Kutz M. Handbook of materials selection

Подождите немного. Документ загружается.

1352 USING COMPOSITES

Table 2 Properties of Commonly Used Structural Fibers

Fiber Type

Tensile Strength

(ksi)

Tensile Modulus

(Msi) Density (gm / cm

)

Type E-glass (electrical and structural

applications)

500 10.5 2.54

Type S-glass (highest mechanical

properties)

650 12.4 2.49

Aramid (Kevlar 49) 550 19 1.45

High-strength graphite, aerospace type

(AS-4 or T-300 type)

570 32 1.79

Intermediate modulus graphite, aero-

space type (IM-6)

740 40 1.76

Commercial graphite (Panex 33) 550 33 1.81

ﬁbers contains a waxy, semisolid resin and two powders. The powders are stirred

into the molten resin to produce a consistent mix under controlled conditions.

Subsequently, the ﬁbers are forced into a sheet of molten resin to form the

prepreg. Such a product could never be made on the shop ﬂoor.

Weight and load issues are intertwined because the ﬁber type and ﬁber form

determine how much material is needed to carry loads. Strength-density com-

parisons for ﬁber tyes are shown in Table 2.

Fiber form (conﬁguration) has a signiﬁcant impact on the mechanical prop-

erties of the laminate. Although generalizations are risky, some basic rules are:

●

Loads are carried by the ﬁber while the resin transfers load from ﬁber to

ﬁber.

●

Continuous ﬁbers in the load path yield higher properties than short ﬁbers.

●

Unidirectional tapes yield higher properties than fabrics do because fabrics

have ﬁbers in the off-axes, which may not be needed to carry load.

●

Highly loaded, structurally efﬁcient parts have ﬁbers in four directions:

●

longitudinal—main load path

●

90⬚ transverse—secondary load path

●

25⬚–45⬚ off-axis plies in pairs (Ⳳ)—carry shear

●

To control warpage, the lay-up should be symmetrical about a centerline.

●

Fibers best carry in-plane loads when ﬂat, so thin fabrics produce higher

properties than thick fabrics, due to less ﬁber waviness or kinking during

weaving.

●

Thin tapes and fabrics can be accurately matched to the load requirements

to produce a lighter but more expensive structure.

●

Commercial ﬁbers have ﬁber bundles 4–16 times as large as the corre-

sponding aerospace grades, so these ﬁbers are preferred for tape and

chopped ﬁber applications where optimum weight is not required.

●

Plied laminates have the lowest properties in the thickness direction.

To improve the impact resistance of panels and the joint between the skin

and ﬂanges of co-cured stiffeners, an increase in the normal tension strength is

6 SELECTION OF MATERIAL AND MANUFACTURING CONCEPT 1353

Table 3 Typical Properties of Composites Produced by Different Methods

Process Material Construction

Tensile

Strength

(ksi)

Modulus

(psi ⫻ 100,000)

Density

(lb/ in.

)

Chpped ﬁber spray 30 –50% glass / polyester 9–18 8–18 0.6

Compression mold SMC 25% glass / polyester

SMC 50% glass / polyester

50% glass mat / polyester

8–12

24–30

25–30

9–20

0.76

0.65

0.66

Filament winding 40% glass /epoxy 80–150 40–90 0.1

Pultrusion 40–80% glass roving/ epoxy 60–150 40–60 0.1

35–50% glass mat / polyester 12–30 10–25 0.8

Autoclave molding

prepregs

7781 glass fabric/ epoxy 68 34 0.1

Commercial unidirectional

graphite/ epoxy

145 100 0.063

Intermediate modulus unidi-

rectional graphite / epoxy

(50% in load path)

180 120 0.063

Graphite fabric /epoxy 60 70–100 0.063

Note: Data for Comparison Only—Not to be Used in Design

required. Since stitching the uncured plies puts ﬁbers in the z direction, a me-

chanical engineer was sent to a sewing machine shop in Manhattan’s garment

district. He found a deep throat, manually controlled machine formerly used to

make shoes and handbags. Kevlar thread with a breaking strength of 150 lb was

used to stitch the prepreg prior to cure. In the skin to stiffener ﬂange joints, a

stitch was found to be structurally equivalent to a fastener, but much cheaper to

install. The development of the resin injection process permits the use of sewn

dry fabric, and three-dimensional (3D) woven or braided preforms to produce

composites with excellent normal tension properties.

