40207_15

CARBON-CARBON COMPOSITES 15 John D. Buckley

15.1 INTRODUCTION n

Carbon-carbon (CC) materials are a generic class of composites similar to the graphite/epoxy family of polymer matrix composites. These materials can be made in a wide variety of forms, from one-dimensional 1 -D 2-D

to n-dimensional, using unidirectional tows, tapes, or woven cloth (Fig. 15.1). Because of their multiformity, their mechanical properties can be readily tailored (Table 15.1). Carbon materials have high strength and stiffness potential as well as high thermal and chemical stability in inert environments. These materi- 3-D n-D als must, however, be protected with coatings and/or surface when used in an OXi- dizing environment.

Fig. 15.1 Multiformity and general properties of carbon-fiber and carbon-matrix composites.

Table 15.1 General properties of carbon-carbon composites

Ultimate tensile >276 MPa >40 000 psi strength Modulus of >69 GPa >lo7 psi elasticity Melting point >41OO0C 7412°F Thermal ~11 .5 W m-' K-' 6.64 conductivity h ft "F

Linear thermal ~1 .1 x 10"OC 6.1 x W7"F expansion Density <2990 kg m-3 186.6 lb/ft3

Handbook of Composites. Edited by S.T. Peters. Published in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7

The development of CC materials began in 1958 and was nurtured under the US Air Force space plane program, Dyna-Soar and NASA's Apollo projects. It was not until the Space Shuttle Program that CC material systems were intensively researched. The criteria that led to the selection of CC composites as a thermal protection system were based on the following requirements:

1. maintenance of reproducible strength levels at 1650°C (3002°F);

2. sufficient stiffness to resist flight loads and large thermal gradients;

3. low coefficient of thermal expansion to minimize induced thermal stresses;

4. oxidation resistance sufficient to limit strength reduction;

334 Carbon-carbon composites

5. tolerance to impact damage; 6. manufacturing processes within the state of

Carbon-carbon composites consist of a fibrous carbon substrate in a carbonaceous matrix. Although both constituents are the same ele- ment, this fact does not simplify composite behavior because the state of each constituent may range from carbon to graphite. Crystallographic carbon, namely graphite, consists of tightly bonded, hexagonal arranged carbon layers that are held together by weak van der Waals forces. The single crystal graphite structure is illustrated in Fig. 15.2 (Bokros, 1969). The atoms within the layer

the art.

plane or basal plane (a-b direction) have a covalent bond strength of =524 kJ/mol (Kanter, 1957), while the bonding energy between basal planes (c direction) is =7 kJ/mol (Dienes, 1952). The result is a crystal that is remarkable in its anisotropy, being almost isotropic within the basal plane but with c direction properties that differ by orders of magnitude. On a larger scale, carbon, in addition to its two well-defined allotropic forms (diamond and graphite), can take any number of quasicrystalline forms ranging continu- ously from turbostratic (amorphous, glassy carbon) to a highly crystalline graphite (Fig. 15.3) (Bokros, 1969).

C

E a Reference directions

Fig. 15.2 Tightly bonded, hexagonally arranged carbon layers held together by weak van der Waals forces.

I LC

:r do02

Fig. 15.3 Comparison of (a) carbon turbostratic structure with (b) 3-D graphite lattice (Bokros, 1969).

Carbon fibers 335

The anisotropy of the graphite single crystal encompasses many structural forms of carbon. It ranges in the degree of preferred orientation of the crystallites and influences porosity, among other variables. A broad range of properties is the result of this anisotropy. In CC composites, this range of properties can extend to both constituents. Coupled with a variety of processing techniques that can be used in the fabrication of CC composites, great flexibility exists in the design of and the resul- tant properties obtained from CC composites.

The wide range of properties of carbon materials can be shown when comparing the tensile modulus of commercially manufac- tured carbon fibers that range from 27.6 GPa (4 x 106 psi) to 690 GPa (100 x lo6 psi). In fabrication, the fibers can be used in either continuous or discontinuous form. The directionality of the filaments can be varied ranging from unidirectional lay-ups to multidirectional weaves. Fiber volume fraction constitutes another variable. The higher the volume fraction of a specific high-strength fiber in a matrix, the greater the strength of the composite. The matrix can be formed via two basic approaches: (1) through the carbonization of an organic solid or liquid, such as a resin or pitch, or (2) through the chemical vapor deposition (CVD) of carbon from a hydrocarbon. A range of carbon structures can be obtained by either approach. Finally, heat treatment of the composite material at graphitization temperatures offers additional variability to the properties that can be obtained. Typically, there is an optimum graphitization temperature at which the high- est strength can be obtained for a given composite composition of fiber and matrix (Edie et al., 1986; Stoller et al., 1974).

