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A Comparison of an Integrally Stiffened Beaded Panel Made From Stretch Broken Carbon Fiber and a Honeycomb Core Sandwich...

Permanent Link: http://ufdc.ufl.edu/UFE0041283/00001

Material Information

Title: A Comparison of an Integrally Stiffened Beaded Panel Made From Stretch Broken Carbon Fiber and a Honeycomb Core Sandwich Panel
Physical Description: 1 online resource (59 p.)
Language: english
Creator: Nadel, Adam
Publisher: University of Florida
Place of Publication: Gainesville, Fla.
Publication Date: 2009

Subjects

Subjects / Keywords: aircraft, bead, broken, carbon, composite, core, corrosion, fiber, honeycomb, panel, sandwich, sbcf, skin, stretch
Mechanical and Aerospace Engineering -- Dissertations, Academic -- UF
Genre: Aerospace Engineering thesis, M.S.
bibliography   ( marcgt )
theses   ( marcgt )
government publication (state, provincial, terriorial, dependent)   ( marcgt )
born-digital   ( sobekcm )
Electronic Thesis or Dissertation

Notes

Abstract: Honeycomb core sandwich panels are commonly used in aircraft construction. They provide a buckling-resistant, lighter-weight alternative to the classic skin and stringer aircraft construction. However, corrosion and durability issues plague honeycomb sandwich panel construction. An alternative to both is a carbon composite beaded panel. The beads provide a geometric advantage to resist buckling and increased strength capability over flat panels of similar material thickness. Making carbon composite beaded panels has its own challenges associated with manufacture and laminate quality of such panels. Stretch Broken Carbon Fiber (SBCF) provides an opportunity for composite designs to use higher modulus material and more complex geometries with high laminate quality and stiffness, without the common manufacturing concerns often associated with continuous unidirectional carbon fiber. The formability of the material allows it to stretch and form complex geometries. SBCF is one of several types of materials commonly referred to as a discontinuous aligned fiber material. SBCF is capable of being quickly formed into beaded panels. The unique forming process removes the need to locally cut and overlap splice plies to conform to the geometry thereby reducing touch labor costs per unit, improving the laminate quality due to reduced ply bridging and removing the environmental concerns associated with honeycomb. Stretch Broken IM7 carbon fiber in an 8552 epoxy resin matrix has similar mechanical properties to the continuous IM7/8552. The ultimate purpose of this research was to show that an Integrally Stiffened Beaded Panel (ISBP) made of SBCF could replace a traditional honeycomb sandwich panel made with continuous IM7/8552 uni-directional carbon composite face sheets and HRH-36-.125-4.5 Kevlar Core. This construction was chosen as the baseline because it is being used on a Navy helicopter program currently in development.
General Note: In the series University of Florida Digital Collections.
General Note: Includes vita.
Bibliography: Includes bibliographical references.
Source of Description: Description based on online resource; title from PDF title page.
Source of Description: This bibliographic record is available under the Creative Commons CC0 public domain dedication. The University of Florida Libraries, as creator of this bibliographic record, has waived all rights to it worldwide under copyright law, including all related and neighboring rights, to the extent allowed by law.
Statement of Responsibility: by Adam Nadel.
Thesis: Thesis (M.S.)--University of Florida, 2009.
Local: Adviser: Sankar, Bhavani V.

Record Information

Source Institution: UFRGP
Rights Management: Applicable rights reserved.
Classification: lcc - LD1780 2009
System ID: UFE0041283:00001

Permanent Link: http://ufdc.ufl.edu/UFE0041283/00001

Material Information

Title: A Comparison of an Integrally Stiffened Beaded Panel Made From Stretch Broken Carbon Fiber and a Honeycomb Core Sandwich Panel
Physical Description: 1 online resource (59 p.)
Language: english
Creator: Nadel, Adam
Publisher: University of Florida
Place of Publication: Gainesville, Fla.
Publication Date: 2009

Subjects

Subjects / Keywords: aircraft, bead, broken, carbon, composite, core, corrosion, fiber, honeycomb, panel, sandwich, sbcf, skin, stretch
Mechanical and Aerospace Engineering -- Dissertations, Academic -- UF
Genre: Aerospace Engineering thesis, M.S.
bibliography   ( marcgt )
theses   ( marcgt )
government publication (state, provincial, terriorial, dependent)   ( marcgt )
born-digital   ( sobekcm )
Electronic Thesis or Dissertation

Notes

Abstract: Honeycomb core sandwich panels are commonly used in aircraft construction. They provide a buckling-resistant, lighter-weight alternative to the classic skin and stringer aircraft construction. However, corrosion and durability issues plague honeycomb sandwich panel construction. An alternative to both is a carbon composite beaded panel. The beads provide a geometric advantage to resist buckling and increased strength capability over flat panels of similar material thickness. Making carbon composite beaded panels has its own challenges associated with manufacture and laminate quality of such panels. Stretch Broken Carbon Fiber (SBCF) provides an opportunity for composite designs to use higher modulus material and more complex geometries with high laminate quality and stiffness, without the common manufacturing concerns often associated with continuous unidirectional carbon fiber. The formability of the material allows it to stretch and form complex geometries. SBCF is one of several types of materials commonly referred to as a discontinuous aligned fiber material. SBCF is capable of being quickly formed into beaded panels. The unique forming process removes the need to locally cut and overlap splice plies to conform to the geometry thereby reducing touch labor costs per unit, improving the laminate quality due to reduced ply bridging and removing the environmental concerns associated with honeycomb. Stretch Broken IM7 carbon fiber in an 8552 epoxy resin matrix has similar mechanical properties to the continuous IM7/8552. The ultimate purpose of this research was to show that an Integrally Stiffened Beaded Panel (ISBP) made of SBCF could replace a traditional honeycomb sandwich panel made with continuous IM7/8552 uni-directional carbon composite face sheets and HRH-36-.125-4.5 Kevlar Core. This construction was chosen as the baseline because it is being used on a Navy helicopter program currently in development.
General Note: In the series University of Florida Digital Collections.
General Note: Includes vita.
Bibliography: Includes bibliographical references.
Source of Description: Description based on online resource; title from PDF title page.
Source of Description: This bibliographic record is available under the Creative Commons CC0 public domain dedication. The University of Florida Libraries, as creator of this bibliographic record, has waived all rights to it worldwide under copyright law, including all related and neighboring rights, to the extent allowed by law.
Statement of Responsibility: by Adam Nadel.
Thesis: Thesis (M.S.)--University of Florida, 2009.
Local: Adviser: Sankar, Bhavani V.

