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Composite Design Section 2 of 3 AMTS-SWP-0048-F-2011

AMTS STANDARD WORKSHOP PRACTICE _________________________________________

Composite design Section 2 of 3: Composite design guidelines

Reference Number:

AMTS-SWP-0048-F-2011

Date:

October 2011

Version:

Final


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Contents 1 Scope ............................................................................................................................ 2 2 Technical terms.............................................................................................................. 2 3 Primary references ......................................................................................................... 3 4 Fundamentals of Composite design decisions. .............................................................. 5 5 Composite laminate advantages and disadvantages ..................................................... 5

5.1 Advantages ............................................................................................................. 5 5.2 Disadvantages ........................................................................................................ 6

6 Guidelines for Composite design ................................................................................... 8 6.1 Fundamental laminate design guidelines ................................................................ 8

6.1.1 Guideline 1: Lay-up symmetric about mid-surface ................................................ 8 6.1.2 Guideline 2: Balance Laminates ............................................................................ 9 6.1.3 Guideline 3: Do not Extrapolate test data .............................................................. 9

6.2 Fibre dominated un-notched laminates ................................................................. 10 6.2.1 Guideline 4: Fibre dominated laminates if at least 10% of plies in basic 4 directions. .................................................................................................................... 11 6.2.2 Guideline 5: Primarily loaded plies internal .......................................................... 12

6.3 Stability of fiber dominated laminates .................................................................... 12 6.3.1 Guideline 6: Lay-up symmetric about mid-surface ............................................... 12 6.3.2 Guideline 7: ±45° Plies on Exterior ...................................................................... 12

6.4 Thermal response of laminates ............................................................................. 13 6.4.1 Guideline 8: CTE must be considered in designs ................................................ 13 6.4.2 Guideline 9: 10% 0° and 90° plies to avoid thermal expansion. ........................ 13 6.4.3 Guideline 10: Use most ductile Resin satisfying conditions ................................. 13 6.4.4 Guideline 11: Max operating temp 10°C below Tg ........................................... 14

6.5 Stacking sequence and inter-laminar free edge stresses ...................................... 14 6.5.1 Guideline 12: Edge stress controlled in design .................................................... 14 6.5.2 Guideline 13: Limit layer thickness within laminate less than 0.5mm ................... 14

6.6 Poison’s ration mismatch between laminates and bonded or co-cured stiffeners .. 15 6.6.1 Guideline 14: Poisson mismatch of skin & bonded stiffener <0.1......................... 15

6.7 Holes, cutouts, impact damage ............................................................................. 16 6.7.1 Guideline 15: Assume hole dia. 6mm anywhere in composite ............................. 17 6.7.2 Guideline 16: <60% Plies at cutouts and bolted joints ......................................... 17 6.7.3 Guideline 17: Maximum 60% plies in any direction .............................................. 17 6.7.4 Guideline 18: Reinforcement around cutout should be interspersed .................... 18

6.8 Joints .................................................................................................................... 18 6.8.1 Guideline 29: 35% ±45° plies at bolted joints .................................................... 18 6.8.2 Guideline 20: Balance the Membrane stiffness of the adherents. ....................... 19 6.8.3 Guideline 21: Tapering ends for minimizing of Peel stresses in Thick joints. ....... 19 6.8.4 Guideline 22: Scarf /step lap thick joints .............................................................. 19 6.8.5 Guideline 23: most ductile adhesive that satisfies requirements .......................... 19 6.8.6 Guideline 24: Co-cure step lap joints ................................................................... 20 6.8.7 Guideline 25: Ensure adhesive/laminate cure cycles are compatible................... 20 6.8.8 Guideline 26: Design repairable joints. ................................................................ 20 6.8.9 Guideline 27: Correct surface preparation of adherents are essential ................. 20 6.8.10 Guideline 28: Corrosion barrier between graphite and aluminum ...................... 20

6.9 Tapering of skins and flanges bonded to skins ...................................................... 20 6.9.1 Guideline 29: Taper drop-offs ............................................................................. 20 6.9.2 Guideline 30: Angle ply pairs should be dropped off together .............................. 20 6.9.3 Guideline 31: Outer plies should cover all the other drop-offs............................. 21 6.9.4 Guideline 32: Stiffeners and Beam flange edges taper <10:1. ............................ 21

6.10 Damage tolerance, durability, and certification. ..................................................... 21 6.10.1 Guidelines for Certification purposes ................................................................. 21


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1 Scope A sophisticated analysis plays an important role in the development of aerospace structures.

