Kutz M. Handbook of materials selection

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1138 SPACECRAFT APPLICATIONS OF ADVANCED COMPOSITE MATERIALS

Fig. 2 Longitudinal CTE for typical graphite reinforced laminates.

governed by the matrix properties and typically ranges from 20 to 35

␮

␧/⬚ C.

Most applications of composites have plies oriented at various angles, producing

a net CTE in between the longitudinal and transverse values, depending on the

speciﬁc layup. Shown in Fig. 2 are values for the CTE of carbon ﬁber/epoxy

composite laminates for various layups. For laminates with plies at various an-

gles, ﬁber stiffness signiﬁcantly affects the total laminate CTE since the rein-

forcement ﬁbers are much stiffer than the polymer matrix. In general for angle

ply laminates, as ﬁber stiffness increases, the net laminate CTE will decrease.

This is a major reason why ultrahigh modulus ﬁbers are often selected for pre-

cision spacecraft components.

Although the low CTE of composite laminates is an extremely valuable char-

acteristic, it can lead to new design problems, such as CTE mismatches in metal-

to-composite bonded interfaces. Careful attention must be given to thermal

strains to prevent bond failure or part deformation. Furthermore, because space-

craft often use a combination of composites and metals, the thermal expansion

of the complete structure must be analyzed to ensure that stress levels and de-

formations stay within allowable limits over the entire temperature range.

High Speciﬁc Stiffness and Strength

The most common reason for using advanced composites is mass reduction.

Unidirectional high modulus carbon ﬁber composites can have a longitudinal

stiffness greater than 550 GPa. This is 8.5 times greater than the speciﬁc stiffness

(stiffness/density) of most engineering metals, all of which have approximately

the same speciﬁc stiffness (see Table 2). Because unidirectional laminates have

low transverse strength and stiffness, they are not commonly used in real appli-

cations. Instead, laminates containing plies oriented at various angles are de-

signed according to the principle and secondary loads. Although the speciﬁc

2 USE OF ADVANCED FIBER-REINFORCED COMPOSITES IN SPACECRAFT 1139

[0/+45/-45/90]S

[0/+60/-60]S

[0/+45/-45]S

[0/+15/-15]S

[0]

Aluminum

Normalized Specific Stiffness

Fig. 3 Comparison of speciﬁc stiffness of various IM7/ 8552 laminates normalized

to speciﬁc stiffness of aluminum.

Table 3 Comparison of Typical Pitch Fiber Thermal Conductivities

with Aluminum

Fiber

Fiber Conductivity

(W/m⬚C)

Quasi-isotropic Laminate

Conductivity (W/m⬚C)

Amoco K1100 1100 322

Mitsubishi K13C 640 188

NGF YS-90A 500 151

Amoco P55 120 37

Aluminum 167 167

strength and stiffness of angle-plied laminates is less than that of unidirectional

laminates, it is still superior to metals. Shown in Fig. 3 is the speciﬁc stiffness

of IM7/8552 carbon ﬁber/epoxy laminates for various lay-ups. Note that since

the speciﬁc stiffness is greatest for a unidirectional lay-up, careful selection of

lay-up is key to achieving maximum mass savings.

Thermal Conductivity

Recent interest in using composites for electronics enclosures has led to higher

thermal loads being placed on composite structures. In fact, in electronics en-

closure applications for spacecraft, the thermal demands are often more impor-

tant than mechanical considerations.

Modern pitch-based ﬁbers that have

thermal conductivity exceeding that of aluminum have been developed to help

meet these new demands (Table 3).

The thermal conductivity of composite laminates is similar to their mechanical

characteristics in that they are highly directional and sensitive to matrix selection

1140 SPACECRAFT APPLICATIONS OF ADVANCED COMPOSITE MATERIALS

and lay-up. For typical carbon ﬁber composites, the transverse and through-the-

thickness thermal conductivity is three to four orders of magnitude less than the

longitudinal thermal conductivity. Although this is often viewed as a disadvan-

tage of composites, it can be exploited to design speciﬁc thermal paths within

component modules.

Often a quasi-isotropic lay-up is chosen to simplify ther-

mal and mechanical analysis, but they are generally not optimal for structural

efﬁciency. Unfortunately, analytical prediction of the thermal conductivity prop-

erties for various laminates has proven difﬁcult and quasi-isotropic lay-ups are

therefore selected to simplify design.