15.2 CARBON FIBERS

The properties of carbon fibers can vary over a wide range depending on the organic precursor and processing conditions used. At present, graphite fibers are produced from

three precursor materials: rayon, polyacryloni- trile (PAN) and petroleum pitch. Fibers having a low modulus (27.6 GPa (4 x 106 psi)) are formed using a rayon precursor material that may be chemically pretreated by a sequence of heating steps. First, the fiber is heated to >400"C (752°F) to allow cellulose to pyrolyze (decomposition or chemical change during thermal conversion of organic materials to carbon and graphite). Carbonization (continued heating of organic material to >lOOO"C (1832°F) to initiate ordering of the carbon structures produced by pyrolysis) is completed more rapidly at >lOOO"C (1832°F). Upon completion of carbonization, the fiber is graphitized (continued heating of carbonized organic materials to the 2000-3000°C (3632-5432°F) range of produce 100% graphite-ordered crystal structure) by heating to >2000"C (3632°F); the fiber is now, for all practical purposes, 100% carbon. High-modulus carbon fibers from rayon precursors are obtained by the additional process of stretching the carbon fibers at the final heat treatment temperature. Nigh-modulus (344 GPa (50 x 106 psi)), high-strength (2.07 GPa (300 x lo3 psi)) carbon fibers are typically made from PAN or, in some cases, mesophase pitch precursors. These fibers are processed similarly in a three-stage operation (Fig. 15.4) (Diefendorf, 1987). The PAN fibers are initially stretched from 500-1300% and then stabilized (cross-linked) in an oxygen atmosphere at 200°C (392°F) to 280°C (536°F) under tension. Carbonization of the fibers is conducted between 1000°C (1832°F) and 1600°C (2912°F). Finally, graphitization is accomplished at >2500"C (4532°F). Mesophase pitch fibers undergo the same processing procedure as PAN fibers but do not require an expensive stretching process during heat treatment to maintain preferred alignment of crystallites (Fig. 15.4) (Diefendorf, 1987). Control of fiber shape has resulted in improved fiber strength (4.1 GPa, 600000 psi) (Cogburn et al., 1987) when produced from melt-spun, mesophase petroleum pitch (Fig. 15.5) (Cogburn et al., 1987). Round fibers using the same method


PAN process

Pitch process 1 t z z I I z i m Carbonize Graphitize

Spool Epoxy sizing Surface treatment

Fig. 15.4 Carbon fiber production using PAN and pitch processes (Diefendorfer, 1987).

Hydraulic piston

Cartridge housing

Heating collar

Spinnerette

Melt-spun carbon filaments

Melted-pitch precursor

Wind-up bobbin

Melt-pressure indicator

Fig. 15.5 Melt spinning apparatus used to produce noncircular carbon fibers (Cogbum et al., 1987).

Carbonfibers 337

had a strength of 2.1 GPa (300 x lo3 psi) (Edie et al., 1986). Of the shapes studied, the C-shape and hollow fibers were found to be superior in strength to round solid and trilobal cross sec- tions (Edie et al., 1986; Cogburn et al., 1987).

15.2.1 CARBON FIBERS IN CARBON MATRIX

Addition of a matrix to carbon fiber, either through the carbonization of an organic precursor or by the deposition of pyrolytic carbon, is conducted at 800°C (1472°F) to 1500°C (2732°F). Subsequent heat treatment of the composite material may involve temperatures to 3000°C (5432°F).

15.2.2 DISCONTINUOUS FIBER COMPOSITES

Fabrication of discontinuous fiber composites uses short carbon fibers combined with either a pyrolytic carbon or pyrolyzed organic matrix. This approach to CC composites gen- erally does not have true fiber reinforcement as an objective. Rather, discontinuous fiber substrates have been used to:

1. increase fabrication capability of large-scale structures;

2. achieve a more nearly isotropic material; 3. increase the composite interlaminar tensile

strength; 4. along with continuous filament substrates,

obtain a stronger composite by providing additional nucleation sites that serve to reduce composite porosity.