It is required that many specialities are combined on an overall problem in any complex design structure. For each of these specialities the specific knowledge must be known by the designer. [1]

This SWP is part 2 of three SWP’s which covers the following information: SWP 42: Composite design Section 1 of 3 - Composite definition

- Composite classification - Basic terminology - Lamina Theory - Laminate Analysis Theory - Static strength life of composites

Theory - Beam Analysis theory

SWP 48: Composite design Section 2 of 3 - Fundamentals of Composite design

decisions - Advantages / Disadvantages - General guidelines for composite

designs (Unidirectional tape laminates)

SWP 49: Composite design Section 3 of 3 - Introduction to Composite analysis

software - Comparison of various software tools - Quick start guide to using LAP

2 Technical terms Adhesive:

Substance applied to mating surfaces to bond them together by surface attachment.

Balanced laminate:

Any laminate that contains one ply of minus theta orientation, with respect to the principle axis of the laminate, for every identical ply with a plus theta orientation (e.g. a laminate with a principal axis of 0º combined with an equal number of plies that have -45º and +45º orientations).

Buckling:

Failure mode usually characterized by unstable lateral deflection, rather than breakage, under compressive force.

Carbon fibre:

Produced by pyrolysis of an organic precursor fibre, such as PAN (polyacrylonitrile), rayon or pitch, in an inert atmosphere at temperatures above 982ºC/1800ºF. “Carbon” is often used interchangeably with “graphite” but carbon fibres are typically carbonized at about 1315ºC/2400ºF and contain 93% to 95% carbon while graphite fibres are carbon fibres submitted to graphitization at 1900º to 2480ºC (3450ºF to 4500ºF) after which they contain


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more than 99% elemental carbon.

Composite: Three-dimensional combination of at least two materials differing in form or composition, with a distinct interface separating the components. Composite materials are usually manmade and created to obtain properties that cannot be achieved by any of the components acting alone.

Cure cycle: The specific sequence of temperatures, pressure and time used to cure a

specific matrix system.

Epoxy: A thermosetting polymer containing one or more epoxide or oxirane groups, curable by reaction with amines or alcohols, used as a resin matrix in reinforced plastic products and as the primary component in certain structural adhesives. Cured epoxy resin is highly resistant to chemicals and water and its performance properties are relatively unaffected by extreme temperatures.

Fatigue: Failure or deterioration of a material’s mechanical properties as a result of

repeated cyclic loading or deformation over time.

Impact strength:

A material’s ability to withstand shock loading as measured during a test in which a specimen is fractured.

Lap joints:

A joint made by overlapping two parts and bonding them together.

Matrix: Material in which reinforcing fibre of a composite is embedded. Matrix materials include thermosetting and thermoplastic polymers, metals and ceramic compounds.

Pre-cure: Full or partial hardening of a resin or adhesive before pressure is applied.

Reinforcement: The key element that, when combined with a matrix to make a composite,

provides the required properties (primarily strength). Reinforcement forms range from individual short fibres to complex braided, woven or stitched textile using continuous fibres.

Resin: A solid or pseudo-solid polymeric material, often of high molecular weight,

which exhibits a tendency to flow when subjected to stress, usually has a softening or melting range, and usually fractures conchoidally. As composite matrices, resins bind together reinforcement fibres and work with them to produce specified performance properties.

Void: Any pocket of enclosed gas or air within a composite.

3 Primary references The main sources used for this document are indicated below. At the time of publication, the editions indicated were valid. All standards are subject to revision, and parties to agreements based on this document are encouraged to investigate the possibility of applying the most recent editions of the standards indicated below:


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[1] MARSHALL,A.C. 1994. Composite basics fourth edition. Marshall consulting: USA.

Chapter 9,10.

[2] BAILIE, J.A., LEY, R.P. & PASRICHA, A. 1997. A summery and review of composite

laminate design guidelines, Task 22 NASA Contract NAS1-19347. Military aircraft

Systems Division.

[3] Whitehead, R. S., “Lessons Learned for Composite Aircraft Structures Qualification,”

Proc. 1987 Aircraft/Engine (ASIP/ENSIP) Conference, San Antonio, 1-3 December

1987

[4] McCarty, J. E. and Horton, R. E., “Damage Tolerance of Composites,” paper

presented at 15th Congress, International Council of Aeronautical Sciences, London,

England, September 1986.

[5] McCarty, J. E. and Johnson, R. W., “Durable and Damage Tolerant Composite

Commercial Aircraft Structure Design Approach,” J. Aircraft, January 1978, pp. 33-

39.

[6] Kan, H., Whitehead, R. S., and Kautz, E., “Damage Tolerance Certification

Methodology for Composite Structures,” Proc 8th DOD/NASA/FAA Conference on

Fibrous Composites in Structural Design, Norfolk, VA, November 1989.

