Fiber-Reinforced Composite Materials Design

2015 Park Avenue, Suite 8 Redlands, California 92373 (909) 307-6218

Fiber-Reinforced Composite Materials Design & Analysis Primer Prepared By Pedro G. Morales

Introduction The design of fiber-reinforced composite structures relies heavily in the proper assessment of material characteristics and range of application, failure mechanisms, manufacturing processes and tool design. In addition to a good assessment of material and processes, but in the author opinion to a lesser extent, a judicious structural design including fiber architecture is needed in order to take advantage of the material high specific strength, stiffness, and thermo-mechanical capabilities. The purpose of this report is to document the author thoughts on the subject of fiber-reinforced composite structures material and processes as well as a design and analysis methodology. This report is organized as follows:

1.

1.

Material Selection

2.

Failure Mechanisms

3.

Fabrication Methods and Tooling Considerations

4.

Design Guidelines and Analysis Methodology

5.

Composite Materials Analysis Programs

Material Selection

Composite materials are commonly used when high specific strength and/or stiffness is needed to meet requirements or when the need for enhanced equipment performance outweighs its possible high cost. Most likely, and because of economic and time constraints, the engineering group will opt to replace the baseline metal design with a quasi-isotropic equivalent. The fiber system will be used to tune the stiffness and strength, and the resin choice will depend on environmental requirements.

1


To take full advantage of fiber-reinforced composite materials one must consider Fiber Type, Resin System, Form, Grade/Style, and Material Architecture. The Fiber Type is usually chosen on the basis of stiffness, strength, thermal expansion, thermal conduction, and electrical resistivity properties. Other conditions that affect the final choice include material availability, cost, and strain to failure. Ultra High Modulus (UHM) fibers (70-140Msi) have the greatest carbon content and therefore offer the highest thermal conductivities but they are also costly and because of their low strain to failure property can not be formed into complex high curvature geometric shapes nor can the final form adequately resist impact loads. A few high modulus PAN fibers, but mostly PITCH fibers fall into this category. Industry favorites include Amoco K-1100 (900-1,000 Wm/K)1and Mitsubishi K13C2U (600-700Wm/K)1; Aluminum thermal conductivity is 150-200 Wm/K. These fibers are primary candidates in thermal management applications such as electronic enclosures, space radiators, and satellite battery sleeves. In general UHM fibers are too costly (>$1000/lb) to be used in large structural applications. High (50-70Msi) to Standard (33-35Msi) modulus PAN, PITCH and Glass fibers are more commonly used and the cost per lb is more palatable as shown in the fiber property comparison table below. Fiber Property Comparison Table 1

2

Fiber Name

Manufacturer

3

4

5

6

7

8

9

10

Tensile Modulus, Msi

Tensile Strength, Ksi

Thermal Conductivity, W/mK

CTE, in/in/F

Electric resistivity, 10-6 ohm.m

Strain to Failutre, %

$/lb (100 lbmin)

Year

K1100

Amoco

140

450

1100

-.8e-6

1.2

0.25

1,750

1999

K13C2U

Mitsubishi

130

550

620

-.8e-6

1.9

0.42

1,160

1997

YS-90A

NGF

130

510

500

-.8e-6

3

0.30

1,050

1997

M60J

Toray

86

571

75

0.70

720

1997

M55J

Toray

77

585

54

XN-50A

NGF

75

560

180

-.8e-6

0.80

360

1997

7

0.70

370

1997

M46J

Toray

64

613

26

M40J

Toray

55

642

24

-.4e-6

9

1.0

185

1997

0.0

10

1.2

160

1997

T300

Toray

34

514

9

.2e-6

18

1.5

S-2

Owens Corning

12

620

Spectra 1000

Allied Signal

15

435

1.1

.9e-6

20

2001

5.5 2.2

3.0

Columns three and four of the fiber property comparison table show the typical variation of strength with stiffness. As modulus increases strength decreases. Higher fiber stiffness is a result of the high temperature carbonization process aim to align carbon planes along 1

