Kutz M. Handbook of materials selection

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1148 SPACECRAFT APPLICATIONS OF ADVANCED COMPOSITE MATERIALS

sors (strain gauges, optical-ﬁber-based grating sensors, etc.) on these materials

due to difﬁculties in placing the sensor and maintaining the connecting wire or

ﬁber during composite fabrication. For wired sensors, the connection requires

slits or holes in the layers, or it must extend across the layer to the edge of the

lay-up. Either method is tedious to produce and introduces ﬂaws or potential

ﬂaws into the composite material, because structural integrity is compromised.

Flaws can be minimized through the use of very ﬁne wires, but they are delicate

and often break during lay-up, which is also a problem in the use of optical

ﬁber. To avoid these difﬁculties, the U.S. Ofﬁce of Naval Research (ONR)

funded development programs for wireless sensors that can be powered and

queried without physical connections.

ONR had several high-level requirements for the ARRQEM sensor system.

The primary requirement was the measurement of various physical phenomena

(e.g., strain, pressure, temperature) within a composite structure. Other require-

ments included:

●

Low recurring cost (including installation costs and the use of consumable

components such as sensors)

●

Usable lifetime commensurate with the instrumented structure (50 years

in some cases)

●

Interrogation possible from a minimum distance of several inches without

physical contact

●

Accommodation of a varying number of different or similar types of sen-

sors in close proximity

●

No signiﬁcant effect on the instrumented structure (e.g., do not exceed

the critical ﬂaw size)

●

Correlation of the measured effects with true effects in the observed struc-

ture

●

Safety in handling and use during structure processing and service life

Several of these requirements (e.g., lifetime, ﬂaw size, safety) preclude the use

of embedded batteries.

In the program an embeddable strain gauge rosette (shown in Fig. 8) was

successfully developed and limited testing was conducted. Further development

in this area will lead to the application of this technology to satellites and other

complex composite structures.

Integrated Electronics. Traditional electronic systems designs provide for

one or more printed wiring boards (PWBs) that may be surface mounted on

primary structures or enclosed by a separate structural housing. Although this

method is simple and time tested, mass and volume are not minimized because

PWBs are usually nonstructural elements. Composite materials can be used to

create printed wiring boards that function as both electrical and structural com-

ponents. This then eliminates the need for a separate support structure, i.e., the

housing, leading to drastic reductions in the mass and volume of electrical sys-

tems—an important factor for small spacecraft design.

3 EXAMPLE SPACECRAFT APPLICATIONS 1149

Fig. 8 ARRQEM system concept. Wireless interrogation of inexpensive embedded sensors

are conducted by external reader device. Digitally-represented sensor readings are transferred

to the device for storage or analysis.

Table 4 Considerations in Spacecraft Material Selection

Property

Truss

Members

Primary

BUS

Deployable

Structures

Electronics

Enclosures

Optical

Benches

Mirrors/

Reﬂectors

Mass ⻫⻫⻫⻫⻫⻫⻫⻫⻫⻫⻫⻫⻫⻫⻫⻫

CTE thermal

stability

⻫⻫⻫⻫⻫⻫⻫⻫⻫⻫⻫⻫⻫

Thermal conductivity ⻫⻫ ⫻ ⻫⻫⻫⻫⻫

Electrical

conductivity

⻫⻫ ⫻ ⻫⻫⻫⻫ ⫻

Cost ⻫⻫⻫⻫⻫⻫⻫⻫ ⫻

Stiffness ⻫⻫⻫⻫⻫⻫⻫⻫⻫⻫⻫⻫⻫⻫⻫⻫

Strength ⻫⻫⻫⻫⻫⻫⻫⻫

Note: ⫻ : not critical, ⻫: somewhat important, ⻫⻫: very important, ⻫⻫⻫: mission critical.

3 EXAMPLE SPACECRAFT APPLICATIONS

Because the total allowable mass of complete spacecraft systems is limited by

the launch vehicle constraints, the minimization of structural mass is required

to maximize the allowable payload mass. This minimization can lead to the

extension of spacecraft operating life because more fuel can be carried, or ad-

ditional scientiﬁc payloads can be accommodated. Use of advanced composite

materials is a key factor in reducing the structural mass of spacecraft but also

provides other advantages such as low thermal expansion and manufacturing

ﬂexibility. Table 4 shows the most common spacecraft structures and the relative

importance of structural properties to each.