[7] Paul, P. C., Saff, C. R., Sanger, K. B., Mahler, M. A., and Kan, H., “Out of Plane

Analysis for Composite Structures,” Proc 8th DOD/NASA/FAA Conference on

Fibrous Composites in Structural Design, NASA CP 3087, 1989.

[8] Anonymous, “Advance Composite Design Guide,” AFML, 1973.

[9] Nemeth, M. P., “Buckling Behavior of Long Symmetrically Laminated Plates

Subjected to Shear and Linearly Varying Axial Edge Loads,” NASA TP 3659, 1997.

[10] Jones, R. M., Mechanics of Composite Materials, McGraw Hill, 1975.

[11] Wang, A. S. D. and Crossman, F. W, “Initiation and Growth of Transverse Cracks

and Edge Delaminations in Composite Laminates, Part 1: An Energy Method,” J.

Comp. Matls. Supplement, Vol. 14, 1980, pp. 71-87.


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4 Fundamentals of Composite design decisions. A Composite design can be affects by some of the following design decisions [1]:

Fibre and load directions is equal

Fibres will only carry loads in the same direction in which the fibres are placed. Fibres that are part of the woven fabrics in other directions, such as crossing yarns, serve other secondary purposes.

Sandwich facings carry loads in facing directions

Sandwich structure facings carry loads only in the direction that lies within the facings’ plane. The core caries any other load or component that runs in another direction, even though the load has to go through the facing to get to the core.

Woven fibres vs. Parallel laid fibres

Woven fabric has a crimp which results from going over and under each other. This structure make the load that it can carry compared to parallel laid fabric with crimp, much smaller. This can result in a strength reduction of at least 5 to 10 % of the fabrics

Resin weak and soft vs. Fibres

When compared to the tensile strength and stiffness of fibres, resins a weak and much softer. Any excess resin in the part can thus result in a lower strength to weight ratio.

Voids degrades strength of laminates

Voids (Air bubbles) trapped in composites do not carry loads and can significantly degrade the laminate’s or sandwich’s strength. Two to three % of voids can cause a reduction of strength of up to 10 to 20 %.

5 Composite laminate advantages and disadvantages

5.1 Advantages The following characteristics are explicit to a well-designed composite laminate [2]:

Good strength and stiffness to weight ratios.

Good Fatigue properties

Design requirements can be met by the tailor-ability of stiffness and strength.

Good Corrosion resistance compared to most aluminium / metal alloys


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5.2 Disadvantages

The following characteristics are considered to be weaknesses of composite designs [2]:

Low inter laminar tension strength

Composite laminates are vulnerable to small out-of-plane eccentricities and loads due to these low strengths. A few examples are given in table 1. [2]

Nonlinear, Most polymer resins have rate-dependant response

Loads in resin matrix-dominated directions may results in the laminate creeping. This is mainly true at elevated temperatures and can result in highly nonlinear stress-strain behaviour. [2]

Micro cracking of polymer matrixes

Structural degrading is not an immediate result of resin matrix cracks, but these crack’s propagation may lead to fuel tank or pressure cabin leaks. This can also result in damage tolerance and degraded durability. Ingression of liquids is also allowed by such cracks. Cyclic freezing/thawing of liquids may accelerate these crack propagations. [2]

Coefficient of thermal expansion (CTE) order of magnitude differences

Laminates that warp upon cooling down from curing is a result of non-symmetric lay-ups. Unacceptable warping and thermal stresses is due to CTE mismatch and transverse to the fibre orientation. Assembly problems and other functional deficiencies are a result of these mismatches. [2]

Strength reduction due to impact-induced damage

The impacted surface will usually not show any damages. These non-visible damages are a big concern for the durability and damage tolerance certification of composite structures. [2]

Polymer resins environmental sensitivity

Exposure to elevated temperatures and moisture may result in the reduction of strength of resin matrix-influenced properties such as shear and compression strengths. High performance military aircrafts have limiting factors of strengths at the Elevated Temperature, Wet (ETW) design conditions.

Reduced electrical and thermal conductivity relative to commonly used structural metals.

Composite structures have a lower thermal/electrical conductivity relative to commonly used structural metals. This can result in the presence of higher thermal gradients and may cause unacceptable thermal/structural responses.

The lessened electrical conductivity influences the response to lightning strikes in two ways:

Direct effects primarily affecting structures

Direct effects are the effects where the energy density is large enough to cause local structural damage or failure. Damage may range from superficial burn marks to skin penetrating on lightning strike attachment points on exterior aircraft. Additionally, overheating problems at joints/fasteners can arise due to low conductivity. Charges built up with fasteners and dumps or arcs from sharp points to adjacent fasteners, this arching may, for example, ignite vapours in fuel tanks as illustrated in Figure 1.


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In-direct effects primarily affecting electronics and electrical subsystems.