The most widely used starting materials are a carbonized, rayon felt substrate with a pyrolytic carbon matrix and short, chopped fibers in a pitch-based matrix. Felt is produced through the mechanical carding of viscous rayon fibers to produce a continuous web of fibers. The webs are folded one on top of another to produce a batt. The batts are then cut, stacked and needled to produce the required felt. The rayon felt is subjected to a controlled carbonization cycle in an inert atmosphere or vacuum; the maximum temperature determines such factors as shrinkage, weight loss and chemical composition of the felt. A maximum carbonization temperature of 1200°C (2192°F) is a nominal standard; the length of the carbonization cycle and rate of temperature rise are dictated by the thickness of the felt. Carbon content in the fibers is ~98%. Carbon-arbon composites have also

Fig. 15.6 Models of fiber arrangements for four short-fiber fabrication techniques: (a) flocking lay-up, (b) pulp molding, (c) isotropic casting, and (d) spray lay-up (Cook, Lambdin and Trent, 1970; Lambdin, Cook and Marrow, 1969; Lambdin and Cook, 1971).


been fabricated from short carbon fibers using isotropic casting, flocking lay-up, spray lay-up and pulp-molding techniques (Fig. 15.6) (Cook, Lambdin and Trent, 1970; Lambdin, Cook and Marrow, 1969; Lambdin and Cook, 1971.). The rationale for using these short fibers is to reduce composite anisotropy (Lambdin, Cook and Marrow, 1969).

15.3 CONTINUOUS FIBER COMPOSITES

Continuous filament substrates reflect the properties of high-strength filaments or achieve a high degree of preferred orientation on the macroscale of the matrix. The fabrication complexity for continuous-filament substrates is determined by two parameters: (1) the directionality of the filaments, and (2) the amount of layer interlocking achieved in the substrate. Filament winding of unidirectional tapes can be used to achieve a highly

oriented substrate, usually with no interlocking between layers. Woven fabrics are used to form a two-dimensional laminate with no interlocking between layers. Helical filament winding, which is directional, results in continuous, adjacent layer interlocking. Multilayer locking is achieved through complex weaving patterns or yarn placement resulting in 'multidirectional' substrates (Fig. 15.7).

15.4 CHEMICAL VAPOR DEPOSITION

The CVD of carbon from a hydrocarbon gas within a substrate is a complex process. Various techniques have been applied to infiltrate various fiber substrates including isothermal thermal gradient (Pierson, 1968), pressure gradient (Kotlensky and Pappis, 1969) and pressure pulsation (Beatty and Kipplinger, 1970). The first two have been the

Fig. 15.7 Interlocking approaches of continuous filament substrates: (a) tape wrapped, shingle; (b) filament wound, helix; and (c) multidimensional.

Carbonized organic composites 339

most extensively used. The isothermal technique is illustrated in Fig. 15.8. The substrate is radiantly heated by an inductively heated sus- ceptor so that the gas and substrate are maintained at a uniform temperature. Infiltration is normally accomplished at 1100°C (2012°F) and at reduced pressures (6 kPa (50 torr)) with the flow rates primarily determined by the substrate surface area. This technique produces a crust on the outer surfaces of the substrate, thus requiring machining and multiple infiltration cycles.

In the thermal gradient technique (Fig. 15.9), the part to be infiltrated is supported by a mandrel that is inductively heated. Therefore, the hottest portion of the substrate is the inside surface, which is in direct contact with the mandrel. The outer surface of the low-density substrate is exposed to a cooler environment and results in a temperature gradient through the substrate thickness. Surface crusting is eliminated because the deposition rate is greater on the heated fibers near the mandrel, whereas the cooler outer fibers receive little or no deposit. Under proper infiltration conditions, the carbon is first deposited on the inside surface and, in a continuous process, progresses radially through the substrate as the densified substrate itself becomes inductively heated. Infiltration is normally accomplished at atmospheric pressure with a mandrel heated to approximately 1100°C (2012°F) (Theis et al., 1970).

15.5 CARBONIZED ORGANIC COMPOSITES

Carbonized organic composites have fabrication procedures that are similar to those of conventional fiber-reinforced, resin-laminat- ing techniques. The starting material is usually a prepregged fabric or yarn (a fabric impregnated with a matrix material in a tacky state). These precursor materials are staged nomi- nally at approximately 100°C (212°F) to achieve the desired degree of tack and flow of the resin. A laminate is then constructed and cured under pressure. Curing temperatures

Hydrocarbon 2 (-Carrier Original fiber gas gas substrate

Fig. 15.8 Isothermal chemical vapor deposition to infiltrate fibrous carbon substrate.

