d x 

NASA Contractor Report 180863 

Stress Rupture Behavior of Silicon 

Carbide Coated, Low Modulus 

CarbonKarbon Composites 

Gary A. Rozak and John F. Wallace 

Case Western Reserve University 

Cleveland, Ohio 

January 1988 

s 

Prepared for 

Lewis Research Center 

Under Grant NAG3-464 

National Aeronautics and 

fBASA-CB- 180863) STRESS BUPTUBE BERBVIOB OF N88-14150 

SILICON CBR3IDE COATED, LOW HODULUS 

CARBOI/CARBON COMPOSITES H, S, Thesis [Case 

Uestern Reserve Univ.1 98 p CSCL 11D Unclas 

Uxcn &-iminiQtratlM 

\ G3/24 0 1 I6683 ~ 
TABLE OF CONTENTS 

. 

1NlRODUCTION ................................................. 

Background .............................................. Fabrication cjf Carbon/Carbon ............................ Oxidation of Carbon ..................................... Stress Effects on Oxidation of Carbons .................. Oxidation Protection of Carbon/Carbon ................... Mechanical Testing of Carbon/Carbon ..................... 

PROCEDURES AND MATERIALS ..................................... 

Experimental ............................................ Materials ............................................... Procedure ............................................... Characterization ........................................ 

Stress Rupture Testing .................................. Mechanical Testing ...................................... 

RESULTS AND DISCUSSION ....................................... 

Characterization ........................................ Mechanical Tests ........................................ Elevated Temperature Flexure Testing .................... Stress RupturE Testing .................................. Discussion .............................................. Flexural Testing ........................................ Stress Rupture Testing .................................. 

CONCLUSIONS .................................................. 

REFERENCES ................................................... APPENDICES ................................................... 

A . Procedural Details .................................. B . Tables .............................................. C . Strain Veasurement .................................. 

1 

2 

4 

7 

9 

1G 

13 

17 

17 

17 

20 

20 

22 

23 

24 

24 

27 

29 

32 

37 

37 

42 

44 

46 

62 

62 

83 

95 

. 

i 
The disadvantages of c/c, in addition to the oxidation problem, 

are low thermal expansion, expensive fabrication procedures, and poor 

off axis properties. The strength perpendicular to the fiber 

directions in the composites are as low as 360 psi (2.5 MPa) [2!. The 

shear strength of the composite is also as low as 2000 psi (14 MPa) 

[ZJ. The fabrication of c/c involves long processing times at high 

temperatures and consequently is expensive. The thermal expansion of 

c/c is small and makes it difficult to apply a successful coating 

because the thermal expansions of most coating materials are greater 

than that of c/c. The disadvantages of expensive fabrication and low 

off axis properties will become secondary if a successful coating 

systeni is devel oped. 

The following subjects are reviewed: 1) background of c/c, 2) 

fabrication of c/c, 3) oxidatiori of carbon, 5) oxidation protection 

methods for c/c, and 6) mechanical testing of c/c in flexure. 

Background 

The first carbon matrix composite was inadvertantly produced in 

the late 1950's during ablation studies of composites with oxide 

fibers and a polymeric matrix [l]. 

yielded a carbon char as the matrix when a crucible was accidentally 

covered. Later, the oxide fibers were replaced by a carbon fiber in 

the composite and the first c/c was the result. 

c/c in the succeeding years was directed by its usage in aerospace 

The pyrolization of this composite 

The development of 

[I! 0 

2 
Aerospace appl ications of c/c began with single 1 ife mission 

missiles and evolved into multimission components on the space 

shuttle. A c/c replaced bulk graphite as the material for the nose 

cone and exhaust nozzle structure on single mission missiles for 

better performance [l]. 

modulus c/c, is employed on the space shuttle for such applications as 

the nose cone and leading edges [31. 

greatest heat in re-entry from orbit and require good ablation 

resistant properties r31. 

A reinforced c/c (RCC), low strength and low 

These parts experience the 

The reuse of the shuttle requires an oxidation protection system 

for the c/c to extend its life. The RCC material has a silicon 

carbide coating for protection [3]. The coating is appl ied to the 

parts by a pack cementation conversion of the surface to silicon 

carbide. The coating is further enhanced by a surface layer of tetra 

ethyl ortho silicate (TEOS), which is brushed onto the surface and 

then cured to form a si1 icon dioxide layer. 

The improvement in the mechanical properties of c/c was the next 

evolutionary step. Low modulus carbon fibers were replaced by higher 

modulus fibers and resulted in an Advanced c/c (ACC). 

ACC is three times greater than RCC and the modulus is six times 

greater. 

and hypersonic aircraft [33. 

engine exhaust nozzle 1 iners [31 was successful for ground based test 

in a FlOO turbine engine. 

pursuit of coated ACC materials as liners for the two dimensional 

(2-D) exhaust nozzle 1 iner program. 

The strength of 

Applications of ACC are being pursued in the space shuttle 

The appl ication of coated ACC as jet 

The success of this program prompted the 

3 
Oxidation problems, experienced on the coated ACC materials in 

the 2-D exhaust nozzle program, resulted in an increase of research in 

inhibited c/c (ICC). These c/c incorporate chemicals, viz. boron, and 

silicon, in the matrix that inhibit the oxidation of carbon by 

blocking oxidation sites. 

Usages of c/c are not limited to aerospace applications. Carbon/ 

carbon has been used successfully as biomedical implants because of 

the biological chenrical inertness of carbon. Carbon/carbon is f iriding 

use as an asbestos replacement in vehicle brakes because of its 

friction and wear characteristics. 

described in a recent review article by Fitzer [41. 

Further applications of c/c are 

Fabrication of Carbon/Carbon 

The fabrication of c/c requires a choice of composite precursor 

materials for the fibers and matrix. 

entails the lay up, impregnation, carbonization, and densification. 

More details are found in a review by Stoller et a1 [5]. 

The composite fabrication 

One method of forming the matrix of c/c is the pyrolization of 

organic precursors. The avail ab1 e polymeric precursors to the matrix 

consist of high char forming polymers, e.g. phenolic resins, furfural 

and pitches [6]. 

carbonization cycle. 

3GOO"F (2000°C) will convert the matrix to graphite in the case of the 

pitch and alcohol precursors, whereas the phenolic resins remain 

amorphous after this hcat treatment [51. 

The matrix is amorphous carbon after the 

Further elevated temperature processing above 

Deposition of pyrolitic graphite by chemical vapor deposition 

(CVD) methods is the alternative route of matrix formation and 

4 
involves a reaction of a hydrocarbon gas that deposits carbon on the 

matrix. The properties of this graphite is dependent on the 

deposition temperature and rate. This method can also be eriiployed to 

densify the matrix of charred matrix composites. 

Comercially available carbcn fibers are derived from different 

organic precursors -- rayon, pitch, and polyacrylonitrile (PAK) -- all 

c?f which undergo similar processing steps to arrive at the final 

properties. Other polymers, such as phenolics, phenol formaldehyde, 

resins, and polyimides can be used as fiber precursors but are not 

commercially avail ab1 e. The processing of the fibers invol ves the 

spinning of the precursor into fibers and yarns, stabilization, and 

carbonization. 

temperature of 300" to 500" (150" to 260°C). 

additional heat treatment in a controlled ctmosphere to decompose the 

precursor and yield carbon fibers. 

elasticity of the fibers can be improved along the axis of the fiber 

by a stretching procedure during production. 

modulus of a rayon fiber from 4 Msi (28 GPa) to above 50 Msi (345 

GPaj; strength is increased from 100 ksi (700 MPa) to over 300 ksi 

(2100 MPa) [Sf. Donnett and Bansal [71 give greater details on fiber 

fahrication. 

The fiber yarns arc stabilized with a cure at moderate 

Carbonization is an 

The strength and modulus of 

This can increase the 

Carbon/carbon is fabricated in many fiber orientations. Some 

applications employ chopped fibers in random orientation or carbon 

felts as t.he base preforms. Woven fabrics of carbon yarns are common 

for aerospace appl ication. 

is faster [8]. The shear strengths in the plane of the ply is better 

Cloths are easier to hand1 e and the lay up 

5 
with a woven fabric than unidirectional plies of yarns which are laid 

up orthogonally [93. 

orthogonal to each other. The yarn directions are called the warp and 

fill and the warp yarns are stitched, i.c. a crossover, by the fill 

yarns. Consequently, the warp yarns are generally straighter anti 

stronger than those in the fill direction. 

Some of the available weave types are the plain, five and eight 

The carbon fabric consists of yarns running 

harness satin, and unbalanced. The plain weave is fabricated such 

that fill yarns are stitched over every other warp yarn. 

satin weave refers to the degree of stitching, e.g., one stitch cver 

every eight warp yarns eight harness satin weave. 

weave contains a significantly greater number of warp yarns and is 

generally referred to according to the ratio of warp to fill yarns, 

e.g., six to one [ZI. 

The harness 

The unbalanced 

The final properties of the composite are controlled by the 

direction of the fiber layup. 

indicates all the warp fibers are aligned in the same direction. 

balanced or 0-90" 'layup indicates that the fabric in every other ply 

is rotated by ninety degrees. 

in which the fabrics are laid up at angles off the major directions, 

e.g., a +/-45" ply. 

A warp aligned layup or 0" layup 

A 

An off axis layup refers to composites 

The processing step after fabric layup is the matrix formation. 

Forming a matrix with polymeric precursors will yield a polymeric 

precursor processing composite with two distinct sides, the bag and 

mold. 

processing, the mold is placed in a plastic impregnation bag. 

The carbon fabric is laid up on an open mold and, in 

A 

6 
vacuum is applied to this bag arid then the fabric is impregnated with 

the precursor. The side against the mold has a smooth texture of the 

mold. The bag side is not as smooth because the composite surface is 

not restrained by the bag. 

test results. 

These differences may influence mechan 

After precursor impregnation are the cure and carbonization. 

cure crosslinks and stabilizes the polymer. 

decomposes the polynier to a carbon char at high temperature and 

controlled atmosphere. 

The carbonization cyc 

Several cycles of impregnation and 

cal 

The 

e 

carbonization are required to produce a composite with significant 

integrity . 

Another method to densify the composite is a CVD deposition of 

pyrolitic graphite to the matrix. This method requires heating the 

layup to high temperatures and passing a carbon containing gas through 

the fibers [l]. The carbon is deposited to the matrix from this gas. 

Oxidation of Carbon 

The oxidation of carbon and graphite has been analyzed by a 

number of researchers [lo-141, and some with a particular reference to 

c/c [15-171. The oxidation of carbon is influenced by porosity, 

impurity content and crystal1 inity [ll-141. 

researchers report different temperatures at which carbon oxidation 

exceeds one weight percent per twenty-four hours from 570°F (300°C) 

E131 to 840°F (450°C) [14]. However, temperature is the greatest 

factor in determining the rate of carbon oxidation [13,141. 

reaction rate below 1110°F (600°C) is controlled by the reaction of 

This may explain why 

The 

7 
iictive sites on or below the surface [10,121. These sites include 

dislocations, basal plane edges, and impurities [19]. 

(600°C) and 1470°F (800°C) the reaction rate is controlled by the 

Between 1110°F 

transport of reaction gases, 02, C02, and H20, to the active sites. 

Above 1470°F (800°C) the rate is cmtrolled by the transport of 

reaction products across a stagnant film at the surface of the carbon. 

The reaction kinetics of c/c are similar to those just discussed fur 

carbon [15,181. 

The oxidation rate of carbon can be reduced by interfering with 

the operating oxidation mechanism [13]. 

controlled by the number of active sites, and these sites can be 

reduced by decreasing the porosity or chemically blocking the sites 

from oxidation. 

carbon by acting as sites for oxidation [lo, 11, 13, 171. Therefore, 

these impurities must be minimized to decrease the rate. A few 

elements have an inhibiting effect on oxidation, such as boron, 

silicon, and phosphorus [ll]. 

oxidation can be reduced by decreasing the percent of reactants in the 

gas or reducing the flow rate across the specimen [14]. 

will reduce the oxidation rate at all temperatures by effectively 

blccking the sites on the surface [13, 141, but cracks in the coating 

can serve as paths to oxidation sites. 