This affects the aircraft performance rather than the structure. Composite structures provide less electromagnetic shielding, than metal equivalents, of the interior electronic components. These components may be seriously damaged by static discharge.

Galvanic incompatibility between graphite fibres and metals like Aluminium alloys.

Corrosion damages occur with the incompatibility of materials .The use of carbon-fibre composites with steel structures requires a deeper understanding of the galvanic corrosion theory. This will be properly discussed in section x.

Figure 1: Arcing between unprotected metal fasteners. [2]

Table 1: Composite structure's typical out-of-plane loads. [12]

Configuration Description Application

Flatwise Tension

Skin/stiffener separation due to normal pressure loading

Wing skin/spar flatwise tension due to internal fuel pressure

Transverse Tension

Stiffener web splitting due to transverse

tension loading

Wing skin/spar transverse tension due to chord wise loads

Lateral Bending

skin/stiffener separation due to lateral stiffener bending

Lateral spar bending due to asymmetrical fuel pressure

Post buckling

skin/stiffener separation due to post buckling deformation and loads

Stiffened panels subjected to compressive and/or shear buckling

Curved Panel Bending

inter-laminar stresses due to panel "beam-column" effects

Fuselage skins and frames subjected to bending loads

Thickness Transitions

inter-laminar stresses due to combined loading and local bending

Ply drop offs, build-ups,

and doublers

Irregular Loading inter-laminar stresses and local bending due to axial

Joggles and kinks


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loading in the presence of eccentricity

Bonded Joints

inter-laminar stresses due to local bending arising from eccentricity

Single and double lap bonded joints

6 Guidelines for Composite design The four strengths of composites can easily be out weight by the eight weaknesses with improper designs. To prevent this, a set of guidelines were formulated to take advantages of composite structures. To understand the guidelines, there is certain important theory that needs to be understood which is explained in Section 1 of 3 of the SWP’s on Composite design. (All the guidelines are extractions of [2] BAILIE, J.A., LEY, R.P. & PASRICHA, A. 1997. A summery and review of composite laminate design guidelines, Task 22 NASA Contract NAS1-19347. Military aircraft Systems Division. )

6.1 Fundamental laminate design guidelines

6.1.1 Guideline 1: Lay-up symmetric about mid-surface

It is always good practice to design composites to be symmetric about their middle surfaces due to the following reasons:

To uncouple bending and membrane response

To prevent warping under thermal loading

This guideline is not always easily enforced on sections of thickness tapering, but any asymmetry, due to manufacturing constraints, should be minimized. An unsymmetrical laminate example of complications is the measurement of laminate open-hole tension (OHT) strength. The OHT of most symmetric, fibre dominated laminates can be characterized using a relatively small number of coupon tests.

Figure 2: An example of coupling introduced by unsymmetrical laminates. [2]


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The primary design guideline for composites is to make it symmetrical. For the following reasons:

Simplifies analysis

Simplifies testing

Definition of allowable is improved

Simplify manufacturing.

6.1.2 Guideline 2: Balance Laminates

This guideline suggests that any angled plies, other than 0º and 90º, should occur only in balanced pairs. This implies that every +45º should be accompanied by a -45º for any laminate family of 0/±45/90. A typical example is shown in figure 3.

Figure 3: Balanced and unbalanced symmetric laminates. [2]

A balanced laminate is significant for the following reasons:

The membrane coupling between the in plane normal and shear behaviour is removed since the stiffness coefficients A16 and A26 are both zero.

The laminate bending response may also be simplified.

While adherence to this guideline is prudent is the vast majority, there is however 1 exception; in the case of aero-elastic design of wings. There is mostly a clear advantage to using unbalanced lay-ups to produce extension/compression-shear coupling in the skins of wings. (See chapter 4.2 in reference [2] for more information on this)

6.1.3 Guideline 3: Do not Extrapolate test data

There have been spent a lot of time and effort on the development of aerospace laminated composite structures. Despite this, there is still many applications where original concepts cannot be designed an analyzed with sufficient confidence to flight vehicle certifications. Design development testing must always be used to validate these concepts, of which the following are significant:

1. Typical design guidelines specify fibre orientations in four directions namely: 0º, ±45º, 90º, any other requirement for more directions must be tested.


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2. Failure strains are almost exclusively for uni-axial loadings. Failure criteria for notched laminates subjects to combined loadings needs to be tested.

3. The general composite database is for fibre and matrix specific. Any new materials applied should go through thorough testing

4. The service life and strength reduction on materials or parts, used at elevated environmental conditions, like elevated temperatures or moisture, are very dependent on the situation. Any of these conditions should be tested.