See Resin systems for its effect on thermal conductivity

2


the fiber axis. Aligning the planes reduces fiber extensibility and therefore strain to failure. Column six shows the attractive negative axial CTE property that carbon fibers have. This property allows the analyst to tailor geometrically stable structures under hot and cold environments. This property together with that of high thermal conductivity (columns 5 & 6) is particularly advantageous when designing space structures such as space-based optical telescopes, optical benches, and various satellite components. Military equipment design concerns can be strength, stiffness, ballistic protection and radar transparency. Here is where use or exclusion of composites must be determined at the initial design stage. Radar transparency is accomplished by designing with low electric resistivity fibers. Ballistic protection is best accomplish with the use of high strain to failure fibers, allowing for greater energy dissipation. The fiber property comparison table shows that low resistivity is accomplished at the expense of a reduction in the energy absorption capability (columns 4 & 7). Thus, carbon, Kevlar, S-glass, Spectra 1000 fibers or similar are often used to create hybrid baseline materials (with occasional 1/8” Al plate substitutions). In this manner energy absorbing panels or ballistic/radar invisible structures are created and used in some military vehicles and lately to safeguard aircraft cockpits against firearms. The fighter/tactical military aircraft must sustain extreme aerodynamic and inertia loads repeatedly while maintaining different degrees of radar transparency. Graphite fibers together with the proper resin system can meet the strength, and stiffness requirements as well as offer increased endurance over metal upon redistributing loads as fibers break and bridge cracks to keep them from growing during the equipment life span. A typical composite fatigue behavior chart is shown below

3


The figure above shows that fiber composites offer more time to catastrophic failure and lower damage growth rates. Thus, the increased interest in using carbon composites in the design and manufacturing of rotary components such as helicopter, wind tunnel, and energy-producing wind mills fan blades.

The fiber system(s) thermo-physical response as well as its cost and manufacturing processes should be part of the selection process. Financial considerations include a baseline material characterization program running in parallel with the design and analysis process if a customer approved database is not available. Thus, the incentive to present the customer with alternatives based on previously characterized materials so that only component-level testing is needed. As a rule of thumb is better to stay away from exotic fiber systems as an extensive material characterization program will be required on most likely an expensive fiber system and cost and schedule constraints will force artificial strength and stiffness reductions because of the lack of test data. One must also not forget the rule of supply and demand. Exotic fibers, sometimes independent of price, will have a longer delivery time and this could potentially become a bottleneck in the design process. Manufacturability is a function of form, fabrication method, as well as fiber and in some cases resin system. Subjects to be discuss shortly. But basically, all of the above must be evaluated in order to order material, design tooling and create partnerships. Its sometimes wise to keep track of the qualified vendor current jobs because 1) if they carry too many projects in parallel they will tend to take care of the higher profit, recurring jobs first. In addition, because sometimes several parts are cured simultaneously and not all cure cycles are the same our parts may end up “tagging along” with their prime customer jobs. Sometimes a custom cure cycle must be design in order to ensure proper fiber saturation, or reduce spring back. 2) If the vendor is having financial problems they may end up cutting down on the labor force and unqualified people may end up doing the cutting and placing of tape or cloth or controlling sensitive equipment such as fiber winding or tape placement equipment, using wrong solvents, primers, or misusing abrasion tools. The most common Resin systems associated with fiber composite materials are epoxies, phenolics, cyanate esters (CE’s), bismaleimides (BMI’s), and polyimides. Each formulation has its distinct attributes but additives such as catalysts, flame-retardants, and plasticizers are sometimes use to improve their properties. Epoxies are probably the most commonly used resins and usually come in 250F and 350F cure formulations. Epoxies sometimes suffer from porosity and plasticizers are commonly added to increase toughness. Phenolics are primarily used in ablative and flame-resistant applications such as rocket nozzles in order to reduce char yield and military helicopter bulkheads and commercial aircraft interiors to protect against fire. Phenolics tend to release water vapor during cure creating a greater degree of porosity, and thus tend to be mechanically outperformed by others resin systems.

