Kutz M. Handbook of materials selection

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1158 SPACECRAFT APPLICATIONS OF ADVANCED COMPOSITE MATERIALS

Composite

Facesheets

Metal Insert

Potting Adhesive

Honeycomb

Core

Fig. 13 Cutaway view of typical hardpoint insert in composite sandwich panel.

core or discrete rib stiffeners. Facesheet laminates for these panels have been

designed to meet optimal stiffness, coefﬁcient of thermal expansion (CTE), and

thermal conduction requirements with minimal mass. High modulus or superhigh

modulus carbon ﬁbers are often used in facesheet laminates to achieve the de-

sired properties.

Stiffness requirements for optical benches are usually described by specifying

a minimum natural frequency, typically 30 Hz or greater. Boundary conditions

(ﬁxity) and mounting conditions must be considered during the design process.

The total mass of the bench and ﬁttings is usually limited to less than one

third of the mass of the instruments that it supports.

Design and Manufacturing Considerations

Several important physical constraints are common to composite optical bench

designs. Because thin composite facesheets cannot support threaded fasteners,

separate hardpoints or ﬁttings must be used. A common technique for providing

component mounting points involves the postbonding of aluminum or titanium

inserts into the bench sandwich structure, shown generically in Fig. 13. The

insert must be bonded to both the core material and the facesheet to ensure good

structural performance. In honeycomb core structures, the bonding usually re-

quires the use of a strong paste adhesive for the insert-to-facesheet bond and a

low-density potting adhesive for the insert-to-core bond. In addition to discrete

inserts, metal edge close outs are sometimes used as hardpoints for attachment

to the primary spacecraft structure or other instruments. The bonding of metal

parts into composite structures must be done cautiously because differences in

thermal expansion between composite and metals may result in large stresses

within the adhesive layer, sandwich structure, and metal components.

3 EXAMPLE SPACECRAFT APPLICATIONS 1159

Outgassing and volatile evolution are factors that must also be considered

when using polymer matrix composites for optical bench designs to be used in

space. As discussed in Section 2.1, polymer matrix materials are subject to

outgassing caused by moisture desorption or the evolution of volatiles within

the material. The evolved gasses can form clouds or ﬁlms that can impact the

performance of sensitive optical instruments. Polycyanate matrix materials are

preferred over epoxies for optical bench elements because they exhibit much

lower moisture uptake and therefore less outgassing. Special preﬂight handling

procedures such as nitrogen gas purge treatments are often speciﬁed to minimize

outgassing.

When using low CTE composites for optical benches, allowances must be

made to accommodate dimensional changes from thermal expansion in the

mounting of metal components onto composite optical benches. For example,

aluminum instruments mounted on a low CTE composite bench will exhibit

dimensional changes that are much greater than those of the deck itself. Instru-

ment mounting must be designed so these thermally induced dimensional

changes do not exceed acceptable distortion or stress levels.

3.5 Antennas, Reﬂectors, and Mirrors

Composites have been widely used to construct low-mass antennas and reﬂec-

tors. For example, the high-gain communications antenna on the Voyager space-

craft employed honeycomb core sandwich construction to reduce mass while

retaining structural requirements. Glass or ceramics materials traditionally used

for mirror construction and mirrors using these materials have areal densities of

25–30 kg/m

Equivalent composite designs can have areal density as low as

1–3 kg/m

resulting in an order of magnitude mass reduction over glass and

ceramic designs.

Carbon ﬁber-reinforced composites have additional properties that make them

useful in reﬂector applications. These include:

●

Near-zero CTE

●

High structural efﬁciency

●

High strength

●

High fracture toughness

●

Design and manufacturing ﬂexibility

Proper functioning of mirrors and reﬂectors requires the sustained mainte-

nance of proper form or ﬁgure during service life. Because the thermal extremes

encountered by space structures are quite large, thermal stability of these struc-

tures is a critical requirement. Furthermore, mirror form and ﬁgure must also be

maintained over many thermal cycles. The thermal stability criteria are easily

met if the unique thermal expansion properties of advanced composites are em-

ployed.