5. Composite laminates are extremely sensitive to out-of-plane loads such as those due to eccentric load paths. Extensive testing will be needed for any eccentric load paths on even the simplest laminate.

6.2 Fibre dominated un-notched laminates Theoretically ply fraction orientations may be specified as anywhere from 0 -100%. For example, one may be tempted to specify a laminate with 100% 0° plies with uniaxial loading

along the x axis. Using figure 4 the following properties are derived:

“These properties are typically very orthotropic, with small strength and stiffness with respect to transverse (y-direction) and shear loadings. This is undesirable for the following reasons:

1. While the primary loading is dictated to be uniaxial, secondary loadings in other directions often exist that are not accounted for in the design. Consider the following example witnessed by one of the authors. A strut designed primarily to transmit axial load for a communication satellite payload failed catastrophically during a thermalvac test meant to expose the delicate electronics payload to the space environment— not to test the structure. Failure occurred when the atmospheric pressure contained inside the strut split it due to hoop stresses. Had only a single hoop ply been provided, this very expensive failure would have been avoided.

2. For non-zero, in-plane shear forces, a situation similar to that of item 1 above develops. The shear strength and stiffness of such panels is small so cracks in the resin matrix can develop easily.

3. The stress-strain relationships of laminates having matrix-dominated characteristics can be highly nonlinear. This is illustrated by the shear stress-strain curve shown in Figure 6-1. The nonlinear stress-strain behavior complicates the prediction of structural response. The in-plane normal stress-strain curve in the resin matrix dominated 2

direction is similarly nonlinear (see Figure 6-1). Hence, any laminate containing only 0° plies would exhibit a significantly nonlinear response to transverse or shear loading. This adds considerably to the complexities of the analysis and the experimental characterization of such laminates. There are strong pragmatic reasons for keeping most major structural global response linear, at least up to design limit load. Foremost among these reasons is the increase of resources consumed in the analysis and testing of significantly nonlinear structures. Local nonlinearity is acceptable in specific situations, such as thin skins in stiffened panels designed to operate into the post buckled range. However, even in this situation, the overall load-displacement relation of the stiffened panel is still more or less linear. Further examples of the complexities arising due to nonlinear, resin matrix-dominated behaviour are discussed by Eckstrom and Spain [Ref. 11].

4. Loadings in the transverse (2) direction of laminates with 100% 0° plies could result in creep. Excessive creep can result in unacceptably low fatigue life.


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5. Laminates with plies in only one direction are susceptible to crack propagation. Once a crack develops in the resin matrix, resistance to its propagation is minimal. A great contributor to the excellent fatigue lives exhibited by fibre-dominated laminates is that the bridging of the fibres across the crack significantly hinders its propagation. In axial tension testing of laminates with saw-cuts, it was found that failure due to splitting (cracking along the main load-bearing fibres,

parallel to the principal load direction) occurs whenever the percentage of 0° plies exceeded 60%. The minimum percentage of 0° plies at which splitting occurs is a strong

function of matrix toughness. 6. Resin matrix micro cracks also allow fluid and gas leaks that are unacceptable for pressure

cabins and fuel tanks. Also, moisture ingress can cause structural damage during the

freeze/thaw cycles that occur as the vehicle altitude changes. For these reasons, Guideline 4 was established.” Abstract from [2]

Figure 4: Room temperature Poisson’s ratio (uxy) for high strength. [8]

6.2.1 Guideline 4: Fibre dominated laminates if at least 10% of plies in basic 4 directions.

There is no documentation substantiating this guideline, but this guideline has been followed by a number of productions with good results. The use of smaller subset of designs, result in usable, more robust laminates which are less susceptible to the weaknesses associated with highly orthotropic laminates. The guideline to 10% plies is sometimes interpreted differently with regards to 45° plies. Two

variations on the rule are:

At least 10% +45° and 10% -45° plies

6% 90° plies provided that there is at least 20% ±45° plies.


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6.2.2 Guideline 5: Primarily loaded plies internal

This guideline states to keep a reasonable number of primary load carrying plies away from the outer surface. This guideline keeps minor impacts from damaging critical plies. The outer plies of thick laminates are damaged by the impact and thus absorbing all the shock and sparing the critical laminates. This guideline does not only safe the total destruction of the structure, but also the pilot.

6.3 Stability of fiber dominated laminates

6.3.1 Guideline 6: Lay-up symmetric about mid-surface

Specific valid guidelines for combinations of plan forms, lay-ups, and loadings are not easily identified, due to the large number of parameters involved in buckling-resistant composite panel designs.

The desirability for symmetric, balanced laminates in buckling-critical structures may result in:

Decreased buckling loads.

Larger plates always have larger buckling decreases in case of uniaxial compression.

Shear buckling loads are decreased.