4


Resin Property Comparison Table 1

2

3

4

5

6

7

8

9

10

Resin Type

Resin Name

Manufacturer

Tensile Modulus, Msi

Tensile Strength, Ksi

Thermal Conductivity, W/mK

CTE, in/in/F

Tg, F

Viscosity, poise

Moisture Absorption, %

0.45

230

65-95

2.8

302

1,250

.04

Epoxy

826

Shell

10

.173

38e-6

Phenolic

Typical

Typical

7

.297

22e-6

CE

EX-1515-1

Bryte

10

BMI

RS-8

YLA

Polyimide

High Tg

UBE

.62

19e-6

12

563

17

27e-6

644

2

Low moisture absorption, which translates into minimum outgassing2, is the primary advantage that Cyanate esters have over other resin systems. In addition, and as shown in the resin property comparison table above CE’s provide excellent strength and Tg properties as well as superior electrical properties. Thus, applications range from radomes to space antennae. The glass transition temperature mechanical property is an index of characteristic retention at high temperatures. BMI’s and Polyimides are formulated to posses high Tg’s and therefore are mostly used in high-temperature applications. However, special CE formulations are use as well. The author’s experience with the theater high altitude area defense (THAAD) missile kill vehicle, and the divert attitude control system as well as the air borne laser pressure recovery system, and various launch vehicles confirm that BMI’s as well as polyimides are the best candidates for such applications as they offer hot/wet service temperature to 450F and 700F. Environmental knockdowns usually imposed by the customer and/or applied as a good design practice range from 15% to 5%. The magnitude of these knockdowns is usually a function of the lack of available material data. The engineering team could potentially start the design process with a 30% knockdown on material strength and stiffness, i.e. (chemicals = 0.85)x(moisture = 0.90)x(temperature = 0.95). These knockdowns are applied to the fiber-reinforced composite material (resin + fibers), but in reality environmental conditions only effect the resin composition and because resin provides the means to bind fibers together, redistribute loads upon local failure, and protect against environmental effects material strength and stiffness are directly affected. Thus, resin systems including appropriate additives must be selected at the beginning of the project so that the cumulative material knockdown is not as severe. Sometimes is best to work with customer supplied material data in order to avoid extensive testing that may delay the project because of unexplainable data scatter, changes in architecture, and so on. The resin property comparison table columns 6 thru 9 show other influential parameters to consider: thermal conductivity, CTE, Tg, and viscosity. First note that resins behave as thermal insulators so that the cured composite net thermal conductivity property will be 2

Release of solvents and moisture from composite parts under the hard vacuum of space

5


50% to 70% of the fiber depending on selection and architecture. This hurdle has been somewhat circumvented by Amoco’s ThermalGraph panels formed by fused high thermal conductivity fibers that can be interspersed among other less heat conductive layers and used as space radiators or to transfer heat to other parts of a spacecraft. As previously mentioned CTE is an important parameter in the design of advanced spacebased optical systems and although carbon fibers offer negative CTE’s, proper tailoring is required as resins systems posses large positive CTE’s. A likely architecture will place a large percent of low CTE high thermal conductivity fibers in the optical direction and intermediate modulus fibers in the hoop direction providing consolidation and strength. As parts cure resin cross-linking (crystallization of polymer molecules) takes place and continues up to the glass transition temperature, Tg, where viscosity is so high that no further dimensional change occurs. A common design practice is to stay away from the Tg by 60F in order to avoid any changes in resin mobility. Viscosity is most important when selecting the material form (tape, fabric, braids/weaves), grade/style (weight/tow size), and fabrication method (hand lay-up, fiber winding, RTM, pultrusion). Proper fiber saturation is very important in order to yield maximum finished part thermo-mechanical properties, and as one can imagine the degree of resin viscosity can play a major role in the manufacturing process. For example, while hand laid tapes and fabrics are more forgiving when it comes to resin viscosity, fiber wound and specially RTM’s (resin transfer molded) components may require specific formulations in order to secure proper preform saturation. Some times improper saturation in out-of-the mold RTM parts is obvious because of visible porosity and lack of “shine”. Improper saturation in the case of fiber wound parts can sometimes be hard to detect visually or by ultrasound techniques and end up surfacing during proof testing. Much lower than expected failure loads and excessive audible fiber brakeage during loading are some of the symptoms of improper fiber saturation. As briefly mentioned fiber-reinforced composite Forms include basic fibers (spools), prepregs (tape, fabrics), and braids/weaves. Basic fibers (spools) are primarily used for fiber winding purposes. These spools are place on fiber winding machines where fiber tows go through a resin bath and are then placed on the mandrel. Mandrel geometry is an important parameter used to determine fiber type because of strain to failure constraints. In many instances broad goods as well and precured parts are used to circumvent problems. The author’s experience with the 601HP S/C carbon fiber-reinforced Xenon pressure vessel, the AeroAstro PA-X rocket project, and others indicate that particular attention should be paid to proper tow tension and composite wet-out (lack of fiber resin saturation) because such conditions can reduce strength, stiffness, and provide a conduit for delamination and a change in the catastrophic failure mode. Thus, a proper line of communications between the manufacturing engineer and the shop technician is of outmost importance. Prepregs are perhaps the most commonly used form of fiberreinforced composite materials. Prepregs, usually available in limited widths and Bstaged (tacky condition) are cut to shape and hand laid or mechanically placed on the tool. Formability is the main tape limitation because of strain to failure and possible fiber drift during placement. Fabrics can better conform to high curvatures and add a throughthe-thickness strength component with a reduction in stiffness. Thus, fiber-reinforced