Low temperature oxidation is 

Most elements catalyze the oxidation reaction of 

At higher temperatures the rate of 

A coating 

The oxidation of carbon arid c/c was examined by a few researchers 

[16, 20-221. 

pitch matrix c/c at the temperatures of 1110°F (600°C) and 1470°F 

(800°C). Preferential oxidation of the matrix over the fiber is 

han et a1 [16] analyzed the oxidation in a PAN fiber, 

8 
reported. 

furfuryl alcohol miitrix c/c , between the temperatures of 1200°F 

(650°C) and 1560°F (850°C). The oxidation attack is reported to begin 

at the fiber matrix interface and then progress successively to the 

grain boundaries in the matrix, the graphitic portions of the matrix, 

and then to the optically anisotropic regions of the matrix. The last 

part of the matrix to oxidize is the glassy region of the matrix. 

This preferential oxidation of the crystalline carbon over the glassy 

carbon is in conflict with the reports that glassy carbon oxidizes 

more readily than graphitic carbon [231. 

Yasuda et a1 K151 examined the oxidation of PAN fiber, 

Stress Effects on Oxidation of Carbons 

The effect of stress on the oxidation rate has been examined by a 

few researchers with contradictory results [24-273. Stroud and 

Runnnler measured the effect of stress on the oxidation rate of a 

coated RCC for space shuttle re-entry simulation tests. 

at 2480°F (1360°C) 1251 and stress level of 1.2 ksi (8.17 MPa) 

exhibited a six percent increase in mass loss over an unstressed 

speciirien; at 1160°F (625°C) C241 and 1.0 ksi (6.8 MPa), no significant 

increase in mass loss occurred. Krefold et a1 [26] determined that a 

tensile stress on graphite increases the oxidation rate significantly 

in argon with small amounts of water vapor at 1830°F (1000°C). 

Thrower and Marx [27] conducted a similar experiment using compressive 

stresses. An effect on the oxidation rate was not found. 

Examination 

The effect of oxidation on the degradation of c/c mechanical 

properties was examined by Zhao et a1 [28]. The c/c examined was a 

Y 
semirandom chopped PAN fiber with phenol, furfural resin char and CVD 

densificat on. 

in flowing air. Catastrophic effects on the mechanical properties 

occurred, e.g., for a twenty percent weight loss, Young's modulus 

decreased by 75% and flexure strength by 64%. 

composite did not appreciably change with a 20% mass loss, indicating 

that oxidation occurred internally. 

The oxidation tests were conducted at 1110°F (600°C) 

The volume of the 

Oxidation Protection of CarbonjCarbon 

Numerous oxidation resistant coating and coating methods have 

been attempted for c/c. 

nitrides, silicides, borides, and metals have been applied to c/c 

substrate with pack cementation, CVD, plasma spray, sputtering, and 

el ectro-deposition methods [Z]. Many coating systems are unsuccessful 

because the mismatch of the thermal expansion coefficients causes 

coating failures. 

induces tensile strains in the coating because the thermal expansion 

of the c/c is less than that of the coating [291. Ceramic coatings 

respond to this strain by cracking, leading to the oxidation of the 

substrate. Platinum and other ductile metal have been tried in an 

attempt to overcome this brittle failure mechanism, but failures have 

occurred with pinhole oxidation C3O1. Room temperature deposition of 

coatings has also been attempted to avoid the tensile strains and 

cracks, but these coatings lacked sufficient adherence to the 

substrate. 

Coating systems with various carbides, 

Cooling from the coating deposition temperature 

10 
Some ceramic materials are good candidates for coating c/c, but 

their success depends on their ability to close or seal the cracks at 

temperatures below 800°F (425"C), the temperature for the initial 

oxidation of carbon. The cracks in the coating are closed with 

thermal expansion forces on heating the composite. The temperature 

where the cracks close is called the crack closing temperature. If 

the coating system behaves perfectly elastically, this temperature is 

the coating deposition temperature. This assumption may not be valid 

for the c/c system because the substrate has low shear strength. 

this low strength the c/c could accommodate some of the thermal 

expansion strains with plastic deformation. 

substrate reduce the amount of thernial strain recovery and 

subsequently reduce the crack closing temperature. 

may be reduced by application of a slurry of ceramic powders to the 

coating. These powders fill the cracks and decrease the width of the 

crack. A smaller crack width reduces the amount of strain recovery 

and subsequently reduces the crack closing temperature. 

With 

P1 astic strains in the 

This temperature 

A coating system with a silicon base seals the cracks with a 

glassy layer. 

into the cracks to block the oxidation. The glass layer is formed by 

the oxidation of the si1 icon based coating. The melting temperature 

of silica is high and is lowered by applying glass modifiers as an 

overlay. 

chemistry of the glass and a successful overlay activates around 

1400°F (760°C). Other glass compositions activate at lower 

temperatures, but at the expense of higher temperature protection. 

This coating system operates because the glass flows 

The activation of the glass fluidity depends on the 

11 
The lower temperature activated glasses are too fluid at the higher 

temperature and vaporize or flow off the coating. Therefore, the use 

of the higher activated glass is necessary, and an oxidation window 

exists between 800" and 1400°F (425" and 760°C). 

Another oxidation protection method involves chemically 

inhibiting the oxidation of the c/c substrate. 

the oxidation of carbon are silicon, phosphorus and boron. 

[161 found that silicon carbide was an effective inhibitor up to 

1110°F (600°C). McKee et a1 [221 examined the effects of phosphorus 

compounds on the oxidation rate between 1110°F (600°C) and 1560°F 

(850°C). The rates were different by an order of magnitude between 

the treated and untreated materials, but the apparent activation 

energies for oxidation were the same. This indicates a lower number 

of oxidation sites in the inhibited material with no change in the 

oxidation mechanism. Zhao et a1 E281 employed phosphorus as an 

oxidation inhibitor in c/c for tests at 1110°F (600°C) but did not 

report on the oxidation characteristics. The effectiveness of the 

phosphorus compound at inhibiting c/c above 1650°F (900°C) is 

questionable. The oxidation tests conducted by McKee [22] at a 

constant rising temperature found a 10% oxidation weight loss for a 

phosphorus inhibited material in the temperature range of 1650°F 

(900°C) to 1830°F (1000°C). 

Elements that inhibit 

Han et a1 

The inhibition of c/c oxidation with boron was examined by a 

number of researchers [20, 211. McKee determined that a borate 

additive was effective at retarding the oxidation up to 1830°F 

(1000°C). Above this temperature, it was less effective because of 

12 
i 

the volatility of borate glasses. 

boron oxide in both small and large amounts effectively decreased the 

oxidation rate. This testing was conducted at temperatures below 

1520°F (827°C) to avoid any possible loss of the boron oxide inhibitor 

through volatilization. Small amounts of boron oxide acted to block 

oxidation sites; large amounts acted as a diffusion barrier [21]. 

Ehrburger et a1 [211 found that 

Future oxidation protection systems for c/c will incorporate both 

a coating and an inhibitor. 

because the glass overlay can seal the cracks in the coating on cool 

down until the glass becomes too viscous to flow. The inhibitor is 

active after the glass is ineffective to block oxidation. 

concern will be the loss of the inhibitor because of volatility. 

However, a coating may reduce the loss of inhibitor. 

This cculd yield the best properties 

The largest 

Mechanical Testing of Carbon/Carbon in F1 exure 

The mechanical properties of tensile, compression, and shear 

strengths and the modulus of c/c are a concern for designers of c/c 

components. The high temperature test methods have been developed for 

a vacuum environnient C311. 

the strengths is a great concern. 

composites in an oxidizing atmosphere are difficult to conduct because 

of the problems with gripping C321. 

determining the strength and overcoming the gripping problem. 

However, the analysis of the results is more complicated. 

compressive, and shear stresses develop during a flexural test, and a 

pure stress state is required for accurate property measurements. A 

The effect of an oxidizing environment on 

Tensile tests of high temperature 

Bend testing is a method of 

Tensile, 

13 
discussion of the advantages and disadvantages of the flexure test 

method, particularly for brittle materials, is found in a number of 

reports 133, 341. 

Flexure tests on coated c/c were conducted by researchers 12, 31, 

331; a few of these reports are reviewed here. 

[351 conducted flexural tests on a high modulus, coated c/c at room 

and elevated temperatures in air. The flexural strengths were not 

reported but relative flexural loads for these c/c materials showed a 

higher flexural strength for coated specimens compared to uncoated 

specimens. The flexural failure 1 oads increased from room temperature 

to 2400°F. 

inches per minute, were employed in the flexure tests with no apparent 

affect on the strength of the composite. 

Quinn and Stiglich 

Two crosshead displacement velocities, 0.02 and 0.005 

Vought Corporation conducted mechanical tests on RCC material 

with a proprietary diffusion coating of silicon carbide and examined 

the effect of temperatures on the mechanical properties [361. The 

data were presented in graphical form and the results were not 

summarized. The deflection for the modulus calculation was measured 

at two-thirds of the failure load. The modulus of a 19 ply RCC 

decreased from 2.35 to 2.15 Msi from room temperature to 500°F and 

then increased with temperature to 2.70 Msi at 2500°F (1370°C). The 

flexure strength decreased from the room temperature value to 8.0 ksi 

(55 MPa) at 1000°F (538°C) and increased to 9.0 ksi (62 MPa) at 2500°F 

( 1370°C). 

The application of mechanical loads to a silicon carbide coated 

c/c can cause the formation of cracks in the coating. An oxidation 

14 
pathway is opened by these cracks if the glass overlay does not seal 

the cracks. The glass overlay will be ineffective if the glass is too 

viscous to flow into the cracks or the cracks are too large for the 

glass to fill. The ability of coating to seal the cracks has been 

examined in a stress rupture tests [35,371. Quinn and Stiglich [35] 

conducted stress rupture testing on eleven silicon carbide coated, 

high modulus, c/c specimens at the temperatures of 1830", 2190", and 

2370°F (lOOO", 1200", and 1300°C) with significant scatter in the 

results. Most of the specimens that failed in the stress rupture test 

failed within two hours, while others at the same or greater 

normalized loads survived over a thousand hours. This inconsistency 

of failure times was experienced on a si1 icon carbide pack cementation 

coatings on high modulus materials by another researcher 1371. Static 

fatigue tests of monolithic ceramics exhibited a similar failure 

distribution 1381. The possibility of a stress rupture limit is 

indicated [35] but an insufficient number of tests were conducted. 

The combined loading and oxidation tests at 1600" and 2000°F (871" and 

1093°C) are examined in a report [391. 

permanent deformation was observed after four hours of testing. 

Failures are not reported but 

Carbon/carbon composites retain their mechanical integrity to 

high temperatures and therefore, have potential in aeropropul si on 

applications. The good thermal qualities of c/c have already been 

employed in the space shuttle leading edges. Further implementation 

of c/c is restricted by the absence of inherent oxidation resistance 

and expensive fabrication procedures. The fabrication expense results 

from long process times to form the matrix by either CVD deposition of 

15 
carbon or charring of an organic precursor. The expense of c/c 

becomes minor, if a successful oxidation protection system is found. 

Oxidation of c/c occurs significantly above 800°F (425°C) and is 

reduced by a surface coating or impregnating the c/c with chemicals 

that inhibit carbon oxidation. Many coating systems have been 

attempted on c/c and have failed because thermal expansion mismatch 

strains crack the coating. 

failure mode by the flow of a glass surface layer into cracks to block 

oxidation. This coating type is more successful at higher 

temperatures than 1 ower temperatures. Mechanical 1 oads can a1 so crack 

the coating and allow the oxidation of c/c. This failure mode is 

examined jn the stress rupture test. The present research employed 

the stress rupture test to examine the effects of stress on the 

ability of a silicon carbide coating to protect a low modulus c/c. 

Silicon based coating can prevent this 

16 
PROCEDURE AND MATERIALS 

Experimental 

The primary purpose of this research was the examination of the 

abil ity of si1 icon carbide coating on a low modulus c/c to maintain 

oxidation protection when mechanically 1 oaded. The tests were 

conducted in fast fracture and stress rupture modes. 

measured were the shear and flexural strengths, and density. The 

coating thickness was measured, and the coating was characterized by 

chemical analysis. Differences among the three coating batches and 

between the bag and mold sides were determined in the characterization 

segment. The mechanical tests on coated c/c were conducted at 

elevated temperatures of 1000" , 1500" , 2000°, and 2400°F (538", 816", 

1093", and 1316°C). Flexural strengths and modul i were determined in 

a four point bend test at each temperature. The stress rupture tests 

were conducted with the same test geometry as the flexure test and 

were statically loaded. 

one hundred hour time period because this is a goal for the 

implementation of c/c in aerospace applications. Times to failure in 

the stress rupture tests are reported at various stress levels for 

each test temperature. 