Fig. 15.9 Thermal gradient chemical vapor deposition.


AprepreK+ Cut, lay-up,

3 t o 5 times I cure I

Fig. 15.10 Fabrication steps involved in manufacture of 2-D carboniarbon part impregnated with tetraethylorthosilicate (TEOS).

normally range from 125°C (257°F) to 175°C (347°F) with curing pressures on the order of 2.76 MPa (400 psi). The reinforced resin laminate is then post-cured at 200°C (392°F) to 275°C (527°F). As pyrolysis is initiated, shrink- ing occurs as the organic phase decomposes. Simultaneously the release of vapors from pyrolysis expands the composite material. A slow release of these volatile by-products is required to minimize structural damage to the char. Finally, as higher temperatures are reached, thermal expansion of the carbon char itself occurs after pyrolysis is complete. After the initial carbonization, the material is then subjected to a series of reimpregnation and carbonization cycles until the desired density or the maximum density is achieved. The reimpregnation process is usually conducted under vacuum and pressure to aid in maxi- mizing the pore filling. If graphitization is desired, the high-temperature heat treatment may be used after each carbonization step or at the end of the reimpregnation and recar- bonization cycles.

To summarize, a typical manufacturing cycle of a 2-D CC part is shown in Fig. 15.10. First, a woven graphite fabric that is preimpregnated with phenolic resin is laid up as a phenolic- graphite laminate in a mold and is autoclave cured. Once cured, the part is pyrolyzed to form a carbon matrix surrounding the graphte fibers. The part is then densified by multiple furfural alcohol reimpregnations and pyrolyza- tions. The resulting CC part then is ready for use in inert or vacuum environments. This process is very time consuming. A single pyrolysis may take >70 h in a low-temperature, inert-atmosphere furnace.

Although CC materials can withstand temperatures >3000"C (5432°F) in a vacuum or in an inert atmosphere, they oxidize and sublime when in an oxygen atmosphere at 600°C (1112°F). To allow for use of CC parts in an oxidizing atmosphere, they must be com- pounded with materials that produce oxidation-protective coatings through thermo- chemical reaction with oxygen at >2000"C (3632°F) (Buckley, 1967) or they must be coated

Manufacturing 341

and sealed to protect them (Strife and Sheehan, 1988). For applications such as the Space Shuttle CC leading edges and nose caps, surfaces are converted to silicon carbide in a high-temperature diffusion-coating process (Fig. 15.10). Because of differences in thermal expansion between the silicon carbide and the CC part, the coating develops microcracks when the part is cooled from the coating temperature. To maintain oxidation protection on space vehicles such as the Space Shuttle, cracks are impregnated with tetraethylorthosilicate (TEOS). The TEOS process leaves silica in all of the microcracks, greatly enhancing the oxidation protection of the CC substrate. Current improvements being developed for oxidation protection of the CC Space Shuttle components are additions of low-temperature glass formers that enhance the sealing capability of the exist- ing coating-TEOS system.

15.6 MANUFACTURING

The fabrication process of the Space Shuttle Orbiter nose cap and wing leading edge components (Fig. 15.11) (Curry, Scott and Webster, 1979) is a multi-step process typical of the technology used to produce CC composites. The process steps are illustrated in Figs.

Initial material lay-up is similar to conventional practices with fiberglass-reinforced plastic parts. Square-weave graphite fabric impregnated with phenolic resin is laid-up in an epoxy/fiberglass mold cavity shaped to the desired configuration (Fig. 15.12) (Curry, Scott and Webster, 1979). Lay-up thickness for these components varies from 19 plies in the exter- nal skin and web areas to 38 plies in the attachment locations. Upon completion of lay- up, the part is vacuum-bagged and cured in an autoclave to 150°C (300°F) for 8 h (Fig. 15.12). The cured part is rough trimmed, X-rayed and ultrasonically inspected for irregularities following the cure cycle. Post-cure of the component involves placing the part in a graphite restraint fixture loading it into the

15.12-15.16.

RCC Seal strips (22) LH (22) RH Wing L.E. RCC panels

Nose cap (1) RCC Seal strip (1) LH, (1) RH and (3) Lower RCC ExDansion seal (1) LH, (1) RH and (3) Lower

Fig. 15.11 Leading edge structural subsystem (Curry, Scott and Webster, 1979).

furnace and submitting it to a 7-day cycle during which it is taken to 260°C (500°F) very slowly to avoid distortion and delamination (Fig. 15.12).