Record Information

Source Institution: UFRGP
Rights Management: Applicable rights reserved.
Classification: lcc - LD1780 2009
System ID: UFE0041283:00001


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A COMPARISON OF AN INTEGRALLY ST IFFENED BEADED PANEL MADE FROM STRETCH BROKEN CARBON FIBER AND A HONEYCOMB CORE SANDWICH PANEL By ADAM IAN NADEL A THESIS PRESENTED TO THE GRADUATE SCHOOL OF THE UNIVERSITY OF FLORIDA IN PARTIAL FULFILLMENT OF THE REQUIREMENTS FOR THE DEGREE OF MASTER OF SCIENCE UNIVERSITY OF FLORIDA 2009 1

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2009 Adam Ian Nadel 2

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To my Grandfather who inspired me to be an Engineer and started my love of aircraft 3

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ACKNOWLEDGMENTS I would like to acknowledge the support of the Naval Air Warfare Center (NAVAIR) and Dr. Raymond Meilunas. This research was completed under contract N00421-08C-0017. I would like to acknowledge my employ er, Aurora Flight Science, of Manassas Virginia for giving me the opportunity to conduct this study. I would also like to acknowledge the contract pr ime, Hexcel Corporation, and particularly Principal Investigator, Dr. Gnther Jacobsen. I woul d also like to acknowledge Dr. Gregory Dillon and Mr. Don H. Stiver both of Penn State Applied Research Laboratories and Mr. David Cox of Aurora Flight Sciences for their colla borative contributions to this effort. I would like to thank my advisor Dr. Bhav ani Sankar for his continued support of me. His patience and guidance has always been appreciated. Finally I would like to thank my parents, sister and girlfriend. Without their love and backing none of this would have been possible. 4

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TABLE OF CONTENTS Page ACKNOWLEDG MENTS .................................................................................................. 4 LIST OF TABLES ............................................................................................................ 7 LIST OF FI GURES .......................................................................................................... 8 LIST OF ABBR EVIATION S ........................................................................................... 10 ABSTRACT ................................................................................................................... 11 CHAPTER 1 INTRODUC TION .................................................................................................... 13 Honeycomb Sandw ich Panels ................................................................................ 13 Integrally Stiff ened Beaded Pane ls ......................................................................... 16 Bead Pattern Selectio n ........................................................................................... 20 2 MANUFACTURING PROCESS SUMMARY ........................................................... 24 Honeycomb Sandw ich Panels ................................................................................ 24 Integrally Stiff ened Beaded Pane ls ......................................................................... 24 3 STRUCTURAL ANALYSIS ..................................................................................... 28 Material Pr operties .................................................................................................. 28 General Model Set-Up ............................................................................................ 29 Boundary Conditi ons Loads .......................................................................... 31 Boundary Conditions Constraints .................................................................. 35 Honeycomb Sandwich Panel Layups .............................................................. 37 SBCF ISBP Lay-up ........................................................................................... 39 4 RESULT S ............................................................................................................... 42 Deflecti ons .............................................................................................................. 42 Honeycomb Sandw ich Panels .......................................................................... 42 Integrally Stiff ened Beaded Pa nel .................................................................... 44 Ply Fail ure ............................................................................................................... 45 Honeycomb Sandw ich Panels .......................................................................... 46 Integrally Stiff ened Beaded Pa nel .................................................................... 47 Discussio n .............................................................................................................. 48 5 CONCLUS IONS ..................................................................................................... 50 APPENDIX A: MECHANICA L PROPERTI ES ............................................................... 52 5

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LIST OF RE FERENCES ............................................................................................... 57 BIOGRAPHICAL SKETCH ............................................................................................ 59 6

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LIST OF TABLES Table page 3-1 Mechanical proper ties of IM7/8552 ..................................................................... 29 3-2 Material properties of HRH36-.125-4.5 Honeyco mb Core ................................. 29 3-3 8 Ply Lay-up Sequence for Honeycomb Face Sheet s ........................................ 37 3-4 16 Ply Lay-up Sequence for Honeycomb Face Sheets ...................................... 38 3-5 8 Ply lay-up sequence for the beads .................................................................. 39 3-6 8 Ply lay-up sequence under beads ................................................................... 39 3-7 16 Ply lay-up sequence of areas bet ween beads ............................................... 40 3-8 16 Ply lay-up sequence of areas surroundi ng beads .......................................... 40 4-1 Weight of Panels ................................................................................................ 42 4-2 Summary of panel deflecti ons ............................................................................ 42 4-3 Summary of maximum ply failure index .............................................................. 45 7

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LIST OF FIGURES Figure page 1-1 Structural benefits of h oneycomb sandwich panel .............................................. 13 1-2 Corrosion of an Aluminum honeycomb sandwich panel fr om an HH60J ............ 15 1-3 Impact damaged honeycomb panel .................................................................... 16 1-4 Graphical represent ation of SBCF ...................................................................... 17 1-5 Original SB CF Mach ine ...................................................................................... 19 1-6 Panel made of SBCF with 11 bea ds ................................................................... 20 1-7 Geometry of ISBP ............................................................................................... 22 1-8 Alternate bead c onfigurati ons ............................................................................. 23 2-1 Manufacturing Process ....................................................................................... 25 2-2 Bead geometry shown with broken fibers ........................................................... 26 2-3 ISBP mold tool .................................................................................................... 26 2-4 Result of an 8 ply forming tria l at Penn State ARL sho wing ............................... 27 3-1 Pictures of Finite Element Models ...................................................................... 30 3-2 Honeycomb Sandwich Panel Applie d Loading ................................................... 33 3-3 ISBP Applied Loading ......................................................................................... 34 3-4 Honeycomb Sandwich Panel Constr aints ........................................................... 36 3-5 ISBP Constr aints ................................................................................................ 37 3-6 Lay-up Regions of Ho neycomb Core Panel ....................................................... 38 3-7 Lay-up regions of the SBCF ISBP. ..................................................................... 41 4-1 Honeycomb Panel Deflection for Load Case 1 Outwar d Pressure .................. 43 4-2 Honeycomb Panel Deflection for Load Case 2 Inward Pressure ..................... 43 4-3 Honeycomb Panel Deflecti on for Load Case 3 Shear ..................................... 44 8

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4-4 SBCF ISBP Deflection for Load Case 1 Outwar d Pressure ............................. 44 4-5 SBCF ISBP Deflection for Load Case 2 Inward Pressure ............................... 45 4-6 SBCF ISBP Deflection for Load Case 3 Shear ................................................ 45 4-7 Honeycomb Panel Max Failure Index fo r Load Case 1 Outward Pressure ...... 46 4-8 Honeycomb Panel Max Failure Index fo r Load Case 2 Inward Pressure ........ 46 4-9 Honeycomb Panel Max Failure I ndex for Load Case 3 S hear ......................... 47 4-10 SBCF ISBP Max Failure Index for Load Case 1 Outwar d Pressure ................ 47 4-11 SBCF ISBP Max Failure Index for Load Case 2 Inward Pressure ................... 48 4-12 SBCF ISBP Max Failure Index for Load Case 3 Shear ................................... 48 5-1 Scaled up test piece shown the cross pattern multiplied .................................... 51 9