Properties 

The stress rupture tests were conducted in a 

Materials 

The two major components of c/c specimens in the test sequence 

are the substrate and coating. The low strength, low modulus 

substrate is composed of low modulus fibers and a charred polymeric 

17 
matrix. The coating is a silicon carbide conversion of the composite 

surfaces, plus an overlay of glass sealers. The processing procedures 

of the substrate and coating are proprietary to Hitco of Gardena, 

California; therefore, only 1 imited details about the production are 

expl ained here. 

The precursors to the c/c substrate are low modulus carbon fibers 

for reinforcement and a phenolic resin to form the matrix. The fibers 

are rayon based and the yarns are woven into a balanced eight harness 

satin weave. 

fiber is 100 ksi (690 MPa); the modulus is 6 Msi (4.1 GPa). Each yarn 

contains approximately 720 fibers. The phenolic resin is carbonized 

at elevated temperatures. 

carbonization time and temperature, and are proprietary. 

The fabric designation is 62206. The strength of the 

The properties of the char depend on 

The fabrication of the c/c substrate involves layup, 

impregnation, carbonization, and cutting into specimens. 

panel is molded to the dimensions of 17 by 12 by 0.24 inches thick (44 

x 21 x 0.6 cm.) during the layup procedure. The G2206 fabric is cut 

to size and nineteen plies are incorporated in the fabrication. All 

the warp fibers are aligned in the same direction, and the crossovers 

are nested. 

the crossover side of each ply contacts the crossover side of the 

adjacent ply. A bleeder cloth is set on the exterior surface of the 

molded panel and absorbs excess resin during impregnation. 

Initially a 

Nesting is accomplished by laying up the fabric so that 

Impregnation and carbonization are the next fabrication steps. 

The mold and layup are bagged using standard composite fabrication 

procedures. The phenolic precursor is vacuum impregnated to form the 

18 
panel and cured prior to carbonization. The bleeder cloth is 

subsequently consumed in the carbonization cycle and impresses the 

texture of the cloth to the bag side of the composite. 

The panel is cut into specimens after the first carbonization 

cycle. A one-quarter inch strip is cut from the edges of the panel to 

remove possible edge effects. The dimensions of the specimens cut 

from this panel are 7.75 inches long by 0.6 inches wide (20 x 1.5 cm). 

The specimen thickness is approximately 0.24 inches (0.6 cm). The 165 

specimens for this test sequence are cut from nine panels. After the 

specimens are sectioned, two additional impregnation and carbonization 

cycles are conducted on the specimens to complete the processing. 

The low modulus c/c substrate is coated with a pack cementation 

conversion to silicon carbide plus an overlay of glass sealers. 

Twenty-eight of the 165 specimens were not coated. This particular 

coating was developed by Vought Corporation for the ACC material and 

is moderately successful at preventing the oxidation of this 

substrate. The Vought designation for this coating is type one, mod 

one. This coating technology is licensed to Hitco, the fabricators of 

the specimens. 

The cementation coating procedures require pl acing the specimens 

in a pack of powders in a controlled atmosphere furnace. The pack of 

powders consists of a blend of mostly inert aluminum oxide and silicon 

with minor additions of silicon carbide and boron oxide. The c/c 

substrate is packed in this powder and heated at a controlled rate. 

The outer surfaces of the specimen are converted to silicon carbide. 

A glass overlay is applied to the coating surfaces after the 

19 
cementation process and is based on the organometallic compound, tetra 

ethyl ortho silicate. Three coating batches were required to coat the 

137 specimens in the test sequence. Two batches of 50 and one of 37 

specimens may lead to inconsistencies in testing results. Therefore, 

all tests inclLided at least one specimen from each coating batch. 

Procedure 

The experimental test procedures are divided into three 

sections: the characterization, the flexure tests, and the elevated 

temperature stress rupture tests. The characterization was employed 

to distinguish any inherent differences among the three coating 

batches, and differences between the sides of the specimens. 

and elevated temperature flexure tests were conducted with similar 

procedures, and established the strengths at the test temperatures. 

The stress rupture tests were conducted with the same testing geometry 

as the flexural tests but with the specimen statically loaded at 

stress levels less than the strength. The detail procedures for each 

test are presented in Appendix A. 

The room 

Character; zat i on 

The physical properties of density and coating thickness are 

useful in distinguishing differences between coating batches and 

specimen sides. 

determined by a wax immersion technique for porous material [401. 

substrate density was determined with four coupons, 0.6 x 0.6 x 0.24 

inches (1.5 x 1.5 x 0.6 m), which were cut from three uncoated 

flexural specimens. Each specimen originates from different panels. 

20 

The density of the coating and substrate is 

The 
Imea s ur em 

the coat 

E. 

The 

The density of the coating was measured by cutting coupons from 

both the bag and mold sides. These coupons are prepared by procedures 

explained in Appendix A. Four coating coupons were used for density 

nts, two each from the bag and mold sides. 

ng and substrate are recorded in Tables B and C of Appendix 

The density of 

thickness of the coating was measured at ten equidistant 

positions on a micrograph with a magnification of 50x. The thickness 

of the glass overlay was measured simultaneously. The mean, standard 

deviation, maximum and minimum values of thickness are reported for 

each coating batch with bag, mold, and edge sides. One micrograph was 

made with the warp yarns running parallel to the coating and another 

with fill yarns parallel to the coating. The edge micrographs were 

similar except that the fibers were perpendicular to the edge coating. 

The coating thickness values are recorded in Table D of Appendix B. 

Mass chromotography and x-ray diffraction were used to fndicate 

differences between coating batches and between the sides of the 

specimen. The chrornotography techniques yielded an el eniental analysis 

of the coating; the x-ray diffraction determined the crystallographic 

phases of the coating. The chromotography was employed for analysis 

of coating and ash on each side of a specimen from each coating batch. 

X-ray diffraction phase analysis of the coating was done on the bag 

and mold side from each coating batch. The results of this analysis 

are recorded in Tables E, F, and G of Appendix B. 

The x-ray radiography shows differences between batches and 

internal defects within a specimen, such as delamination or 

21 
deleterious coat 

initial specimen 

perpendicular to 

ng penetration. The radiograph is a record of the 

condition. Radiographs were taken parallel and 

the plane of the specimens. 

Mechanical Testing 

Flexure and shear test established the base line room temperature 

properties. 

fixture of stainless steel. 

the support span was varied to yield either a flexure or shear 

failure. 

flexural mode and the shorter span of 1 1/2 inches (13.8 cm) failed in 

shear. 

(0.13 cm/min.). 

chart recorder. 

significantly. Strengths were calculated at the maximum load. The 

elastic modulus was determined at the deflection corresponding to one 

half of the maximum load. This deflection was corrected for the 

compliance of the test frame. The strength and modulus are reported 

for six coated specimens and eight uncoated specimens in Appendix By 

Tables H and I. 

The properties were measured with a four point bend 

The load span is one inch (2.5 cm) and 

The longer span of 6 1/8 inches (15.5 cm) yielded in a 

The cross head deflection rate was 0.05 inches per minute 

The load deflection curve was recorded on a strip 

The deflection was stopped when the load dropped off 

A sintered alpha silicon carbide fixture was employed for testing 

the coated c/c specimens. The fixture had the same test geometry as 

the room temperature four point bend fixture. The tests were 

conducted in an Applied Test System (ATS) silicon carbide resistance 

furnace, using an Instron universal test frame to apply the load. 

The specimen and fixture were aligned and bonded with an adhesive 

22 
before being placed in the load train. The furnace was equilibrated 

at test temperature for one hour before testing. The flexural 

strength and proportional limit were calculated and are recorded in 

Tables L to 0 in Appendix B. The modulus was calculated with the 

method described in the room temperature tests. 

Stress Rupture Testing 

The stress rupture tests of coated specimens were conducted at 

the same temperature and test geometries as the elevated temperature 

flexure tests. An ATS compression-creep test frame was employed to 

apply direct weight loading to the specimen. A three point 

extensometer with a linear variable differential transformer measured 

the re1 ative deflection, and the data acquisition system calculated 

the strain. This test was conducted with procedures that are similar 

to the flexure test. The specimen was bonded to the test fixture and 

then placed in the load train. 

temperature for one hour before the load was applied. 

extensometer was brought into contact with the specimen prior to 

loading. The test lasted for 100 hours or to failure. This duration 

was chosen because 100 hours is an implementation goal for coated c/c 

in aerospace applications. 

The specimen was equilibrated at 

The 

23 
RESULTS AND DISCUSSION 

Characteri zat iori 

The density of the c/c substrate is 1.38 grams/ml ; that of the 

coating is 2.63 grams/ml. 

batch are presented in Tables B and C of Appendix B. 

density differences do not occur among the three batches or between 

the bag and mold sides of the coating. 

among the three coating batches is less than 4%. 

The results for specimens from each coating 

Significant 

The maximum density difference 

The average coating thickness for the three coating batches was 

0.014 inches (0.035 cm.) on the ply side of the composite and 0.030 

inches (0.075 cm.) on the edge of the composite. The coating 

thicknesses on the bag and the mold sides were similar within the same 

specimen and are averaged in Table I. 

coating thicknesses is listed in Table C on Appendix B. The thickness 

of a specimen from coating batch 1 was 33% greater than the thickness 

of one from coating batch 2 and 23% greater than one from batch 0. 

A detailed summary of the 

The average thickness of the specimen from coating batch 1 does 

not indicate that this coating batch is thicker than the other two 

because thickness measurement of another specimen from coating batch 1 

and a statistical test indicate otherwise. The average thickness of 

another specimen from coating batch 1 was 0.013 inches (0.033 cm.) 

which is similar to the other batches. A two tailed Z-test on the 

thickness data ot batch 1 and 2 had a Z value of 0.53, which suggests 

that differences between the two coating batches are insignificant. 

24 
TABLE I 

Coating Thickness Summary 

Coating 

Coating Thickness Standard 

Number in. cm. Coating Thickness 

0 0.013 0.033 0.003 0.008 

1 0.016 0.041 0.004 0.010 

2 0.012 0.030 0.002 0.005 

Batch on ply Deviation of 

Average 0.014 0.035 

Coating Thickness 

on edge 

in. cm. 

0.032 0.081 

0.029 0.074 

0.028 0.071 

0.030 0.075 

The chemical analysis of the coating and the ash of the substrate 

after coating are tabulated in Tables E and F of Appendix C. 

as SiOp is the major component and sodium as Na20 is the second most 

commonly occurring element. 

greater than in the coating. Sodium catalyzes the oxidation reaction 

of carbon [121 and has a significant impact on the corrosion of 

silicon carbide [41]. 

ortho silicate glass overlay and acts to reduce the viscosity. The 

other minor elements in the coating and ash are aluminum, boron, 

calcium, and iron. Aluminum and boron are constituents of the pack 

cementation process. 

in the silicon carbide. 

Silicon 

The percent of sodium in the ash is 

Sodium is a constituent of the tetra ethyl 

The calcium and iron are present as impurities 

X-ray diffraction phase analysis of the coating reveals beta 

silicon carbide 6s a major phase. 

tabulated in Table G of Appendix B and are ordered in decreasing peak 

intensity. The bag and mold sides in each of the coating batches 

exhibit the same peaks. 

The x-ray diffraction peaks are 

25 
The results of the x-ray radiography indicate density or coating 

thickness differences by the intensities on the radiographs. The 

thinner or less dense areas of a material appear as brighter areas on 

the radiographs while denser or thicker areas are darker. The 

radiographs for the coated c/c are taken with the specimen parallel 

and perpendicular to the ply side of the composite. 

reveal no delamination or coating penetration into the coated 

specimens. The variation of radiographic intensity that occurs among 

the specimens and within a specimen are exhibited in Figures 1 and 2. 

The intensity of the radiograph varies across the length of many 

specimens. These variations are attributed to variations in the 

coating thickness and not the density. 

thickness measurements support this conclusion. The densities of the 

coatings vary less than 5.5% among the batches or within a specimen. 

The thickest coating measurements correspond to the darkest radiograph 

of the three specimens examined. This variation of the intensity 

observed in both parallel and perpendicular radiographs may originate 

with differences in the environment of the pack cementation furnace. 