The next step is initial pyrolysis as shown in Fig. 15.13 (Curry, Scott and Webster, 1979). Pyrolysis tooling composed of graphite restraining fixtures containing the part are loaded into a steel retort that is packed with calcined coke. The retort and its contents then undergo a 70 h pyrolysis cycle at 815°C (1500°F) converting the phenolic resin to a carbon state. During pyrolyzation, the resin forms a network of interconnected porosity for the escape of volatile matter. This stage is extremely critical since, during controlled char- ring of the cured resin matrix, the parts are weak and delamination can easily occur if adequate escape paths and time are not ensured. After this initial pyrolysis cycle, the carbon is


2250 I

1650

w 1100 j

2 485 ; g :% I

8 200 r , h I

0

2 540 I ,

u1 320 UJ 260 ;

I

150' z I \ : n I

M A T E R I A L CUT & 0 BAG ( M Y L A R ) 0 AUTOCLAVE 0 REMOVE BAG L O A D I N POST CURE COLD I LAY.UP 0 APPLYVACUUM CURE 0 ROUGH TRIM RESTRAIN- STORAGE I CLOTH 0 CHECK FOR D R I L L H O L E S I N G

j PLIES LEAKS 0 X.RAY AND/OR FIXTURE I & ULTRASONIC I DEBULK

Fig. 15.12 Lay-up and cure cycle (Curry, Scott and Webster, 1979).

2250 - 2205

1650

g i i o c

540 F 485 5 430

320 ~ 260

C3 150

40 -2c

0 370

a zw x 95

LOADIN PYROLYSIS 0 CLEAN I RETORT INSPECT PACK WITH GRAPHITE CALCINED COKE

LOAD IN AUTOCLAVE CURE msi CURE j VACUUM CHAMBER IMPREG WITH FURFURYL ALCOHOL

IMPREGNATION CYCLE ITMREE 131 TIMESI

Fig. 15.13 Initial pyrolysis (Curry, Scott and Fig. 15.14 Densification impregnation and cure Webster, 1979). (Curry, Scott and Webster, 1979).

Manufacturing 343

designated reinforced carbon-carbon-0 (RCC-0), a state in which the material is extremely light and porous with a flexure strength of 21-24 MPa (3000-3500 psi).

15.6.1 DENSIFICATION

Densification for these shuttle parts is accomplished in three impregnation and pyrolysis cycles (Fig. 15.14) (Curry, Scott and Webster, 1979). Each part is loaded in a vacuum chamber impregnated with furfural alcohol followed by a 2-hour cure period in the autoclave at approximately 150°C (300"F), followed by a post-cure for 32 h to 200°C (400°F). This cycle is followed by a 70-hour 815°C (1500°F) pyrolyzation that is shown in Fig. 15.13. After three impregnation/pyrolyzation cycles, the material is designated RCC-3 with an increased flexure strength of =124 MPa (18000 psi) at room temperature.

15.6.2 COATING

To allow for use of CC composites at elevated temperatures above 2000°C (3632°F) in an oxidizing atmosphere, it is necessary to apply protective coatings to structural components. The oxidation inhibition process consists of two steps: (1) diffusion coating the CC component and (2) applying a sealer to the surface. The coating process (Fig. 15.15) (Curry, Scott and Webster, 1979) used in protecting the CC shuttle components starts with the blending of the constituent powders: 10% alumina, 30% silicon and 60% silicon carbide. This mix is packed around the CC structural component in a graphite retort and loaded into a vacuum furnace where it undergoes a 16-hour cycle that includes drying at 315°C (600°F) and the coating reaction to 1650°C (=3000"F) in an argon atmosphere. The powder characteristics, constituents, formulations and the manner in which the powders are packed around the part are important factors that gov- ern the chemical reactions at the high processing temperatures, the degree of consol-

w

P F

5 w w a s

2250

1650 , I

J I

540 485 430 370 320 260 200 150 95 40 -20

CLEAN COATING ' CLEAN : PREPARATION INSPECT 1 , -X.RRY INRETORT SURFACE -X.RAY ! -ULTRASONIC :e FINALTRIMIDRILL

COATING - ULTRASONIC - DIMENSIONAL

1. DIM. INWECT I

Fig. 15.15 Coating cycle (Curry, Scott and Webster, 1979).

