.2 and 5.1.3.

5.1.2. Ground Loads

Ground loads were calculated by simply analyzing the aircraft in a static position with thefront and rear landing gears remaining stationary on the ground. The reaction loads at the landinggear positions were determined using static analysis and then bending moments were calculated using integral approximations across the length of the forward fuselage section. The normal stresses were then calculated using Equation Set 1 in Appendix F. Normal stresses in the skin arereported in Table 1. These calculations along with computational results helped in the material selection for the skin in Section 6.1.1.

5.1.3 Normal Unaccelerated Flight Loads

A similar process was used to analyze the flight loads experienced by the forward fuselage. Simple beam theory was applied with boundary conditions determined by the aircraft inflight. The normal stresses were calculated and are also reported in Table 1 of Appendix F. Thesecalculations did not consider longerons and their added bending resistance; therefore, the stress estimates can be viewed as over estimates of the actual normal stress in the forward fuselage skin. These calculations aided in the material selection for the skin, which is discussed in Section6.1.1.

5.2. Primary Components

5.2.1. Skin

The initial aerodynamic design for the forward fuselage was a conventional circular or nearly circular cross section. The given production requirement of 60 units per month provided a basis for manufacturing the skin of the forward fuselage with a simplified, low-cost process. However, the center fuselage selected a triple bubble configuration had been selected by the center fuselage team. Various options were considered as a result: 1) a continuation of the triple bubble into the forward fuselage, 2) a horizontal ellipse was considered for the cross sectional area, or 3) maintaining a circular cross section at the nose of the forward fuselage and transitioning to the triple bubble cross section by means of fairings to cap off the outer two bubbles.

Commercial aircrafts typically use circular or relatively circular cross sections, mostly due to pressurization effects. The pressurization loads are best taken by the circular cross sections, which have a resultant net force that is normal to the circumferential direction of the cylinder, allowing for equal stress distribution in all directions. An ellipse does not respond to pressure loads in the same way. For an ellipse, the upper and lower surfaces parallel to the major axis try to deform faster than the corner surfaces, those parallel to the minor axis. This creates a bending load in addition to the circumferential load already present due to pressurization.

The relative strengths of both cross sectional areas were tested by submitting them to internal pressure loading conditions. Loading conditions were determined using similar methods used in sections 5.1.1 through 5.1.3. A detailed analysis is provided in Appendix G. The results determined that the triple bubble configuration would not hold up to internal pressure loads as efficiently as the horizontal ellipse cross section. Stress concentrations for the elliptical cross section reduce dramatically between locations on the major and minor axes of the ellipse, as shown in the figure below. However the triple bubble cross section not only has more spread-out stress concentrations but those found at the bubble intersections were found to be considerably high. High enough that additional structural support would be required which would then increase a significant amount of weight to the overall structure.

Figure 3: Internal pressure analysis of an elliptical cross section

In the end, the decision was made to utilize a horizontal ellipse cross section. This cross section not only leads to a more unified overall aircraft than an circular cross section with space

and material wasting fairings, but it also allowed for a more efficient stress containment without necessitating unnecessary supports. Furthermore, it allows for the forward fuselage to allot leftover space to cargo capacity, which the center fuselage lost substantial space for in utilizing atriple bubble cross design.

5.2.2. Internal Structural Configuration

Internal frame design focused on reducing the number of components while adequately supporting the aircraft at extreme conditions. Conventional frames oriented along the longitudinal and circumferential directions were considered against slanted frames alone. Analysis of conventional versus varying degrees of slanted frames is given in detail in Appendix H. The final selection for the internal structural configuration was frames that are angled 20 degrees from vertical. This configuration proved to withstand the bending loads of the aircraft in flight without even the need for longerons. The lack of these internal structures reduces the weight of the forward fuselage and increases the ease of manufacturing, assuming proper tooling.The maximum stresses experienced in the 20 degree configuration were 0.4 GPa. The best frame orientation slant angle was decided by an iterative process which involved analyzing various angle orientations. This analysis is described in Appendix H along with specific details on meshing and element quality. Final internal structural configuration can be viewed in Figure 8 of Appendix E.

5.3 Secondary Components

5.3.1 Forward Pressure Bulkhead

The forward pressure bulkhead in the forward fuselage is the forwardmost point of the aircraft. The forward pressure bulkhead provides structural rigidity and protects the avionics components typically housed directly behind the forward bulkhead. Additionally, the forward pressure bulkhead is susceptible to high velocity impacts from bird strike or other foreign objects. Penetration of the bulkhead would result not only in structural damage to the airframe but also to antennae and other critical avionics components. With such critical functions as these,the forward pressure bulkhead must be designed to withstand cabin pressure, flight, and possible high impact loading.

Conventional metallic or composite bulkheads feature many components with spars or ribs joined to the bulkhead membrane for additional stiffness. These stiffeners, while effective, add weight and increase overall costs to the aircraft. With the team’s design philosophy in mind, the selected forward pressure bulkhead design reduces part count and weight.