The perpendicular radiographs of the c/c specimens reveal that 

The radiographs 

The density and coating 

the specimens were distorted in all coating batches, as illustrated in 

Figure 2 for batch 2. 

the residual stresses originating from the coating process. The 

uncoated specimens are flat. 

the thermal expansion mismatch between the coating and substrate. The 

substrate is low modulus and consequently is easily deformed at low 

stress 1 eve1 s. 

The warpage of the specimens is attributed to 

The stress is developed as a result of 

26 
The radiography shows texture differences between the specimens 

from panels 8 and 11 than those from other panels. The texture in the 

radiograph of panels 8 and 11 consists of small rounded circles with a 

close ordered pattern. The radiographs of the other panels exhibit 

larger circles with a greater distance between the circles. 

textures are exhibited in Figure 1. 

Both 

The perpendicular radiographs for specimens show differences 

between panels. Most specimens have a dark line about the centerline 

of the specimen as observed in Figure 2. 

the specimen in the radiograph and is assumed to be a higher density 

area. The specimens from panel 8 do not show this line. Panel 11 has 

a lighter line sbout the centerline, and this may indicate an area of 

1 ower density . 

This line runs the length of 

Mechanical Tests 

The final and most significant results in this study are the 

mechanical testing of coated and uncoated c/c samples. Flexure and 

shear tests measured the room temperature strength and modulus in four 

point bending. 

MPa) with an average of 13.6 ksi (94 MPa). Figure 3 is a typical room 

temperature 1 oad deflection curve of uncoated c/c. The proportional 

limit is not apparent in the load deflection curve. Therefore, the 

modulus is calculated with the deflection corresponding to one half 

the maximum load. The average modulus value is 2.3 Msi (16 GPa). The 

measured moduli vary from 2.0 to 2.7 Msi (14 to 19 GPa). 

mechbnical properties for each uncoated specimen are recorded in Table 

The strength ranges from 12.8 to 14.1 ksi (88 to 97 

The 

27 
H of Appendix B. 

substantial load decrease. The fractures were initiated on the 

tension side and then propagated to the compression side. Two 

specimens from panels 3 and 6 fractured into two separate pieces. 

remaining specimens failed almost as catastrophically. 

The room temperature flexure specimens failed with a 

The 

Six low modulus coated c/c specimens yielded an average strength 

of 12.8 ksi (88 MPa) and a range of values from 12.1 to 13.5 ksi (83 

to 93 MPa) at room temperature. A typical load deflection curve for a 

coated c/c at room temperature is plotted in Figure 4. A proportional 

limit is not observed so the modulus is calculated from the deflection 

at one half the maximum load. The data for these flexural tests on 

the coated c/c are recorded in Table I of Appendix B. The average 

modulus for the coated specimen at room temperature is 3.0 Msi (20 

GPa) with the values ranging from 2.7 to 3.4 Msi (19 to 23 GPa). 

The flexural test characteristics on coated c/c are discussed 

lvith regard to the load deflection curve, deflection to failure, and 

type of failure. All of the specimens failed on the tension side of 

the maximum stress zone between the load points. 

single crack across the width of the specimen that did not propagate 

to the compression side except in one specimen. 

load carrying capability. One specimen from panel G failed 

catastrophically into 2 pieces while the other one from the saine panel 

did not. The amount of deflection to maximum load experienced by 

these c/c flexure tcsts has an average value of 0.18 inches (0.46 cm). 

lhese deflection values were not correlated with any of the mechanical 

properties. 

The failure was a 

Most retained some 

28 
The typical load-deflection curves for the coated and uncoated 

c/c specimens is shown in Figures 3 and 4. The failure load decrease 

is greater for the uncoated material than the coated. The 

load-deflection curve is rounded near the maximum load for the coated 

specimens and indicates that additional failure mechanisms are 

operating. This rounding off may indicate debonding. 

The shear strength of the c/c material is apparently so high that 

shear failures could not be obtained on the uncoated or coated 

material. 

shear stress failures occurred in surface crushing rather than in 

shear. Based on these observations it is believed that the shear 

strengths in the coated and uncoated conditions are greater than 2600 

arid 2400 psi (18 and 17 MPa) respectively. The data for the shear 

tests are listed in Tables J and K of Appendix B for uncoated and 

coated c/c respectively. 

When the specimens were tested in short support span, the 

Elevated Temperature Fl exural Testing 

Results of the flexural test that was conducted at room 

temperature and lOOO", 1500", 2000", and 2400°F (538", 816", 1093", 

1316°C) are summarized in Table 11. 

tests are listed in Tables L, M, N, and 0 in Appendix B. 

that the flexural strength first decreases slightly then increases 

with increasing temperature. The proportional 1 imit is observed at 

an elevated temperature and increases with increasing temperature. 

The proportional limit is considered as the point where the load 

versus deflection curve deviates from linearity. The modulus of 

The results on the individual 

It is noted 

29 
elasticity is taken as the initial slope of this curve and is 

significantly higher at elevated temperatures compared to the room 

temperature Val ue. 

TABLE I1 

Summary of Flexural Test Data 

Test 

Temperature 

"F "C 

RT RT 

1000 538 

1500 816 

2000 1093 

2400 1316 

F1 exural 

Strength 

ksi MPa 

12.6 87 

12.4 85 

12.4 85 

13.3 52 

14.6 101 

Proportional 

L imi t 

ksi MPa 

4.0 28 

4.6 32 

6.3 43 

7.3 50 

F1 exural 

Modulus 

Msi GPa 

3.0 21 

6.0 41 

7.1 48 

6.9 48 

6.3 44 

The influence of temperature on the mechanical strength of a low 

modulus, coated c/c composite has been plotted in Figure 5 with the 

data listed in Table 11. The modulus of elasticity (as measured from 

the slope of the initial linear portion of the load-deflection curve) 

is plotted in Figure 6 with the data in Table 11. 

limits are plotted in Figure 7 with the data 'in Table 11. 

The proportional 

The average flexure strength decreasesd 3% from its room 

temperature value, to the value at lOOO"F, the average modulus of 

elasticity increased by loo%, and the proportional limit was first 

observed at this temperature. At 1500°F as compared to 1000°F, the 

average fl exural strength remained the same, the el astic modul us 

increased by 17X, and the proportional limit increased by 14%. 

the bend tests at 2000°F as compared to 1500"F, the average strength 

For 

30 
increased 7%, the average modulus decreased by 1%, and the 

proportional limit increased by 37%. Fcr the flexure at 2400°F as 

compared to the 2OOO"F, the average strength increased by lo%, the 

elastic modulus decreased by 7%, and the average proportional limit 

increased by 16%. 

Three types of load versus deflection curves were observed during 

the flexure tests. The most commonly occurring curve is illustrated 

in Figure 8 and shows significant load decrease after the maximum 

load. The second type of curve, which is demonstrated in Figure 9, 

shows more rounding at peak loads than the other curves. The third 

type of curve shows only a small decrease in load when it falls off 

from the peak load, as indicated in Figure 10. 

occurred in each test is indicated in data tables in Appendix B. 

The type of curve that 

All of the load-deflection curves at room temperature, 2000", and 

2400°F exhibited the rapid decrease in load from the maximum load 

illustrated in Figure 8. One of the curves at 1000°F exhibited a 

rounded portion at the maximum load as demonstrated in Figure 9. 

was the only specimen to fail in this manner. 

attributed to the structure of the c/c composite and the occurrence of 

small internal cracking. The small decrease in load from the maximum 

load, shown by the load-deflection curve demonstrated in Figure 10, 

only occurred at 1500°F and is believed to be associated with the 

behavior of the coating. 

This 

Its behavior is 

When failure occurred on these specimens it followed two modes. 

In one case, the s'pecimen broke completely through its cross section, 

and in the other case considerable internal cracking occurred but the 

31 
specimen remained intact. The initial fracture appears to have 

multiple cracks in the coating that are perpendicular to the major 

stress and strain axis. The failure cracks always occur on the 

tension side and eventually coalesce into a failure. 

The composite separation of the flexure specimen into two pieces 

occurred with some specimens at all temperatures except at 1500" and 

2400°F. At 1500" and 2400°F all of the fractures were of the partial 

type with some load carrying characteristics remaining. lables 0 

through R in Appendix B indicate the fracture modes of the individual 

specimens. 

specimen. 

failure followed about a 30" angle from the vertical in transversing 

across the specimen. 

All failures started on the tension side of the bend 

In the case of the complete separation fracture, the 

This indicated a stepwise type of fracture. 

Stress Rupture Testing 

The results of the stress rupture tests that were conducted at 

the elevated temperatures of lOOO", 1500", ZOOO", and 2400", (538", 

816", 1093", and 1316°C) indicate the existence of a stress rupture 

limit in the silicon carbide coated, low modulus c/c composites. The 

stress rupture 1 imit is defined as the appl ied stress load to the c/c 

that will not cause failure of the specimens within 100 hours. 

stress rupture limit increases with increasing temperature and is 

The 

listed in Table 111. The times 

stress rupture 1 imit are 1 isted 

increased as the applied stress 

individual test is tabulated by 

of Appendix B. 

to failure at stress levels above the 

in Table IV. Times to failure 

was decreased. The test data for each 

temperature in Tables P, Q, R, and S 

32 
I TABLE 111 

Stress Rupture Summary 

Stress Rupture Limit 

Stress Rupture 

"F "C ksi MPa 

Temperature Limit 

1000 538 3.0 21 

1500 816 6.3 44 

2000 1093 7.9 54 

2400 1316 9.2 62 

\ TABLE IV 

I 

Stress Rupture Sumnary 

Times to Failure 

Times to Failure 

( hours) 

4\45] (34) (22) 

1000 538 7.3 12 39 

1500 816 1. L 26 - . 

2000 1093 2.1 4.0 - - L 

2400 1316 1.5 - - - - 

- indicates no failure or not test conducted. 

A specimen in the stress rupture test failed after the 

application of a static flexure stress. This stress cracked the 

coating and created a path for the oxidation of the substrate. The 

glass overlay flowed into the cracks, if the glass was fluid, and 

33 
blocked the oxidation path. 

the coating that were too large for the glass layer for fill. 

Subsequently, the substrate was oxidized and the composite failed. 

This minimum stress level is just above the stress rupture limit. 

A minimum stress level induced cracks in 

Four types of specimen fracture occur during the stress rupture 

tests. 

at all temperatures and is characterized by one dominant crack in the 

coating in the high stress area. 

second type of fracture, a split fracture, and is characterized by one 

major crack on one specimen edge and two separate cracks on the other 

edge. The coating between these two cracks frequently breaks off 

during handling. The third fracture type is shown in Figure 13 and is 

a ply shear fracture with multiple cracks in the coating. This 

fracture pattern has one major crack on the mold side of the coating 

and one crack through the thickness of the specimen. The mold side 

crack propagates through the thickness of the coating and then runs 

parallel to the mold side in a mode similar to shearing. 

illustrated in Figure 13. Figure 14 is an example of the fourth 

fracture type and is a wide fracture with multiple cracks in the 

coating. No major thickness crack is observed in this failure. 

Figure 11 is a photo of a single crack fracture which occurred 

Figure 12 is an example of the 

This is 

Significant scatter occurred in the stress rupture test data. 

The times to failure at a constant stress level and temperature varyed 

greatly and were not correlated with the number of cracks in the 

coating, initial strain rate, or fracture type. Some specimens failed 

in stress rupture while others do not fail at the same or greater 

stress level. The scatter of the failures and nonfailures is 

attributed to variations in the specimens. 

34 
The influence of temperature on the stress rupture limit of the 

5ilicon carbide coated, low modulus c/c has been plotted in Figure 15 

from the data listed in Table 111. 

is 100% greater than the value at 1000°F. 

a 26% increase of the stress rupture limit compared to the value at 

150@"F. The stress rupture liniit at 2400°F is 18% greater than the 

value at 20OO0F. 

The stress rupture limit at 1500°F 

The tests at 2000°F exhibit 

The influence of applied stress on the time to failure is plotted 

for the 1000°F stress rupture test in Figure 16 with data listed in 

Table P of Appendix C. This plot is similar in shape to the other 

temperature curves except for higher stress rupture limits and the 

occurance of failures beyond the 26 hours. Only one test at 1500°F 

failed in excess of 26 hours. This specimen failed in 58 hours and 

the strain was constant from the 8th hour to the 48th hour. The 

strain then increased and failure occurred within 10 hours after 

strain initiation. 

resulted in failure was 3.7 hours; at 2400"F, 2.2 hours for stress 

levels just above the stress rupture limit. 