idation and sintering of the powders. During the process, the outer layers of the CC substrate are converted to silicon carbide. The silicon carbide-coated CC composite part is removed from the retort, cleaned and inspected. During cool down from 1650°C (3000"F), the silicon carbide coating contracts slightly more than the carbon substrate, caus- ing crazing (coating fissures). This crazing together with the inherent material porosity provides paths for oxygen to reach the carbon substrate. To obtain increased useful life of this CC structural component, it is necessary to add an additional oxidation inhibitor. The final process used to provide oxidation protection to this type of CC structure involves impregnating (Fig. 15.16) (Curry, Scott and Webster, 1979) this component with TEOS. The part is covered with a mesh, placed in a vacuum bag and the bag is filled with liquid TEOS. A 5-cycle TEOS impregnation is then performed with the bagged part. After the fifth TEOS cycle, the part is removed from the bag


W

9 320 9 260

8 150 w 95 U 40

5 200

W

8 -20

45 45 45 45 2 .5 I H R S TO TO T O TO T O H RS 6 o w m 6 a 2% MIN MIN M I N MIN HRS

F IVE CYCLES . EVACUATE AND

COATED WEIGHT/ COVER INSTALL RECORD wlm - VACUUM PART

PERSPEC VERIFY MESH FITTINGS 206.741 - S R i - F I L L

COVER

VACUUM

- PRi TUEES

WITH MESH

BAG

CONNECT- BACKFILL REMOVE 0 CURE NDE - % HR @ W ' F - X.RAY VACUUMAND WITHTEOS BAG

F ILL LINES M I X CLEAN - 6HR B W F - EDDY MIXTEOS OVEN CURE PART CURRENT I L I W I D I 225" + 5O'F -ULTRASONIC AND FILL I 4 5 6 0 MINI 0 WEIGH/ RESERVOIR FEED TEOS RECORD

M I X A5 R E Q D TO COVER 5 T H CYCLE CURE U TO 2x H R MlNl COOL T O 1WoF

Fig. 15.16 TEOS impregnation (Curry, Scott and Webster, 1979).

and oven cured at 315°C (600"F), liberating all of the hydrocarbons. This procedure leaves silica (SiO,) in all of the microcracks and fissures greatly enhancing the oxidation protection of the CC structure.

labeled Space Shuttle material, is the strength level of the reusable carbon-carbon (RCC) material used in the Space Shuttle thermal protection system. Even though this material is made with low-strength carbon fibers, its strength efficiency is superior to both superal-

15.7 MECHANICAL PROPERTIES

The extreme thermomechanical requirements of the Space Shuttle have been the impetus for evaluating properties of low-density CC. The use of CC on the nose cap and leading edges of the Space Shuttle makes it imperative to know as much as possible about all the characteristics of this material. The effect of temperature on the ratio of tensile strength to density for several classes of high-temperature materials is shown in Fig. 15.17. The major advantage of CC materials for high-temperature applications is that they do not lose strength as the use temperature is increased. This property is in contrast to other materials such as superal- loys and ceramics. Figure 15.17 shows three levels of CC strength efficiency. The first,

loys -and ceramics at >lOOO°C (1832°F). Development of advanced carbon-carbon (ACC) composites has produced a material that is twice as strong as the CC composite first put on the Space Shuttle. The ACC material is made using woven carbon cloth. When unidirectional carbon fiber tapes are interplied with woven cloth to create a hybrid ACC, strength in at least one direction can be increased by >345 MPa (>50 000 psi). Current data on thermomechanical and thermochemi- cal properties of some of the advanced CC systems show that material composition, oxidation resistance, processing, joining and fiber architecture are producing noticeable improvements in CC materials and structures (Curry, Scott and Webster, 1979; Buch, 1984; Rummler and Sawyer, 1984; Ransone and

Thermal properties 345

Temperature, OC

I I I **%O 0 ,550 1100 1650

- High-strength carbon-carbon <- 160

800 1600 2400 3200 4000 Temperature, OF

Fig. 15.17 Strength-to-density ratio for several classes of high-temperature materials.

Ohlhorst, 1984; Webb, 1985; Gray and Engle, 1985; Johnson and Finley, 1985; Sawyer and Moses, 1985; Maahs and Ransone, 1985; Ohlhorst and Ransone, 1985).