The selected forward pressure bulkhead will feature a deformable, composite dome unlike the traditionally flat, rigid bulkhead design. The dome design handles the pressure loads more efficiently than a flat stiffened panel and the dome membrane deform plastically under impact loads due to bird strike7. The dome deflection increases the impulse time over which the impact occurs and thus transfers less stress to the rest of the aircraft structure5. After deformation, the membrane returns to its undisturbed state unless the impact is great enough to permanently deform the bulkhead. In these instances, a bulkhead repair is necessary although penetration will not occur and as a result, important avionics components behind the bulkhead will remain protected. The forward pressure bulkhead is also canted forward at a specific angle

(typically 5-7 degrees). Canting the bulkhead increases the likelihood of foreign object impact occurring directly into the bulkhead instead of at an angle. Illustrations of the selected bulkhead design are shown in Figures 1 and 2 of Appendix H.

In addition to a deformable, domed forward pressure bulkhead that efficiently dissipates the energy from foreign object impact; the material used for the bulkhead membrane is a high impact smart material. The composite material utilizes the concept of shear thickening which allows the material to remain relatively soft and malleable until impacted violently6. This material increases the effectiveness of the domed bulkhead design by further dissipating energy within the material itself and transferring less stress to the adjoining airframe. This design can beseen in Figure 3 of Appendix H.

5.3.2. Landing Gear

The landing gear box is composed of a thin-walled box section reinforced by T-beams aligned vertically along all four sides. I-beams form the top of the box and extend beyond the width of the box section, attaching to the passenger floor support structure. The vertical T-beams attach to the frames and help distribute the load into the frames and skin.

5.3.3. Interface

The interface is a function of the center fuselage’s triple bubble configuration and the forward fuselage’s elliptical cross section. In preliminary design, some consideration was given to continuing the center fuselage triple bubble on though the forward fuselage. The idea was discarded as it was deemed too complicated a design, and did not align well with our design philosophy. This led to the problem of joining two dissimilar cross sections.

The interface is essentially a cast titanium vertical frame that follows the inner contour ofthe forward fuselage on one side and follows the inner contour of the center fuselage triple bubble on the other. The vertical areas where the cross sectional profiles are discontinuous are filled with a similar thickness of titanium, forming one solid piece. The sections of skin flush with the interface flanges will be attached by means of hi-lok type fasteners through the skin and flange. In addition to attaching directly to the skin, stringers will be run from the aft frames and into the interface. The center fuselage will also be able to terminate their stringers at the interface.

5.3.4. Radome

For the radome, a combination of E-glass and syntactic foam is used. Syntactic foam is made up of microspheres in a cyanate ester resin11. The casing for the radome is a sandwich composite layup, with E-glass as the laminate material and the syntactic foam as the core material. The casing has a single sandwich layer throughout, with a second reinforcing layer in front of the radar for added impact protection. The geometry of the radome is shown above in Figure 1 of Appendix H.

Impact strength is not the only important factor in deciding the materials used for the radome. The radome is where the radar equipment is housed, which means it must be protected by impact, and lightning strikes, but also have clear signal passage. E-glass was chosen as the laminate material because it is an ideal material for protecting devices that must transmit signals.

For lightning strike protection, lightning diverter strips are placed radially over the skin of the radome in a sunburst pattern. The purpose of these strips is to attract lightning and divert it away from the casing and the radar equipment inside. The current in the diverter strips is then directed to the aircraft skin, which has a protective mesh12. The radome does not have this protective mesh because the conductive materials typically used also create problems with signal blocking. The diverter strips still produce some of that risk, but it is lessened because the strips cover muchless surface area than the protective mesh.

5.3.5. Windshield

Figure 4: Windshield lay-up10

The cockpit windshield is designed to be load-bearing, as it is a large section of the forward fuselage with properties that differ from the skin. What makes the windshield load-bearing depends largely on the material. In order to be used as a windshield, a material must have as low a refractive index as possible, preferably approaching a value of unity, to make the pilots’ field of view as clear as possible. The material must also have high impact strength properties, as it is at the front of the aircraft where impact risk is very high. To solve this issue, a composite windshield is used. The primary load bearing window is made up of a composite sandwich, with the laminate being a transparent glass-resin composite and the core a transparent epoxy resin. The glass-epoxy sandwich mimics the refractive properties of tempered glass, but is much more ductile and is less dense. The laminate layers are integrated into the aircraft skin, allowing loads applied to the window to be transferred to the aircraft skin, and vice-versa. A thin layer of tempered glass serves as a failsafe. The tempered glass is designed to shatter upon impact and release some of the impact energy. It also serves as a scratch-resistant surface, as wellas an insulation layer. For anti-icing, an electroconductive face ply is applied between the tempered glass and the primary structural window on the outer surface of the outer lamina10.

5.4. Tertiary Components

5.4.1. Passenger Seating, Fuel Tanks, Avionics Bay

The initial design for the forward fuselage was about 6 meters long from nose to the aft section. This allowed for galley space and lavatories as well as crew seating to be in the forward section. Upon receiving the change order, major modifications had to be made in order to fulfill the request. About ten percent of the total passenger capacity is now located in the forward fuselage in the form of business class seating. In order to accommodate the additional seating, another 6 meters was added to the length of the section.

Due to the shape of the cross section, this extension resulted in large amounts of unused space. Upon communication with the traveling aircraft teams, it was discovered that the center fuselage team was short on cargo space and the wing team was looking for more fuel storage.