6


composite parts tend to be a mixture between tape and fabric. Adding fabric layers for formability, damage protection and durability, to improve bonded surface adhesion, to provide added bolt bearing strength, and to prevent fiber breakout. Tape is used to add stiffness and/or to provide a more efficient thermal path. Braids can look more or less like cloths on the finished part except that they can have a much more intricate weave design enhancing through-the-thickness properties such as damage tolerance. In addition, complex structural components such as stiffening ribs can easily be woven into preforms. The use of preforms and automated tape laying techniques add the element of repeatability, which is highly desirable in the composite manufacturing business. Hand lay-up is prone to human mistakes due to errors in cutting, placing, and lay-up sequence. The author uses the labels Grade and Style to describe aerial weight and fabric tow size (fiber bundle). Grade/Style is a short but important subject as it can excite a failure mechanism in the later stages of the design process and whose understanding proves beneficial to help the engineering team select the proper product. Commonly used cured ply thickness range from about 5mils for tapes and 7.5mils for fabrics. Smaller but not commonly used fiber diameters can result in cured ply thickness in the order of 2mil. The author’s experience with the MIT (geosynchronous microwave imager) GEM reflector study, and Harris Ka-Kv band reflector program indicates that ply thickness in the order of 2mil are needed for the design of space reflectors in order to ensure the proper fiber architecture and tiling scheme resulting in the most uniform and smallest dimensional change while maintaining good thermal properties. Similarly, fabrics and braids can be woven from various tow sizes. Typically one can find fabrics woven from 1k, 3k, 6k, and 12k fiber bundles (1k = 1000 filaments per bundle). Clearly there are advantages and disadvantages in choosing a tow size over another. Larger tow sizes result in faster thickness build up but depending on the number of such layers proper wet-out may be harder to achieve unless a less viscous resin formulation is used and/or the cure cycle modified. In addition, larger tow sizes mean larger valleys and this situation can create adhesion problems when one of this layers functions as a bonding surface. Valley Tow

The problem arises because of the surface preparation procedure that commonly involves abrasion, solvent wipe, and surface priming. Aggressive abrasion can result in damage to the plies below while inadequate abrasion can leave a porous surface whereby only 60 to 70% of the bond surface has been properly abraded (depending on tow size).

7


The author’s experience has been that a 4K tow is already too large for proper hand abrasion and a peel-ply3 is a better alternative to surface preparation. Material Architecture is perhaps one of the most important design steps after fiber and resin selection. Baseline material (laminate) thermo-mechanical characteristics are typically evaluated in this step. Classical lamination plate theory (CLPT) is commonly used to determine in-plane stiffness and strength properties used in the initial stages of the design process. In short CLPT neglects through-the-thickness effects and is based upon the assumption of thin laminates built up from plies having orthotropic material properties with their principal or fiber direction oriented at an angle,  (as shown below) to the material axis, x, and integrated across the thickness for all plies 1,2,3,…N resulting in a relation between loads and laminate response [1].

{

Lamina

1 2 3

. .