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LIST OF ABBREVIATIONS SBCF Stretch Broken Carbon Fiber ARL Pennsylvania State University Applied Research Laboratory ISBP Integrally Stiffened Beaded Panel OML Outer Mold Loft/Line/Layer IML Inner Mold Loft/Line/Layer SBS Short Beam Shear NIAR Wichita State University Nati onal Institute of Aviation Research FEA Finite Element Analysis FEM Finite Element Model SPC Single Point Constraint GPa Giga-Pascal MPa Mega-Pascal msi Million pounds/square inch ksi Thousand pounds/square inch Poisson Ratio E Elastics Modulus G Shear Modulus F1t/c Tension or compression strength of a material in the direction F2t/c Tension or compression strength of a material in the direction F12 Shear strength of a material RTD Room Temperature Dry DOF Degree of Freedom Measure of Angle in degrees relative to a defined Coordinate System 10

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Abstract of Thesis Pres ented to the Graduate School of the University of Florida in Partial Fulf illment of the Requirements for t he Degree of Master of Science A COMPARISON OF AN INTEGRALLY ST IFFENED BEADED PANEL MADE FROM STRETCH BROKEN CARBON FIBER AND A HONEYCOMB CORE SANDWICH PANEL By Adam Ian Nadel December 2009 Chair: Bhavani Sankar Major: Aerospace Engineering Honeycomb core sandwich panels are comm only used in aircraft construction. They provide a buckling-resistant, lighter-w eight alternative to the classic skin and stringer aircraft construction. However, corrosion and durability issues plague honeycomb sandwich panel construction. An alternative to both is a carbon composite beaded panel. The beads provide a geometric advantage to resist buckling and increased strength capability over flat panels of similar material thickness. Making carbon composite beaded panels has its own challenges associated with manufacture and laminate quality of such panels. Stretch Broken Carbon Fiber (SBCF) provides an opportunity for composite designs to use higher modulus material and more complex geometries with high laminate quality and stiffness, without t he common manufacturing concerns often associated with continuous unidirectional car bon fiber. The formability of the material allows it to stretch and fo rm complex geometries. SBCF is one of several types of materials commonly referred to as a discontinuous aligned fiber material. 11

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12 SBCF is capable of being quickly form ed into beaded panels. The unique forming process removes the need to locally cut and overlap splice plies to conform to the geometry thereby reducing touch labor cost s per unit, improving the laminate quality due to reduced ply bridging and removing the environmental concerns associated with honeycomb. Stretch Broken IM7 carbon fiber in an 8552 epoxy resin matrix has similar mechanical properties to the continuous IM 7/8552. The ultimate purpose of this research was to show that an Integrally Stiffened Beaded Panel (ISBP) made of SBCF could replace a traditional honeycomb s andwich panel made with continuous IM7/8552 uni-directional carbon composite face sheet s and HRH-36-.125-4.5 Ke vlar Core. This construction was chosen as the baseline bec ause it is being used on a Navy helicopter program currently in development.

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CHAPTER 1 INTRODUCTION 1 Chapter 1 Honeycomb Sandwich Panels Sandwich panel construction uses thin, l oad-bearing skins or face sheets bonded to a light weight honeycomb core or foam. The face sheets can be made from metal or a carbon fiber/epoxy resin material system. This type of construction has excellent stiffness and a high strength to weight ra tio, and therefore has been used in many airframe applications for fix ed and rotary wing aircraft. Figure 1-1 shows a comparison and benefits of a honeycomb sandwich panel over a solid sheet of a given thickness [1]. Figure 1-1. Structural benefits of honeycomb sand wich panel. Source: Hexcel Composites HexWeb Honeycomb Sandwich Design Technology 2000. Generally the honeycomb core carries the out-of-plane or transverse shear forces on the panel. The in-plane and bending loads are carried through membrane loads in 13

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the face sheets. The honeycomb sandwich panel can be modeled as two dimensional if the shear deflection is neglig ible. Modeling the honeycomb core with three dimensional elements is also possible. The honeycomb co re should be given an effectively zero inplane modulus. FEA requires that some val ue be given for the elastic modulus. The shear modulus are given based on the ribbon and warp directions of the core. Those directions are defined by the way core is manufactured [1]. Honeycomb sandwich structures have proven costly to maintain due to limited durability. Perhaps the largest single li mitation has been corrosion. The Navy and Coast Guard operate their aircraft on the open seas in one of the harshest aircraft environments known. In addition to direct salt spray exposure, the maritime carrier aircraft environment features ex posure to sustained high leve ls of humidity. Although improvements in aluminum core have been made, aluminum honeycomb does not withstand the naval aviation environment and has been all but eliminated from new designs due to high recurring maint enance costs to repair corroded and damaged panels. Figure 1-2 shows a picture of a corroded al uminum honeycomb core sandwich panel from a Coast Guard HH60J helicopter [2]. 14

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Figure 1-2. Corrosion of an aluminum honey comb sandwich panel from an HH60J. Source: Ted Wiesner HH-60J Seahawk Helicopter Corrosion Issues Army Corrosion Summit 2003 http://www.armycorrosion.com/past_su mmits/summit2003/download1.cfm?fn ame=wiesner.pdf (Accessed 10/2009) Non-metallic core such as HRH-10 Nomex or HRH-36 Kevlar/Korex do not corrode directly, however they remain suscept ible to other failure modes. The very weight saving benefit (thin, highly loaded, structurally efficient skins) offered by sandwich construction also has some weakness. These thin skins are more susceptible to impact and puncture damage. This can lead to moisture and fluid intrusion, which can then degrade the adhesive bond lin e and/or the honeycomb core. Navy rotary wing aircraft sandwich structures are particularly susceptible to this damage, because many sandwich panels feature very thin, minimum gage skins. The operating limitations of carrier based and forward based Marine Corps ai rcraft results in a higher probability of impact to the skins of the aircraft. 15

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A B Figure 1-3. Impact damaged honeycomb panel A) Whole honeycomb sandwich panel B) Close up of impact indentation Corrosion, impact damage, manufacturing complications and other issues can lead to a debond between a face sheet and the honeycomb core. The debonding of the face sheet can cause the panel to buckle under compressive, compression due to bending, and shear loads. Sankar and Nara yanan investigated the buckling of a honeycomb sandwich panel with debonded face sheets under compressive loads using Finite Element Analysis (FEA) and experimentati on. They characterized the failure of the core and face sheets [3]. The geometries that can be construct ed using honeycomb sandwich panels are also limited. Honeycomb sandwich panels can not have high curvature even with fairly constant thickness face sheets. The very thickness of the honeycomb that provides the stiffness and strength benefit hinder construction of a curved panel. Processes have been developed to form honeycomb core into hi gher curvature parts; however they are extremely expensive and time consuming. Integrally Stiffened Beaded Panels Bead-stiffened panels are frequently used in metallic construction, and have been used in the past for light to moderately loade d composite structures. However, bead 16