Woven truss core has been incorporated into the walls of large radomes to

carry heated air for deicing. Braided structures are integrated assemblies having

one or more ﬁber types intertwined in a 3D conﬁguration. Complex shapes such

as sections of an automobile chassis have been produced and have met structural

requirements.

The mechanical properties of laminates using various ﬁber/resin combina-

tions are given in Table 3. The table is intended to illustrate the factors affecting

laminate strength and stiffness discussed previously.

However, one additional factor must be considered in selecting design con-

cepts: the panel stiffness requirement, which can be met by using either a solid

composite panel or a sandwich structure.

The design and fabrication of composite sandwich panels, with plywood,

balsawood, various types of rigid foam or honeycomb cores, is well established.

Walls of large truck trailers, made of panels having core and ﬁberglass plywood

skins, have excellent resistance to both corrosion and puncturing. Sandwich pan-

els are widely used in marine and aerospace applications to save weight of panels

requiring high stiffness.

Having selected potential materials and conﬁgurations, several manufacturing

issues must be considered. The ﬁrst is production rate. Supporting the building

1354 USING COMPOSITES

Table 4 Process Sequence Flow Times

Operation Time (min)

Pultrude one foot of channel 0.1

Injection mold a small part 0.2

Hot stamp/trim part from TP sheet 0.75

Press mold using rapoid curing polyester

SMC

2.0

Filament wind 1# on mandrel 0.1–2.0

Spray chopped ﬁber/resin to form a bathtub 10.0

Lay-up a small part with fabric prepreg 30.0

Autoclave cure a large panel 540

Hydroclave cure a large rocket nozzle 4800

of 20 Boeing 737s per month requires a different mind set than producing 30,000

identical parts per month for a popular sedan. The times to complete processes

are in Table 4. Data for several ﬁber-reinforced thermoplastics are included.

Their high output rates are widely accepted in automobile design. In contrast,

thermosets are limited in high-production applications to press molding of sheet

molding or bulk molding compounds because their process times are as short

as 2 min.

A second issue, facilities, is very complex. If the facilities, personnel, and

experience are available, the project probably will proceed smoothly. Often,

equipment, ﬂoor space, and utilities must be acquired for a new program at an

estimated capital expenditure that is relatively easy to deﬁne. The difﬁculty is

in evaluating the beneﬁt to the company. Intangibles include the cost of entering

a new business, anticipated proﬁt from the speciﬁc operation, stabilizing em-

ployment, and possibly using available space. Even the most realistic evaluation

is meaningless in times of corporate austerity, or if management is in an expan-

sionist mood and willing to accept risk.

Another option is to buy a complex part with many spec requirements from

a specialist who has the expertise, an available facility, and a good track record.

In the boom times of the 1970s, a every large aerospace company had a com-

posites shop. Even so, all the precision nose radomes, except for one type, were

made in Marion, Virginia, because the supplier there was extremely capable and

efﬁcient. However, not all vendors are as competent as this one, a factor that

must be included in make or buy decisions.

Table 5 is an overview of composite manufacturing processes. Except for

pultrusion, all other processes have a common limitation on part shape: Simply

stated, the part must be easily removable from the mold. While it is possible to

use segmented (breakaway) tooling, the combination of assembly, disassembly,

extra cleaning after use, and higher tool repair costs, constitutes an undesirable

cost center. Many parts are made on washout or frangible cast plaster mandrels,

simply because the part geometry prevents removal of any other type of tooling.

For limited production of a complex shape, a composite costs less than its metal

equivalent because the higher cost of the metal forming tools cannot be amor-

tized over enough units. Also, the metal part must be an assembly of individually

made components while the composite part can be molded as an assembly.