Laminate

Go back to Analysis Approach

x

N

Go to FEM guidelines C5

z

Go to FEM guidelines C6.

y

Go back to Analysis Approach D1 D2

1

2



x

Note that in order to determine laminate properties such as effective thermal conductivity, CTE, stiffness, strength, and so on lamina properties must first be characterized. That is, longitudinal (0-deg), transverse (90-deg), and rail (45-deg) material characterization tests must first be performed if reliable data is not available. In addition, because CLPT cannot be used to estimate trough-the-thickness material characteristics such as flat-wise interlaminar tension and shear strength, further material 3

A removable outside fabric ply molded onto the surface of a laminate to provide a chemically clean surface for bonding when it is removed.

8


testing is still necessary. So although CLPT is used for initial material characterization, a material characterization test program that includes through-the-thickness effects should be in place in the early stages of the design process and run parallel to it. Material architecture involves the selection of the fiber type, resin system, form, grade/style, and lay-up or ply stacking sequence to create the baseline material. The subjects of fiber, resin, form, and grade/style have previously been discussed in an individual basis. Their influence in the manufacturing and thermo-mechanical performance of the product will be discussed upcoming sections. The significance of the stacking sequence is the main subject for discussion in this section as a form of material selection. Similarly, stacking sequence will be discussed in sections to come to cover the effect on manufacturing and thermo-mechanical performance of the product. The lay-up selection depends on the basic understanding that the function of the fibers is to transfer the load, heat, acoustical energy, and so on, while that of the resin is to bind the fibers together, provide environmental protection, and provide a secondary load path upon fiber breakage. Thus, if the requirement is to transfer heat most high thermal conductivity fibers should be aligned in the direction of the desired heat flow. Similarly, if stiffness is required most high stiffness fibers should be lined up in the reinforcing direction. If torsional stiffness is sought, most fibers should follow a 45-deg path relative to the surface of revolution generator. These lay-up schemes may seem intuitively obvious, but the difficulty usually arises when is time to determine the percent orientation for each group. For example 40% 0-deg 40% 45-deg 20% 90-deg or (40/40/20), their location through the laminate thickness, for example ply 1 thru 2 45-deg, ply 3 thru 4 0deg ply 5 thru 6 90-deg and so on. This is because only on rare occasions would one select a single orientation for the whole stacking sequence, one example being that of bonding high stiffness unidirectional tape to the flange of an H extrusion to add bending stiffness as shown below UHM Uni-tape 7075 Al extrusion

In most cases a balance between strength, stiffness, dimensional stability and thermal performance must be achieved. Thus, the reason for initial characterization using CLPT resulting in what are called carpet plots. A typical carpet plot is shown below

9


A carpet plots typically shows a laminate material response as a function of percent of various material orientations. The plot above shows the x-direction effective unnotched laminate strength as a function of percents in the lay-up family shown. Most initial carpet plots are not as complicated as the one shown but the basic objective is similar. Such plots form the basis for initial sizing and material selection implying an iterative process. Moreover, such preliminary plots are theoretical in nature and must be updated as test results become available. In addition, because failure actually starts at the lamina level and because of unknowns in the laminate failure mechanism, residual characteristics are commonly evaluated at the lamina not laminate level. Thus, such information is used for initial sizing and not for detailed analysis.

2.

Failure Mechanisms

Fiber-reinforced composite materials are in essence microstructures consisting of a binding compound and reinforcement in the form of continuous or “chopped” filaments.

Chopped Fiber Composite

Fiber-reinforced Continuous Filament Composite

10


At the constituent level damage4 initiates at the fiber/resin interface because of poor adhesion, and at resin micro cracks because of unavoidable porosity. In addition, a “hidden” damage initiator can be moisture because it can induce blistering under certain thermal conditions. During the lay-up and build up process tape drift, improper tow tension (winding), improper splicing, error in lay-up schedule, wet-out deficiencies, the use of bad resins (resin with lower than specified viscosity because of excessive moisture resulting in poor fiber/resin adhesion) can lead to damage initiation during equipment operation. At the laminate level differences in ply-to-ply material properties, including fiber orientation, can lead to delamination and/or warping because of the consolidation process or upon thermo-mechanical loading. Failure may also occur because the service load has exceeded the material endurance limits. At the manufacturing and assembly level failure may occur because of drilling resulting in fiber break-out, improper bonding surface preparation, excessive fastener torque, improper countersunk hole depth resulting in cracks or fastener pull-thru at lower than expected levels, surface or internal damage because of installation impact loads and others. During operation damage and failure can occur because of conditions such as excessive transverse loads leading to delamination, warping because of CTE and/or drastic changes in stiffness, de-bonding because of excessive peel loads in an adhesively bonded joint, and others. The previous short paragraphs present some of the possible conditions that can lead to damage and possible failure in fiber-reinforced composite structure including panels, chopped fiber molded composites, or fiber-reinforced laminated composites. The table below shows the most common sources of damage initiation known to the author that could result in component failure, the associated failure mode(s), likely indication of the failure mode(s), and possible preventive measure(s).