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stiffened fabrication using conventional composite manufacturing methods poses significant challenges. Typically, bead geom etries feature compound curvature with tight radii. These features ar e difficult, if not impossible, to form using conventional unidirectional tape or fabric prepreg without unacceptabl e wrinkles. In many cases, the pre-impregnated carbon fiber material is locally cut or overlap spliced to provide forming compliance, which requires additional overlappi ng and patches to re store the necessary laminate strength. This adds weight as well as labor cost. With co nventional ply-by-ply manual lay-up methods, every ply must be labori ously and skillfully laid up into multiple compound curvature bead cavities. The formability of SBCF material in the pre-impregnated ca rbon fiber state significantly simplifies the manufacture of parts like ribs, skins and bulkheads with integral stiffening beads. The process not only requires less hand labor than ply-by-ply manual lay-up into a mold, but also ultimately facilitates automation, since the flat ply lay-up can readily be performed using autom ated tape lay-up. SBCF is made from standard continuous carbon fiber. A graphical representation of SBCF is shown below in Figure 1-4 Figure 1-4. Graphical representation of SBCF 17

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The filaments of tow ar e stretched and broken between two revolving wheels spinning at different speeds. Each filam ent stretches and breaks between the pinch points of the two rollers at its weakest point. The average break length is determined by the distance between the pinch points. The load is transferred between adjacent filaments by shear at the break points. The process by which the material is made is detailed for the original AS4/M73 material system in Reference 4. The original machine which produced the AS4/M73 material is shown below in Figure 1-5 The machine was originally used by the textile industry to stretch materi als such as cotton before being made into clothes or other products. T he IM7/8552 material is produced on a newer machine. 18

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Two pinch point locations between small rollers Wheels spinning at progressively faster speeds Figure 1-5. Original SBCF Machine. S ource: Abdallah, M.G., Hansen, N. W. and Jacobsen G. Development of Stretc h Broken Carbon Fiber Materials. Proceedings of the 50th International SAMPE Symposium and Exhibition, Long Beach, California, May 1-5, 2005Jacobsen, G. Mechanical Performance Characterization Of Stre tch Broken Carbon Fiber Material., Proceedings from the 54t h SAMPE Technical Conf erence, Baltimore, Maryland, May 18 21, 2009. While the ability to make single surfac e bead-stiffened parts like ribs, spars and bulkheads is useful, typical stressed skin ai rframe structure requires internal stiffened skins, in which the exterior side of the ski n panel is smooth and continuous. The bead stiffened panel concept is structurally e fficient, impact resist ant and lightweight. Pennsylvania State University Applied Research Lab (ARL) has developed a concept for manufacturing an integrally bead-sti ffened panel which consists of a smooth, continuous outer skin, as required on an aircraft outer mold line (OML), and a beaded inner skin. This method relies upon the formability of SBCF pre-impregnated carbon 19

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fiber material, and a vacuum forming proce ss to form the beaded inner panel. The cured inner panel is then bonded to a cured out er skin using film adhesive. Figure 1-6 below shows a panel made of SBCF with 11 beads [5]. The panel was made as a demonstration for a helicopter; each bead is appr oximately 26.7 cm (10.5 in) and 6.4 cm (5.25) across. Figure 1-6. Panel made of SBCF with 11 bea ds. Source: Dillon, G. and Stiver, D.H. III. Development Of Enabling Auto mated Forming Technology For Stretch Broken Carbon Fiber (SBCF) Materials, Proceedings from the 54th SAMPE Technical Conference Baltimore, Maryland, May 18 21, 2009 Bead Pattern Selection Rotary wing aircraft and fixed wing aircra ft designers and fabricators recognize the potential offered by an integrally stiffened unitized structure. The basic intent was to design an effective unit cell of reinforced stru cture that might ultima tely be extended as a series of repeat units to stiffen a large area or replace a group of panels with one more efficient piece. This is consistent with the building block approach to the process development that has been successfully appli ed in prior phases of this work. The 20

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selected bead pattern would act similar to a traditional honeycomb sandwich panel by providing stiffness and strength in multiple directions. The design geometry was also chosen to showcase and evaluate the materials ability to form. This ISBP design select ed produced an inner skin with a complex 25.4 mm (1.0 in) deep bead geometry designed to carry a load similar to a 25.4 mm (1.0 in) thick honeycomb panel. This panel has bead st iffeners in the 0, 90, +45, and -45 directions to compliment the selected quasi-isotropic lay up and give it the universal load capacity of the traditional honeycomb panel. See Figure 1-7 for the geometry of the replacement beaded panel. 21

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A B Figure 1-7. Geometry of I SBP. A) Dimensioned drawing of ISBP B) Isometric view of ISBP 22

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The selected geometry and lay-ups chosen in this research have not been optimized. The selection of bead patterns and ply lay-ups can be tailored to the loading of specific components thereby reducing weight further. Figure 1-8 shows two panels which are designed to carry load in a specific direction: A) load in one direction along the beads B) shear only load. A B Figure 1-8. Alternate bead configurations A) Loading in one direction B) Loading primarily in shear 23

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CHAPTER 2 MANUFACTURING PROCESS SUMMARY 2 Chapter 2 Honeycomb Sandwich Panels A common way of manufacturing honeycomb sandwich panels is to lay-up the under core face sheet on a tool matching the OML of the aircraft. The core is then placed on the under core laminate and then the ov er core lay-up is put down. Vacuum is drawn at several steps during this process to compact the laminate. This process is normally done by hand in a process called hand lay-up. The locations of ply drop-offs under and over the core need to be coordinated to assure that the co re does not bridge over the change in face sheet thickness and t hat the core can conf orm to the geometry. Several processes including heat forming the honeycomb core before placement and machining the honeycomb core are possible; however both are expensive and can lead to a higher scrap rate of the honeycomb co re detail. The part is then cured. Honeycomb core is also susceptible to cr ushing during an autoclave cure process. Integrally Stiffened Beaded Panels The following is a summary of the manufac turing process developed at Penn State ARL by which an ISBP can be made [5]. The entire inner skin laminate is laid up on a flat tool. The lay-up is then transferred ont o a pre-heated tool shaped to the inner mold line (IML) of the final part. Vacuum is drawn on the lay-up forcing the laminate into the female mold of the tool. Vacuum holes in the tools assist the forming. The heat from the tool causes the resin in the laminate to becoming less viscous allowing each of the SBCFs to elongate along its ax is by enlarging the gaps bet ween the filament segments, allowing the fibers to conform to the mold. A graphic descripti on of this process is given in Figure 2-1 A separate tool that is shaped to t he OML of the part is used to form the 24