1355

Table 5 Design Limitations Imposed by Manufacturing Processes

Process/ Material Cure Method

Inherent Restrictions on Design

(Not including equipment

limitations or special tooling

concepts needed to remove

parts from mold) Comments

A. Spray chopped ﬁber/ resin mix on mold

(may be supplemented by laid-up mat)

Ambient temperature cure or

press molding

● Low properties ● Good for complex shapes

● Low equipment / material cost

B. Hand lay-up of ﬁber resin combinations

on mold (mat, woven yarns and rovings, un-

itapes, wet resin, or prepregs)

All (determined by resin type)

● Ergometrics

● Elapsed time before resin be-

comes unworkable

● Labor intensive, but some automation is available

C. Sheet or bulk molding compound (SMC /

BMC)

Press molding

● Low properties ● Good for complex shapes and high production rates

● Material costs can be low

D. Resin injection into closed mold or vac-

uum bag containing reinforcement

Ambient temperature or oven

or press cure

● Resin ﬂow limits part size

● Molds must be clamped shut

to resist injection pressure

● Good for complex shapes

● Some resins cure quickly

● Compatible with advanced dry ﬁber preforms such

as braided or sewn assemblies that optimize ﬁber

placement in load paths

E. Filament winding of resin impregnated

bands of ﬁbers (may be supplemented by

local fabric patches)

Ambient temperature or overn

cure

● Limited to surface of revolu-

tion shapes

● High throughput

● On-site production of lare structures possible

● Moderate startup costs

● Prepregs yield better properties

F. Pultrusion (mat, fabrics, braided pre-

forms)

RF or heat cure in a continu-

ous process

● Limited to constant cross

sections

● Design restrictions deﬁne

conﬁguration/ wall thickness

● High output

● Moderate capital requirements

● Braided stock raises properties

G. Automated lay-up of carbon ﬁber tape /

fabric

Autoclave cure

● Tape layers limited to gentle

contours

● Fiber placement systems for

complex contours

● N/ C programs are complex

● Small plies not produced efﬁciently

● High material utilization of tape

● High capital costs

● N/ C programs critical to cost

● Automated ply cutting and

laser projectors to guide ply

placement by hand

● Applicable to both tape and fabric

● High output for moderate capital costs

● N/ C programs control material cost

1356 USING COMPOSITES

Part removal is also facilitated by using a springback factor in right-angle

corners (tool 91.5

⬚ to get a 90⬚ corner). Other limits on part size are due to the

equipment used in the curing process. Heat curing is used for several purposes:

●

To rapidly cure the resin as in press molding of SMC (sheet molding

compound)/BMC (bulk molding compound) with 2-min cures.

●

To cure resins having service temperatures about 200⬚F, which are almost

totally unreactive at ambient temperature.

●

To produce low-void composites from dry materials through resin ﬂow

into the microscopic gaps between ﬁbers and the plies.

●

In these processes, the resin cure requires heating in an oven, press, or

autoclave (pressure vessel), which introduces limitations of size and ther-

mal capability.

The last process variable is curing pressure, which is used to densify the

laminate by suppressing porosity. Techniques for applying pressure include:

●

Contact molding—apply paint or rubber roller to wet lay-up to roll out

some of the entrapped air.

●

Wrap with shrink tape to apply pressure during oven cure.

●

Vacuum bag molding—seal the part with a plastic ﬁlm and evacuate to

13–15 psi; cure at ambient temperature or in an oven.

●

Autoclave molding—place the vacuum-bagged part in a heated pressure

vessel and pressurize with inert gas to 45–150 psi (high recurring and

facility costs).

●

Hydroclave molding—place the vacuum-bagged part in a pressure vessel

ﬁlled with water. The pressure is developed by thermal expansion (high

recurring and facility costs); 250–500 psi.

●

Press molding (rapid part production possible; 50–1000 psi.