The longest length of a test at 2000°F that 

The influence of temperature on the time to failure at a constant 

stress has been plotted in Figure 17 with the data listed in Table IV 

at the stresses of 8.0 and 9.5 ksi (5.5 and 6.5 MPa). The tinie to 

failure with a 9.5 ksi load decreased by 50% with an increase in 

temperature to 1500°F from 1000°F. The failure time at 9.5 ksi is 30% 

greater at 2000°F compared to the 1500°F test. The time to failure is 

30% less at 2400°F than the time at 2000°F. The increase in the time 

to failure at the 2000°F is attributed to a statistical variation. 

35 
One specimen at 2000°F and 9.5 ksi had a high failure tine and 

influenced the average. The time to failure increase between 1500°F 

and 2000°F was not observed at the 8.0 ksi stress level and indicates 

that the increase at 9.5 ksi was not characteristic. 

All of the fractures in the stress rupture test at 2400°F were a 

single crack fracture mode as seen in Figure 11. 

is observed zt a11 temperatures and is the most frequently encountered 

at 2000°F. The split fracture mode is seen in Figure 12 and occured 

in the stress rupture tests at 1000, 1500 and 2000°F. Most of the 

1500°F stress rupture test specimens fractured in this mode. The ply 

shear and the wide fracture modes were encountered only in the 1000°F 

stress rupture test and are seen in Figures 13 and 14. These modes 

are characterized by many cracks in the coating. 

This fracture mode 

The oxidation of the composite in the stress rupture test is 

examined on random specimens. One failed stress rupture specimen from 

each test teniperature was prepared for an optical examination along 

with a room temperature flexure specimen. These specimens were vacuum 

impregnated, cut longitudinally, and polished. 

are presented in Figure 18 in ascending order by temperature, 1.e. the 

room temperature specimen at the bottoni and the 2400°F (1316°C) 

specimen at the top. The upper sample is at the bottom of this 

figure. 

The 1000°F sample is different than the other higher temperature 

samples due to a depression area in the central portion of the 

composite. 

which is the controlling oxidation mechanism. 

These cross sections 

The four oxidized samples have void areas under the coating. 

This area is believed to be evidence of internal oxidation 

The oxidation of the 

36 
higher temperature specimens appears to occur by surface oxidation in 

a pore under a crack in the coating. 

Discussion 

Characterization 

Some correlation of the mechanical properties to x-ray 

radiographic inspection are indicated. The differences between 

specimens from panels 8 and 11 and other panels are previously 

described. The specimen strength from these two panels are the lowest 

two values, 13.1 and 12.8 ksi (90 and 88 MPa), respectively. The 

moduli of these two panels are also the lowest values, 2.2 and 2.0 Msi 

(15 and 14 GPa). Removing these values from the calculations of the 

averages increases the average strength to 13.9 ksi (95 MPa) and the 

modulus to 2.4 Msi (16 GPa). The lower mechanical strength and 

modulus is observed at 1000" and 2400°F for the specimens from Panel 

8. The strength and modulus of Panel 11 is among the highest at 

2400°F. 

properties is inconclusive. The radiographic differences had no 

effect on the stress rupture test data. 

The effect of radiographic texture on the mechanical 

F1 exural Testing 

Results show a decrease of 6% in flexural strength for the coated 

compared to the uncoated specimens. 

thickness of the coating compared to the specimen have been made to 

determine the cause of this different strength. 

approximately 12% of the specimen thickness and 10% of the width. 

Calculations based on the 

The coating is 

37 
These calculations indicate that if the coatings were not bearing some 

of the load, the strength would decrease by 30%. 

the coating is carrying some of the stresses in bending. 

the modulus of elasticity of the coating is somewhat higher than the 

value for the c/c material. This results in nioving the neutral axis 

towards any side bearing a higher proportion of the stresses. Because 

of the presence of cracks in the coating that occur more readily on 

the tension side, the neutral axis moves toward the compression side. 

The cracks in the coating on the tension side lower the samples load 

carrying capabil ity and produce higher stress in the composite 

material directly under the coating. Observations of this failure 

type show that the cracks occur in this region. 

This indicates that 

In addition, 

The behavior of the flexural strengths of the coated c/c 

composites with temperature is similar to results obtained on silicon 

carbide coated RCC-3 oiaterial by Vought [36]. The strength of a 19 

ply RCC-3 is 14% greater than the value in this test sequence. The 

strength of the RCC-3 is reported to decrease first with temperature 

and then increase to 2% greater than the room temperature value at 

2000°F. The strength increased further at higher temperatures. The 

proportional 1 iniit was not reported. 

The strength decrease of a coated c/c from room temperature to 

1000" and 1500°F (538" and 816°C) is insignificant as determined by a 

two sample t test at a 0.05 confidence level. 

that the strength increases at 2000" and 2400" (1093" and 1316°C) were 

insignificant. 

above room temperatures [43]. 

This test also showed 

The strength of the fiber is greater at temperatures 

38 
The influence of the coating on the flexural strength and modulus 

is expected; however, the actual modulus increase is unexpectantly 

large. The elastic behavior of the uncoated, low modulus c/c with 

temperature is not reported in the literature. Information [2J on 

elastic properties of uncoated, high modulus material s indicates that 

the modulus is constant with temperature in an inert atmosphere. 

Also, the modulus of the fiber does not increase with temperature 

[421. 

the 1 oad path at higher temperatures. 

These two facts indicate that the coating must be dominating 

The modulus increase phenomena is also observed on the RCC-3 

material, but the magnitude of the increase in the RCC-3 is less than 

the c/c examined in this test sequence. 

RCC-3 material in the flexural test increased with temperature. The 

modulus of the coated RCC-3 material increased 45% at 1000°F and 70% 

at 200OOF. The difference between the behavior of the two moduli is 

attributed to coating thickness differences. The coating thickness on 

the RCC-3 material is 0.006 inches (0.015 cm.) and is less than half 

the thickness of the coating in this test sequence. 

The modulus of the coated 

The coating carried tensile and compressive stresses in the 

elevated temperature flexural test. In the room temperature fl exural 

test the coating carried only compressive stress because of the 

presence of cracks as previously described. The coating carried 

tensile stress at elevated temperature because the cracks in the 

coating were closed by thermal expansion niismatch forces. The tensile 

stress in the coating is indicated by the proportional limit, the 

slope of the load-deflection curve, and the increase in the elastic 

39 
modulus. The proportional limit is attributed to the failure of the 

coating in tension and subsequently the shedding of all of the tensile 

stresses. The load is then carried by the substrate and the coating 

in compression. This load distribution is similar to the load 

distribution in the room temperature flexural test. The slope of the 

load-deflection curve above the proportional limit indicated this lozd 

distribution because this slope W~S similar to the room temperature 

slope. The modulus at elevated temperatures was more than double the 

room temperature value. The modulus of c/c is constant with 

temperature. Therefore, the additional load carrying capacity can 

only be attributed to the tensile loads in the coating. 

The coating carried tensile loads because the cracks in the 

coating were closed by thermal expansion mismatch forces. The thermal 

expansion of the coating is greater than the thermal expansion of the 

composite. 

closing temperature. 

thermal expansion forces at this temperature. In a perfectly elastic 

system this temperature is the coating deposition temperature. 

specimens in this sequence do not behave perfectly el astical ly because 

the crack closing temperature is less than 1000°F (538°C) while the 

coating is deposited in the temperature range of 2200" to 2400°F 

(1200" to 1300°C). The lower crack closing temperature is attributed 

to the overlay of glass sealers applied at room temperature. 

slurry decreased the crack width and aided the heal ing of the cracks. 

Additional plastic deformation was believed to occur because of the 

warp character of the specimen as evidenced by the radiograph in 

This system yields a teniperature known as the crack 

All the cracks in the coating are closed by 

The 

This 

40 
Figure 2. This deformation was part of the inelastic behavior that 

can decrease the crack closing temperature. 

deformation was not indicated by microscope investigation. 

Evidence of this 

The modulus of the coating at elevated temperature was influenced 

by the thermal behavior and bonding within the coating. The modulus 

increased significantly from room to elevated temperatures. This 

increase is attributed to the coating carrying tensile loads. The 

weak links in the coating are the cracks in the coating. 

these cracks is required for the transfer load in the coating. The 

healing occurs by thermal mechanical closing of the cracks and 

bonding. At 1000°F (538°C) only compressive stresses and thermal 

energy assist the healing of the cracks. Above 1400°F (760°C) a 

liquid phase bonding mechanism improves the healing of the cracks and 

result in a higher moduli at 1500°F (816°C) and above. 

the modulus at 2400°F (1316°C) is attributed to the decrease in the 

modulus or' the coating at that temperature. 

Healing of 

The decline of 

The proportional limit of the c/c specimens increased with 

temperature because the amount of compressive stress in the coating 

increased with temperature. Compressive stress was induced in the 

coating with thermal expansion forces by heating a specimen to 

above the crack closing temperature. 

in the coating is proportional to the temperature above the crack 

closing temperature. The initial loading in the flexural test 

decreased this compressive load in the coating before a tensile stress 

was induced in the coating. 

temperature and required more load to overcome this stress. 

The amount of compressive stress 

This compressive stress increased with 

41 
Subsequently, the proportional limit increased with temperature. The 

proportional limit is believed to be influenced by other factors, but 

not as strongly. 

Stress Rupture Testing 

The increase of the stress rupture limit with temperature 

increases is attributed to two factors: the fluidity of the glass 

overlay and compressive stress in the coating. The compressive stress 

in t.he coating increased with temperature increases as previously 

described for the proportional limit. The higher compressive stress 

increased the minimum load to cause a sufficiently large crack in the 

coating and allow the oxidation of the substrate. 

higher loads were employed at higher temperatures to obtain a stress 

rupture failure and stress rupture limit. A specimen failed in a 

stress rupture test because the glass layer flowed too slowly into the 

stress induced crack; consequently, the substrate oxidized. The 

fluidity of the glass increases with temperature. Therefore, the flow 

of the glass, to block the oxidation, improved with increasing 

temperature. 

rupture 1 imit with increasing temperatures. 

Subsequently, 

This improved oxidation protection increased the stress 

The results of this testing sequence indicate that this coating 

system is successful s.t preventing oxidation of the low modulus c/c. 

Compressive stress is induced in the coating upon heating and is 

desirable because the cracks .in the coating are closed. 

compressive stress arises from one or both of two mechanisms: 

overlay of glass sealers at rooin temperature reduces the crack width, 

or plastic deformation within the composite accommodates the thermal 

The 

the 

42 
expansion mismatch strains. If the latter is the dominating mechanism 

then the coating retains the compressive stress on thermal cycling and 

the coating is a successful oxidation barrier. If the first mechanism 

is dominating, then with thermal cycling cracks are formed in the 

coating and the coating is an unsuccessful oxidation barrier. The 

thermal cycling testing is important because this environment is 

experienced in propulsion systems. 

stress rupture and thermal cycling tests to examine both of these 

affects on a coating ability to prevent oxidation. 

Future examinations will combine 

43 
CONCLUSIONS 

The following conclusions are based on the results obtained in 

this work on a silicon carbide coated, low modulus carbon/carbon (c/c) 

composite. 

The stress rupture data indicates the existence of a stress 

rupture limit at each temperature; below this limit, failure does 

not occur. The stress rupture 1 imit increases 100% between 1000" 

and 1500°F. This increase is attributed to the activation of 

glass sealers above 1400°F to block oxidation. 

rupture limit increase is linear from 1500" to 2400°F for a total 

of 48%. 

in the 100 hour time limit while others failed at equal or lower 

calculated stresses. The scatter of the failures is attributed to 

variations in the material. The scatter of times to failure 

is greater at lower temperature arid stresses; this behavior is 

attributed to the number of oxidation sites and cracks in the 

coating . 

The fl exural strength at elevated temperature compared to room 

temperature initially decreases and at 2400°F increases to 15% 

greater than the room temperature strength . A proportional 

limit is not observed at room temperatures but occurs at elevated 

temperatures. This proportional 1 imit increases with increasing 

temperature. The elastic modulus of the composite at room 

temperature is 30% higher for the coated than the uncoated 

specimens. 

100% at 1000°F compared to the room temperature value. 

The stress 

Some tests above the stress rupture liniit did not fail 

The modulus of elasticity increases by greater than 

Further 

44 
smaller increases are observed with increasing temperature until 

it decreases at the highest temperature, 24OO0F. 