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outer skin. The outer skin is generally a smooth shape for aerodynamic purposes. It can be made from SBCF or continuous uni-dir ectional carbon fiber. The two skins are then bonded together in a second cure to form the final part. B A C Figure 2-1. Manufacturing Proc ess. A) Inner Skin Laminate Lay-up Flat B) IML Tool C) Laminate drawn into tool forming IML Surface of Part Figure 2-2 below shows the geometry of the bead shape. There is also a notional stretch broken fiber shown. As the bead is wi der than 5.1 cm (2.0 in) there are at least 25

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two breaks of the fiber across the bead. The tool used to create the panel is shown in Figure 2-3 One break Second break 76.20 mm [ 3.00 in ] R 31.75 mm Individual fibe r R 9.525 mm [ R 0.375 in ] Figure 2-2. Bead geometr y shown with broken fibers Figure 2-3. ISBP mold tool Forming trials began with 4 plies [0/90/+45/ -45] of SB IM7/8552 uni-directional prepreg tape. It only took thr ee iterations to make the n eeded modifications to the cure tool and to select ply orientation before mo ving to the quasi isotropic 8-ply skin. Three additional 8-ply iterations were done wit h positive results. The part shown in Figure 3-2 26

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was created via the process developed by Penn State ARL [5]. A B Figure 2-4. Result of an 8 ply forming tria l at Penn State ARL showing A) Tool side surface B) Bag side surface 27

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CHAPTER 3 STRUCTURAL ANALYSIS 3 Chapter 3 To compare the honeycomb sandwich pane ls and the Integrally Stiffened Beaded Panels finite element models of both were made. FEMAP v9.31 was used as the pre and post processor. NEi Nastran v9.2 was used as the solver. Material Properties Testing on cured IM7/8552 panels made fr om SBCF has shown that the SBCF material has statistically equivalent modulus and strength to the continuous fiber materials in almost all categories. The details about this SBCF fiber and prepreg are found in Reference 6. This testing was comp leted at NIAR under contract to Hexcel. The material properties for t he IM7/8552 are shown below in Table 3-1 [7 & 8]. The only property not tested and documented in Reference 7 is the 90 compressive strength. That proper ty was taken from the 8552 Resin Data Sheet [8]. The properties are taken from the SBCF properties and used fo r both the continuous and stretch broke material. This was done because the differenc es are statistically insignificant and only one sample for the continuous material wa s tested. The density for the carbon fiber/epoxy resin system is 2.06 kg/m3 (0.057 lbs/in3). The material properties of the HRH-36-.125-4.5 Kevlar honeycomb core with 3.2 mm (0.125 in) cell size weighing 72 kg/m3 (4.5 lbs/ft3) are shown below in Table 3-2 The baseline material properties [9] provided by the manufacturer ar e for 12.7 mm (0.5 in) thick core. The core used is 25.4 mm (1.0 in) thick which requires a reduction of 12% based on the of the core thickness [10]. The values include the knockdown. All properties are r oom temperature dry (RTD). 28

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Table 3-1. Mechanical proper ties of IM7/8552. Sources: Jacobsen, G. Mechanical Performance Characterization of Stre tch Broken Carbon Fiber Material, Proceeding from the 54th SAMPE Technical Conference Baltimore, Maryland, May 18-21 2009 and Hexcel Cor poration, HexPly 8552 Epoxy Matrix Product Data 2005 Property Value E11 163GPa (23.64 msi) E22 8.9 GPa (1.29 msi) G66 4.63 GPa (672 ksi) 12 0.34 F1t 2681 MPa (389.9 ksi) F1c 1665 MPa (241.5 ksi) F2t 72.2 MPa (24.5 ksi) F2c 304.7 MPa (44.2 ksi) F12 92.6 MPa (13.4 ksi) Table 3-2. Material properties of HRH-36-.125-4.5 Honeycomb Core. Sources: Hexcel Corporation, HexWeb HRH-36 Product Data 2004 and Hexcel Composites, TSB120 Hexweb Honeycomb Attributes and Properties 1999 Property Value Gxz 200 MPa (29.0 ksi) G y z 77.2 MPa (11.2 ksi) General Model Set-Up The composite laminates for bot h the honeycomb panel and the beaded panel were modeled using CTRIA3 and CQUAD4 elements. Those elements are isoparametric membrane-bending plat e elements with three and f our nodes respectively. The properties were applied by the use of a PCOMP card. The material properties were assigned using a MAT8 card. The core was modeled using solid elements; CHEX8 elements for the full thickness areas of the core, CWEDGE6 elements for the core ramp areas and CTET4 elements for the ramp corners. Those element s are iso-parametric solid elements with eight, six, and four nodes respectively. This corresponds to the edges of the core being machined to an angle of 20.0. Figure 3-1 shows views of the two Finite Element Models used for the analysis. 29

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A B Figure 3-1. Pictures of Finite Element Models A) ISBP Panel B) Honeycomb Sandwich Panel According to the testing completed by Hexcel the properties of continuous and stretch broken IM7/8552 are statistically identica l [7]. Therefore, the same material card was used to model the continuous and Stre tch Broken carbon fiber material forms. Thinning is experienced by the panels during tr ial forming. Several measurements were taken from the test panel and it was measured that the reduction in thickness for the laminate was up to approximately 12% [11]. For a conservative analysis this reduction 30

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was increased to 15% in ply thickness. T he nominal ply thickness is 0.1524 mm [0.006 in]. The corresponding reduced ply th ickness is 0.1295 mm [0.0051 in]. No shear and bending interaction is desired. Therefore, a symmetric is desired for this application. Furthermore, it was desir ed to have a quasi-isotropic lay-up. This reduces the effect of a fi ber misalignment which could further reduce manufacturing costs. An 8-ply lay-up is the minimum r equired number of plies to achieve a symmetric, quasi-isometric lay-up using uni-directional fi bers. This is therefore considered the minimum gauge for both the honeycomb sandwich panel face sheets and the beaded panel. Other criteria may be imposed which would increase the minimum thickness of the laminate further. Those are not considered as they are assumed to apply to both panels equally. Tsai-Hill Failure theory was used to com pute ply failure index. The equation of Tsai-Hill Failure Inde x is shown below as Equation 1 When the Failure Index exceeds 1, the material has failed [12]. F1, F2, and F12, are the material strength allowables in the specified material direction. The allowables are shown above in Table 3-1 The tension allowable is used if the calculate stress in the ply in that direction is in tension. For example if 1 is positive then F1 is F1t. NEi Nastran v9.2 will calculate a failure index of a given ply for the user. Th is feature was used to calculate the failure of the plies. 2 12 2 12 2 1 21 2 2 2 2 2 1 2 1FFFF FI Equation 1. Tsai-Hill Failu re Theory Equation [12] Boundary Conditions Loads Three load cases were modeled to simulate loads an aircraft skin panel would experience during flight. They are: 31

