Composite Materials for Aircraft Structures: A Brief Review

Composite Materials for Aircraft Structures: A Brief Review

Composite Materials for Aircraft Structures: A Brief Review of Practical Application Jared W Nelson, PhD Candidate Department of Mechanical and Industrial Engineering Montana State University ME 480 Introduction to Aerospace, FALL 2011 Introduction

Composite materials are used more and more for primary structures in commercial, industrial, aerospace, marine, and recreational structures Design and Analysis of Aircraft Structures 13-2 From Last Time Composite parts used for aircraft applications are defined by Material, process, and manufacturing specifications.

Material allowable (engineering definition). All of these have a basis in regulatory requirements. Most efficient use of advanced composites in aircraft structure is in applications with Highly loaded parts with thick gages. High fatigue loads (fuselage and wing structure, etc). Areas susceptible to corrosion (fuselage, etc). Critical weight reduction (empennage, wings, fuselage, etc). Use must be justified by weighing benefits against costs.

Design and Analysis of Aircraft Structures 13-3 Composition of Composites Fiber/Filament Reinforcement High strength High stiffness Low density Carbon,

Glass, Aramid, etc Matrix Good shear properties Low density Thermoset & Thermoplastic Epoxy, Polyester, PP, Nylon, Ceramics, etc.

Composite High strength High stiffness Good shear properties Low density Anisotropic! Design and Analysis of Aircraft Structures 13-4 Overview

Micromechanics Study of mechanical behavior of a composite material in terms of its constituent materials Ply Mechanics Study of mechanical behavior of individual material plies based on variations from global coordinate system Macromechanics Study of mechanical behavior utilizing ply mechanics of a homogenized composite material

Failure Theories Design and Analysis of Aircraft Structures 13-5 CADEC: Introduction Compliment to text: Barbero, EJ. Introduction to Composite Materials Design; Taylor & Francis, 1999. Software free onlinesearch keywords CADEC & Barbero

Design and Analysis of Aircraft Structures 13-6 Micromechanics: Assumptions Lamina: Macroscopically homogeneous

Linearly elastic Macroscopically Orthotropic Initially stress free Fibers: Homogeneous

Linearly elastic Isotropic/Orthotropic Regularly spaced Perfectly aligned Carbon/epoxy (AS4/3501-6) composite (Vf=.70) Matrix:

Homogeneous Linearly elastic Isotropic Assumptions in Micromechanics of Composites Design and Analysis of Aircraft Structures 13-7 Micromechanics: Rule of Mixtures Vf,max approximately 78% Common range = 55-67%

Design and Analysis of Aircraft Structures 13-8 Micromechanics: Determining Properties Design and Analysis of Aircraft Structures 13-9 Micromechanics: Rule of Mixtures (E1)

Design and Analysis of Aircraft Structures 13-10 Micromechanics: Determining Properties Design and Analysis of Aircraft Structures 13-11 Micromechanics: Rule of Mixtures (E2)

Design and Analysis of Aircraft Structures 13-12 Micromechanics: Determining Properties Design and Analysis of Aircraft Structures 13-13 Micromechanics: Rule of Mixtures (12)

Design and Analysis of Aircraft Structures 13-14 Micromechanics: Determining Properties Design and Analysis of Aircraft Structures 13-15 Micromechanics: Rule of Mixtures (G12)

Design and Analysis of Aircraft Structures 13-16 Micromechanics: Other Methods & Strengths Design and Analysis of Aircraft Structures 13-17 Micromechanics: Halpin-Tsai (E2)

Halpin-Tsai: Semiempirical (1969) version to obtain better predictionBarbero empirical curve fitting parameter, commonly 2a/b Design and Analysis of Aircraft Structures 13-18 Micromechanics: Determining Properties Design and Analysis of Aircraft Structures

13-19 Micromechanics: Longitudinal Tensile Strength Design and Analysis of Aircraft Structures 13-20 Micromechanics: Determining Properties Design and Analysis of Aircraft Structures

13-21 Micromechanics: Thermal & Electrical Cond Design and Analysis of Aircraft Structures 13-22 Ply Mechanics So what happens if we vary the fiber direction angle away from the 1-direction?