3) The radiographic examination of the coated c/c revealed two 

different textures. The mechanical properties are not 

significantly different between the specimens of the two types of 

textures. The chemical and physical characterization of the 

composite did not yield significant differences among the coating 

batches and between the bag and mold side of the composite. 

45 
1. 

2. 

3. 

4. 

5. 

6. 

7. 

8. 

9. 

10. 

REFERENCES 

D. L. Schmidt, "Carbon/Carbon Composites", SAMPE Journal, 8 [3] , 

pp. 9-19, 1972. 

T. E. Schmid, "Oxidation Resistant Carbon/Carbon Composites for 

Turbine Engine AFT Sections", AFWAL-TR-82-4159, pp. 74-77, 105, 

168, 1982. 

A. J. K1 ein , "Carbon/Carbon Composites" , Advanced Material s and 

Processes , 130 [51 , pp. 64-68, 1986. 

E. Fitzer, "The Future of Carbon-Carbon Composites" , Carbon 25 

[Z], pp. 163-190, 1987. 

H. M. Stoller, B. 1. Butler, J. D. Theis, and M. L. Lieberman, 

"Carbon Fiber Reinforced Carbon Matrix Composites" , in 

Composites: 

1971, J. W. Weeton, and E. Scala, eds., pp. 69-136, 1974. 

J. W. Warren and C. D. Coulbert, "An Introduction to Carbon 

Composite Material SI', Super-Temp Company; Santa Fe Springs , CA. 

Report, pp. 3-14, 1971. 

J. B. Donnet and R. C. Bansal, Carbon Fibers, Marcel Dekker, Inc. 

New York, pp. 7-10, 1982. 

Ibid, Ref. 7, p. 226. 

M. K. Piggot, Load-Bearing Fibre Composites, Pergamon Press, New 

York, p. 59, 1980. 

H. W. Chang and S. K. Rhee, "Oxidation of Carbon Derived from 

Phenolic Resin", Carbon, 16 [l], pp. 17-20, 1978. 

State of the Art, Proceedings AIME Fall Meeting, 

46 
11. F. M. Lang and P. Magnier, "Action of Oxygen and Carbon Dioxide 

Above 100 Millibars on "Pure8' Carbon", in Chemistry and Physics 

of Carbon, P. L. Walker ed., Marcel Dekker, New York, pp. 

121-209 , 1968. 

12. W. S. Horton, "Oxidation of Pyrolytic Graphite" in Proceedings of 

the Fifth Conference on Carbon, Vol . 2, MacMillan Company, New 

York, pp. 223-241, 1963. 

13. T. J. Clark, R. E. Woodley, and D. R. DeHalas, "Gas-Graphite 

Systems'', in Nuclear Graphite, R. E. Nightengale, ed., Academic 

Press, New York, pp. 387-441, 1962. 

14. A. J. Kennedy, "Initial Assessment of Graphite as a Structural 

Material in Conditions of High Thermal Flux", NATO, 1960. 

15. E. Yasuda, S. Kimura, and Y. Shibusa, "Oxidation Behavior of Carbon 

Fiber/Gl assy Carbon Composite" , Transactions of the Japanese 

Society of Composite Materials, 6 [l], pp. 14-23, 1980. 

K. H. Han, H. Ono, K. S. Goto, and G. R. St. Pierre, "Rate of 

Oxidation Carbon Fiber/Carbon Matrix Composites with an Anti 

16. 

Oxidation at high Temperature", Journal of the El ectrocheniical 

Society: Sol id-State Science and Technology, 134 L41, pp. 

1003-1009, 1987. 

17. H. W. Chang and R. M. Rusnak, "Oxidation Behavior of 

Carbon/Carbon Composi tes'l , Carbon , 17 [SI, pp. 407-410, 1979. 

C. L. Mantel, Carbon and Graphite Handbook, Interscience 

Publishers, New York, p. 350, 1968. 

G. R. Hennig, "Surface Oxides on Graphite Single Crystal", 

Proceedings of the Fifth Conference on Carbon, Vol . 1, MacMillan 

Company, New York, pp. 143-146, 1962. 

18. 

19. 

47 
20. 

21. 

22. 

23 

24. 

25. 

26. 

27. 

28. 

29. 

D. W. McKee, "Borate Treatment of Carbon Fibers and Carbon/Carbon 

Composites for Improved Oxidation Resistance", Carbon, 24 [SI, 

pp. 737-741, 1986. 

P. Ehrburger, P. Baranne, and J. LaHaye, "Inhibition of Oxidation 

of c/c Composites by Boron Oxide", Carbon, 24 [4], pp. 495-99, 

1986. 

D. W. McKee, C. L. Spiro, and E. J. Lamby, "The Inhibition of 

Graphite Oxidation by Phosphorus Additives", Carbon, 22 [3], pp. 

285-290, 1984. 

J. B. Donnet and A. Voet, Carbon Black, Marcel Dekker, New York, 

p. 147, 1976. 

C. W. Stroud and D. R. Rummler, "Mass Loss of a TEOS-Coated 

Reinforced Carbon/Carbon Composite Subjected to a Simulated 

Shuttle Entry Environment", NASA TM 81799, 1980. 

C. W. Stroud and D. R. Rummler, "Mass Loss of TEOS-Coated RCC 

Subjected to the Environment at the Shuttle Wing Leading Edge", 

NASA TM 83203, 1981. 

R. Krefold, G. Linkenheil, P. Glaud and W. Karcher, Carbon '72, 

Baden-Baden, p. 216, 1972. 

D. R. Marx and P. A. Thrower, "An Investigation of Stress Effects 

on Graphite Oxidation", Carbon, 21 C41, pp. 451-459, 1983. 

3. K. Zhao, R. C. Bradt, and P. L. Walker, "Effect of Air 

Oxidation at 873°K on The Mechanical Properties of a 

Carbon/Carbon Composite" , Carbon, 23 [l], pp . 9-13, 1985. 

D. L. Chang and L. Rubin, "Mechanical and Thermomechanical 

Responses of Si1 icon Carbide Coated Composites" , SD-TR-83-33, 

1983, p. 33. 

48 
30. 

31. 

32. 

33. 

34. 

35. 

36. 

37. 

38. 

R. Keiser, "Oxidation Protection for High Strength CarbonKarbon 

Composites" , AFWAL-TR-82-4060, 1982, p . 149. 

C. D. Pears, H. S. Starrett, G. W. Driggers and D. L. VanWagner, 

"Test Methods for High Temperature Material Characterization", 

AFML-TR-79-4002, p. 335 , 1979. 

3. Z. Gyekenyesi, and J. H. Hemann, "High Temperature Tensile 

Testing of Ceramic Composites", NASA CR in pub1 ication, 1987. 

J. S. Evangelides, R. A. Meyer and 3. E. Zimner, "Investigation 

of the Properties of CarbonKarbon Composites and Their 

Re1 ationship to Non-destructive Test Measurements", 

AFML-TR-70-213, 1971. 

Y. M. Tarnopolskii and T. Kincis, "Methods of Static Testing for 

Composites" in - Failure Mechanics of Composites, eds. G. C. Sih 

and A. M. Skudra. Elsevier Science Publishers B. V. Amsterdam, 

Nether1 ands , 1985, pp. 266-267. 

G. D. Quinn and J. J. Stiglich, "Mechanical Testing of Coated 

Carbon/Carbon Composite at Elevated Temperature in Air", MTL Th 

86-15, 1986. 

T. W. Smith, et a1 . , "Leading Edge Structural Subsystem 

Mechanical Design Allowable", Yo1 . 11, Vought Corporation report 

221RP00439, 1976. 

K. Tossi, Personal Communication. 

G. D. Quinn, "Review of Static Fatigue in Silicon Nitride and 

Si1 icon Carbide", Ceramic Engineering and Science Proceedings, - 3, 

1-2, pp. 77-98, 1982. 

49 
39. J. Fullerton, T. E. Schmid, and R. E. Warburton, "Carbon/Carhon 

Composites for Advanced Turbine Exhaust Norzl est' , AFWAL 

TR-80-2104, 1980, p. 54. 

40. "Bulk Density of Porous Refractories by Wax Imnersion", C914, 

Annual Book of ASTM Standards, Section 15, Volume 15.01. 

J. L. Smialek and N. S. Jacobson, "Mechanism of Strength 

Degradation For Hot Corrosion of a-Sic," Journal of the American 

Ceramic Society, 69 [lo], pp. 741-752, 1986. 

41. 

42. Personal Communication Greg A. Underwood, Hercules Aerospace 

Company, Magna, Utah. 

43. R. M. Gill, Carbon Fibers in Composite Materials, The Plastic 

Institute, Butterworth and Co., Ltd., 1972, p. 72. 

50 
ORIGINAL PAGE IS 

OF POOR QUALITY 

. 

Figure 1. Typical radiograph of coated c/c perpendicular to the plies 

for the specimens 2-43 to 7-50. 

51 
Figure 2. Typical radiograph 

the specimens 2-37 

of coated c/c para1 le1 to the pl ies for 

to 2-50, 

52 
8 

Figure 3. Typical load-deflection curve of an uncoated carbon/carbon 

composite in four point flexure at room temperature. 

Figure 4. Typical load-deflection curve of a coated carbon/carbon in 

four point flexure at room temperature. 

53 
18 

v) :F 12 

0 

0 

D a 

Figure 5. Plot of the flexure strength versus temperature. 

0 

0 0 

a 

Figure 6. Plot of elastic modulus versus temperature. 

54 
Figure 7. Plot of the proportional limit versus temperature. 

Figure 8. Type I load-deflection curve of a coated carbon/carbon at 

240G"F (1316°C) in four point flexure. 

55 
Figure 9. Type I1 load-deflection curve of a coated carbon/carbop at 

100C"F (538°C) in four point flexure. 

EFLECTIQ 

Figure 10. Type 111 load-deflection curve of a carbon/carbon at 

1500°F (816°C) in four point flexure. 

56 
Figure 11. Typical single crack fracture mode in stress rupture, 

specimen 2-34 at 1000°F (538°C) and 6500 psi (45 MPa). 

Figure 12. Typical split crack fracture mode -in stress rupture 

specimen 0-13 at 1500°F (816°C) and 6500 psi (45 MPa). 

57 
Figure 13. Typical ply shearing fracture rode in stress rupture 

specimen 2-02 at 100G"F (538'C) and 8000 psi (55 MPa). 

Figure 14. Typical wide fracture mode in stress rupture specimen 2-07 

at 1000°F (538°C) and 3600 psi (25 MPa). 

58 
I a- / 

Figure 15. Plot of the stress rupture limit versus temperature. 

0 10 w) W eo 100 120 

1 Iff 10 fAI1URt. ta 

Figure 16. Plot of time to failure versus applied stress at 10cIO"F. 

59 
0 

0 

KS I 

0 6.0 

0 9.5 

0 

0 

0 

Figure 17. Plot of time to failure versus temperature at stress 

levels of 8.0 and 9.5 ksi. 

60 
- ,- 

2400'F 

200OoF 

1 5OO0F 

1000°F 

ROOM 

TEMPERATURE 

Figure 18. Longitudinal crossection of stress rupture specimeps and 

room temperature flexure specimen. 

61 
Appendix A 

Procedural Details 

Density Measurement Procedure 

The density measurements of the c/c substrate and silicon carbide 

coating are guided by a wax immersion technique, ASTM C914. 

coupons are cut from room temperature flexural specimens with the 

coupons coming from both sides of the failure zone. The substrate 

coupons originate from the uncoated specimens and the coated coupons 

require additional cutting to obtain the coated specimen. 

mold sides of the coating are cut away from the substrate with a 

diamond saw. 

removed with an oxidation treatment at 1020°F (550°C) for 20 hours. 

The ash residue is removed from the coupon. 

ultrasonically cleaned in ethanol for 20 hours and then dried at 250°F 

(120°C) for 12 hours. 