The excess space was used to add cargo space in the forward section, as well as a forward cargo door. Two 930 gallon fuel tanks were also added to the forward fuselage, to help with the fuel demands of the aircraft. The elevated flight deck left approximately one meter of space between the passenger floor and the flight deck floor. This space was allocated for avionics components.

5.5 Engine Selection

PW1127G engines were chosen in final selection over the CFM56-7B. The geared turbofan (GTF) was preferred due to greater fuel efficiency. According to the manufacturer’s website, the PW1127G is capable of fuel efficiency increases of up to 16% versus today’s best engines22. Fuel efficiency was the greatest concern when selecting engines. Aside from better fuel economy, other added benefits of a GTF include reduced carbon emissions and reduced noise pollution22.

5.6. Mass Properties

Using an average density of 1.51 g/cm3 for most materials and a fuel density of .773 g/cm3, the calculated mass of the model comes to about 20,000 kg for the entire section. The initial target mass value mentioned in PDR2 was 5400 kg, based on Raymer’s weight estimate for a forward fuselage being 10% of the overall aircraft weight24.

This initial estimate was overshot by almost a factor of four. There are two main contributions to the inaccuracy of the initial estimate. The first is the addition of the business class section to the forward fuselage. The second is the additional weight due to the fuel tanks in the forward section.

5.7. Primary Component Materials

No one company is optimized to purchase and house the equipment for production of all the parts of an aircraft. Outsourcing is common for a Tier-1 level supplier. Therefore, the focus of this project is on the production of the primary components of the section assigned, the forward fuselage. All other components--secondary, tertiary and so forth--are assumed to be outsourced. Further discussion of the logistics of the supply chain management involved in this process of integrating parts can be found in section 6.2.

5.7.1. Composite vs. Metal

One aspects outlined in the forward fuselage´s design philosophy is an emphasis on usinglightweight materials. In the figure below is a decision tree used for the material selection, beginning with the composite-metal decision.

Figure 5: Decision flow for material selection of primary components

Composites are prized not only for their exceptional specific strength. A great deal of research has been invested in the use of composites in the aircraft industry. Many aircraft manufacturers have at least one primary component made from composite material: the Gulfstream G650’s nacelles, the Airbus A350WB’s fuselage. There is still a fair amount of skepticism concerning the performance and longevity of composites13. However, funding for the research and development for composites is only anticipated to increase in the future. Composites were chosen for the primary components of the forward fuselage.

5.7.2. Thermoplastics vs. Thermosets

Thermoplastics also have an uphill battle for their claim in the aerospace territory. Only thermosets have been used for primary components of commercial aircraft thus far. This can be attributed partly to previous inability to mass produce them as well as differences in mechanical properties. Unreinforced thermoplastics are notorious for higher creep and lower strength properties when compared to thermosets14. However, these can often be combated with techniques such as adding stiffeners and other reinforcements. Thermoplastics boast many advantages: higher impact toughness, lower weight, good FST (flame, smoke and toxicity) properties, no expiration date and the ability to be reworked/ recycled. The aforementioned advantages are further explained below:

1. Higher impact toughness not only guards better against bird-strike but also reduces the likelihood of part scrapping during the manufacturing process. This is a legitimate concern in an industry where many machinists are still adjusting to working with composite parts. Something as simple as a machinist accidently dropping a tool onto a composite part can require hours of inspection alone to determine the effects on the composite part.

2. Lower-weight materials for primary structures mean a substantial reduction in the overall weight of the plane. This is attractive to future customers due to the anticipated fuel savings.3. In order to meet FAA requirements for fire, smoke and toxicity performance, additional layers of insulation are necessary to guard thermoset materials15. Nearlyall thermoplastics perform well in these three areas and thus do not require additional layers, which add weight and cost to the final design.4. Theoretically, thermoplastics have an infinite shelf life. Therefore, unlike thermosets, they do not require temperature-controlled storage before curing. Thisreduces transport costs (and complications), storage costs and costs associated with expired materials.5. Recyclability of a material not only means that trimmings can be reused rather than wasted; it also implies a level of re-workability. A level of degradation to theresin and/or fiber can sometimes be anticipated but otherwise the re-workability of a thermoplastic means the ability to correct any flaws incurred by the manufacturing process. Porosity is an excellent example of a flaw that can be eliminated by reprocessing the thermoplastic part20.

Methods for manufacturing thermoplastics have been in development for many years.

The U.S. government made a great investment in thermoplastic R&D in the 1990s. At that time, feasible manufacturing methods were not to be obtained13. However, a recent resurgence of interest in thermoplastics has resulted in qualification for various thermoplastic materials and manufacturing processes. The Thermoplastic Affordability for Primary Aircraft Structures (TAPAS) program is an example of collaboration between various companies, led by Airbus, to increase technology readiness levels of thermoplastic manufacturing processes. Boeing is also involved in a similar program, the ThermoPlastic Research Center which is based in the Netherlands.

5.7.3. PPS and PEI

Polyphenylene Sulfide (PPS) is an aerospace-grade resins others that are offered commercially for composite systems. Poly ether ether ketone (PEEK) was one of the first thermoplastics used in composites for aerospace applications but PPS has greater thermal stability and is typically less expensive to obtain19. PPS also has a lower processing temperature than PEEK, which reduces the residual stresses incurred during processing as well as reduces energy-consumption costs19.