Speciﬁc applications and examples of composite material use in space struc-

tures will be described in the following sections. The space structures are clas-

siﬁed as either primary or secondary structures, and are further grouped by

speciﬁc component classes. For spacecraft, primary structures are those that are

critical to achieve the basic mission, for example, surviving launch and on-orbit

1150 SPACECRAFT APPLICATIONS OF ADVANCED COMPOSITE MATERIALS

station keeping maneuvers. Secondary spacecraft structures are any that are

related to the performance of speciﬁc mission functions, such as instrument

support structures. The information in this section is intended to provide an

overview of the existing state of the art of composite spacecraft structures and

to address the associated advantages and limitations.

3.1 Primary Spacecraft Bus/Chassis Structure

Advanced composites are being used more frequently for primary spacecraft bus

structures to reduce mass and increase precision in large, one-of-a-kind space-

craft and smaller high-production spacecraft. Typical designs employ multiple

struts that support stiffened panels that serve as decks for mounting instruments.

Examples of both types are described below. Issues associated with the use of

advanced composites for primary spacecraft bus structures will also be ad-

dressed.

Example: FORTE

An example of modern, low-cost, low-mass composite spacecraft primary struc-

tures is the fast on-orbit recording of transient events (FORTE) spacecraft, de-

veloped jointly by the Los Alamos National Laboratory and Composite Optics

Inc.

The compressed development schedule of this small spacecraft dictated a

simple, modular design with a minimum number of unique parts. Bulk fabri-

cation methods may be efﬁciently employed for the production of multiple parts,

and cost and complexity are reduced. Because the same tooling is used to pro-

duce multiple units, the resulting parts are dimensionally consistent and thus

interchangeable.

The spacecraft structure consists of three decks and three octagonal frame

cages (trusses). The decks are composed of carbon ﬁber/epoxy laminates and

aluminum honeycomb core that are bonded to form a sandwich panel. Copper

foil is co-cured onto the external surfaces of the sandwich panels to provide

positive grounding and EMI shielding. All of the sandwich panel facesheets are

0.762 mm thick and have a ply layup of [0/60/120]

. Aluminum-threaded inserts

are bonded into each deck arranged with hole patterns and hardware sized spe-

ciﬁcally for each component. A total of 25 different components are mounted

on the three decks.

For simplicity and expedience, all of the frame parts are cut out of 1.22-mm-

thick [0/45/90/135]

ﬂat carbon ﬁber/epoxy laminates. A computer-controlled

waterjet cutter was used to cut out the frame parts from two large ﬂat laminates.

The parts were then ‘‘snapped’’ together using mortise and tenon joints. Finally,

the parts were placed in tooling ﬁxtures and bonded together. The use of this

streamlined manufacturing process provides rapid and repeatable fabrication of

the structures and eliminates the need for multiple unique and expensive tools.

The original design of the FORTE spacecraft speciﬁed an all-aluminum struc-

ture. The use of composite materials for this design led to a mass reduction of

approximately 50% over the aluminum structure. Additionally, the low-mass

composite design improved the inertial properties of the FORTE spacecraft. Fi-

nally, because of the utilization of advanced design and manufacturing technol-

ogy, the cost of the composite structure was comparable to that of the equivalent

aluminum structure.

3 EXAMPLE SPACECRAFT APPLICATIONS 1151

Fig. 9 Composite strut structure of MSN spacecraft (reprinted for Ref. 16).

Example: Midcourse Space Experiment

The Midcourse Space Experiment (MSX), a large sensing platform, has precise

pointing and tracking requirements that are critical for mission success.

The

spacecraft is divided into three sections: an electronics platform/launch vehicle

interface, a thermally stable truss section, and a temperature-controlled instru-

ment section. The truss section forms the center of the spacecraft and provides

the primary support for the Space Infrared Imaging Telescope III (SPIRIT III).

Stiffness, rotational stability, minimization of mass, low coefﬁcient of thermal

expansion, and low thermal conductivity were critical design requirements that

could only be satisﬁed by the use of composite materials for the truss members.

The truss section is shown in Fig. 9. It consists of 20 carbon ﬁber/epoxy

I-beams connected with titanium ﬁttings and carbon ﬁber/epoxy plates. Three

different sizes of I-beams are used, classiﬁed according to their location: lower

diagonals, upper diagonals, and verticals. Customized titanium ﬁttings join the

I-beams, provide mounting interfaces for the SPIRIT III instruments, and con-

nect the truss section to the upper and lower spacecraft sections.