The 

The bag and 

Some c/c is retained on the "coating coupon" and is 

The coating coupons are 

The weight of the coupon (CW) is recorded before it is coated 

with wax. 

the coupon is coated by dipping it into the wax. The weight of the 

wax coated specimen is recorded in air (WCW) and imrriersed in ethynol 

(WCI). 

and the density is calculated with the following equation: 

Parafin wax is liquified with a double boiler technique and 

The temperature of ethanol is recorded to a tenth of a degree 

cw Density = 

mWW0 P-AGE BLANK NOT FLMm 

-- wcw - WCI - wcw - cw 

LD WD 

Eq. A1 

63 
CW = Coupon Weight 

WCW = Wax Coated Weight 

WCI = Wax Coated Immersed Neight 

LD = Ethanol 

WD = Wax Density 

The density of the wax is 0.905 g/cc. The density of the 

ethanol is determined by interpolation of the temperature, density 

chart in the CRC handbook*. The data is recorded in Appendix By 

Tables B and C. 

A density measurement of the pack cementation silicon carbide 

coating requires the removal of the glass overlay from the coupon. 

attempt to measure this density was made. 

of the coating was placed in a 20% hydroflouric acid solution to 

dissolve the glass overlay. The coating coupon was completely 

zttacked and only a fine powder remained in the acid after a twelve 

hour soak. 

An 

A glass coupon consisting 

F1 exural Test Procedures 

The details on the determination of the room and elevated 

temperature flexural properties are presented. Both room and elevated 

temperature tests were similarly conducted and are described together. 

The facil ities 

and data reduction are summarized for the flexure test at room and 

specimen preparation, cal ibration, testing procedures 

* CRC - Handbook of Chemistry and Physics, R. C. Weast, ed., 60th 

edition, 1979-80, pg. F-6. 

64 
elevated temperatures of IOOO", 1500", 2000°, and 2400"F, (538", 816", 

1093", and 1316°C). 

The flexural test system is seen in Figure 1A and consisted of a 

load frame, load cell, water cooled coup1 ings, si1 icon carbide suppcjrt 

ana push rods, as silicon carbide resistance heated furnace, and the 

test fixture. The room temperature tests were conducted in this load 

train configuration. A 1000 pound capacity load cell was employed to 

measure the load. The output of the load cell was recorded on a strip 

chart. This chart a1 so recorded displacement of the crosshead. 

lhe room and elevated temperature test fixtures have the same 

testing geometry, but were fabricated from different materials. 

load fixtures were also different. 

fabricated from stainless steel as seen in Figure A2 and thc elevated 

temperature fixture from sintered alpha silicon carbide as Figure A3 

illustrates. The room temperature loading fixture is loaded at a ball 

bearing point. The elevated temperature fixture is beveled to insure 

The 

The room temperature fixture was 

width of the room 

de fixtures used for 

the point loading of 

temperature fixture 

the high temperature 

the specimen. The overall 

s half of the silicon carb 

tests. 

The testing geometry of this test 1:s four point bending. The 

support rollers are 1.00 inch in diameter (2.54 cm) and the long 

support span for flexural failures is 6.12 inches (15.5 cm). 

support span in the room temperature test is decreased to 2.00 inches 

(5.08 cm) with the addition of spacer blocks for the shear tests. 

load rollers are 0.44 inches (1.11 cni) in diameter and are in a load 

span Gf 1.00 inch (2.54 cm). 

The 

The 

65 
The test specimens are coated and uncoated low modulus c/c 

composites with the approximate dimension of 7.75 long by 0.60 wide by 

0.24 inches thick (20 x 1.5 x 0.60 cm). 

dimensions of each specimen are measured in the high stress zone with 

vernier calipers to within k.001 inches (.003 cm). 

eight uncoated specimens are tested at room tenrperature and five 

specimens at each of the elevated temperatures. 

specimen from each coating batch are tested at each temperature. 

The width and thickness 

Six coated and 

At least one coated 

Calibration 

Cal ibration was required for the load cell, chart speed, cross 

head deflection, and the compliance of the load train. 

capacity load cell was cal ibrated with weights having cal ibration 

traceable to the National Bureau of Sta.ndards. The procedure begins 

after the system warms up for a half an hour. 

disassembled and the load cell is placed on the crosshead. 

was electronically zeroed and balanced on the twenty pound scale. 

The 1000 pound 

The load train was 

The system 

The 

fifty pound sca 

set on the load 

chiirt recorder. 

scale. The cal 

e was selected and fifty pounds of static weights were 

cell. This full scale reading was calibrated on the 

The reading was verified on the one hundred pound 

bration was complete and the load train was 

reassembled. 

beginning of every test day. 

This cal ibration procedure was conducted at the 

The calibration of the chart recorder was a simple procedure. 

The chart rate was selected, 1.0 inch per minute (2.54 cmimin.). The 

chart recorder and timer are started simultaneously and stopped after 

66 
an interval of two minutes. 

actual distance travelled. This procedure was conducted once. 

The expected result was compared with the 

The calibration of the crosshead deflection was conducted with a 

Starret dial indicator for displacement measurements mounted on a 

magnetic base. The indicator is graduated at 0.0005 inch (.0013 cm) 

ivcrements. The magnetic base was mounted to the test frame and the 

spindle of the indicator contacted the crosshead. The dial indicator 

was zeroed. The crosshead deflection rate was set at 0.05 inches per 

minute (0.13 cm/min.), and deflected for two minutes. The expected 

deflection was compared to the measured deflection. This calibration 

method was conducted once. 

The complimce of the load train was measured with the dial 

indicator that was employed in the previous calibration, and the load 

which was measured on the strip chart recorder. The silicon carbide 

test fixture without the load fixture was placed on the support rod. 

The upper push rod was lowered until it contacted the fixture. 

chart recorder was balanced without a load applied. 

load were applied to the load train by manually lowering the crosshead 

as indicated on the chart recorder. The dial indicator was zeroed. 

The crosshead was manually lowered 0.0005 inches and the load and 

deflection were recorded. 

The 

Four pounds of 

Subsequently, the load train was lowered in 0.0005 inch 

increnients and the loads were recorded. The results from this 

procedure are as follows: 

67 
Deflection 

Inches 

0.0000 

O.OOO5 

0.0010 

0.0015 

0.0020 

0.0025 

0.0030 

0.0035 

0.0040 

0.0045 

0.0050 

0.0055 

Load 

1 bs. 

4.0 

5.3 

- 

7.4 

9.2 

11.6 

14.9 

18.9 

24.8 

34.1 

45.1 

56.8 

69.5 

This compliance measurement was conducted once at room 

temperature and was applied in all modulus calculations. 

compliance values at elevated temperatures were assumed to be similar. 

The 

Test Procedures 

The test procedure for the conduction of the four point bend 

flexural tests of coated and uncoated c/c coinposi tes invol ved 

a1 ignnient, heating the furndce, preloading, and calculating the 

f 1 exural properties. 

(0.05 inches per minute) (0.13 cm/rnin), chart speed (1.0 inches per 

niinute) (2.5 cm/min) and lGad cell scale (100 pounds) were set at the 

beginning of the test. 

The testing variables of the crosshead speed 

68 
IThe alignment of the specimen on the support rollers was obtained 

optically. The mold side of the specimen was in tension. 

specimen was centered lengthwise and across the width of the roliers. 

In the elevated test, the support rollers were bonded to the fixture 

with an adhesive. The specimens were similarly bonded to the rollers. 

This maintained the alignment in the transport of the fixture. At 

elevated temperatures, the adhesive decomposed and the bond was 

broken. 

load fixture between the support rollers. 

to the specimen with the adhesive for the elevated temperature test. 

The test fixture was placed in the load train. 

positioned in the maximum stress area. 

to within 0.25 inches (0.6 cm) of the load fixture. The furnace was 

heated to temperature and was soaked at the test temperature for one 

hour. The pen on the chart recorder was balanced and a preload of 

four pounds was imposed on the fixture by manually lowering the 

crosshead. 

sign i f icantl y . 

The 

The alignment fixture in Figure A3 was employed to center the 

The load fixture was bonded 

Thermocouples were 

The upper push rod was lowered 

The test was stopped when the load decreased 

The flexure properties of proportional 1 imit., strength, and 

modulus were calculated after the testing is complete. The values 

from the chart recorder were used in the calculations. The equations 

employed in the stress and modulus calculations are based on beam 

theory. 

constants are homogeneous. 

homogeneous and invalidate the equations. 

These equations assume that the cross section and elastic 

Coated and uncoated c/c composites arc not 

In lieu of a more rigorous 

69 
calculation, the values obtained were used as a first approximation 

and for comparison within this test sequence. 

The test geometries and nomenclature are summarized in Figure 4A. 

The loads on the strip chart which correspond to the maximum and 

proportional limit were used in the strength calculations. The slope 

of the load-deflection curve was measured from the pre-load to the 

proportional 1 imit and used to calculate the modulus. If no 

I 

I 

proportional 1 imit was apparent, then the slope from the pre-load to 

one-ha1 f of the maximum 1 oad was recorded for the niodul us calcul ation. 

I 

The stress was calculated with the following equation: 

3PKL 

2 

0 = -. Eq. A2 

where u was the strength or proportional limit; P the load; w the 

width; and t the thickness. The KL value was 2.56 inches (5.5 cm) for 

I the flexure test. 

The shear strength ,T, was calcul ated with the foll owing 

equation : 

3P 

4wt 

=-- 

The modulus, E, was calculated with the following equation: 

E=- - [KL3/3 f (KL*)LS/;! + (KL)LS2/81 

wt3 6 

Eq. A3 

*Eq. A4 

70 
where LS is the load span and P/6 is the load-deflection slope the 

load span is 1.0 inch (2.54 cni) and the equation reduces to: 

55.3 P 

ZT E= Eq. A5 

The deflection value, 6, was corrected for the compliance of the 

load train that was measured. 

Stress Rupture Testing Procedure 

The flexural stress rupture properties of coated low modulus 

carbon/carbon (c/c) composites were examined at the elevated 

temperature of lOOO", 1500", 2000", and 2400°F (538", 816", 1093", and 

1316°C). The facil ities, cal ibration, specimen preparation, and 

procedures of this test are sumiarized in detail. 

consisted of two compression creep test frames, silicon carbide 

resistance heated furnaces, silicon carbide test fixtures, 

extensometry and a data acquisition system (DAS). The compression 

creep test system were fabricated by Applied Test System and are shown 

in Figure A5. The test fixture in the load train is pictured in 

Figure A6 and this fixture was used for the flexure test at elevated 

temperatures. 

The test facilities 

* Pears, C.D., H. S. Starret, G. W. Driggers, and D. L. Van Wagoner, 

"Test Methods for High Temperature Material Characterization , AFML 

TR 79-4002, 1975. p. 54. 

71 
Ext en some t ry 

The extensometry system of the compression creep test system was 

designed to measure a relative displacement which is described in 

Appendix C. 

The probe tips contact the specimen and the relative displacement is 

transferred through the a1 umina rods to the extensometer housing. 

outer two rods sit on the inner housing and center rod sits on the 

spring supported core rod. A linear variable differential transformer 

(LVDT) deflects with the outer rods and measured the displacement. 

Only the relative displacement of the inner rod to the outer rods 

affects the output. 

Figure A7 illustrates the schematic of the extensometry. 

The 

Data Acquisition System 

The data acquisition system (DAS) consisted of a IBM PC XT with 

an interface which was fabricated by Advanced Peripheral s Incorporated 

(API). The API equipment consists of a type "R" thermocouple, voltage 

input, and a relay modules. Each module contains eight channels for 

data input or control. These modules are interfaced with the IBM PC 

XT by an integrated circuit board. 

The DAS software was programmed it] Basic, and it controlled test 

parameters and assisted in the calibration of the extensometer. The 

software controlled the temperature on the extensometer and the flow 

of air into the test system. The extensometer housing was maintained 

within 0.5"C of 36°C by a therinocouple and relay channels on the DAS. 

A heat tape is wrapped around extensometer housing and is connected in 

series with a relay channel on the DAS. A type 'IT" (copper 

72 
. 

constantan) thermocouple is placed inside the housing and connected to 

a channel on the thermocouple module. The software read the 

temperature of the housing every minute and then sent a signal to the 

relay module to open or close the relay which sent the power to the 

heat tape. 

A similar thermocouple and relay pair were employed to maintain 

the temperature of the control box to above 34°C. 

thermocouple is placed in the control box and two 75 watt light bulbs 

are connected in parallel and then in series with a relay of the DAS. 

The pair of light bulbs is placed at the base of the control box. The 

temperature of the control box is checked every minute by the DAS. 