There are a limited number of engineering thermoplastics with mechanical properties suited for aerospace applications. Although there is a possibility for the development of other thermoplastics in the years to come, these materials have been certified and used for other programs. Qualification can be an expensive and lengthy process so this as an important consideration for the imminent use of thermoplastic materials. Commercial availability is also very important in terms of supply chain management for this project: if a material is not easily to obtain, it is unlikely that suppliers will be able to keep up with the high volume production rate requirement. Although PPS was originally chosen as the material of choice, after full-scale testing* it was determined that the strength characteristics were not enough for the design. At that point, a decision had to be made: implement strategies such as increasing part thickness to recover

strength or to choose a different resin. Changing resins implies a loss in investment on all the testing done for PPS. A similar aerospace-grade thermoplastic resin, Polyetherimide (PEI), was chosen to replace PPS. Due to the similarities of both resins, the capability of reusing tools and equipment used to make the PPS testing models, results in less of a loss from the PPS testing.

*In order to simulate real industry experiences in product development, the sponsor imposed changes upon each team. In this case, the material was to hypothetically have undergone full-scale testing and failed.

6. Design Logistics

Table 2: Manufacturing methods for primary components of the forward fuselage

Component Material

Skin Automated Fiber Placement

Frames Powder-infused Pultrusion

Floor Automated Tape Laying

The leading factors considered when choosing the manufacturing processes include: equipment costs, geometry limitations, cycle time, technology-readiness level and automation. These reflect the design philosophy’s emphasis on simplified manufacturing and part simplicity.

6.1. Primary Components

6.1.1. Skin

In accordance with the design philosophy, efforts to reduce part number and assembly steps were made during the manufacturing process selection of the skin. Pressforming is a common thermoplastic manufacturing processes, favored due to the very quick cycle times. However, quality of parts using pressforming, especially for parts with complex geometries, is questionable. Furthermore, there are size limitations for these parts. Team 1 wished to reduce part numbers of the skin to the bare minimum--one part. Pressforming and other similar thermoplastic manufacturing processes are not capable of this. However, automated fiber placement (AFP) is very well suited for curved parts and large geometries. Automated fiber placement production cycles are not nearly as quick but the anticipated decrease in assembly time, labor hours and extra parts would decrease overall production as well as costs. Furthermore, higher pressures are required for pressforming which results in higher production costs as well as shorter lifetimes for tooling17.

Another thermoplastic manufacturing process was considered: vacuum-assisted resin transfer molding. Although this process requires substantially less capital investment in equipment than the previously mentioned processes, problems remain with wet-out of the fibers. Furthermore, complete wet-out of the fabric is already difficult enough with thermoset resins andeven harder with thermoplastic resins, which require a higher melt temperature. Table 2 in

Appendix F compares these thermoplastic processes for the skin based on the following criteria: equipment cost, geometry limitations, cycle time, technology-readiness level and level of automation.

Automated Fiber Placement consists of a robotic arms wrapping pre-impregnated composite fabric around a mandrel. Since the composite fabric in this instance is a thermoplastic,the material must be heated past the material’s glass transition temperature through methods like laser or induction welding. This particular process is combined with the fabric placement throughan attachment on the head of the robotic arm. The fabric used for this application is unidirectional tape, which has the greatest tensile strength due to its high fiber content and fiber orientation. Placement of the tape at varying angles during AFP gives the skin directional stability in all directions. Additionally, lightning strike and corrosion protection is incorporated in the final outer layers of the skin by use of glass fibers and metal-mesh imbedded carbon-fiber fabric.

6.1.2. Frames

Again, while more traditional thermoplastic processing techniques with faster cycle timesexist, there are still geometric limitations for these parts. Pultrusion also has some geometrical limits but is well-suited for long parts, especially those with a simple and unchanging cross section. Furthermore, pultrusion predisposes anisotropic properties on parts in the feed direction of the work done.

For optimal dispersion of the thermoplastic resin, fibers for pultrusion can be powder impregnated before heating and forming. The die through which the fiber-resin mixture is pulled matches the cross-section chosen for the frames, a T-shape. The anisotropic properties mentionedfor pultruded parts are especially desirable for the frames of the forward fuselage due to the nature of loading. 6.1.3. Floor: Automated Tape Placement

The decision for the floor component followed the same decision process as the skin. Theonly change is the type of automated fiber placement, automated tape placement, which is optimized for parts with little or no curvature13. ATP will be used to lay-up the composite layers. An induction welding system, integrated into the head of the robotic arm that also does the lay-up, heats the material above melt temperature to give it drapability. No curing is necessary as thethermoplastic will return to solid form after cooling from the induction welding process. The inside sandwich layer will use traditional adhesive and honeycomb to join the two composite sections.