CC components on the Space Shuttle are required to have adequate strength at design temperatures to withstand the aerodynamic loads of flight and to continue to do so for the operational life of the component. Minimum mechanical properties are guaranteed through statistical analysis of a data sampling having at least 99% probability and 95% confidence. The primary variables affecting the structural design allowables are temperature, material thickness, coating thickness, biaxial stress conditions and substrate mass loss due to oxidation through the mission life of the component (Table 15.2) (Curry, Scott and Webster, 1979).

Figure 15.18 (Curry, Scott and Webster, 1979) illustrates the typical effect of ply thickness on the allowable stress values for tension, bending, compression and shear used for design. As fabricated, room temperature mod-

ulus values for the TEOS material are shown in Fig. 15.19 (Curry, Scott and Webster, 1979). The effect of temperature on the as-fabricated tension strength properties is shown in Fig. 15.20 (Curry, Scott and Webster, 1979). As shown in Figs 15.17 and 15.20, the strength of CC composite material does not decrease sig- nificantly with temperature. Typically, above 1425°C (2600°F) there is an increase in strength.

The effect of substrate mass loss through oxidation on tensile strength is shown in Fig. 15.21 (Curry, Scott and Webster, 1979). Mass loss results in a significant reduction in design allowable stress, emphasizing the value of the additional oxidation protection provided by the TEOS treatment.

15.8 THERMAL PROPERTIES

15.8.1 THERMAL OXIDATION

A critical requirement when using CC composites is the ability to withstand numerous

346 Carbon-carbon composi tes

Table 15.2

Mechanical properties test TEOS __

Non-TEOS -.

as fabricated conditioned as fabricated conditioned

Flexure Tension Compression Shear Comer flexure Interlaminar tension Interlaminar shear Coefficient of thermal expansion Impact, etc. Simultaneous cycling Mission cycling Tension Flexure Compression

201 196 170 192 20 20 30 10 20 -

3 3 3

221 80 57 79

50 19 6 10

-

-

36 24 27

40 52 46 17 3 5

9

3

-

-

- - -

24 34 28 17 3 5

6

18

-

-

- -

-

Total 868 609 1 75 135

NOTE BREAK IN SCALE

COMPRESSION

TENSION 7

IN-PLANE SHEAR

14 I I 1 I 1 I 15 20 25 30 35 4 0

NUMBER OF PLIES

Fig. 15.18 Design allowables at room temperature as-fabricated (Curry, Scott and Webster, 1979).

Thermal properties 347

z . 0 2 Y

E 35 2

v) 2 w I-

0 -

28 m s - s < 2 1 - 2

35 r TVPICAL SECANT HOOVLUS

___- 28 PLY ---_______________--------- 111 PLV

.

-

u r I . I I I I I 1 I

t E 14 Y

7.0 1 Gw

3.5 1 1 I 1 1 I 1 I5 10 25 30 35 4a

NUMBER OF PLIES

Fig. 15.19 Design allowables at room temperature as-fabricated (Curry, Scott and Webster, 1979).

thermal and thermomechanical loads during re-entries of the Space Shuttle into the earth's atmosphere. Although CC Space Shuttle components have an oxidation-inhibiting silicon carbide coating, they can lose mass over an extended temperature range without apparent surface recession. Photomicrographs of CC specimen surfaces show minute fissures and thermal microcracks, some of which terminate at the coating substrate interface. Specimens exposed to convective and radiant heat trans- fer tests micrographically have shown the presence of voids at the coating substrate surface. Tests to characterize the effects of these

flaws were performed over a wide range of pressures and temperatures in both plasma arc jets and radiant-heating test facilities. Arc jet tests on CC specimens ranged in temperature from 815°C (1500°F) to 1870°C (3400°F) and atmospheric pressures from 0.01 Pa to 0.10 Pa. Radiant-test conditions ranged from 420°C (800°F) to 1425°C (2600°F) and pressures ranging from 0.01 Pa to 1.0 Pa. Mass-loss data for the CC shuttle specimens exposed to the arc jet and radiant-heating tests are presented in Figs 15.22 and 15.23. Figure 15.22 (Curry, Scott and Webster, 1979) shows mass loss at 980°C (1800°F) and 0.05 Pa as a function of exposure

Fig. 15.20 Design allowables as-fabricated (Curry, Scott and Webster, 1979).


n z 19.6 -i - W .x

I 14.7 v) v) s $ 9.8 4 I

4.9

38 PLY 28 PLY 25 PLY

1s PLY

-

-

-

-

8

3 c UJ 1.4

I 1 I 1 J 0 4.9 9.8 14.7 1 U.6 24.5

MASS LOSS - kdm’x 1P

Fig. 15.21 Design allowables conditioned (Curry, Scott and Webster, 1979).