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Outward Pressure of 10,342 Pa [1.5 psi] Inward Pressure of 10,342 Pa [1.5 psi] Shear load of 609.6 N/mm [520.8 lbs/in] The over core honeycomb side and the beaded side are considered the inner surface (IML). The opposing flat surfac e is considered the out er surface or the aerodynamic surface (OML). Load Case 1 was applied against the inner surface outward. Load Case 2 was applied against the outer/aerodynamic surface and inward. The shear forces were applied in-plane. Th e pressure loads were applied with PLOAD4 cards which are pressure loads applied to th e faces of the plate elements. The shear forces were applied at 16 locations around the perimeter. The corner locations had shear loads for both edges applied to them. This was done to simulate the transfer of shear loads across a skin from a piece of mating structure to another such as from aircraft frame to aircraft frame along the leng th of a fuselage. The shear loads were applied with FORCE cards which are static forces applied at a node in a specified direction. The loads are applied around the perimeter of the panel s where the OML and IML faces come together. The loading of the Honeycomb Sandwich Panel is shown below in Figure 3-2 The loading of the ISBP is shown below in Figure 3-3 32

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A B C Figure 3-2. Honeycomb Sandwich Panel Applied Loading A) Outward Pressure B) Inward Pressure C) Shear 33

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A B C Figure 3-3. ISBP Applied Loading A) Outward Pressure B) Inward Pressure C) Shear 34

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Boundary Conditions Constraints Single Point Constraints (SPC) were used to constrain the models. These constraints fix the node it is applied upon in the specifie d degree of freedom (DOF). The constrained DOF is represented by a num ber; 123 for the XY&Z translational DOFs respectively and 456 for the rotational DOFs respectively. The coordinate axes are shown above in Figure 3-2 and Figure 3-3 for the two panels. The locations of the constrained nodes in the model are consist ent with probable fastener locations of such a panel. The constraints are used to proper ly constrain the model or remove the singularity inherent in the finite element analysi s. The constraints were chosen to match the intended load paths for each load case. For the pressure load cases it was assum ed each fastener location will react shear and axial loads. This is due to the fact t hat the pressure load would need to be carried into the mating structure by the fasteners. Therefore, the three translational degrees of freedom were constrained. The SPC locations for load cases 1 and 2 are identical. For load case 3, the shear load case, the fastener locations were constrained in the out-of-plane, or fastener axial load capability. The in-plane fastener constraints (1&2) were not applied as the loading is app lied in those directions simulating the transfer of load from one aircraft frame to an other aircraft frame. The center of the panel was also constrained in the 1, 2, and 5 DO Fs to construct a statically determinant model. As the reaction loads at that point is zero, these constraints do not affect the FEA solution. The SPC locations for can be seen below as blue triangles with the constrained degrees of freedom specified. Figure 3-4 shows the constraints for the Honeycomb 35

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Sandwich Panel. Figure 3-5 shows the constraints on t he SBCF ISBP. Both figures show the OML of the panels. A B Figure 3-4. Honeycomb Sandwich Panel Cons traints A) Pressure load cases (#1&2) B) Shear load case (#3) 36

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A B Figure 3-5. ISBP Constraints A) Pressure load cases (#1&2) B) Shear load case (#3) Honeycomb Sandwich Panel Lay-ups The honeycomb panel was modeled to have 25.4 mm thick HRH36-.125-4.5 Kevlar core. The face sheets and the co re shared nodes to transfer the load between the two. The face sheets above and below the core were modeled with the minimum 8ply quadi-isometric lay-up. The honeycomb panel would use continuous uni-directional carbon fiber. The nominal thickness of ma terial was used. The lay-up sequence is shown below in Table 3-3 The lay-up is shown as blue in Figure 3-6 below. Table 3-3. 8 Ply Lay-up Sequence for Honeycomb Face Sheets # of Plies Material Thickness Orientation () 8 IM7/8552 0.1524 mm (0.006 in) [0/90/45/-45]s The area surrounding the core was model ed with a 16 Ply lay-up. This corresponds to the two face sheets comi ng together around the core. This lay-up 37

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sequence is shown below in Table 3-4 The lay-up is shown below as orange in Figure 3-6 Figure 3-6 also has a cut-away of the h oneycomb sandwich panel showing the core and face sheets. The core is shown as gray. Table 3-4. 16 Ply Lay-up Sequence for Honeycomb Face Sheets # of Plies Material Thickness Orientation () 16 IM7/8552 0.1524 mm (0.006 in) [0/90/45/-45]2sA) B) Figure 3-6. Lay-up Regions of Honeycomb Core Panel A)Lay-up Regions of the Honeycomb Sandwich Panel B) Cut-away of honeycomb sandwich panel Area surrounding core Shared nodes between face sheets Nominal thickness face sheets Honeycomb Core 38

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SBCF ISBP Lay-up The beaded panel was modeled with 25.4 mm tall beads to match the same geometric profile as the honeycomb panel it is intended to replace. The beads have an approximate width of 86.5 mm. See Figure 1-7 for detailed dimensions. The beads are assumed to be hollow. The beaded areas were modeled with the minimum 8-ply quasiisometric lay-up. To account for the thi nning due to stretching of the fibers during the forming process the ply thickness was reduced by 15% [11]. The lay-up is shown below in Table 3-5 The lay-up is shown below as gray in Figure 3-7 Figure 3-7 also has a cut-away view of a bead with the thi cknesses of the laminates shown. Table 3-5. 8 Ply lay-up sequence for the beads # of Plies Material Thickness Orientation () 8 IM7/8552 0.1295 mm (0.0051 in) [0/90/45/-45]s The area under the bead is un-stretched and could be made from continuous material. It has an 8 quasiisometric ply la y-up, too. That la y-up sequence is shown below in Table 3-6 The lay-up is shown below as blue in Figure 3-7 Table 3-6. 8 Ply lay-up sequence under beads # of Plies Material Thickness Orientation () 8 IM7/8552 0.1524 mm (0.006 in) [0/90/45/-45]s The triangular area between the beads is modeled with a 16-ply lay-up. This corresponds to the combination of both t he beaded sheet and the flat outer sheet. The 8 plies of the beaded side area also stretched due to the forming process. Therefore the top 8 plies of the lay-up have their th ickness reduced by 15%. The lay-up sequence can be seen below in Table 3-7 The lay-up is shown below as green in Figure 3-7 39