Mirrors, reﬂectors, and antennas using composites fall into two groups: all-

composite designs, and hybrid designs. Hybrid designs use composite support

structures but combine these with noncomposite elements such as glass or ce-

ramic reﬂective facesheets. All-composites designs use composite elements for

1160 SPACECRAFT APPLICATIONS OF ADVANCED COMPOSITE MATERIALS

Fig. 14 NGST hybrid composite mirror (reprinted from Ref. 28).

all parts of the mirror, including the reﬂective surface. Coatings such as alumi-

num, silver, or gold are applied to all composite mirrors to provide a highly

reﬂective surface. The following section describes hybrid composite mirrors and

all-composite mirrors that represent the current technology.

Hybrid Composite Mirrors

Hybrid designs consisting of composite structures that support glass, ceramic,

or metal clad facesheets have been used for visible wavelength applications.

Although these designs weigh much more than all-composite designs, these hy-

brid mirrors have exhibited the best quality ambient ﬁgure. Additionally, because

traditional optical ﬁnishing techniques are used to achieve the appropriate ﬁgure

in these structures, fewer problems may be encountered than in all composite

designs. An example of a hybrid visible mirror is the Next Generation Space

Telescope Mirror System Demonstrator (NGST NMSD), developed at Compos-

ite Optics Inc.

The NGST NMSD mirror (shown conceptually in Fig. 14) employs a glass

front facesheet with a composite supporting structure and rear facesheet. The

aperature is 1.6 m and the areal density for the mirror assembly is 15 kg/m

11 kg/m

for the mirror substrate and 4 kg/m

for the reaction structure. During

the design process manufacturing techniques were considered along with the

physical and mechanical properties of glass, composites, and adhesives. This

strategy was taken to ensure that maximum performance (low mass) was

achieved at minimum cost.

Ultrahigh modulus carbon ﬁber/cyanate ester resin composites were chosen

to construct the back facesheet and shear web core because of their high stiff-

ness, low density, near-zero thermal expansion properties, and manufacturing

3 EXAMPLE SPACECRAFT APPLICATIONS 1161

Table 6 MLS Composite Primary Reﬂector Performance

Parameter Requirement Actual Value

Dimensions 1.6 ⫻ 0.8 m ellipse Same

Mass 10 kg 8.6 kg

Areal density 7.8 kg / m

6.7 kg / m

Stiffness (natural frequency) 80 Hz 228 Hz (calculated)

Surface accuracy (as-fabricated) 8.5

␮

m rms 4.37

␮

m rms

Surface stability (on-orbit environment) 18

␮

m rms 6.1

␮

m rms

Solar absorptance 0.40 0.43

Absorptance/ emittance (

␣

/␧)1⬍

␣

/ ␧ ⬍ 2 1.3

ﬂexibility. For the front reﬂective facesheet, Zerodur glass was chosen because

it has a small negative CTE and can be shaped using traditional optical ﬁnishing

techniques. Waterjet pocket milling techniques were used to remove material

from the backside of the facesheet leaving an isogrid stiffening pattern. A tough-

ened paste epoxy adhesive (EA 9309) was used to bond the composite shear

web core to the front Zerodur glass facesheet and composite rear facesheet.

Thick adherend shear tests conducted in the operating temperature range (room

temperature to 35 K) proved that the adhesive would withstand the cool-down

thermal strain differentials between the mirror and the core

A subscale test mirror was fabricated with the same materials, construction,

and processing used in the full-scale mirror. A key concern was ensuring that

the thermal expansion of the composite supporting structure was the same as

that of the Zerodur reﬂector. Because small variations in composite character-

istics are unavoidable a thin metallic plating was applied to the composite struc-

ture to ﬁne tune its CTE. This proved to be an effective and predictable method

for CTE management as no measurable change in ambient ﬁgure was detected

in the test thermal ranges after 50 cycles. The development, fabrication, and

testing of this mirror proved that thermal strain management may be achieved

by proper material selection.

All-Composite Mirrors

All-composite reﬂector designs for X-ray, microwave, and radio-frequency (RF)

applications have been successfully produced. An example of an all-composite

design is the Microwave Limb Sounder (MLS) primary reﬂector.