Carbon ﬁber/epoxy gusset plates are used to join the upper and lower diag-

onals (Fig. 10). Gusset plates are commonly used for planar joints of composite

or metal members. When composite beams and/or gusset plates are used, ad-

hesive bonding is the preferred method of joining rather than the use of me-

1152 SPACECRAFT APPLICATIONS OF ADVANCED COMPOSITE MATERIALS

Fig. 10 Example of typical gusset plate joint.

chanical fasteners. This is an important difference between laminated composite

structures and metal structures, and one that must be considered for successful

design.

The carbon ﬁber/epoxy system selected for the MSX I-beam construction

was T50/ERL1962. T50 carbon ﬁber reinforcement was selected for its high

stiffness values coupled with minimal thermal conductivity. The ERL1962 resin

(Amoco) was selected because it is a toughened epoxy system that exhibits good

resistance to microcracking. The I-beam ply lay-up, [0

/60/⫺60]

, results in

maximum axial stiffness. Prior to assembly an eutectic coating was applied to

control moisture uptake, and the beams were then painted with Chemglaze Z306

ﬂat black polyurethane paint. The total mass of the truss assembly is 47.1 kg.

3.2 Deployable Structures

Deployable structures are deﬁned as those that can be stored in a small volume

conﬁguration for launch and that are later deployed for space operation. De-

ployable structures are required elements for many missions to accommodate

the need for large spans or planar components that are impossible to launch in

their extended conﬁguration. The most common deployable structures in use

today are solar arrays, which are folded for launch. The recent delivery of the

additional solar panels to the International Space Station is an example of this

small-volume launch to large-volume expansion. Current and future missions

will require deployable structures to support large baseline optical instruments,

space-based radar applications, experiments that require environmental isolation,

or solar sails.

Deployable structures can be classiﬁed as either rigid or inﬂatable. Inﬂatable

structures require hydrostatic (gas) pressure for deployment, in contrast to rigid

3 EXAMPLE SPACECRAFT APPLICATIONS 1153

deployable structures that use purely mechanical means for deployment. Inﬂat-

able structures may be subdivided into three groups: continuously inﬂated struc-

tures, rigidized inﬂatable structures, and single inﬂation structures.

Inﬂatable Structures

When compared to rigid deployable structures, inﬂatable structures generally

have smaller packaged volume and lower mass. This technology is a signiﬁcant

contributing element in the design of future large solar power systems.

There

are three categories of inﬂatables, which are abbreviated as follows: continuously

inﬂated (CI), rigidized inﬂatable (RI), and single inﬂation (SI).

CI structures

require gas pressurization over the duration of their service life. Inﬂatable rafts

used for recreational water sports are a typical example of CI structures. RI

structures require inﬂation for initial deployment support only until the structure

is rigidized, usually via planned material property changes. SI structures are

designed for one-time, usually short-duration use, e.g., the airbag system for the

Mars Pathﬁnder Lander.

Both the mass of the structure itself and the mass of the deployment gas,

together referred to as the ‘‘wet’’ mass, must be considered in the spacecraft

design process. Since rigidized inﬂatable structures generally have the lowest

values for wet mass they are the best choice for most inﬂatable applications.

Consequently, the majority of inﬂatables research has been dedicated to RI de-

velopment. Some applications, however, such as habitation modules, short-

duration experiments, and landing systems are best solved by utilizing CI or SI

structures.

Continuously inﬂated structures must remain pressurized throughout their en-

tire service life. Some beneﬁts include form deﬁnition and simplicity of tech-

nology required. Since a pressurized membrane is consistently taut, CI structures

have better surface deﬁnition than rigidized inﬂatables. This surface deﬁnition

is particularly useful in antenna and reﬂector applications. Additionally, a CI

structure is technologically simpler than a RI structure because no material trans-

formation is required for operation. However, sufﬁcient inﬂation gas must be

supplied for the entire operational life span of the mission, and this requirement

affects the mass parameters of other mission elements. The total quantity of gas

required varies according to several factors, e.g., inﬂation pressure, estimated

gas leakage, total inﬂated volume, and mission life span. CI structures are there-

fore not ideal for long-term or long life-cycle applications because mission op-

eration would require a prohibitively large amount of inﬂation gas.

Rigidized inﬂatable structures are better suited to long life-cycle applications

because only the amount of inﬂation gas needed for initial deployment is re-

quired. Several approaches to rigidization have been investigated, and polymer

matrix composite materials appear to be the material of choice. Rigidization is

usually achieved through a chemical or phase change of the polymer matrix

(such as curing), which increases its stiffness and therefore the stiffness of the

whole composite structure.