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Table 3-7. 16 Ply lay-up sequence of areas between beads # of Plies Material Thickness Orientation () 8 IM7/8552 0.1295 mm (0.0051 in) [0/90/45/-45]s 8 IM7/8552 0.1524 mm (0.006 in) [0/90/45/-45]s The area surrounding the beads was model ed with a 16-ply lay-up. This corresponds to the beaded sheet and flat sheet coming together around the beads similar to the area surrounding the honeycomb core. The forming process can pinch the plies and prevent thinning due to elongation. Each indi vidual ply can slip on the tooling or the adjacent plies which may also prevent thinning due to the elongation. The plies are therefore all at their nominal thic kness. This lay-up sequence is shown below in Table 3-8 The lay-up is shown below as orange in Figure 3-7 Table 3-8. 16 Ply lay-up sequence of areas surrounding beads # of Plies Material Thi ckness (mm) Orientation () 16 IM7/8552 0.1524 mm (0.006 in) [0/90/45/-45]2s 40

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A B C Figure 3-7. Lay-up regions of the SBCF ISBP A) I nner/Beaded Surface B) Outer Surface C) Cut-away of bead. Area surrounding beads Under Bead Nominal Thickness Between Beads Thinned and Nominal Thickness Bead Thinned Thickness Shared Nodes Between Laminates 41

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CHAPTER 4 RESULTS 4 Chapter 4 The results for the two panels are compar ed below. Failure modes which would occur in a similar fashion on either panel are not considered. These failure modes include failure of t he film adhesive, fastener bearing, and inter-laminar shear failures among others. The weights of the panels are given below in Table 4-1 The weight of film adhesive and other common materials is not included as it is assumed that both panels would have a similar am ount of those materials an d would not influence the weight difference. Table 4-1. Weight of Panels Panel Weight Honeycomb Core 1.908 kg (4.207 lbs) SBCF ISBP 1.432 kg (3.158 lbs) Deflections The following sections summarize the def lections of the two panels. Their deflections are summarized below in Table 4-2 Table 4-2. Summary of panel deflections Load Case Honeycomb Panel SBCF ISBP Outward Pressure 0.655 mm (0 .0258 in) 3.785 mm (0.149 in) Inward Pressure 0.638 mm (0.0251 in) 3.581 mm (0.141 in) Shear 1.834 mm (0.0722 in) 1.626 mm (0.064 in) Honeycomb Sandwich Panels The total magnitude deflection (A) and Z defle ction (B) for the three load cases are shown below in Figure 4-1 Figure 4-2 and Figure 4-3 respectively. The deflection is shown at 10 times the actual. All deflection values shown in the figures are in inches. 42

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A B Figure 4-1. Honeycomb Panel Deflection fo r Load Case 1 Outward Pressure A) Total Translation B) Z Translation A B Figure 4-2. Honeycomb Panel Deflection fo r Load Case 2 Inward Pressure A) Total Translation B) Z Translation 43

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A B Figure 4-3. Honeycomb Panel Deflection fo r Load Case 3 Shear A) Total Translation B) Z Translation Integrally Stiffened Beaded Panel The total magnitude deflection (A) and Z defle ction (B) for the three load cases are shown below in Figure 4-4 Figure 4-5 and Figure 4-6 respectively. The deflection is shown at 10 times the actual. All deflection values shown in the figures are in inches. A B Figure 4-4. SBCF ISBP Deflection for Load Case 1 Outward Pressure A) Total Translation B) Z Translation 44

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A B Figure 4-5. SBCF ISBP Deflection for Load Case 2 Inward Pressure A) Total Translation B) Z Translation A B Figure 4-6. SBCF ISBP Deflection for Load Ca se 3 Shear A) Total Translation B) Z Translation Ply Failure The following sections summarize the ma ximum ply failure index of the two panels. Their indices ar e summarized below in Table 4-3 Table 4-3. Summary of maximum ply failure index Load Case Honeycomb Panel SBCF ISBP Outward Pressure 0.00409 0.0825 Inward Pressure 0.00488 0.0923 Shear 0.103 0.223 45

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Honeycomb Sandwich Panels The maximum ply failure index (A) IM L side (B) OML side of the honeycomb sandwich panel for the three load cases is shown below in Figure 4-7 Figure 4-8 and Figure 4-9 respectively. A B Figure 4-7. Honeycomb Panel Max Failure Index for Load Case 1 Outward Pressure A) IML side B) OML side A B Figure 4-8. Honeycomb Panel Max Failure I ndex for Load Case 2 Inward Pressure A) IML side B) OML side 46

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A B Figure 4-9. Honeycomb Panel Max Failure Index for Load Case 3 Shear A) IML side B) OML side Integrally Stiffened Beaded Panel The maximum ply failure index (A) Bead Side (B) OML Side for the SBCF ISBP three load cases is shown below in Figure 4-10 Figure 4-11 and Figure 4-12 respectively. A B Figure 4-10. SBCF ISBP Max Failure Index for Load Case 1 Outward Pressure A) Bead side B) OML side 47

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A B Figure 4-11. SBCF ISBP Max Failure Index for Load Case 2 Inward Pressure A) Bead side B) OML side A B Figure 4-12. SBCF ISBP Max Failure Index for Load Case 3 Shear A) Bead side B) OML side Discussion The purpose was to compare the ability of both panels to carry the applied loads not which panel can carry a greater load. The weight of each panel is used to determine if the ISBP can be substituted for a honeycomb sandwich panel provided the loads can be carried. In all aircraft, especia lly rotary wing aircraft, reducing weight is essential to improve performance. In this study additional capability is not considered a 48

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benefit as the loads are representative and the panels are constructed from the minimum number of plies. The air load cases do not significantly stre ss either of the ty pes of panels. The combination of the pressure loading with the in-plane loading of the panels would not adversely affect the ability of the panel to carry the in-plane loads Therefore, the results of the pressure load cases (#1 & 2) can be considered inconsequential to this study. Both panels are capable of carrying the loads which would need to be carried by a rotary wing aircrafts primary structure skin. The shear load case (#3) is most relevant as the loads best simulate the critical loadi ng of an aircraft skin. During that load case the ISBP has a higher failure index than the honeycomb sandwich panel but is still a considerable amount of margin remaini ng for both types of panels. The honeycomb sandwich panel is 33% heavier than the ISBP. As the ISBP is also capable of carrying the same load, it would appear to be a more efficient design. It may be possible to reduce the weight of the honeycomb sandwich panel by using a lower aerial weight uni-directional ca rbon fiber tape. Currently both the SBCF and continuous fiber material have equal nominal thicknesses and densities. This would result in a lower thickness of the honeycomb sandwich panel face sheets. This corresponding drop in thickness would results in a high failure index bringing it in line with or higher than the ISBP. 49