6.3.2 Guideline 7: ±45° Plies on Exterior

The maximization of the major bending stiffness’s may influence the stacking sequence depending on the loading direction. By locating the ± plies on the outer surface, the buckling resistance can be maximized. This is explained by figure 5 and the following equation: [10]


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Figure 5: effect of orthotropic parameter b and anisotropy parameters g and d on buckling coefficients for simply supported plates subjected to linearly varying edge loads [9]

6.4 Thermal response of laminates

6.4.1 Guideline 8: CTE must be considered in designs

Severe differences in the CTE in the lamina 1 and lamina 2 directions indicate that laminate CTEs are strong functions of lay-up. During the curing cycle cool down, significant residual stress can be build up. Due to this CTE must be carefully considered.

6.4.2 Guideline 9: 10% 0° and 90° plies to avoid thermal expansion.

The laminate membrane thermal strains, for symmetric, balanced laminates, depend linearly on the CTE of the laminate. The CTE is large for a small percentage of 0º or 90º plies. By enforcing the fibre-dominated laminate theory, excessive values of CTE can be avoided.

For bonded or bolted joints to metal structures, the control of laminate CTE is important, since thermal loading of bonded or bolted joints is sensitive to CTE mismatches.

6.4.3 Guideline 10: Use most ductile Resin satisfying conditions

During the cool-down from the stress-free temperature, resin matrix toughness must be great enough to prevent the occurrence of intra-laminar cracks. It is good practice to look at the intra-laminar thermal stresses generated when 0º and 90º plies are laid up in contact as illustrated in figure 4.


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Figure 6: Intra-laminar thermal stresses in 0º and 90º plies

6.4.4 Guideline 11: Max operating temp 10°C below Tg

The laminate’s maximum operating temperature should at least be 10º below the wet transition temperature. At elevated temperatures, in the presence of moisture, the resin matrix’s compression and shear properties degrades highly. This degration is to the softening (plasticizing) of the hot/wet environment to which the resin matrix is exposed. The environment reduces the ability to support the fibres and increases the likelihood of fibre micro buckling

6.5 Stacking sequence and inter-laminar free edge stresses The previously discussed guidelines only refer to overall ply orientation percentages and the ply distribution about the laminate mid-plane. This chapter will focus on the ply grouping and ply orientation angle in coordination with each other.

To address the stacking sequences, the inter-laminar stress at the free edges of the laminates needs to be investigated. (Refer to SWP 42)

6.5.1 Guideline 12: Edge stress controlled in design

At boundary or edge zones, of roughly one laminate thickness from any free edge, the classical lamination theory is invalid. The differences in the Poisson ratio’s between adjacent plies that have different orientations, and differences in the CTE of these plies, causes inter-laminar stress in the laminates subjected to membrane loading.

6.5.2 Guideline 13: Limit layer thickness within laminate less than 0.5mm

Laminates of the same percentage of plies at each orientation angle, but lower stacking sequences than the other laminates, may fail at lower membrane loadings due to the existence of inter-laminar stresses.

This guideline corresponds to four layers of typical carbon fibre of 0.0125 thicknesses; hence this guideline is sometimes referred to as not more than four plies at the same orientation angle.


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Figure 7: Outer ply as free body to illustrate inter-laminar stresses. [2]

Figure 8: Geometry of a free edge delamination [11]

6.6 Poison’s ration mismatch between laminates and bonded or co-cured stiffeners

6.6.1 Guideline 14: Poisson mismatch of skin & bonded stiffener <0.1

A wide range of Poisson ratios can arise in laminate designs as seen in figure 7.. A stiffener which is bonded or co-cured to a laminate is a design feature where performance is sensitive to Poison’s ratio.

There may sound reasons for high percentage designing of the stiffener with 0° plies and the laminate with ±45°, but the tension loads resulting situation can be seen in figure 6. The

stiffener resists the Poisson contraction resulting in severe stress at the skin/stiffener interface.

Figure 9: Room temperature Poisson’s ratio (xy) of usable high strength graphite epoxy

laminates [8]


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Figure 10: Strain due to Poisson’s ration differences in skin and bonded stiffener. [2]

For example: As an example, the initial designs could be as follows: Stiffener is a [60/20/20] lay-up. Enter Figure 9 assuming a [20/20/60] lay-up since the principal load in Figure 10-1 is in the y

orientation. Hence, xy = 0.21. Skin laminate is a [10/80/10] lay-up. Based on Figure 9, xy = 0.52. Industry experience with these structures has taught that the difference in Poisson’s ratio between a stiffener and a bonded or co-cured skin laminate should be limited to a value of ²0.1. The design of the example obviously fails this criterion. It may be desirable to execute the following redesign:

Stiffener is a [60/30/10] lay-up. Entering Figure 6-6 with a [10/30/60] lay-up, xy = 0.36. Skin

laminate is a [10/70/20] lay-up. Based on Figure 6-6, xy = 0.41. This satisfies the criterion of a Poisson’s ratio mismatch less than 0.1.