The last factor in choosing materials and processes is market value, or the

perception of what a product is worth. This varies with application. In the highly

competitive automobile industry, large volume demands the lowest price possi-

ble, and there are few large structural composite applications in production due

to cost. The cabs of large trucks, made at the rate of 60,000 per year, trade

tooling cost against labor cost, which beneﬁts composites. In contrast, luxury

items such as the Corvette, premium tennis rackets, and composite kit airplanes,

have a cachet that increases the sales price. As an extreme, high-performance

race cars and America’s Cup yachts can command a hefty premium over similar

products made from conventional materials. Price competition in all areas of the

aerospace industry has forced price reductions obtained through production ef-

ﬁciencies, which reduced scrap and rework, and produced cheaper materials that

are easier to use.

7 DETAIL DESIGN

An obvious design requirement is that parts must be easily removable from

tooling without damage. Design details to be avoided include (1) reverse ﬂanges,

7 DETAIL DESIGN 1357

which require segmented tooling; (2) molded-in stiffness perpendicular to the

direction that a part is removed from a mold, except on the bottom of a box;

(3) lack of an adequate taper (draft); and (4) surface buildups, which tend to

‘‘hang up’’ on a mold during part removal.

These principles are illustrated by a press-molded junction box used with

buried telephone lines. Flanges to which the access door is attached project

beyond the box. The box is molded with the ﬂanges up, so the stiffeners on the

opposite face are half-rounds. They are conﬁgured to be horizontal to provide

maximum support to the vertically positioned terminal strips inside the box. The

stiffeners on the remaining four faces of the box are blade conﬁguration, tapering

in both width and height away from the ﬂange. The design is tailored to simplify

removing the part from the tool. The part is molded in 2 min, then is popped

out of the tool quickly.

Five of the eight fabrication schemes involve lay-up of plies on mold, re-

gardless of how the part is cured. In many cases, this operation represents the

majority of labor costs. To become more efﬁcient, both the number of plies and

the design of individual plies must be optimized. First, the number of plies must

be reduced, if possible, by using thicker materials, if they are available. A side

issue is weight penalty. In aircraft ﬁberglass parts, net sections are usually less

than 0.1 in. thick, and often are 0.05 inch thick, especially in sandwich construc-

tion. For these applications, the preferred dry glass fabric is style 181, which

weighs 8.9 oz per yd

and yields a composite thickness of 0.01 in. Boat hulls,

chemical storage tanks, and swimming pools require thicker laminates and are

less weight critical. So woven roving fabrics ranging up to 52 oz per yd

, with

a ply thickness of 0.05 in., are used for these applications. The weight of

chopped ﬁber mat varies from 0.75 to 3.0 oz per yd

. In addition to reduced

labor cost, thicker materials cost less per pound because throughput is increased

at the material supplier’s facility.

A second issue involves tailoring plies to simple shapes, which can be nested

in a cutting pattern, which reduces materials usage. Often, edge plies are used

at the edge of panels for bearing strength. They should be cut as strips. For

high-rate production, automated ply generation should be considered. Die cutting

is applicable for repetitive shapes. In plants producing large numbers of com-

posite parts repetitively, N/C cutting machines are used to cut and identify the

plies or ply segments. Then the plies are kitted and sent to the lay-up ﬂoor.

These N/C ply cutters come from the garment industry. A roll of material up

to 60

ⴖ wide is unrolled on a 20⬘ long table. A computer controlled cutter cuts

out individual plies of any shape. A second head uses ink to identify each piece.

The trick is to locate the pieces close together to avoid wasting material.

A good method to reduce the number of plies is to convert from a monolithic

panel to sandwich construction. Several low-cost methods are available.