An off-axis

parabolic surface provides the front reﬂective surface shape, with a 1.6-m major

axis and 0.8-m minor axis elliptical projected aperture. The structure of the MLS

reﬂector is similar to that of the NMSD discussed previously, except the reﬂec-

tive facesheet is composite instead of glass. Two 24-ply carbon ﬁber/polycyanate

composite facesheets are adhesively bonded to a triangular composite isogrid

core to create the reﬂector. The performance requirements and the actual oper-

ational values of important reﬂector parameters are listed in Table 6. RF reﬂec-

tivity, solar concentration, and thermal absorption requirements dictated a

specialized treatment for the composite surface. An initial precision bead blast-

ing, followed by vapor deposition coating with aluminum, and a ﬁnal layer of

silicon suboxide satisﬁed the performance parameters.

1162 SPACECRAFT APPLICATIONS OF ADVANCED COMPOSITE MATERIALS

A high accuracy coordinate measuring machine was used periodically during

the fabrication process to measure the ﬁgure accuracy of the reﬂector. These

measurements conﬁrmed that the mirror surface directly replicates the mold

surface within a factor of two. This data suggests that higher ﬁgure accuracies

can be obtained through further reﬁnement and polishing of the mold surface.

Despite the small surface errors, the ﬁnal ﬁgure accuracy value, 4.37

␮

mrms,

is well within the 8.5-

␮

m speciﬁcation. This project demonstrated the feasibility

of fabrication low-mass, high-ﬁgure accuracy reﬂectors using composite tech-

nology.

REFERENCES

1. R. M. Jones, Mechanics of Composite Materials, 2nd ed., Taylor & Francis, Philadelphia, 1999.

2. ASM, ASM Engineered Materials Handbook, Vol. 1, ASM International, Metals Park, OH, 1987,

p. 114.

3. R. B. Seymour, History of Fibrous Reinforcements, Proceedings of the Symposium on ‘‘History

of Polymeric Composites,’’ R. B. Seymour and R. D. Deanin (eds.), VNU Science Press, 1987,

p. 59.

4. D. S. Mehoke and P. A. Wienhold, Practical Constraints in Using High Thermal Conductivity

Composite Materials in Spacecraft Applications JHU-APL, Conf. paper presented at SAE Inter-

society Energy Conversion Engineering Conference, Vancouver, BC, Society of Automotive En-

gineers, Inc., August 1999.

5. M. Fan and W. L. Niemeyer, ‘‘Structural Design and Analysis of a Light-Weight Laminated

Composite Heat Sink for Spacecraft PWBS,’’ NASA Goddard Tech Paper #3679, 1997.

6. T. C. Magee and J. C. Roberts, ‘‘Structural and Thermal Optimization of a Composite Electronics

Enclosure for Spacecraft Applications,’’ SAMPE J. Adv. Materials, 33(4), 26–32, October (2001).

7. R. E. Evans, D. E. Hall, and B. A. Luxon, ‘‘Nickel Coated Graphite Fiber Conductive Compos-

ites,’’ SAMPE Quarterly—Society for the Advancement of Material and Process Engineering,

17(4), 18–26, July (1986).

8. John L. Paretti, ‘‘Affordable Lightweight High Conductive Polymer Composite Electronic Pack-

aging,’’ Proceedings of the 44th International SAMPE Symposium and Exhibition, Long Beach,

CA, 1999.

9. E. M. Silverman, ‘‘Space Environmental Effects on Spacecraft: LEO Materials Selection Guide,

Part 1,’’ NASA Contractor Report 4661, 1995.

10. C. T. Golden and E. E. Spear, ‘‘Graphite/ Epoxy Structure of the SPAR Telescope’s Optical

Telescope Assembly,’’ Proceedings of the 29th International SAMPE Symposium and Exhibition,

Reno, NV, 1984.

11. E. Silverman and M. Rhodes, ‘‘Composite Isogrid Structures for Spacecraft Components,’’

SAMPE J., 35(1), 51–58 (1999).

12. J. R. Vinson and R. L. Sierakowski, The Behavior of Structures Composed of Composite Ma-

terials, Martinus Nijhoff Publishers, Dordrecht, The Netherlands, 1986, pp. 239–283.