3.3 Electronics Enclosures

Lightweight advanced composites electronics enclosures have been developed

for many different vehicle applications, including spacecraft. Electronics enclo-

sures are discrete structures designed for the sole purpose of holding printed

1154 SPACECRAFT APPLICATIONS OF ADVANCED COMPOSITE MATERIALS

wiring boards and electrical components in place. Composite enclosures can be

made signiﬁcantly lighter than machined aluminum enclosures and may be pro-

duced at a lower unit cost while possessing equal or better mechanical and

thermal performance. Three groups of functional requirements—mechanical

loading and physical attachment speciﬁcations, thermal control for installed elec-

tronic components, and environmental shielding—must be satisﬁed in the design

process for electronics enclosures.

Both physical support and protection for electronic components (often printed

wiring boards) are provided by electronics enclosures. Design considerations

must include the accurate alignment of printed wiring boards (PWBs) to a moth-

erboard or input/output connection area, and provisions for mechanical attach-

ment to the spacecraft bus. The enclosure must be stiff enough to meet internal

deﬂection speciﬁcations during launch conditions, and alignment and support

must be maintained over thermal extremes. Designs usually include a removable

cover to facilitate installation, removal, and exchange of enclosed components.

Mechanical loads from launch or thermal expansion are the primary drivers

for the structural design requirements. Static equivalent of dynamic loads are

calculated from a curve of power spectral density of random vibration, and

launch loads are generally used to represent loadings for mechanical analysis,

as described by Steinberg.

The walls of the enclosure can then be sized by

analyzing them as individual orthotropic ﬂat plates with simply supported bound-

aries. This method greatly simpliﬁes initial sizing and enables the designer to

create a more optimum structure. This method is superior to the standard ‘‘black

aluminum’’ design process where metals are simply replaced with dimensionally

equivalent quasi-isotropic composites. Roberts

has shown that a reasonable, if

somewhat conservative, approximation of actual static deﬂections and natural

modes can be produced with orthotropic analysis.

The thermal environment inside the electronics enclosure must be controlled

to ensure the safety of sensitive internal electronics, so some provision must be

made to expel heat generated by electronic components. The thermal dissipation

path may consist of several parts. Paths begin at individual electronic compo-

nents, but all must end at the ﬁnal heat sink or thermal reservoir, usually space.

A typical thermal path may be subdivided as follows: electronic component to

PWB, PWB to card guide, PWB and card guide to the enclosure wall, wall to

mounting feet, and feet to base or enclosure wall to space. Conduction is the

chief mode of thermal ﬂow in electronics enclosures with the exception of ra-

diation from the enclosure walls to space.

The composite ﬁber properties rather than the resin properties are the main

contributors to values for in-plane thermal conductivity. Carbon ﬁber composites

can be made that have equal or better in-plane thermal conduction than solid

aluminum.

However, through-the-thickness thermal conduction is limited by

the polymer matrix and is generally several orders of magnitude less than in-

plane conduction. As such, careful attention must be given to the design of

thermal attachments to ensure adequate thermal conduction across joints and

connections. Metallic inserts or coatings are commonly used for thermal load

distribution through-the-thickness of composite panels.

Electronics enclosures must provide radiation and electromagnetic shielding

to preserve and to protect the electronics within the housing. This requirement

has received the majority of attention because composites have much higher

3 EXAMPLE SPACECRAFT APPLICATIONS 1155

Table 5 Comparison of Composite and Lightweight Aluminum Electronic Enclosures

Enclosure Type

Weight

[kg (lb)]

% Weight Savings

from Original Design

Launch Cost at

$10K/ lb

Original machined aluminum 8.4 (18.5) N / A $185,000

Lightweight machined aluminum 14.4 (9.7) 48% $97,000

Cast aluminum 4.5 (9.9) 46% $99,000

Composite 2.6 (5.8) 69% $58,000

electrical resistance than metals. New techniques have been developed and im-

plemented to increase or augment the shielding effectiveness of composite en-

closures. A variety of speciﬁc methods are based on the use of metallic foils,

coatings, or meshes incorporated with the laminate as discussed above.

Examples

The following descriptions of composite electronics enclosure designs represent

current practice and the state of the art.

A National Aeronautics and Space Administration (NASA) study (Fenske et

al.

) described a comparison study of composite versus aluminum electronics

enclosures. A multidisciplinary team produced a set of speciﬁc electrical, ther-

mal, and mechanical requirements (summarized in Table 5) that were used as

criteria for the development of three different new enclosure types. These light-

weight electronics enclosures were independently developed and tested against

a baseline machined aluminum enclosure. Each development team used a dif-

ferent manufacturing method to produce the separate enclosures. The three en-

closures—cast aluminum, lightweight machined aluminum, and composite—

were compared to the baseline standard machined aluminum enclosure

speciﬁcally in terms of cost and performance. Thermal, electrical and mechanical

speciﬁcations were satisﬁed by all three of the enclosures developed.