n 39.0

NONTEOS \ /

P /

- 9 . 8 1 1 W I T H TEOS

I I I I 1 a 2 4 6 8 10

EXPOSURE TIME HOURS

P /

/

d/ 9 8

,P NON TEOS /

/

/

/ ’ , ,

d / ,’ &--e--- B \WITH TEOS

/ ,P NON TEOS /

/ ’ , ,

d / ,’ &--e--- B \WITH TEOS

0 2 4 6 8 10 EXPOSURE T I M E HOURS

Fig. 15.22 Mass loss comparison plasma arc jet environment (Curry, Scott and Webster, 1979).

Fig. 15.23 Mass loss comparison radiant environment (Curry, Scott and Webster, 1979).

Applications 349

14.5 7

11.1-

'F ai.

E B

1 5'8' 8 i' 2.9. a f

\ ,- SPECIFIC HEAT

THERMAL CONDUCTIVITY PARALLEL TO PLY

-THERMAL CONDUCTIVITY PERPENDICULAR TO PLY

+--TOTAL EMISSIVITY

-0 - -28(1 -18 280 538 818 1094 1372 16sO

TEMPERATURE "C

Fig. 15.24 Reinforced carbon-carbon thermal properties (Curry, Scott and Webster, 1979).

thermal conductivity is dependent upon the mass loss experienced by the CC composite, resulting from subsurface oxidation.

Results of thermal conductivity studies for shuttle CC composite shuttle materials are shown in Fig. 15.24 (Curry, Scott and Webster, 1979). To simplify thermal modeling, no differ- entiation has been made for conductivity variation resulting from the number of plies in the substrate. Results for conditioned specimens having a mass loss of 0-5 Pa (0.1 lb/ft2) suggest that thermal conductivity decreases with mass loss. Figure 15.24 also shows that neither specific heat nor emittance was affected by material or mass loss conditioning.

15.9 APPLICATIONS

An example of the state of the art in CC composite applications is a one-piece, bladed turbine rotor that, in service, is coated to pre- vent oxidation. The rotor offers higher temperature performance without cooling; low weight and use of low-cost, non-strategic materials (Miller and Grimes, 1982). Other gas turbine engine applications using CC composites include exhaust nozzle flaps and seals, augmenters, combustors and acoustic panels.

CC material systems using coatings, TEOS and additions to the basic CC recipe have improved the oxidation resistance of products made of CC composites by an order of magnitude. These composites are being used in products such as the nozzle in the F-100 jet engine afterburner, turbine wheels operating at >40 000 rpm, nonwetting crucibles for molten metals, nose caps and leading edges for missiles and for the Space Shuttle, wind- tunnel models and racing car and commercial disk brakes (Klein, 1986).

Pushing the state of the art in CC composites is the piston for internal combustion engines (Miller and Grimes, 1982; Taylor, 1985). The CC piston (Fig. 15.25) would per- form the same way as any piston in a reciprocating internal combustion engine while reducing weight and increasing the mechanical and thermal efficiencies of the engine. The CC piston concept features a low piston-to-cylinder wall clearance; this clearance is so low, in fact, that piston rings and skirts are unnecessary. These advantages are made possible by the negligible coefficient of thermal expansion of this kind of CC (0.54 x

/ O F ) . (Carbon-carbon composites can have a range of thermal expansion coefficients,


Fig. 15.25 Carbon+arbon automotive piston.

depending on the processing techniques.) CC material maintains its strength at elevated temperatures allowing the piston to operate at higher temperatures and pressures than those of a comparable metal piston. The high emittance and low thermal conductivity of the CC piston should improve the thermal efficiency of the engine because less heat energy is lost to the piston and cooling system. The elimination of rings reduces friction, thus improving mechanical efficiency.

Besides being lighter than conventional pistons, the CC piston can produce cascading effects that could reduce the weight of other reciprocating components such as the crank- shaft, connecting rods, flywheels and balances, thus improving specific engine performance (Taylor, 1985).

ACKNOWLEDGEMENTS

The author acknowledges Mr. D.M. Curry of NASA Johnson Space Center and H.C. Scott and C.N. Webster of the Vought Corporation for the data, as referenced, on which a portion of the present paper is based. Acknowledgement is also given to Mrs. H.A. Coombs for her valuable contribution in assisting in the formatting of this paper.

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