CADEC uses Micromechanics results and fiber angle Plane Stress Transform stress/strain Off-Axis Compliance/Stiffness 3D Constitutive Equs Design and Analysis of Aircraft Structures

13-23 Ply Mechanics: CADEC Design and Analysis of Aircraft Structures 13-24 Ply Mechanics: Compliance Plane Stress Design and Analysis of Aircraft Structures

13-25 Ply Mechanics: CADEC Design and Analysis of Aircraft Structures 13-26 Ply Mechanics: Transformations Design and Analysis of Aircraft Structures

13-27 Ply Mechanics: CADEC Design and Analysis of Aircraft Structures 13-28 Ply Mechanics: Off-Axis Stiffness Matrices Design and Analysis of Aircraft Structures

13-29 Ply Mechanics: CADEC Design and Analysis of Aircraft Structures 13-30 Ply Mechanics: Stress-Strain Relationships Stress-Strain Relationship:

ij Cij ij With 3 planes Cij has 81 terms, but since: and: ij ji ij ji only 36 terms Orthotropic material (2 planes of symmetry) reduces to 9 terms:

Design and Analysis of Aircraft Structures 13-31 Ply Mechanics: Orthotropic Material Design and Analysis of Aircraft Structures 13-32 Macromechanics What if there are multiple lamina at differing angles?

CADEC uses Micromechanics and Ply mechanics to determine: Stiffness and Compliance Equations Laminate Moduli Global and Material Stresses and Strains Strains and Curvatures

Thermal and Hygroscopic loads For both Intact and Degraded materials Assumes: Plane sections remain plane Symmetry about a neutral surface No shear coupling

Perfect bonding Design and Analysis of Aircraft Structures 13-33 Shorthand Laminate Orientation Code Tapes or Undirectional Tapes [45/0/-45/902 /-45/0/45 [45/0/-45/90] s

Tapes or undirectional tapes Each lamina is labeled by its ply orientation. Laminae are listed in sequence with the first number representing the lamina to which the arrow is pointing. Individual adjacent laminae are separated by a slash if their angles differ. Adjacent laminae of the same angle are depicted by a numerical subscript indicating the total number of laminae which are laid up in sequence at that angle. Each complete laminate is enclosed by brackets. When the laminate is symmetrical and has an even number on each

side of the plane of symmetry (known as the midplane) the code may be shortened by listing only the angles from the arrow side to the midplane. A subscript S is used to indicate that the code for only one half of the laminate is shown. Design and Analysis of Aircraft Structures 13-34 Shorthand Laminate Orientation Code Fabrics and Tapes and Fabrics [(45)/(0)/(45)] Midplane

Fabrics [(45)/0(-45)/90] Midplane Tapes & Fabrics When plies of fabric are used in a laminate. The angle of the fabric warp is used as the ply direction angle. The fabric angle is enclosed in parentheses to identify the ply as a fabric ply. When the laminate is composed of both fabric and

tape plies (a hybrid laminate). The parentheses around the fabric plies will distinguish the fabric plies from the tape plies. When the laminate is symmetrical and has an odd number of plies, the center ply is overlined to indicate that it is the midplane. Design and Analysis of Aircraft Structures 13-35 Macromechanics: CADEC

Design and Analysis of Aircraft Structures 13-36 Macromechanics: CADEC Design and Analysis of Aircraft Structures 13-37 Macromechanics: Defining Laminate

Design and Analysis of Aircraft Structures 13-38 Macromechanics: Defining Laminate Design and Analysis of Aircraft Structures 13-39 Macromechanics: Material Properties

Design and Analysis of Aircraft Structures 13-40 Macromechanics: CADEC Quirkiness Design and Analysis of Aircraft Structures 13-41 Macromechanics: Review Outputs

Design and Analysis of Aircraft Structures 13-42 Macromechanics: Global Stresses Design and Analysis of Aircraft Structures 13-43 Macromechanics: ABD Matrices

Stiffness of composite where: [A] = in-plane stiffness. [D] = bending stiffness. [B] relates in-plane strains to bending moments and curvatures to in-plane forces bending-extension coupling. [H] relates transverse shear strains to transverse forces.

Design and Analysis of Aircraft Structures 13-44 Macromechanics: ABD Matrices Design and Analysis of Aircraft Structures 13-45 Macromechanics: ABD Matrices

Design and Analysis of Aircraft Structures 13-46 Macromechanics: Stiffness Equations Design and Analysis of Aircraft Structures 13-47 Macromechanics: Stiffness Equations

Design and Analysis of Aircraft Structures 13-48 Macromechanics: Laminate Moduli Design and Analysis of Aircraft Structures 13-49 Macromechanics: Laminate Moduli

Design and Analysis of Aircraft Structures 13-50 Macromechanics: Degraded Material Design and Analysis of Aircraft Structures 13-51 Macromechanics: Degraded Material What is a degraded material?