Composite Optics Inc. used proprietary SNAPSAT technology to manufacture

the composite electronics enclosure. This approach produced an easily scalable

design that can be used to produce enclosures housing between 6 and 14 PWBs.

The enclosure (Fig. 11), consists of a double-wall system that minimizes cost

and mass by limiting use of expensive highly conductive materials. Aluminum

foil was co-cured onto the inner side of the walls to provide shielding but was

later found to be insufﬁcient for grounding requirements. Additional thin copper

strips were later soldered directly to the nickel-plated PWB attachment ﬁttings

to achieve adequate grounding.

Table 5 lists the mass of each enclosure and the corresponding launch costs

assuming a ﬁgure of $10,000/lb to orbit.

An independent research and development team at the Johns Hopkins Uni-

versity Applied Physics Laboratory (JHU-APL) used previous research to de-

velop a composite electronics enclosure that could be used on a variety of

spacecraft, without requiring substantial modiﬁcation or redesign.

The project

goal was to produce an electronics enclosure with the mass reduction and high

stiffness beneﬁts of composites and without high recurring manufacturing costs.

The JHU-APL electronics enclosure design (shown in Fig. 12) consists of a

three-sided box with integral mounting ﬂanges bonded to a baseplate thermal

plane. Unidirectional pitch ﬁber contained in the baseplate thermal plane was

1156 SPACECRAFT APPLICATIONS OF ADVANCED COMPOSITE MATERIALS

Fig. 11 Composite electronics enclosure (from Ref. 23).

Fig. 12 Composite electronics enclosure (from Ref. 24).

3 EXAMPLE SPACECRAFT APPLICATIONS 1157

used to distribute heat over the entire area, to increase thermal transfer efﬁciency.

The outer wall acts as a secondary thermal plane and also contains pitch ﬁber.

The total quantity of expensive high thermal conductivity pitch ﬁber is mini-

mized by making the thermal path from the PWB to the spacecraft or thermal

reservoir as short as possible.

Data gained from this program supports the conclusion that a composite elec-

tronics enclosure can be built with equivalent or better performance than a ma-

chined aluminum design and with lower unit cost. The simplicity of the design

and tooling developed by JHU-APL make the techniques applicable to a wide

range of missions. Approximately 50% mass savings can be achieved by using

composite electronics enclosures over traditional aluminum designs. In a follow-

up study investigating the optimization of design with respect to structural and

thermal loading, it was shown that judicious discrimination in the use of expen-

sive high thermal conductivity materials is key to minimizing the total cost of

the enclosure. The two designs developed in the study had equivalent mass,

thermal, and structural performance, but the difference in cost between the two

was 10-fold, with only one design being equal in costs to the original machined

aluminum standard.

3.4 Optical Benches and Instrument Structures

The advantages in the use of composite materials can be utilized in individual

instrument structures as well as the larger spacecraft structures. Optical benches

that support critical instruments such as telescopes and interferometric optics,

are commonly constructed from carbon ﬁber composites because they offer ther-

mal stability impossible to attain with conventional metal structures. Optical

bench construction usually includes one or more panels or struts that provide

mounting, support, and alignment of optical instruments. Composite optical

benches have been successfully used in applications that range in size from

smaller than this handbook to as large as the Hubble space telescope. An over-

view of the fundamental requirements and basic structures that have been used

for optical benches is presented here.

Typical Optical Bench Requirements

While there are differences and unique requirements for every optical bench, the

elements of dimensional stability, stiffness, and mass are the three main require-

ments common to most designs. As previously mentioned, dimensional stability

is the primary requirement for optical benches. Tight tolerances must be main-

tained throughout manufacturing, launch, deployment, and service life. Taken

together, meeting speciﬁed tolerances is often a challenging task.

Near-zero thermal expansion (in one or two directions) is a key property of

composites and one that is invaluable for optical instrument applications. Strin-

gent dimensional tolerances are often required for the location and alignment of

optical components. These tolerances must be maintained over a wide range of

operating temperatures. For example, Composite Optics Inc. imposes a standard

maximum angular tolerance of 30

␮

rad for all optical components, and even

tighter tolerances are speciﬁed for more advanced instruments.

Stiffened carbon

ﬁber composite panels are the material of choice to satisfy these tolerance re-

quirements. The panels are generally sandwich structures with either honeycomb