6.7 Holes, cutouts, impact damage The design of composite structures is complicated by holes and cutouts. Previous sections have considered cases free of holes, impact damage and manufacturing imperfections, but this is unrealistic when designing airframes. It is thus not important to question if holes or damage exist, but to what level of hole/damage the structure should be designed. The influence of notches on compression strength is more significant than the influence of notches on tension strength, illustrated in figure 11.

Figure 11: Influence of defect or damage type of compression strength of typical fibre-

dominated graphite epoxy laminates. [4]

“In selecting the notch type and size to be designed into the structure, the following four points should be considered:


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1. A majority of the laminates in service today are less than 3/8-inches thick. Hence, a vast majority of fasteners used to join these laminates are approximately 1/4 inches in diameter. Therefore, mechanically-fastened laminates may contain a large number of 1/4-inch holes.

2. The fidelity of non-destructive inspection techniques lead to the conclusion that rogue (undetected) flaws had to be accepted in production flight hardware. These rogue flaws included porosity, damaged fibres, small inclusions, and impact damage from dropped tools.

3. There are many databases for military aircraft containing data on the relationship between strength degradation and damage delectability. These data support the idea that an impact causing barely visible impact damage (BVID) produced a level of strength loss approximating that caused by the presence of a 1/4-inch diameter hole. This similarity of the effect on strength of BVID and a 1/4-inch hole is by no means precise. Laminate thickness, lay-up, plan form, edge support, fibres, resin matrices, impactor shape, impact location, and environment all influence the correlations between the effects on strength of BVID with that of a 1/4-inch hole.

4. Measurements of the strength of laminates containing a 1/4-inch hole are generally repeatable and consistent. Use of impacted laminates to measure strength requires careful control of many more independent variables and results in data exhibiting excessive scatter. Hence, the effects of damage are investigated separately on a case by-case basis.” [2]

6.7.1 Guideline 15: Assume hole dia. 6mm anywhere in composite

The initial designed laminate structure must account for the presence of fasteners holes of typically 6mm in diameter.

6.7.2 Guideline 16: <60% Plies at cutouts and bolted joints

The final design of composite laminates must provide sufficient post-impact strength.

6.7.3 Guideline 17: Maximum 60% plies in any direction

This guideline will prevent the laminate of splitting parallel to the principal loadings axis at holes and cutouts. This splitting failure is typical for bolted joints or removable inspection panels.

Fiber dominated composites are relatively brittle and virtually not ductile up to failure, relative to the structural metals, which makes the laminates sensitive to high elastic stress concentrations. Orthotropic material can also exhibit much higher stress concentrations at notches than metal. The effect of lay-up on stress concentration factors at holes is illustrated in figure 7.


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Figure 12: Stress concentrations at large holes in high strength graphite epoxy laminates. [2]

6.7.4 Guideline 18: Reinforcement around cutout should be interspersed

The pad-up around a cutout of the reinforcing plies, could be created in two primary ways. It could be either laid up with all the plies contiguous or interspersed with the laminate. The latter is preferred as the load is absorbed over many lye interfaces rather than a single one.

Figure 13: Interspersing reinforcing plies around a cut-out. [2]

6.8 Joints

6.8.1 Guideline 29: 35% ±45° plies at bolted joints

Laminates are often designed for the allowance of bolted repairs although initially for mechanical fasteners. This has led to constraints imposed on bolted joint designs, regardless of the initial presence of such joints.

An excessive percentage of 0° plies and deficit of ±45° and 90° plies in uni-axially joints can lead to cleavage and shear out failure at unacceptable low loads. [7] The Failure in highly orthotropic laminates involve splitting along the 0° axis, as illustrated in figure 9.


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Figure 14: Failure of highly orthotropic laminates. [2]

6.8.2 Guideline 20: Balance the Membrane stiffness of the adherents.

This means that the adherent’s membrane stiffness needs to be the same on both sides of the joint. Relatively to their balanced counterparts, the unbalanced designs significantly suffer from major strength loss.