6.2. Joining and Repair: Thermoplastic Welding

The use of thermoplastic welding as the primary joining offers several advantages over the two most common methods of joining:

1. A clear advantage over mechanical fastening methods, such as riveting and drilling. Not only are there less parts to be dealt with, but the high stress

concentrations seen with rivets and bolts are diminished with a continuous weld. Welded joints are less suitable than mechanical fasteners only when the part thickness is very small.2. A clear advantage over adhesive bonding: welding offers quicker cure cycles, less surface preparation, greater bond strength and has no material expiration date21. A chart comparing adhesive bonds to a type of thermoplastic welding--laser welding--that further compares the two can be found in Appendix I.

The most formidable obstacle for thermoplastic welding is one that comes with the material selection. Maintenance, Repair, Operations (MRO) units that operate on in-service airplanes still prefer repair methods best-suited to metal parts. These methods still function for thermoplastic parts but to a lesser degree. Training MRO staff to adapt their methods to a new material may prove to be challenging, especially for airports located in obscure locations that may not have immediate access to all the proper tools20.

Finally, it is important to note that this joining method is consistent with the design philosophy’s emphasis on simpler manufacturing methods, especially in terms of automation. Welding can be done manually or with the use of robotic arms. The current technology for robotic arms continues to improve upon the range and flexibility of these machines. This not only reduces labor costs but decreases variability in the joining operation which thus increases repeatable quality.

6.3. Production

For this project, two major industrial requirements are emphasized: ease of transportationbetween production facilities and high volume production. Ease of transportation depends largelyon the location of the facilities. If facilities are moved outside of the United States, transportationwould likely be via boat or cargo aircraft, depending on the size of a part and the cost to transportit. If facilities remain in the US, transportation would be easiest via freight. Keeping production in the US would likely reduce transportation costs; however the major deciding factor in choosing a supplier would be the quality of the component design and the supplier’s budget for the design.

This project requires a very high production volume compared to current industry rates. A major issue that plagues large projects in the aerospace industry today is backlog. During a visit to the Spirit facility in Kinston, an engineer for the Airbus A350 XWB mentioned that due to constant design revisions and the need to communicate these revisions between suppliers, backlog restricted the increase of production rates. Other factors in this very low production rate are that the facility was just built in 2010, and that the parts are shipped between Kinston and Saint-Nazaire, France during assembly23. In order to meet the high production rates required for this project, it is essential to incorporate methods of reducing backlog as much as possible. One method is to use suppliers that can produce more than one product to be used in assembly, so thatdesign changes can be communicated much more quickly and effectively. Another is to ensure that there are back-up suppliers identified in the likely event of a shortage from the primary suppliers.

As far as actual production and assembly of parts, the aircraft industry is increasing investment in automated processes. Automation, as it involves large and expensive machinery, has a much higher capital costs than processes done manually. Automated production still

requires personnel to verify, monitor, program machinery, but automation reduces the impact of the human error on the component being manufactured. Automated machinery does not completely totally eliminate error but it does produce errors that are more consistent and predictable and thus more easily controlled.

7. Conclusions and Recommendations

The forward fuselage design and manufacturing plans developed by Team 1 strive to meet the requirements given by the sponsor as while maintaining loyalty to the design philosophy identified and developed by the traveling forward fuselage team at the start of the Spirit Global Design Challenge. Further assessment of the economic climate could have been considered to better meet the emerging needs of the aircraft in the coming 15-20 years, but this issomething that changes year by year, month by month. Team 1 did their best to identify and consider possible drawbacks of all nearly all designs and processes. With more time and resources, better results could have been made on many of the requirements. Overall, Spirit Global Design Challenge offered Team 1 a very unique and rewarding experience through the course of the fall semester of 2012. The team members of Team 1 have all profited from workingin this Challenge, regardless of the results.

8. Acknowledgements

Team 1, the traveling forward fuselage, has many people to thank for this unique experience in the Spirit Global Design Challenge 2012. This project would not have been nothing without the advising and support of the individuals and organizations mentioned below:

Spirit Aerosystems and Infosys, for their support and investment in us, not only financially but also through their time and dedication to Spirit Global Design Challenge.

Kevin Brauning and Allison Ronning from Spirit Aerosystems, for their incredible dedication to SGDC, in planning many details of the Challenge itself as well as answering every single question all 48 students participants approached them with.

Infosys Mysore and Indian students, for hosting the entire traveling team in Mysore and providing us with the necessary resources to establish good team collaboration.

Students and faculty at North Carolina State University, for hosting the entire traveling team in Raleigh for a week of research even while school was in session during the entire visit.

Professors at our respective universities:

1. Dr. Charles Yang from Wichita State University, for advising on structures and composites.

2. Dr. Scott Miller from Wichita State University, for general aerospace advising.3. Dr. Kara Peters from North Carolina State University, for advising on structural

health monitoring and composites.4. Dr. Larry Whitman, for coordinating the many details of the Challenge.

Fellow SGDC teammates from the traveling team, for providing an intellectual and competitive experience as well as their patience with inter-team communications.

Industry professionals:

1. Arnt Offringa of Fokker Aerostructures, for his advising on thermoplastics.2. Winand Kok of TenCate Advance Composites, for his advising on thermoplastics

and review of the rough drafted thermoplastic sections of this report.3. Carroll Grant of Aerospace Composites Consulting, for his advising on automated

composite manufacturing.4. Joe Spangler of Toho Tenax, for his advising on thermoplastics suppliers and

repair.