EMI Shielding and Electrical Characteristics

Electromagnetic interference (EMI) shielding is an important concern for most

space ﬂight programs. Shielding mitigates the effect of high-energy solar radi-

ation and is used to isolate instruments from one another. Conventional EMI

shields are made of metals such as aluminum and tantalum. Many studies have

been done evaluating the use of composite materials for electronics enclosures

with the primary aim of reducing the total mass of the enclosures while main-

taining acceptable EMI shielding levels.

A typical shielding level required for spacecraft is 60 dB for a frequency

range from 0.10 to 18 GHz.

Although carbon ﬁbers are inherently electrically

conductive, laminates with carbon ﬁber reinforcements generally do not provide

adequate EMI shielding when used alone. Several approaches to increase the

level of shielding have been studied by various groups. Common approaches to

increase the shielding effectiveness include the addition of metal foil or mesh

to the laminate, coating the ﬁnished laminate with metal, or the use of metal-

coated ﬁbers within the composite plies.

Many research groups have measured the shielding effectiveness of carbon

ﬁber-reinforced composite laminates. The standard test method follows MIL-

STD-285 or some variation thereof. Although this standard was originally in-

tended to test the shielding effectiveness of enclosures used in EMI testing, it

is widely used as a procedure to test materials, gaskets, ventilation panels, and

apertures. The test procedure consists of placing a radiation source and a re-

ceiving antenna on opposite sides of a shielded room with a shielded barrier

between them. The material to be tested is placed over an aperture in the barrier

and the signal strength at the receiving antenna is measured. The signal loss is

then taken as the shielding effectiveness.

Test results consistently indicate that the shielding characteristics of unaug-

mented carbon ﬁber composites are much more similar to metals than to plastics.

By adding solid metal laminations or coatings to composite laminates, the shield-

ing effectiveness is roughly equivalent to that of solid aluminum. For multipiece

shield assemblies, experimental testing has shown that joint surface contact re-

sistance is critical for shielding effectiveness. Consequently, nickel and silver

are sometimes selected over aluminum or copper because of their superior ox-

idation resistance.

Recently, highly conductive polymers have been investigated to increase the

shielding effectiveness of composites and eliminate the need for metal coatings.

2 USE OF ADVANCED FIBER-REINFORCED COMPOSITES IN SPACECRAFT 1141

Additionally, research into using high conductivity boron interculated graphite

ﬁbers (boron placed into the interstices of the graphite ﬁbers) has produced

promising results.

Environmental Durability

The impact of space environmental effects on materials is dependent on the type

of mission, and more importantly, the orbit in which the spacecraft operates.

The orbital space is generally divided into three regions based on orbit altitude:

low-Earth orbit (LEO, up to 1000 km), mid-Earth orbit (MEO, 1000–35,000

km), and geosynchronous Earth orbit (GEO, 35,000 and higher). The types and

intensities of various environmental effects depend greatly on orbit altitude and

inclination. In this section the effects of various environmental factors on ad-

vanced polymer matrix composite materials will be discussed, each of which

produce distinct effects on materials. The relative importance of each of these

effects is a function of the orbital placement and must be evaluated for each

mission.

The predominant environmental factors inﬂuencing spacecraft are (1) atomic

oxygen, (2) ultraviolet radiation/solar exposure, (3) micrometeroid and debris

impact, (4) thermal cycling induced microcracking, (5) contamination, (6)

vacuum-induced outgassing, (7) spacecraft charging, and (8) environmental

synergistic effects. Of these, atomic oxygen, vacuum-induced outgassing, and

thermal cycling–induced microcracking are of unique concern to polymer matrix

composites.

The following discussion of space environmental effects on composites is

based on Silverman,

and the reader is referred to it for further explanation.

Atomic Oxygen Effects. Atomic oxygen erodes organic materials causing

surface recession and degradation of optical and thermal properties. Experimen-

tal results from the long-duration exposure facility (LDEF) clearly demonstrate

that atomic oxygen in LEO will erode all polymeric materials, including those

commonly used on spacecraft for thermal and electrical insulation, as paint ve-

hicles, and as composite matrix materials. The rate of erosion varies for different

materials, and in some cases is a function of exposure time. Although there is

no simple way to predict the susceptibility of certain materials to atomic oxygen

erosion, a broad database has been collected that can be used for preliminary

material screening. Of particular interest is data that conﬁrms that polycyanate

matrices have much lower atomic oxygen reactivity than epoxy matrices.