Flat panels can be laid up and cured in a multicavity press. A widely used

system for movable interior partitions used ﬁberboard skins, paper honeycomb,

and ﬁlm adhesive that cured rapidly. The boat building industry is licensing the

patented SCRIMP process developed by Seaman Composites. The process has

been used to make individual ﬂat sandwich panels 18 ft

⫻ 40 ft, which are

sequentially co-cured into a multisided assembly. The skins are polyester-

impregnated rovings, and the core is foam slabs prespliced with room temper-

ature curing adhesive. The resin is injected into both skins while the assembly

1358 USING COMPOSITES

is vacuum bagged. The resin is pressurized in conventional paint spray pots and

is uniformly distributed within the lay-up by a proprietary method. The panels

have withstood extensive testing with excellent results. The process is attractive

because structural panels of any conﬁguration are produced at low material and

labor costs with minimal facility requirements.

A third way to produce sandwich panels is co-curing in an autoclave. The

core is cut to size, sometimes preformed to the appropriate shape, tapered at the

edges if needed. The core is laid over the adhesive and a toolside skin and is

itself covered with a second adhesive ply before upper (inner) skin is laid up, a

vacuum bag is installed, and the lay-up is cured under positive pressure. For

helicopters operating at temperatures below 180

⬚F, the core is a polyamide foam

and the panel cure is 1 h at 225

⬚F at 45 psi. The secret of the process is that

the skin lay-up, adhesive, and core deform sufﬁciently under the applied load to

ﬁt snugly together to form structural bond lines. Sandwich panels requiring

higher loads and service temperatures are produced using nonmetallic honey-

comb cores. Precautions must be taken, however, in the tool design and part

setup to prevent core crushing from curing loads applied to the edges of the

core. These three approaches to sandwich construction all have some provision

for localized patch plies, so tapered skins are feasible. In most cases, the core

does not have to be processed to accept the extra plies, which provides a sig-

niﬁcant cost reduction.

The vast majority of composite parts require trimming because they are made

oversize intentionally. Generally, it is very difﬁcult to lay up and cure a large

part without edge ﬂaws. Unfortunately, composites are very abrasive and soon

wear out the high-speed steel/carbide cutters used in aluminum fabrication. The

least expensive method of cutting is with a ﬂuid-cooled diamond grit blade. So

the preferred part designs have edges that are straight. Cutouts in panels should

be minimized, but they can be routed using special design cutters operating at

a speciﬁc range of revolutions per minute. Initially, cutouts were net molded

using tooling to cut the prepreg plies accurately, then the hole was net molded

to design conﬁguration during cure. However, improvements in trimming pro-

cesses have made net molding less attractive.

Waterjet systems used for trimming develop a narrow, focused beam of water

at pressures up to 60,000 psi. The water jet travels up to 2800 ft per minute.

The high energy of the water jet produces localized failure of the composite to

yield a clear-cut edge. The cutting efﬁciency of the water jet can be improved

by incorporating powdered garnet into the stream. Both systems have no restric-

tions on shape, and may be either manually or computer controlled. Computer-

controlled routers are also available that trim panels of any shape efﬁciently.

Both methods are suited to high-rate production. So there are three approaches

to trimming, but there is a cost in departing from sawing straight cuts.

Many composite parts go into assemblies such as hulls or decks, bridge repair

panels, or major aircraft sections. The two primary methods of joining separate

parts are mechanical fastening and structural bonding. The selection of assembly

method is limited by the cohesive strength of the available adhesives. At load

intensities above the capabilities of the adhesive, fastening is mandatory. Con-

versely, lightly loaded joints should be bonded due to the complexities of fas-

tening.

7 DETAIL DESIGN 1359

Mechanical fastening of composites is more difﬁcult than riveting or bolting

metallic structures for several reasons. In addition, there are more problems

fastening weight-critical graphite/epoxy structures than ﬁberglass or Kevlar as-

semblies. In fact, structural analysis and manufacturing techniques for highly

stressed bolted joints were the last major design and manufacturing areas to be

fully understood and implemented in production.