10. References

1. Global Design Challenge. Spirit Aerosystems. <http://www.spiritaero.com/GlobalDesign.aspx>

2. “Long-Term Market”. About our Market. The Boeing Company. <http://www.boeing.com/commercial/cmo/fleet_development.html>

3. “A4A Quarterly Cost Index: U.S. Passenger Airlines”. Airlines for America. <http://www.airlines.org/Pages/A4A-Quarterly-Cost-Index-U.S.-Passenger-Airlines.aspx>

4. Harris, Charles. “Opportunities for Next Generation Aircraft Enabled by Revolutionary Materials”. Powerpoint Presentation. AIAA SDM Conference, April 4-7, 2011, Denver, CO.< https://www.aiaa.org/uploadedFiles/About-AIAA/Press_Room/Key_Speeches- Reports-and-Presentations/Charles_Harris_presentation_2011.pdf >

5. Barney B.L. Anderson, J. R. (2010). Patent No. US7766277B2. United States.

6. AZoNano. (2010, July 28). Researchers Invent New Flexible, Lightweight, Impact-Resistant Composite Material. Retrieved December 6, 2012, from AZoNano: <http://www.azonano.com/news.aspx?newsID=18773>

7. Niu, M. C.-Y. (1999). Airframe Stress Analysis and Sizing. Hong Kong: Conmilit Press Ltd.

8. Niu, M. C.-Y. (1988). Airframe Structural Design. Hong Kong: Conmilit Press Ltd.

9. Wang, J. C. (2005). Stress Analysis of an Elliptical Pressure Vessel Under Internal Pressure. Hartford: Rensselaer at Hartford.

10. Nordman, Paul S. (2005). U.S. Patent No. 6,889,938 B1. Renton, WA: U.S. Patent and Trademark Office.

11. Bibin John, C. P. Reghunadhan Nair, Dona Mathew, K. N. Ninan. Foam Sandwich Composites with Cyanate Ester BasedSyntactic Foam as Core and Carbon-Cyanate

Ester as Skin:Processing and Properties.Propellants, Polymers, Chemicals and Materials Entity, Vikram Sarabhai Space Centre, Trivandrum 695022, India. 16 July 2008.

12. Hall, Allen. Thunderstorm Protection of Aircraft Radomes. Lightning Technologies Inc.

International Conference on Lightning and Static Electricity. 19 September 2005. <http://dzgcx.cuit.edu.cn/kcjs/ldkxjc/product_photo/2011512124717688.pdf>

13. Grant, Carroll. Aerospace Marketing Contractor, Aerospace Composite Consulting.

14. Kok, Winand. Manager of Engineering, TenCate Advanced Composites. Personal communication.

15. Offringa, Arnt. Direction of R&D, Fokker Aerostructures. Personal communication.

16. Meyer, Andrea. Research Engineer, Center of Innovation for Biomaterials in Orthopedic Research. Personal Communication.

17. Ticona. "Thermoplastic Composites for Aerospace." Internation Association of Plastics Distribution Apr.-May 2010: p. 2.

18. Tencate, “Advanced Composite Material Selector Guide: Aircraft”: p. 10. <http://az290931.vo.msecnd.net/www.aircraftinteriorsexpo.com/__novadocuments/8390x$query$xvx$eq$x634671499433970000 >

19. Strong, Brent A. “High Performance Thermoplastic Composites”, pgs. 168-175. Fundamentals of Composites Manufacturing: Materials, Methods and Applications. Society of Manufacturing Engineers, 2008.

20. Spangler, Joe.”Thermoplastic Lunch and Learn: Toho Tenax” National Institute of Aviational Research, Wichita, KS. 5 December 2012. Lecture, Q&A.

21. Klein, Jurgen, and Andreas Kraus. Is Plastic Laser Welding Economical? July 2004. <http://www.laserplasticwelding.com/is-laser-plastic-welding-economical.pdf>

22. “Pure Power PW1000G Engine”. Pratt & Whitney. <http://www.purepowerengine.com/index.html>

23. Spirit Aerosystems A350WB Facility Tour. Spirit Aerosystems, Kinston, NC. 19 October 2012.

24. Raymer, Daniel P. Aircraft Design: A Conceptual Approach, 4e 2006. AIAA Education Series.

11. Appedices

Appendix A: General Aircraft Requirements

4.1 General Requirements

Passenger capacity: 170 Cargo capacity: 1300 cu. ft

Target ‘Max Cruise’: Mach 0.85 Target Maximum Range: 6000 nmi

High Volume Production: 60 units/month– Assumed recovery of capital expenditures after 1500 units.

4.2 General Stress

Minimum Weight PhilosophyNo yielding at limit loads

Fatigue Life = 72,000 cycles– 1 Cycle = 1 Ascent, Cruise, Descent

No detrimental deformation Damage tolerance

No catastrophic crack growth Corrosion tolerance

Lightning Strike Birdstrike (per FAA requirements)

Anti-icing Compatibility with ‘PrimaryCommercial Service Airports’ asdescribed in FAA guidelines.

• Limit flight = 2.5G descent/ascent

• Ultimate flight = Limit flightx 1.5

• 9G fwd/6G downemergency landing case

• Max. Take-off Weight to bedetermined as an estimateof structural weights,passenger & cargo capacity,etc.