6.8.3 Guideline 21: Tapering ends for minimizing of Peel stresses in Thick joints.

The weakness of polymer matrix composites laminates that cannot carry significant peel stresses is a design that needs attention. Peel stress can be described as the inter-laminar tension due to the moment generated near the end of a bolted joint. The moment due to the eccentricity of one adherent middle surface to the other is balanced by this moment. Tapering of the ends or peel-resisting fasteners, near the ends, are most frequently the solution for the high peel stresses in bolted joints

6.8.4 Guideline 22: Scarf /step lap thick joints

All but the thinnest laminates’ peel stresses will be minimized with this guideline. Any adherent greater than 2mm requires stepped lap or scarfed joints. Step laps with the composite forming the outer membrane are preferred for composite-to-metal joints, since the composite can be co-cured to a metal part per-machined with external steps. Scarf joints are preferred for composite-to-composite joints.

6.8.5 Guideline 23: most ductile adhesive that satisfies requirements

A singular virtue in joint design is Adhesive ductility. Low joint strength and greater sensitivity to minor details and tolerances are a result of brittle adhesives. (Refer to SWP 12)


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6.8.6 Guideline 24: Co-cure step lap joints

This is advised for the simplifying of manufacturing. Potential tolerance problems can occur when machining a cured laminate to a close tolerance fitted over a stepped member.

6.8.7 Guideline 25: Ensure adhesive/laminate cure cycles are compatible

Laminates and adhesives are subjected to the same curing cycle in co-bonded joints. This cycle thus needs to result in a complete cure for both. It must also be ensured that the adhesive cure cycle, in secondarily-bonded structures, dos not degrade the properties of the pre-cured laminates. If the adhesive requires a curing temperature near the Tg of the composite, it is likely unacceptable. (Refer to SWP 12)

6.8.8 Guideline 26: Design repairable joints.

The joints may be damaged severely in service, thus providence is needed for adequate space and edge distance to install mechanical fasteners.

6.8.9 Guideline 27: Correct surface preparation of adherents are essential

Light abrasion of bonded surfaces is necessary. Tool release agents and removed peel plies can contaminate the surfaces to be bonded resulting in reduced joint strength. (Refer to SWP 21)

6.8.10 Guideline 28: Corrosion barrier between graphite and aluminum

Due to galvanic corrosion, the bonding laminates containing graphite fibres to aluminum, the adhesive needs to be embedded in a glass fabric scrim of cloth that acts as a barrier between the aluminum and composite surfaces.

6.9 Tapering of skins and flanges bonded to skins

6.9.1 Guideline 29: Taper drop-offs

A taper ratio of at least 20:1 is necessary for laminate thicknesses normal to loading directions. For secondary loading directions, the thickness changes will occur at a taper ratio of at least 10:1. This guideline is stated often in terms of ply numbers that may drop over a given horizontal distance.

6.9.2 Guideline 30: Angle ply pairs should be dropped off together

Applying this guideline will prevent the laminate of becoming locally unbalanced.


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6.9.3 Guideline 31: Outer plies should cover all the other drop-offs

As illustrated in figure 11, this will prevent any edge delamination.

Figure 15: Drop-off designs of plies. [2]

6.9.4 Guideline 32: Stiffeners and Beam flange edges taper <10:1.

A discontinuity in lateral bending stiffness is created by un-tapered flanges. A moment from any lateral loading, and the tendency for the skin to peel away, is a product of the resulting eccentricity. Since flanges are normally parallel to the primary load direction and normal to a secondary load direction, a taper of 10:1 is specified. By tapering the flange with the outer ply, covering the inner ply drops, as illustrated in figure 11, the post-impact strength is improved.

6.10 Damage tolerance, durability, and certification.

An aircraft built of composites has complex certification matters. Real military or commercial aircraft structures must operate satisfactory when damaged. Even in severe damage situations, the damage tolerance may be only the completion of flight and not even the injury to passengers and the crew. The required structural capabilities, as a function of damage extent, vary for different aircraft categories.

Damage tolerance is described as the ability of the structure to resist damage initiation and/or growth for a specified length of time. The durability is an economic issue, but a higher durable structure requires fewer inspections and repairs. Durability is shown by fatigue testing.

Due to the large difference in behaviour of composites to metals, the certification is inherently different. The fatigue life of well-designed, fibre-dominated composites can be much higher than for metallic structures. This is however countered by their environmental sensitivity, impact damage and out-of-plane loadings. An added complication is that most “all-composite” structure do contain metal fitting, and metal fittings require interrogation. For this reason the certification of aircrafts via separate treatment for metals and composites are seldom justified in practise. Use full back ground on this topic can be obtained from References 8,9, 10 and 11.

6.10.1 Guidelines for Certification purposes

The following guidelines should be followed for certification purposes:


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Guideline 33: Durability and damage tolerance must be accounted for during all stages of design

Guideline 38: Laminates will be at least thick enough to withstand minor impacts without damage.

Guideline 39: Design for repairability.

Documents

AMTS STANDARD WORKSHOP PRACTICE - AMT · PDF fileAMTS STANDARD WORKSHOP PRACTICE