Design and Analysis of Aircraft Structures 13-52 Macromechanics: ABD Comparison Design and Analysis of Aircraft Structures 13-53 Macromechanics: CADEC Alt Methods

Data can be entered into .DAT and .DEF files Easily reloaded into CADEC More user friendly

Enter laminate Save Open CADEC Load Laminate Run Laminate Analysis Analyze Design and Analysis of Aircraft Structures 13-54 Failure Theories

Many failure criteria, most popular: Maximum stress criterion Maximum strain criterion Tsai-Hill failure criterion Tsai-Wu failure criterion Design and Analysis of Aircraft Structures

13-55 Not Just An Academic Exercise Consequence of Misalignment in Large, Composite Structure Design and Analysis of Aircraft Structures 13-56 Failure Theories: CADEC

Design and Analysis of Aircraft Structures 13-57 Failure Theories: CADEC Design and Analysis of Aircraft Structures 13-58 Failure Theories: Max Stress Criterion

Design and Analysis of Aircraft Structures 13-59 Failure Theories: Tsai-Wu Criterion Design and Analysis of Aircraft Structures 13-60 CADEC Demo

Design and Analysis of Aircraft Structures 13-61 Compression Fixture for Thick Laminates Analysis as performed by Julie Workman, Research Assistant Design Impetus

Design challenges Brute force solution CADEC solution for loading based on 2% strain Design and Analysis of Aircraft Structures Shear Loading Clamping fixtures apply a shear load to the specimen in order to introduce an axial load Tabbing required, very precise tolerances required for coupons, IITRI structure is very massive Can result in end crushing (Adams; 2005)

Design and Analysis of Aircraft Structures End Loading/ Combined Loading End Loading Fixtures apply load axially to fibers Low transverse shear strength can result in end-brooming in a tabbed or un-tabbed specimen Some question in the D 695 Fixture as to whether the fixture carries some of the load (Wegner, Adams; 2000) Design and Analysis of Aircraft Structures Proposed Components

Design and Analysis of Aircraft Structures Challenges Geometric Challenges: Gage section which will accommodate flaws coupon geometry Utilize existing back component with bolt threads and coupon keyway

Match materials, dimensions and bolt axes on existing fixture components Axis of symmetry of alignment rods and coupon in plane with each other Massive structure Design Challenges Coupon must carry all of the load, no fixture interference

zero binding Zero misalignment Design and Analysis of Aircraft Structures Laminated Plate Theory Strains refer to middle surface strains, k-terms are curvature terms Q is the stiffness matrix in terms of layer engineering constants in the coordinates of the material Nx is the resultant force in the x-direction of the laminate obtained by integrating the stress in that direction through

the thickness. Design and Analysis of Aircraft Structures Laminated Plate Theory [Q] is actually [Q-bar] the transformed stiffness matrix. [Q-bar] multiplied by the strain of each layer integrated over the thickness of the laminate is equal to strain multiplied by thickness. Where [A] is [Q-bar]*t ---Assumed all layers with misalignment have same properties

Enter strain, x = .02 CADEC calculates [N] Stress and subsequent force is backed out from there Design and Analysis of Aircraft Structures CADEC Analysis Design and Analysis of Aircraft Structures CADEC Analysis Design and Analysis of Aircraft Structures

CADEC Analysis Design and Analysis of Aircraft Structures Results Figure 1: 20-ply Variable Angle Angle yy Ny

Force in y 0 4.29 5.72 6.435 5

205.24 273.65333 307.86 10 672.6 896.8

1008.9 15 1388.6 1851.4667 2082.9 20

2266.94 3022.5867 3400.41 25 3201.62 4268.8267

4802.43 30 4079.96 5439.9467 6119.94 35

4795.96 6394.6133 7193.94 40 5263.32 7017.76

7894.98 45 5425.62 7234.16 8138.43 Design and Analysis of Aircraft Structures

Concluding Remarks Composite design fairly simple Assumptions lead to simplified analysis Idealized Real-world? CADEC

Begin with component properties Micromechanic, Ply and Macromechanic analysis Apply loads and match against failure criteria Simple structures (Not covered) Software options: COMPRO, MSExcel, Matlab, MathCAD, etc. Composites still require significant analysis and physical testing Parts/Structures are only as good as the manufacturing You can never make good parts with bad materials, but you can easily make bad

parts with good materials! Design and Analysis of Aircraft Structures 13-73

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