4.3 Additional Requirements

Technical Requirements• Manufacturing process• Maintenanceconsiderations• Inspectability

‘Soft’ Requirements• Economic feasibility study• Basic market outlook andoverview• Marketability

– During manufacturing– Fleet Support

Appendix B: Forward Fuselage Specific Requirements

Specific RequirementsMinimum Definition• Primary structuralconfiguration• Aerodynamic configuration• Forward landing gearattachment• Forward Pressure Bulkhead• Passenger entry door• Cockpit visibility• Flight deck exit• Accommodate 10% of totalpassengers as business class

Specific Stress & Loading• 18 psi ultimate/9 psi limitinternal pressure• Birdstrike – no cockpitpenetration• Account for stressconcentrations arounddoors due to internalpressure– Passenger entry door – plugdoor (carries no load)

Appendix C: Rubric Criteria and Report Weighting

Table 1: Rubric Criteria

Technical Elements Soft Skills

Satisfaction of design requirements Ability to work globally

Innovation Utilization of global resources

Technical Validity Utilization of global support

Performance improvement Project Management/Planning

Integration considerations (external) Global communication

Integration considerations (internal) Group Dynamics

Applicable to core business Response to Adversity

Technical Readiness Levels Collaborative Design

Possible Solutions

Economic Analysis

Maintenance Consideration

Manufacturing Analysis

Manufacturing and Processes

Awareness of design risks

Method

Table 2: Report WeightingSubmission Weight 1 2 3 4 5 Possible PointsReview 1 5% Graded at the discretion of the present judges. 25Review 2 5% Graded at the discretion of the final judging panel. 25Review 3 5% Graded at the discretion of the present judges. 25Review 4 5% Graded at the discretion of the final judging panel. 25Final Submission 80% Graded per the attached rubric 400

Appendix D: Weighted Criteria Matrix

Factors

Requirements

Shape

Weight

MaterialAerodynamic efficiency

Structural

Strength

Propulsion System

FuelEfficienc

yPassenger Capacity 140 2 2 0 0 0 1 0Cargo Capacity: 1000 cu.Ft 2 2 0 0 1 0 0Max cruise: M 0.85 2 2 2 2 2 2 1Max range: 6000 nm 1 2 1 2 1 2 2Min weight philosophy 0 3 2 1 1 1 2Fatigue life of 72000 cycles 0 0 2 0 2 0 0Resist lightning strike 0 0 2 0 0 1 0Damage tolerance 1 0 2 0 2 0 0Corrosion tolerance 0 0 2 0 0 0 0Birdstrike 1 0 2 1 2 2 0Anti-icing 0 0 2 0 0 2 0

Totals 9 11 17 6 11 115

Appendix E: Forward Fuselage Drawings

Figure 1: Top view

Figure 2: Side view

Figure 3: Front view

Figure 4: Bottom view

Figure 5: Side view with transparent skin

Figure 6: Landing gear attachment

Figure 7: Isometric view of interface (skin transparent), fuel tanks and cargo space

Figure 8: Frame view

Appendix F: Manufacturing Comparison Tables

Table 1: Manufacturing processes for a single-piece fuselage skin

Initial Costs

Geometry limitations

Cycle Time

Technology Readiness Level

Level of Automation

Vacuum Aided Resin Transfer Molding Low Low High Low MediumAutomated Fiber Placement High Low Medium Medium HighPressforming High Medium Low Medium High

Table 2: Manufacturing processes for T-shaped frames

Initial Costs

Geometry limitations

Cycle Time

Technology Readiness Level

Level of Automation

Vacuum-aided Resin Transfer Molding Low Low High Low MediumPressforming High Medium Low Medium HighPultrusion Medium Low Medium Medium HighHand Lay-up Low Low High Low Low

Appendix G: Governing Equations

Figure 1: Obround vessel under internal pressure, p (Section 5.1.1)9

Equation Set 1: Obround normal stress equations

σ A , B=PRt

±PLCc6 AI

σC , D=P( R+ L)

t±

PLc6 I [3 ( L+2 R )−

CA ]

Where

P = internal pressure

c=t2

I =t 3

12

C=L2 (2+3π )+12 R2

A=2( L2

R )+πR

Equation 2: Longitudinal Stress due to Internal Pressure

σ longitudinal=P Aellipse

C ellipse t

Equation Set 3: Ellipse Geometry Equations

Aellipse=πab

C ellipse≅ π (a+b)[1+ 3( a−ba+b )

2

10+√4−3( a−ba+b )

2 ]a=

12

( Lengt h of Major Axis )

b=12

( Lengt hof Minor Axis )

Eccentricity: e=√1−( ba )

2

=0.7839

I zz=14

πa b3

I yy=14

π a3 b

J O=14

πab (a2+b2

)

Table 1: Normal Stress in Fuselage Skin for Ground and Flight Loads

Ground Loads Flight Loadsx (m) M(x) (kN m) σ (MPa) M(x) (kN m) σ x (MPa) σ6G (MPa)