As stated previously, composites are very abrasive, so carbide drills are used

extensively with special point/ﬂute conﬁgurations, often in conjunction with

motors that advance the drill a precise amount into the workpiece for each

revolution. Even so, many automated drilling systems automatically change drills

as often as after preparing only 40 holes. Drilling can produce structural damage

around the holes, which can affect structural integrity. This damage includes

extensive delamination of the ply where the drill exits the workpiece, localized

delamination around the hole, and ﬁber shredding of the aramid ﬁbers in the

hole region.

Hole preparation requires a gentle touch. To further increase the difﬁculty,

many fasteners require a precision sized hole such as 0.188 in. pin requiring a

0.189 to 0.191-in. diameter hole. The countersinks for the fastener heads must

be tightly controlled to obtain a good ﬁt so the heads can bear uniformly against

the composite.

Fastener selection is a design issue. For ﬁberglass and Kevlar structures, con-

ventional design fasteners may be used. Stainless steel fasteners are required for

wet or corrosive environments.

A whole new family of fasteners is required for graphite/epoxy structures for

several reasons. First, the composite is so cathodic that only titanium fasteners

are corrosion resistant. A-286 fasteners pit, and monel generates heavy corrosion

products that do not appear to reduce joint strength. Also, because carbon ﬁber

composites are more sensitive to high bearing loads than metals, fastener heads,

nuts, and washers should be enlarged to provide as much bearing surface as

possible to prevent pullout. The larger area also spreads out the high clamping

(preload) force desired for good strength and life, so that the laminate is not

damaged by matrix crushing. Fastener edge distance and spacing have been

studied extensively; 3D (D

⫽ fastener diameter) edge distance and 5D spacing

are used widely. Many fasteners are installed with wet sealant as a corrosion

preventative. Note of caution: The design of highly loaded joints and the selec-

tion of fasteners requires specialized knowledge. Fastener hole preparation and

installation also require speciﬁc knowledge and preparation of detailed and un-

derstandable instructions for mechanics.

The last major problem with assembly of composites is their lack of ductility.

Metallic structures can be clamped together to produce an acceptable ﬁt. In one

case, an aircraft’s wing covers were machined in the ﬂat and forced to ﬁt the

curved substructure by hydraulic clamps. Such a procedure is guaranteed to

cause severe structural damage to composites by internal delamination. Thus,

mating composite parts must be made to ﬁt by shimming with precured com-

posite shims and paste epoxy adhesive that cures at ambient. In lightly loaded

joints, the shim adhesive can be reasonably thick if the composite members are

overdesigned, so that bearing failure will not occur. In highly loaded joints, the

shim thickness is reduced because of bolt ﬂexibility. In graphite/epoxy overlap

1360 USING COMPOSITES

specimens pulled in tension, the joint strength was reduced 40% when the shim

thickness was increased to 0.06 in. from 0.03 in. The 0.125-in. titanium bolts

were kinked locally because the shim could not resist the applied load as well

as the composite could.

As a result, the aerospace industry has agreed on the following procedure for

resolving gaps in bolted joints:

Gap Procedure

0–0.005 in. Do nothing

0.005–0.03 in. Use liquid shim

Greater than 0.03 in. Use solid shim to reduce gap to less

than 0.03 in.; then use liquid shim

Good design practice for composite assemblies involves controlling the faying

(mating) surfaces by tooling. The substructure of a torque box assembly is usu-

ally tooled to control the chordal height of the spar and ribs. If the cover tooling

reproduces the substructure contour, the cover will ﬁt reasonably well. The cover

thickness variations are thus reﬂected away from the joint to simply ﬁt up. The

mechanical assembly of a complex graphite/epoxy structure is a challenging,

difﬁcult, and expensive operation. However, through the design, manufacturing,

and quality control interface, it can be made manageable. Also, as the assembly

is understood, adjustments and reﬁnements to the design and manufacturing

process can be made to expedite production and reduce costs.

The bonding option presents two alternatives that manufacturing prefers to

mechanical assembly: (1) structural bonding of individual parts and (2) an in-

tegrally molded assembly produced by curing several lay-ups concurrently.