4

Damage does not constitute catastrophic failure but its cumulative effect can reduce equipment performance below a threshold we shall call failure.

11


Sources of Damage and Preventive Measures, (1/6) Failure Mode Preventive Measure(s) Indication(s) Excessive Delamination because NDE will Choose low moisture resin moisture in resin of thermal shock– show map and adhesives, thermal induced blistering insulation, stitching Cracking because of Visual Protective coating, fiber thermal cycling inspection glass outer layer Improper lay-up Warping of laminate Visual Use symmetric, balanced due to consolidation inspection lay-ups. Modify cure cycle process to reduce spring back. Consider post-cure and fastening of incompatible CTE parts Warping of panel due Visual Above plus schedule plies to consolidation inspection to diminish large process differences in edge-band CTE Source of damage

Failure Mode(s)

12


Sources of Damage and Preventive Measures, (2/6)

Source of damage Improper lay-up

Failure Mode(s) Micro-cracking and delamination during service

Failure Mode Indication(s) Sublaminate separation visible upon dissection or by NDE

Delamination buckling during service, panels

Visual inspection

Step down transition delamination and ply failure during service End fitting transition fracture during service

NDE will show map

Laminate break-out because of drilling

Visual inspection

Impact damage during installation or service

Visual inspection, NDE

Bolted joint bearing, shear, or tension failure during service

Visual inspection

Visual inspection

13

Preventive Measure(s) Do not group more than four plies with the same orientation. Avoid 90-deg plies in highly loaded bends. Monitor flat facesheet buckling strength during analysis, schedule ±45’s to the surface, reduce core thickness In drop-off zones do no more than 2 plies at a time at 0.2in intervals. Try to match fitting and laminate stiffness and CTE at the junction, scarf composite into fitting, consider chopped fiber molded fittings Schedule at least one layer of fabric material on the bounding surfaces Schedule at least on layer of fabric material on the bounding surfaces. Consider using NGF XN-05 or similar amorphous carbon fiber prepregs on bounding surfaces. Schedule at least 50% of ±45’s locally, local reinforcement, larger bolts, multiple bolts


Sources of Damage and Preventive Measures, (3/6) Source of damage Improper bonding procedure

Failure Mode(s) Facesheet separation at ramp during service Bond failure during service Additional Information

Improper material selection

Failure Mode Indication(s) Surface bump, NDE Glossy surface, inconsistent bondline thickness

Fiber breakage or folding during the consolidation process

NDE

Fiber drift during layup

Visual inspection

Fracture during service because of insufficient wet out

Porosity, dryness

14

Preventive Measure(s) Use two layers of film adhesive in region More aggressive surface abrasion, careful solvent wipe, adequate and evenly distributed bonding surface pressure (clamps, clecos), use of peel ply, change adhesive system considering environmental and chemical effects as well as material characteristics Use mold line radius >0.125”, consider fibers with higher strain to failure values In winding: better tow tension control and higher resin viscosity. In common hand or assisted tape lay-up: use fabric and work in to conform to curvature Change to less viscous resin and/or tweak cure cycle, change fabrication method


Sources of Damage and Preventive Measures, (4/6) Source of damage Improper material selection

Improper design configuration

Failure Mode Preventive Measure(s) Indication(s) Structural failure due Catastrophic Phenolic resin plus rayonto excessive heating failure based precursor carbon (aerodynamic, fibers or C-C composite flame) substrates, glass/phenolic shield, TPS . Use molded glass/phenolic brackets and fittings to transition into internal structures Performance ThermoUse higher thermal reduction due to vacuum conductivity fibers to excessive heating chamber data channel heat to colder (radiation) areas of spacecraft Longitudinal rapture Visual At burst longitudinal on proof inspection strain