The amount of surface recession due to erosion is directly proportional to the

atomic oxygen ﬂuence (total integrated ﬂux), hence the recession on a particular

surface is dependent on its location on the spacecraft and the surface attitude

relative to the ﬂight path, spacecraft altitude, orbit inclination, and solar activity

conditions. In selecting materials for spacecraft, the designer must be aware of

atomic oxygen effects and calculate the total amount of surface erosion that will

occur. If this value is considered unacceptable, protective coatings that have low

atomic oxygen reactivity may be indicated. Materials successfully used as coat-

ing materials include extremely thin metals, silicon oxide, aluminum oxide, and

silicone room temperature vulcanization (RTV).

1142 SPACECRAFT APPLICATIONS OF ADVANCED COMPOSITE MATERIALS

Vacuum-Induced Outgassing. Polymer matrix composites tend to outgas

due to moisture desorption or evolution of volatiles within the material. The

expelled gases can be especially detrimental to optical instruments and solar

cells aboard the spacecraft. Surfaces may be obscured by condensed outgassing

products or local clouds may be formed that affect sensitive instrument readings.

Additionally, molecular contamination from outgassing can degrade the perform-

ance of thermal control surfaces, especially those that rely on passive radiative

cooling.

Outgassing of composites is characterized by three quantities based on ASTM

E595: total mass loss (TML), collected volatile condensable material (CVCM),

and water vapor regained (WVR). CVCM is a measure of the potential for

outgassed products to form surface deposits, while TML is a measure of the

potential for the formation of molecular clouds. Typical current design require-

ments are

⬍1.0% TML and ⬍0.1% CVCM, but these are expected to become

more stringent for future spacecraft. Outgassing data on speciﬁc material systems

is available in several databases. Of particular importance is the data indicating

that polycyanate and thermoplastic matrix materials characteristically produce

less outgassing products than conventional epoxy matrix materials.

Thermal Cycle–Induced Microcracking. Thermal cycling–induced micro-

cracking is due to the difference in the CTE of each individual ply parallel to

and transverse to the ﬁber direction. The CTE parallel to the ﬁbers is dominated

by ﬁber properties and can be slightly negative. The transverse CTE is controlled

by the matrix properties and is generally about half of the value for the CTE of

the matrix alone. Consequently, internal stresses are induced due to differential

thermal expansion for any composite lay-ups that have plies oriented at different

angles. During repeated thermal cycles cracks form and grow parallel to the

ﬁber direction of each ply.

Microcracking of composites can cause strain hysterisis and signiﬁcant

changes in the overall CTE, both of which present dimensional stability prob-

lems. However, microcracking is sometimes employed to achieve a desired CTE.

For example, the struts on the metering truss of the Hubble space telescope were

deliberately subjected to microcracking to ‘‘tune’’ them to the desired CTE.

New toughened epoxies and polycyanate matrix systems have greatly reduced

the severity of microcracking, but designers should remain aware of the potential

for thermal cycling–induced microcracking.

2.2 Typical Structures

Several ‘‘standard’’ structural forms are used for composite spacecraft structures:

solid laminates, sandwich laminates, strut structures, and isogrid structures. Each

form has beneﬁts and restrictions ranging from cost to performance. This section

will discuss these forms and their applications.

Solid Laminate Construction

Solid composite laminates are the simplest of the forms used in composite struc-

tures. Several plies of a selected composite material are stacked to create a ﬁnal

laminate of the desired thickness. The laminate may be treated as a continuous,

homogenous material once its properties are determined. As such, structural

2 USE OF ADVANCED FIBER-REINFORCED COMPOSITES IN SPACECRAFT 1143

Fig. 4 Sandwich structure construction. (Figure reprinted with permission of the Society of

Manufacturing Engineers, Fundamentals of Composites Manufacturing: Materials, Methods, and

Applications, Copyright, 1989.)

analysis is simpliﬁed and uncertainties are reduced. Solid laminates are best

suited to applications where the primary loading is in-plane tension or shear.

Examples include primary structural members that must withstand the axial

launch and maneuvering loads, and torsion tubes such as those used to mount

deployable solar cell arrays and antennas.

Sandwich Structures

In applications where bending stiffness and strength are critical, sandwich struc-

tures are often used because they provide high bending strength and stiffness

with low mass. Sandwich structures consist of two (or more) facesheets separated

by a low-density core (Fig. 4). This conﬁguration provides substantially greater

bending strength and stiffness with little mass increase when compared to the

facesheets without cores. Sandwich panels with aluminum facesheets have been

used in a wide variety of aerospace applications, providing much higher bending

stiffness and strength values. The desirable properties of advanced composites

such as low coefﬁcients of thermal expansion and directionality also apply to

sandwich structures with composite facesheets.