0 0 0 0 0 0

0.5 -1.6 0.006 -1.6 0.006 0.0331 -6.4 0.022 -6.4 0.022 0.134

1.5 -14.5 0.050 -14.5 0.050 0.3012 -25.7 0.089 -25.7 0.089 0.535

2.59 -43.1 0.149 -43.1 0.149 0.8973 -11.2 0.039 -58.6 0.203 1.220

3.5 82.6 -0.286 -80.4 0.279 1.6723.73 124.8 -0.433 -91.4 0.317 1.901

4 173.2 -0.601 -105.3 0.365 2.1914.5 259.5 -0.900 -134.6 0.467 2.800

5.05 349.2 -1.211 -172.1 0.597 3.5805.5 418.4 -1.451 -206.8 0.717 4.303

6 491.1 -1.703 -249.7 0.866 5.1966.55 565.8 -1.962 -302.2 1.048 6.287

7 622.8 -2.160 -349.2 1.211 7.2657.0358 627.2 -2.175 -353.1 1.224 7.346

7.45 675.4 -2.342 -400.6 1.389 8.3358 728.5 -2.526 -474.6 1.646 9.875

8.46 766.6 -2.658 -542.8 1.882 11.2949 800.4 -2.776 -633.9 2.198 13.188

9.47 823.4 -2.855 -719.5 2.495 14.97010 838.6 -2.908 -826.8 2.867 17.203

10.5 846.0 -2.934 -935.0 3.242 19.45411 846.6 -2.936 -1050.0 3.641 21.846

11.5 840.5 -2.915 -1171.7 4.063 24.37712 827.7 -2.870 -1300.1 4.508 27.049

12.5 808.1 -2.802 -1435.2 4.977 29.86113 781.7 -2.711 -1577.1 5.469 32.814

13.1 775.7 -2.690 -1606.3 5.570 33.421

Appendix G : Triple Bubble and Horizontal Ellipse Cross Section Analysis

The following is a detailed analysis of the triple bubble cross section under internal pressure loads. The figure below, shows the “weak” regions of the structure where the stress values are the highest. These high stresses indicate regions where the structure is likely to fail first. High tensile and compressive stresses are on the order of 10 MPa.

Figure 1: Von Mises Contour Plot of Triple Bubble Structure

Figure 2: Internal Pressure Loads on the Triple Bubble Configuration

In the two above figures, regions of green to blue color indicate compressive stresses while regions of yellow to red indicate tensile stresses. The circled regions indicate areas where reinforcements would need to be added. This reinforcements could negatively impact the weight of the structure.

Comparison of the horizontal elliptical cross section to the triple bubble configuration will follow:

Figure 3: Von Mises Contour Plot of Elliptical Cross Section

The above figure shows the Von Mises contour plot for an elliptical structure. The structure was dimensioned on a similar scale to the triple bubble and was tested under the same loading conditions. The high stress values can be noted on a magnitude of approximately 3 MPa. These stress values are less than the triple bubble section and thus carry the internal pressure loads more efficiently. The horizontal elliptical cross section was chosen as a result of this analysis.

Appendix H: Internal Structural Configuration and Optimization

The analysis of the internal structural configuration was an iterative process. The frame configurations were tested under the ultimate internal pressure load and a 6G bending load, determined in section 5.1.3. The tested configurations included 1) 45 degree inclined frames, 2) 30 degree inclined frames, 3) 20 degree inclined frames, 4) 20 degree inclined frames with decreased frame spacing, and 5) traditional circumferential frames. The following figures show the Von misses stresses on the different design configurations.

Figure 1: Internal structural configuration: 45 degree inclination

Little changed between the 30 and 40 degree frame inclination configurations. The maximum and minimum stress values remained on the same order of magnitude, however, regions of lowerstress increased. This can be viewed in the figure below where regions near the frames have stress values higher in magnitude than in the 45 degree configuration.

Figure 2: Internal structural configuration: 30 degree inclination

Figure 3: Internal structural configuration: 20 degree inclination

The 20 degree configuration gave much better results than the previously analyzed configurations. The average stress value was 0.4 GPa.

Figure 4: Internal structural configuration: 20 degree inclination with decreased spacing

When the frame density was increased, the stress values decreased further. The average stress value was around 0.3 GPa in the ultimate loading condition. Keeping in mind the strength to weight ratio of the structure, the decreased spacing between frames was not seen as beneficial enough to warrant the added weight in the structure. This iterative process resulted in finding the optimum frames inclination and spacing for the elliptical cross section.

Meshing and element quality were also considered during this analysis. The following two tablesshow the mesh and elements used in the internal structure analysis.

Figure 5: Mesh settings for structural analysis

Figure 6: Element quality settings for structural analysis

The number of elements and the quality depends on the capacity of the computer to process the analysis. The thickness of the skin is 10-12 mm thick so the appropriate element size would be around 2-3 mm . Due to a lack of computational power, this element size would be too small for a timely solution. Because the design analysis was based primarily on an iterative process the element size was increased to 10 mm in order to reduce the computing time necessary for a solution.

Figure 5: Mesh visualization of elliptical configuration

Appendix I: Bulkhead and Radome Figures

Figure 1: Radome and forward pressure bulkhead geometry5

Figure 2: Forward pressure bulkhead design5

Figure 3: High impact composite material layered on bulkhead membrane5

Appendix I: Joining and Bonding

Figure 1: Comparison of adhesive bonding versus welding