The manufacturing objective for structurally bonded assemblies is to produce

an acceptable unit with a minimum of effort with little rework or scrap. To

achieve these aims, the process must provide for:

●

Cleaning the mating surfaces to prepare them for bonding

●

Good ﬁt up of mating surfaces to glue-line tolerances

●

Locating the parts in the proper orientation during adhesive cure

●

Using a thin ﬁberglass/nylon cloth (scrim) to guarantee a continuous bond

line

●

Obtaining a consistent bond-line thickness within speciﬁed limits.

The surfaces of individual parts are contaminated with mold release to prevent

them from sticking to the mold, dust from trimming, and ﬁngerprints. Surfaces

can be prepared either by abrasion followed by a solvent wipe, or removal of a

sacriﬁcial external peel ply. Peel ply is preferred for large areas and local areas

at the center of a large area. Highly stressed aircraft parts are usually made with

a peel ply to avoid damage to the ﬁbers.

Part ﬁt up to bond-line tolerances is mandatory for structural integrity and

fatigue resistance. The optimum process consists of a balanced lay-up to prevent

warpage and geometrically accurate tooling to control the contour of the faying

7 DETAIL DESIGN 1361

surfaces. However, these preferences cannot always be granted. Often, for cos-

metic purposes, the tool surfaces must be away from the joint, as in the case of

boat hulls. Consider the example of a very large ocean-going ﬁberglass/polyester

hull approximately 3.5 in. thick. Thickness variations could be almost

Ⳳ0.25

in. thick. Using surveyors’ techniques, shipyard workers measured the wall-to-

wall distance and laid out the trim line on the bulkhead 0.125 in. undersize. The

⫻ 30 ft bulkhead was lifted into the hull at the desired station and local

interferences sanded until the bulkhead was seated in the desired position. A

gusset was cured in place on one face of the bulkhead while it was supported.

After cure the scaffold was removed, and the gap between the hull and bulkhead

was ﬁlled with glass fabric and resin before the other gusset was laminated in

place. This is a large-scale example to explain the concept of sanding or patching

parts to obtain the proper glue line for the adhesive being used.

A factor in joint strength is width. Peak stresses occur at the edges and decline

toward the center. The average allowable stress decreases with length at a certain

point because added area contributes little load-carrying capability. While man-

ufacturing prefers narrow joints to ease the ﬁt up problem, engineering also

beneﬁts through weight reduction. More importantly, manufacturing prefers bond

lines that are in straight-line elements on ﬂat planes. Bonding over bumps is

risky due to the high potential for improper bond-line thickness. Antipeel fas-

teners reduce bond-line stresses in highly stressed or complex-geometry bonded

joints.

Two categories of adhesives are used in composite structures:

●

Paste or liquid adhesives, typically mixed on the shop ﬂoor and cured at

ambient (bond-line thickness 0.002–0.02 in.)

●

Calendered ﬁlm adhesives that are cut to size, placed in the joint, and

cured at elevated temperature under applied loading (bond-line thickness

0.001–0.01 in.)

The paste adhesive tolerates a thicker glue line with larger tolerances than

the ﬁlm adhesive. Thus, if the analysis, and weight considerations, can tolerate

paste adhesives, they are preferred because parts can have larger tolerances in

the ﬁt up.

In addition to maintaining an adequate bond line by satisfactory ﬁt up, a thin

cloth is used to prevent adhesive squeezeout. Individual parts also must be po-

sitioned relative to each other as speciﬁed by the engineering drawing. Small

assemblies are bonded in a holding ﬁxture. As assemblies become larger, ﬁxtures

become more cumbersome. An alternative is to provide coordinating holes in

the design so parts can be pinned together during bonding. These holes can be

subsequently used for antipeel fasteners, installation of systems, or to make joints

fail-safe with structural fasteners.

Despite the precautions needed to assemble bonded structures, bonding is still

much easier for manufacturing than fastening.

The concept of co-cured structure is an attempt to eliminate the assembly

problems discussed previously. The idea is to place several lay ups on the mold

at once to produce a panel. The application to foam sandwich and honeycomb