13. A. Brent Strong, Fundamentals of Composites Manufacturing: Materials, Methods, and Appli-

cations, Society of Manufacturing Engineers, Dearborn, MI, 1989.

14. M. Obal and J. M. Sater, ‘‘Multifunctional Structures: The Future of Spacecraft Design?’’ 5th

International Conference on Adaptive Structures, Sendai, Japan, 1994.

15. T. C. Thompson, C. Grastataro, B. Smith, K. Krumweide, and G. Tremblay, ‘‘Development of

an All-Composite Spacecraft Bus for Small Satellite Programs,’’ Proceedings of the 8th AIAA /

Utah State University Annual Conference on Small Satellites (A95-33901 08-12), Utah State

University, Logan, UT, August 29–September 1, 1994, pp. 1–17.

16. W. E. Skullney, H. M. Kreitz, Jr., M. J. Harold, S. R. Vernon, T. M. Betenbaugh, T. J. Hartka,

D. F. Persons, and E. D. Schaefer, ‘‘Structural Design of the MSX Spacecraft,’’ JHU-APL Tech.

Digest, 17(1), 59–76 (1996).

17. B. Derbes, ‘‘Case Studies in Inﬂatable Rigidizable Structural Concepts for Space Power,’’ Pro-

ceedings of the 37th AIAA Aerospace Sciences Meeting and Exhibit, Reno, NV, January 11–

14, 1999.

18. L. A. Roe, ‘‘Inﬂation Systems for Near-Term Space Missions,’’ Proceedings of the 41st AIAA

Structures, Structural Dynamics, and Materials Conference and Exhibit, Atlanta, GA, April 2000.

REFERENCES 1163

19. D. S. Steinberg, Vibration Analysis for Electronic Equipment, 2nd ed., Wiley, New York, 1988.

20. J. C. Roberts, in ‘‘Analytic technique for sizing the walls of advanced composite electronics

enclosures,’’ Composites: Part B, Elsevier, Oxford, UK, 1999.

21. B. G. Carkhuff and J. C. Roberts, ‘‘A Test Fixture for Mapping Thermal Conduction in Com-

posites under Transient and Steady-State Conditions,’’ Experimental Tech. J., 21(3), 12–14, May/

June (1997).

22. J. C. Roberts, M. H. Luesse, and T. C. Magee, ‘‘A Technique for Locally Increasing Surface

Heat Spreading and Through-Thickness Thermal Conductivity of Graphite /Epoxy Laminates,’’

Proceedings of the 9th American Society for Composites, Technical Conference, Univ. of Del-

aware, Newark, September 20–22, 1994 (A95-20803 04-24), Technomic Publishing Co., Inc.,

Lancaster, PA, 1994, pp. 1098–1105.

23. M. T. Fenske, J. L. Barth, J. R. Didion, and P. Mul. ‘‘The Development of Lightweight Electronics

Enclosures for Space Applications,’’ SAMPE J., 35(5), 25–34 (1999).

24. P. D. Wienhold, D. S. Mehoke, J. C. Roberts, and E. D. Schaefer, ‘‘The Design and Fabrication

of a Low Cost Spacecraft Composite Card Cage,’’ Proceedings of the 42nd International SAMPE

Symposium and Exhibition, Anaheim, CA, 1997.

25. T. C. Magee and J. C. Roberts, ‘‘Structural and Thermal Optimization of a Composite Electronics

Enclosure for Spacecraft Applications,’’ SAMPE J. Adv. Materials, 33(4), 26–32, October (2001).

26. K. J. Dodson and J. E. Rule, ‘‘Thermal Stability Considerations for Space Flight Optical

Benches,’’ Proceedings of the 34th International SAMPE Symposium and Exhibition, Reno, NV,

1989.

27. E. P. Kasl and D. A. Crowe, ‘‘A Critical Review of Ultra-lightweight Composite Mirror Tech-

nology,’’ American Institute of Physics: Space Technology and Applications Forum, DOE Conf-

980103, Department of Energy, Washington, DC, 1998.

28. E. P. Kasl, G. V. Mehle, J. E. Dyer, H. R. Clark, S. J. Connell, and D. A. Sheikh, ‘‘Recent

Developments in Composite-Based Optics,’’ SPIE Conference on Space Telescopes and Instru-

ments V, Kona, Hawaii, 1998, pp. 735–746.