The most commonly used core material for space applications is aluminum

honeycomb, although other core materials such as aramid honeycomb and syn-

tactic foam are sometimes used. Typical honeycomb cores for space applications

range from 12.7 to 50.8 mm in thickness, and cell size ranges from 3.18 to 12.7

mm. Important properties that affect the selection of various types of sandwich

structures are core venting, core CTE, and the need for discrete hardpoints that

are used to attach the panel to the structure and to mount instruments to the

panel.

Venting. A very important consideration that must not be overlooked is that

the core must allow venting of gases within the core. If this is not done, high

pressure inside the core can lead to facesheet debonding or core explosion in

the space vacuum environment. The cell walls in aluminum honeycomb are

usually perforated for venting requirements.

CTE. Thermal expansion characteristics of the core must also be considered

for successful sandwich structure utilization. The in-plane coefﬁcient of thermal

expansion (CTE) value is usually provided by the manufacturer and is assumed

1144 SPACECRAFT APPLICATIONS OF ADVANCED COMPOSITE MATERIALS

Cylindrical Cross Section

Tapered ends for

truss connections

Fig. 5 Typical composite truss member.

to be constant for all in-plane directions. However, recent experience has shown

that the CTE of aluminum honeycomb core may be different in the ribbon and

ﬁll directions, therefore caution must be used when designing dimensionally

stable sandwich structures. Additionally, the through-the-thickness CTE is dic-

tated by the material that the honeycomb is composed of (e.g., aluminum) so

careful attention must be given to the design of end close-outs and ﬁttings to

ensure compatibility with the sandwich structure.

Hardpoints. Because the facesheets in sandwich structures are very thin,

discrete hardpoints must be integrated into the sandwich panel to provide loca-

tions for mechanical fasteners. The hardpoints prevent fastener pullout and lo-

calized panel crushing by distributing the mechanical loads over a larger area

of the panel. These hardpoints are most often individual metal inserts bonded to

the sandwich panel. While the need for hardpoints introduces additional com-

plexity and expense, careful planning for fastener locations and fastener quan-

tities will minimize the increased fabrication cost.

Truss Structures

Truss structures composed of many individual struts are common in spacecraft

to provide stiff, large structures with low mass. Advanced composites are ideally

suited for strut applications because they carry predominantly uniaxial loads.

This allows the directional properties of ﬁber-reinforced composites to be ex-

ploited, producing a very efﬁcient structure. The advantages of composite use

for mass reduction are augmented by the thermal stability inherent in composite

truss structures, which can be ensured by the speciﬁcation of a wide range of

values for the coefﬁcient of thermal expansion. A typical composite strut is

shown in Fig. 5. Aluminum or titanium end ﬁttings are generally bonded to the

composite strut to provide attachment points.

2 USE OF ADVANCED FIBER-REINFORCED COMPOSITES IN SPACECRAFT 1145

Fig. 6 Isogrid stiffened panel (reprinted from Ref. 11).

Isogrid Structures

Isogrid construction is a new alternative method of fabricating structures that

have high bending strength and stiffness at low mass. Originally developed for

compression-loaded metallic panel designs, the isogrid concept incorporates a

unique triangular pattern of stiffener elements instead of a core structure to

achieve required stiffness.

Several new programs have been initiated to fabri-

cate advanced composite materials structures using the isogrid concept.

An isogrid stiffened panel is composed of three sections, each constructed

from the same base composite material. The three sections are the primary face-

sheet, the web, and the ﬂange stiffeners, as shown in Fig. 6. Because isogrid

stiffened panels have only one complete facesheet, more radiating area is ex-

posed than in conventional sandwich panels. Consequently, isogrid stiffened

structures offer improved heat dissipation over that of honeycomb core struc-

tures. The additional radiating area is especially advantageous in the design of

solar arrays where the efﬁciency of the array is strongly dependent on operating

temperatures. Isogrid construction has the potential to reduce the cost of complex

structures through reduced tooling and assembly costs.

Joints

The design of composite structures must include joining considerations. Lami-

nated composites tend to have poor bearing strength and toughness when com-

pared to metals. Consequently, bonded joints are strongly preferred over the use

of discrete mechanical fasteners for joining composite materials. In contrast to

joints using discrete mechanical fasteners, bonded joints provide a distributed

load transfer that greatly reduces stress concentrations. The performance of

bonded or bolted composite material joints for a variety of loadings and ge-

ometries has been evaluated in many studies.

Results indicate that adhesively

bonded joints offer superior strength and reliability when compared to joints that

use mechanical fasteners such as bolts and rivets.