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CHAPTER 5 CONCLUSIONS 5 Chapter 5 The data shows that beaded panels can compete with honeycomb panels. While the total load capability is lower the ISBP has a lower weight. Both panels were able to carry the necessary loads with margin remaining. The bead height and laminate thickness can be adjusted to produce any desired results given adequate design clearance. The mechanical property data al so must be develop ed for an elevated temperature wet environment and evaluated. The assumption of symmetric quasi-isot ropic lay-ups for both panels should be considered a starting point. Additional work needs to be done to compare the ISBP with a honeycomb panel made from woven fabric carbon fiber. The analyst can assume that woven carbon fiber has quasi-isotropic lamina properties. This would make it possible to have a symmetric quasi-isotropic face sheet with three fabric plies. However the fabric is often thicker, weaker and has a lower elastic modulus compared to the unidirectional material. The long term advantage is that the hollow beaded panels would not have the corrosion and damage issues associated with sandwich panels. The beads could have drain holes that allow moisture and contaminates to be flushed out of the bead cavity. This will help eliminate some of the core related maintenance problems that currently exist. These holes could be placed at lightly loaded locations to minimize any effects of the hole. The possibility of reduced m anufacturing costs is also attractive. Further work is being done is to scale up the bead design. Figure 5-1 shows a formability test specimen where the cross bead pattern has been repeated. This was 50

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done to test the formability of the material over a larger mold. The geometry shown is intended to transfer the shear and longitudina l loads from one airc raft frame to another aircraft frame. The vertical beads have been removed indicating this panel is not intended to carry load in that direction. Figure 5-1. Scaled up test piece shows the cross pattern multiplied. Source: Cox, D and Nadel A.I. Design, Analysis and M anufacturing Transition for Stretch Broken Carbon Fiber (SBCF) Materials. Proceedings of the 54th International SAMPE Symposium and Exhibition Baltimore, Maryland, May 18 21, 2009. 51

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APPENDIX A MECHANICAL PROPERTIES 6 Appendix A Below in the following charts are the mec hanical properties for the SBCF and the continuous fiber material. The properties pres ented are: 0 Tensile Strength, 0 Tensile Modulus, 0 Compressive Strength, In-Plane Shear Strength, In-Plane Shear Modulus, 90 Tensile Strength, 90 Tensile Modulus, and Poisson Ratio. The data is taken from Reference 7. 52

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53

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54

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Below is the 90 Compressive Strength. This is a primarily matrix driven property. Therefore it is acceptable to use this proper ty. Note the values are in Standard units. These are from Reference 8. 55

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56 Below are the properties for the HRH-36-.125-4.5 honeycomb core [9].

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LIST OF REFERENCES 1) Hexcel Composites HexWeb Honeycomb Sandwic h Design Technology 2000. 2) Ted Wiesner HH-60J Seahawk Helicopter Corrosion Issues Army Corrosion Summit 2003 http://www.armycorrosion.com/past_ summits/summit2003/download1.cfm ?fname=wiesner.pdf (Accessed 10/2009). 3) Sankar, B.V. and M. Narayanan "Finite Element Analysis of Debonded Sandwich Beams under Axial Compression", J. Sandwich Structures & Materials 3(3):197-219 2001. 4) Abdallah, M.G., Hansen, N. W. and Jacobsen G. Development of Stretch Broken Carbon Fiber Materials Proceedings of t he 50th International SAMPE Symposium and Exhibition, Long Beach, California, May 1-5, 2005Jacobsen, G. Mechanical Performance Characterization Of Stretch Broken Carbon Fiber Material., Proceedings from the 54th SAMPE Technical Conference Baltimore, Maryland, May 18 21, 2009. 5) Dillon, G. and Stiver, D.H. III. Development Of Enabling Automated Forming Technology For Stretch Broken Carbon Fiber (SBCF) Materials, Proceedings from the 54th SAMPE Technical Conference Baltimore, Maryland, May 18 21, 2009. 6) Jacobsen, G. and Schimpf, W. Process Development and Characterization of Stretch Broken Ca rbon Fiber Materials, Proceedings from the 54th SAMPE Technical Conference Baltimore, Maryland, May 18 21, 2009. 7) Jacobsen, G. Mechanical Performance Characterization of Stretch Broken Carbon Fiber Material, Proceeding from the 54th SAMPE Technical Conference Baltimore, Maryland, May 18-21 2009. 8) Hexcel Corporation, HexPly 8552 Epoxy Matrix Product Data 2005. 9) Hexcel Corporation, Hex Web HRH-36 Produc t Data 2004 10) Hexcel Composites, TSB120 Hexweb Honeycomb Attributes and Properties 1999 11) Cox, D and Nadel A.I. Design, Anal ysis and Manufacturing Transition for Stretch Broken Carbon Fiber (SBCF) Materials. Proceedings of the 54th 57

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58 International SAMPE Sym posium and Exhibition Baltimore, Maryland, May 18 21, 2009. 12) Jones, R.M., Mechanics of Composite Materials Taylor & Frances Inc, Philadelphia, PA 1999.

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BIOGRAPHICAL SKETCH Adam Nadel was born in 1983 in New Hyde Park, New York. He grew up in Franklin Square, New York attending John Street Elementary School and H. Frank Carey High School; graduating in June 2001. So me of his activities included playing tuba in high school bands, various varsity spor ts, Model UN, and skiing. He also earned the rank of Eagle Scout in the Boy Scouts of America and the Vigil Honor in the Order of the Arrow. In the fall of 2001 Adam began undergraduate studies in mechanical engineering at The George Washington University. He founded a student chapter of the American Institute of Aeronautics and Astr onautics, serving as its president for the chapters first two years. He graduated magna cum laude in May 2005. Upon graduation Adam moved to Florida to work for the Northrop Gr umman Corporation. While in Florida he became a certified SCUBA Div er and started graduate school in the Fall of 2006 at the University of Florida in the Department of Mechanical and Aerospace Engineering where Dr. Bhavani V. Sankar agreed to be his advisor. In the spring of 2007 Adam moved to Virgin ia to work for Aurora Flight Sciences which he continues to do currently. His primary function is the Stress Analysis Lead for the Sikorsky CH-53K Main Rotor Pylon whic h is a primarily Carbon Fiber/Epoxy structure in addition to work on other projects including SBCF. He graduated with a Master of Science in Aerospace Engineering from the University of Florida in December 2009. Adam is son of Frank and Alane Nadel. He has a younger sister; Erica. 59