29. P. B. Willis, J. Dyer, and S. Dummer, ‘‘Fabrication and Thermo-optical Properties of the MLS

Composite Primary Reﬂector,’’ SPIE Conference on Materials and Applications, Denver, CO,

1999, pp. 200–205.

1165

CHAPTER 37

SELECTION OF MATERIALS FOR

BIOMEDICAL APPLICATIONS

Michele J. Grimm

Bioengineering Center

Wayne State University

Detroit, Michigan

1 INTRODUCTION 1165

2 ORTHOPEDIC MATERIALS:

TOTAL HIP ARTHROPLASTY 1166

2.1 Function 1167

2.2 Biocompatibility 1174

2.3 Current Material Selection 1178

3 BLOOD-CONTACTING

BIOMATERIALS: VASCULAR

PROSTHESES 1180

3.1 Function 1181

3.2 Biocompatibility 1183

3.3 Current Material Selection 1186

4 SPACE-FILLING BIOMATERIALS:

BREAST IMPLANTS 1188

4.1 Function 1188

4.2 Biocompatibility 1188

4.3 Current Material Selection 1191

5 SUMMARY 1191

BIBLIOGRAPHY 1192

1 INTRODUCTION

Materials have been used for medical implant applications for centuries. Starting

even before George Washington’s famous wooden teeth, before the use of co-

conut shells in the 1800s by South Seas natives to replace missing portions of

the skull (Sanan and Haines, 1997), humans have attempted to use materials

from biological and inorganic sources to replace diseased or damaged tissues.

In fact, some examples of biomaterial implants, particularly in the form of gold

and silver, date back to prehistoric times (Sanan and Haines, 1997). These ma-

terial selections of the past, based more on the availability of materials and

anecdotal evidence than scientiﬁc method, resulted in varying degrees of success.

The development of aseptic surgical technique in the midnineteenth century by

Lister provided the basic medical tool needed for the more widespread and

successful use of biomedical implants (Park and Lakes, 1992). Since the twen-

tieth century, as knowledge of the biological mechanisms behind the interactions

of implanted materials and tissues has increased, the selection of materials for

medical implants has been based on progressive improvements and experimental

evidence.

When selecting a material for use in a medical implant application, two gen-

eral considerations need to be taken into account: the functional requirements

HandbookofMaterialsSelection.EditedbyMyerKutz

1166 SELECTION OF MATERIALS FOR BIOMEDICAL APPLICATIONS

Fig. 1 Anatomy of the hip.

of the implant and how it will interact with the body. The function of the implant

includes the physiologic role that it will replace, as well as the length of time

that it is designed to fulﬁll that role. The interaction of the body and the im-

planted material must be examined from two perspectives—the effect of the

biological environment on the material properties and the effect of the material,

and any degradation that may occur, on the local and systemic physiology of

the body.

It is not possible within the scope of this chapter to discuss the optimum

choice of materials and the supporting rationale for every medical implant ap-

plication. This task is made even more difﬁcult by the fact that the development

and selection of materials is an ever-changing ﬁeld. In light of this fact, the

following pages will use three implant examples to serve as the background for

a discussion of the considerations in the selection and evaluation of materials

for medical implants. Each section will ﬁrst be organized around the functional

requirements of the implant, with the resulting concerns regarding material–

tissue interaction being discussed for the materials that would ﬁrst meet the

functional demands. Geometrical considerations and overall implant design will

also play a major role in the success of an implant; however, this chapter will

focus on the selection of biomaterials as a discrete step in the design process.

2 ORTHOPEDIC BIOMATERIALS: TOTAL HIP ARTHROPLASTY

The hip (Fig. 1) is one of the most commonly replaced joints, due to the high

incidence of both osteoarthritis and osteoporotic hip fracture within the popu-

lation. As with the other joints in the skeletal system, the hip has two main

functions: (1) to transfer load from one bone to another and (2) to allow for

motion between the bones. The loads on the hip have been found to vary from

0.5 to 8 times body weight during activities of daily living, including brisk

walking (Paul, 1999). The loads can be expected to reach even higher levels

during events such as a stumble. The hip is essentially a ball-and-socket joint

2 ORTHOPEDIC BIOMATERIALS: TOTAL HIP ARTHROPLASTY 1167

Fig. 2 Example of total hip implant, including femoral and acetabular components.

with 6 degrees of freedom constrained by the bony and ligamentous structures

of the joint. These functional constraints have generally deﬁned the goals for

design of hip implants over the past half-century.