1146 SPACECRAFT APPLICATIONS OF ADVANCED COMPOSITE MATERIALS

2.3 Manufacturing

No discussion of the application of composite materials is complete without an

overview of the wide range of manufacturing methods available. In fact, the

manufacturing process is integral to the effective and cost-efﬁcient use of com-

posites and must be considered throughout the entire design process. The man-

ufacturing method that is best suited to the application may be dictated by a

combination of total cost constraints, production time, production quantity, and

the physical properties required by the designer. It is impossible to provide a

concise description of all possible manufacturing methods here. A comprehen-

sive text such as Fundamentals of Composites Manufacturing provides speciﬁcs

on the various composite manufacturing methods/processes.

Some of the

unique aspects of composite structure manufacturing will be discussed here, with

a primary focus on cost reduction.

Reduction in component cost and complexity are always welcome attributes

for spacecraft. Because composites are built up and formed by material addition

instead of material removal, consolidation of many individual components into

one entity can be achieved. This consolidation reduces the part count (and there-

fore cost) and results in more compact and efﬁcient structures. Reduction of the

size and mass of entire systems has become more important as spacecraft be-

come smaller. Part consolidation is a key technology to achieving this goal.

New communication and observation systems that use constellations of mul-

tiple small satellites are being designed. Spacecraft design for these programs

represents a departure from the traditional design philosophy because many sim-

ilar or identical satellites must be produced in a costly manner. The use of

advanced composite structures instead of metals may potentially reduce recur-

ring manufacturing costs signiﬁcantly. Most metal parts for spacecraft are pro-

duced by the removal of material from a larger entity to achieve the ﬁnished

form. This machining constitutes the majority of the cost in the production of

metal parts. The use of composite components on the other hand can drastically

reduce recurring machining cost and material waste because a single reusable

mold can be used to produce many near-net shape parts.

An assortment of cost reducing automated or computer-assisted processes are

available that may minimize labor expenses. Commonly used manufacturing

processes include ﬁlament winding, automated tow placement, automated ply

cutting, laser-projected lay-up templates (guides), pultrusion, matched die mold-

ing, and resin transfer molding. Not all processes work equally well for every

structure; instead one or two processes are usually far superior for a particular

structural conﬁguration.

Multifunctional Structures

Mechanical parts, sensors, wiring, and electronic components can be combined

in single composite structures to form integrated, or multifunctional, structures.

Because composite materials are manufactured or built up from many thin in-

dividual plies, individual component elements may be embedded within the lay-

ers as they are applied (referred to as packaging). The resulting multicomponent

integration is a necessary design element in the success of future small satellite

missions. The embedding of passive or active electronic components provides

2 USE OF ADVANCED FIBER-REINFORCED COMPOSITES IN SPACECRAFT 1147

Fig. 7 ARRQEM sensor package.

new opportunities and venues for integrated component diagnostics, sensing, and

structural health monitoring.

Multifunctional structures are broadly deﬁned as structures that support

additional tasks that may be unrelated to basic mechanical load carrying.

Examples of functions included in this deﬁnition are damping enhancement

elements, passive electronics and circuits, sensors, actuators, and active elec-

tronics. Multifunctional structures may include different levels of complexity

that range from simple passive thermal management or electromagnetic wave-

guides to advanced fully integrated adaptive structures containing sensors,

actuators, and active electronics. Research groups have commonly adapted a

stepping stone approach toward the achievement of the ultimate goal of highly

integrated structures with minimum mass and volume. The vast body of work

in this area will not be covered here, but a brief overview of some of the more

mature technologies is provided below.

Embedded-Sensor Technology. Substantial work has been done to develop

techniques for embedding sensors in composite structures and in testing their

response. Much of this research has been focused on the embedding of various

types of strain sensors in composite structures to create integral, real-time dis-

tributed sensing of local strains. The data provided by these sensors is useful

for the determination of the status of deployable structures, in the design of the

shape or form of spacecraft components, in the measurement of vibration levels,

and for damage detection.

Starting in 1994, a program of research began into remotely queried embed-

ded microsensors (RQEM) sponsored by the Naval Research Laboratory. RQEM

developed a working prototype of the sensor and solved many of the difﬁcult

research issues. In 1997 a follow-up program, Applied Research in Remotely-

Queried Embedded Microsensors (ARRQEM), reﬁned the system concept (Fig.

7) and solved many of the difﬁcult engineering issues.

The primary target materials for the installation of remotely queried sensors

were the various aerospace composites, including glass and carbon-ﬁber rein-

forced thermoset and thermoplastic resins. It is undesirable to use ‘‘wired’’ sen-