The total joint replacement for the hip is divided into three components: the

femoral stem and femoral head, which may be integrated or modular, and the

acetabular cup (Fig. 2). The combination of the femoral head and the acetabular

cup provide the bearing surface for the joint. The femoral stem transfers the

load to the femur and provides resistance to the bending moment caused by the

anatomy of the joint. Total hip replacement began in the 1930s, with stainless

steel used for both the acetabular cup and the femoral head, which was bolted

onto the natural femoral neck. In the 1950s, the general design of the implant

was expanded by McKee to include the stemmed femoral component familiar

today. In 1959, Charnley introduced a plastic acetabular component, which ar-

ticulated with the metallic femoral component. This idea for low-friction arthro-

plasty remains the basis for the predominant implant designs used today

(Swanson, 1977).

2.1 Function

Load Support

Because of its primary role in mechanical support, the stem of a femoral pros-

thesis can realistically be manufactured from a metal, a ceramic, or a composite

material. For this discussion, composite materials will not be included, as they

have not yet been utilized in commercially available total joint replacements

(they are available for bone plate applications). However, as development of

1168 SELECTION OF MATERIALS FOR BIOMEDICAL APPLICATIONS

Table 1 Representative Properties of Cortical

Bone

Compressive strength (MPa) 131–224 longitudinal

106–133 transverse

Tensile strength (MPa) 80–172 longitudinal

51–56 transverse

Shear strength (MPa) 53–70

Source: Data from Cowin (1989).

composite materials for this application continues, they must meet the same

requirements to be discussed for ceramic and metallic materials.

Strength. The ﬁrst property to be considered for a load-bearing implant is

its mechanical strength. The loading of the femur is dynamic and, while esti-

mated to range up to 8 times body weight, is difﬁcult to determine precisely.

Therefore, as the implant will be loaded in essentially the same way as the

natural bone, it is reasonable to assume that a material that will provide the same

or greater load-bearing capacity as bone will meet the necessary mechanical

requirements. The mechanical properties of cortical bone (Table 1) form the

lower limit for properties of materials to be selected for the femoral stem. If the

neck of the femoral stem is designed to be longer than the normal range of

femoral neck lengths [approximately 9 cm, from the edge of the greater tro-

chanter to the apex of the femoral head (Center et al., 1998)], then the increased

bending moment in this region should also be considered. The primary mode or

modes of loading that will be seen for a particular implant application will also

be important factors in the determination of whether a material meets the con-

straints regarding mechanical strength. The hip will be loaded primarily in bend-

ing and compression, due to its unique geometry. Tensile and shear strength of

any replacement material therefore become of great importance.

In addition to the yield or ultimate strength of an implant material, the fatigue

strength is also important for structures that will be cyclically loaded over an

extended period of time. At the current time, hip replacements are designed to

last for approximately 20 years. As an individual typically loads and unloads

the joint thousands of times per day during normal daily activities, fatigue-

induced failure at this location is of paramount concern. In contrast to the ma-

terials that might be chosen to replace the original tissue, healthy bone has the

capacity to repair any microfractures that may occur through continuous cycling.

This normal remodeling function helps to eliminate the occurrence of fatigue

fractures in normal bone.

Based on the cyclic loading that a material is likely to undergo when im-

planted in the body, it is reasonable when evaluating selections to compare the

endurance limit of the materials under consideration to the experimentally de-

termined strength values for bone. One further complication in this process,

however, is that materials will fatigue differently in a typical, air-based labora-

tory environment than they will in the ionic soup of the human body. Because

of this, it is important to measure the fatigue performance of materials in an

environment that closely mimics that of the implant location—including the

ionic composition, temperature, and pH. Substantial work has been conducted