A type "T" 

The DAS controlled the flow of air into the furnace with relay 

connections to solenoid valves. The flow of compressed air into the 

furnace was initiated at the start of the test. The air flowed into 

the furnace at a rate of 10 ml per minute. At failure, the DAS 

stopped the flow of compressed air and began a purge of nitrogen into 

the furnace chamber at a rate of 1.0 liter per minute. 

flow cools the furnace and minimizes the amount of oxidation of the 

c/c substrate after failure. The purge continued until the 

temperature was less than 800°F (425°C). 

The nitrogen 

The function of the DAS is the acquisition of data from the 

stress rupture test. 

13% rhodium) thermocouples and the output of the LVDT. 

thermocouple module converted the mill ivol t reading to temperature by 

means of charted values. 

from the output of the LVDT as explained in the equations of Appendix 

The DAS reads two type "R" (platinum, platinum - 

The 

The strain on the specimens was calculated 

73 
C. 

the hard disk. The frequency of data storage depended on the test 

segment. On heat up, soak and cool down, the data was recorded every 

ten minutes. At the end of the soak period, the LVDT output was 

zeroed and load applied with procedures described in a later 

paragraph. 

minute for the first ten minutes, then every two minutes for the next 

fifty minutes. At the end of this first hour an algorithm controlled 

the recording frequency, and is based on the percent increase of 

strain over the previous recorded value. 

the specimen is straining, as rapidly as once per minute. 

The time, temperature, and strain were recorded to a data file on 

Immediately after loading, the data was recorded every 

The frequency is higher when 

Extensometry Cal ibration 

A DAS software routine allowed the calibration of the 

extensometer. 

extensometer and the cal ibration micrometer, while the LVDT output 

voltage is displayed only the screen. The calibration fixture is 

shown in Figure A8 and the Starret micrometer is graduated in 0.0001 

increments. The DAS and electronics of the extensometry required a 

twenty minute warm up period before operations. The upper push rod 

was removed from the load train and the calibration fixture was 

inserted . 

The procedures involved mechanical manipulations of the 

Extensometer adjustments and fixture preparation were required at 

the beginning of cal ibration. The adjustment screws and micrometer on 

extensometer housing were lowered to a niinimum position. The 

alignment clip was centered and bonded to the fixture. The test 

74 
fixture is then set into the load train with the probe tips of the 

extensometry extending through the cl ip. A downward pressure was 

exerted on the center tip to check for a springy response. 

response indicated proper a1 ignment in the extensometer. 

inch thick spacer b8rs were placed on each side of the test fixture. 

The calibration fixture is set on the spacer bars. The center probe 

tip contacted the center of the calibration micrometer. 

micrometer was positioned where the spindle face was at the same level 

as the cal ibration fixture (0.9125 inches). 

This 

One quarter 

The 

The next routine determined the level of the adjustment screws 

for all the stress rupture testing. At this level, the probes will 

not droop during a test and yield erroneous data. 

the extensometer housing is raised 0.25 inches (0.6 cm), and the 

adjustnients screws were evenly increased until the outer probes 

contact the calibration fixture. A downward force was exerted on the 

micrometer in the extensometry housing and then released. The 

position of the outer probe was examined. If the probes did not 

contact the calibration fixture, then the adjustment screws were 

raised and the response to the downward pressure on the micrometer was 

rechecked. This iteration continued until a consistent response of 

probe tips contacting the calibration fixture was obtained. 

level was marked on the adjustment screws at the entrance to the 

housing with an indel ible marker. 

The micrometer in 

This 

The mechanical calibration of the LVDT began with the selection 

of "Lisplay LVDT Voltage" routine from the DAS Main Menu. 

electronic zero was obtained before the mechanical cal ibration. 1 he 

An 

75 
toggle switch on the LVDT control panel was set to calibrate and the 

zero knob is on the control panel was rotated until a zero output was 

displayed. The toggle switch was returned to the operate position. 

The micrometer in the extensometer housing was rotated until the 

output voltage reads zero volts. The LVDT was then calibrated at 

one-ha1 f full scale deflection. The cal ibration micrometer was 

deflected one-ha1 f full scale (0.015 inches) ( .038 cm) and the output 

voltage was 5.00 volts for perfect calibration. This voltage was 

obtained by rotating the range knob on the LVDT control panel. The 

cal ibration micrometer was returned to the zero position by exceeding 

the zero value and then approaching it with a downward direction of 

the micrometer. A zero voltage was not displayed at the zero level 

and was obtained by rotating the micrometer in the housing until zero 

volts is displayed. The calibration started again with a half scale 

deflection and an iteration of deflections and range knob adjustments 

were required to cal ibrate. 

Stress Rupture Testing 

The stress rupture test procedures involved the alignment and 

loading of the specimen prior to the operations of the test. The 

a1 ignment of the specimen, 1 oad train and extensometry are required 

before heating the furnace. At the beginning of the aligning 

procedures, the data acquisition system (DAS) and extensometry power 

was turned on and allowed to warm up. 

The alignment of the specimen in the test fixture was 

accomplished with a ruler and an epoxy adhesive. This bonding was 

76 
conducted to maintain alignment during handling. 

and the bond was destroyed at elevated temperatures. 

rollers were bonded with epoxy to the fixture. The mold side of the 

specimen was placed downward on the tension side. The length-wise 

alignment of the specimen on the support rollers was obtained 

optically by approximating an equal overhanging distance of the 

specimen over the support roller. Alignment of the specimen across 

the width involved maneuvering the specimen until the distance between 

the specimen edge and the edge of support roller was equivalent on 

either side of the specimen. This distance was measured with a ruler. 

The specimen was then bonded to the rollers. 

The epoxy decomposed 

The support 

Adjustments to the extensometry were necessary to minimize the 

possibility of fracturing the probes while inserting the fixture into 

the lcad train. The inner housing of the extensometry was lowered to 

a minimum position by decreasing the adjustment screws and micrometer 

head at the base of the extensometry housing. Misalignment of the 

center tip in the extensometry commonly occurred and the proper 

alignment was needed before inserting the test fixture. A downward 

force was applied to the center tip and the resistance to the force 

was either springy or rigid. A rigid response indicated 

non-alignment. Alignment was obtained by raising the micrometer head 

and maneuvering the center rod until the response was springy. 

Mechanical adjustments to the test frame and extensometry were 

required to position the test fixturc in the load train. The upper 

push rod is raised and the test fixture was placed in the load train 

with the extensometry probes extending through the fixture. The 

77 
alignment of the center tip was rechecked with the previously 

described procedure. The inner housing is raised with the adjustment 

screws until the center probe contacted the specimen. The "Display 

LVDT Voltage" routine was selected from the DAS menu. The output 

voltage was adjusted to a value between negative four and five volts 

by increasing the micrometer head. The DAS was then returned to the 

DAS main menu. 

The load fixture was aligned on the specimen after the 

thermocouples were placed in the maximum stress area. 

fixture was placed on the specimen and an alignment device, Figure A9, 

centered the load fixture between the support rollers. 

fixture was aligned across the width by the grooves machined in the 

alignment pins. The upper push rod was lowered to one quarter inch 

above the load fixture. 

selecting the "Initiate Test" routine from the DAS menu. The test 

variables were input as the DAS prompts and heat up begins. 

The load 

The load 

The furnace heat up was initiated by 

Once at test temperature, some adjustments to the test system 

were performed 

load fixture. 

DAS menu. The 

quarter of an 

The upper push rod was lowered until it contacted the 

The "Display LVDT Voltage" routine was selected on the 

adjustment screws on extensonietry were raised one 

nch and the micrometer head is raised until the output 

voltage was between negative three and four volts. 

returned to the main menu. 

beforc! loading the specimen. 

The DAS was then 

Thc furnace equilibrated for one hour 

Preparation of the test system for the application of stress to 

the specimen began with calculating the weight required to load a 

78 
ORIGINAI; PAGE rS 

OF POOR OTT4LITY 

specimen io a desired stress level. 

the load train was added to the weight platen. The "Display LVDT 

Voltage'' routine was chosen from the DAS menu. The extensometry 

adjustment screws were raised tcr the level marked on the screws during; 

the calibration procedures. The micrometer head was adjusted until 

the output was within one-tenth of a volt from zero. 

This weight less the weight of 

The "Zeroing" routine was selected from the "Display Voltage" 

routine. 

voltage. 

prior to cclculating the strain. 

specimen by lowering the bolt that supports the weight platen at ten 

seconds before the next reading. 

controlled by the DAS as previously described. 

The DAS read and recorded the initial voltage as the zero 

Subsequent LVDT readings were reduced by the zero voltage 

The load was then applied to the 

The remainder of the test was 

Figure fil. Flexure test load frame. 

79 
Figure A2. Room temperature flexure text fixture. 

Figure A3. Elevated temperature flexure test fixture with centering 

fixture. 

80 
LOAD 

I 

I I--I4 t t 

r 

0 '-+ -1 LS SPAN I-- I 

I- SUPPORT SPANFigure A4. Flexure geometry ncmenclature. 

Figure E. Stress rupture test frame. 

81 
Figure A6. Test fixture in load train. 

LOAD FIXTURE -, , LOAD ROLLLRS 

r SPtCI WN 

ROLLtRS 

FIXTURE -. 

‘PIN ALIGNMENT CLIP 

EXTENSWTRV RODS /d.jjL 

r 

/ 

INNER LVDT HOUSING 

r 

LVDT , LVDT CORE > 

SPRING ADJUSTER SCREW 

LVDT HOUSING 

CORE ROD 

SPRINGS 

‘-RICRCWTER HEAD 

Figure A7. Extensometry for stress rupture test. 

ORIGINAL PAGE 1s 

QUALITY: 

82 
Figure A8. Cal ibraticn fixture. 

Figure A9. Centering fixture. 

83 
Appendix B 

TABLES 

TABLE A 

Designation of Specimen Number, Panel Number, Coating Batch Number 

Carbon /Carbon 

Coating Batch 0 Panel Number Specimen Numbers 

Coating Batch 1 

Coating Batch 2 8 

9 

9 

10 

11 

0-01 to 0-18 

0-19 to 0-38 

0-39 to 0-49 

0-50 

1-02 to 1-07 

1-09 to 1-14 

1-16 to 1-18 

1-19 to 1-21 

1-23 to 1-28 

1-30 to 1-35 

1-37 to 1-43 

2-01 to 2-08 

2-50 

2-09 to 2-26 

2-27 to 2-45 

2-46 to 2-49 

TABLE B 

Density of Low Modul us Carbon/Carbon Substrate 

Density Values g/cc A B C D Average Specimen 

19 1.382 1.377 1.374 1.374 1.377 

89 1.379 1.379 1.380 1.382 1.380 

1.381 1.378 1.377 1.376 1.378 159 

Overall Average 1.378 g/cc 

84 
TABLE C 

Density of Silicon Carbide Coating 

Batch Specimen Density 

Number Side glcc g/cc 

0 Bag 2.61 2.61 

0 Mold 2.61 - 

1 Bag 2.57 2.70 

1 Mold 2.58 2.60 

2 Bag 2.63 2.71 

2 Mol d 2.61 2.75 

Overall Average 

Average 

Density 

2.61 

2.61 

2.64 

2.59 

2.67 

2.69 

2.63 g/cc 

85 
TABLE D 

Summary of the Coating Thickness Measurements 

Number Average 

of Coating Standard Maximum 

Data Side Thickness Deviation Value 

Points Type (ins.) (ins.) 

Batch '0 Warp Yarns Parallel to Coating 

12 Edge 0.024 0.004 0.031 

0.014 0.004 0.022 

22 Mold 0.013 0.002 0.017 

24 Bag 

Batch '0 Fill Yarns Parallel to Coating 

12 Edge 0.032 0.005 0.042 

0.011 0.001 0.012 

0.003 0.018 

12 Ba 9 

12 Mol d 0.012 

Batch '1 Warp Yarns Parallel to Coating 

8 Edge 0,023 0.004 0.029 

24 Ba 9 

24 Mold 0.014 

0.017 0.003 0.024 

0.004 0.020 

Batch '1 Fill Yarns Parallel to Coating 

10 Edge 0.029 0.004 0.037 

12 0.016 0.003 0.023 

12 Mol Bag d 0.016 0.003 0.020 

Batch '2 Warp Yarns Parallel to Coating 

12 Edge 0.023 0.003 0 028 

24 Bag 0.012 0.002 0.019 

24 Mold 0.012 0.002 0.019 

Batch '2 Fill Yarns Parallel to Coating 

9 Edge 0.028 0.003 0.032 

0.011 0.002 0.014 

0.003 0.016 

12 Bag 12 Mol d 0.012 

(ins.) 

0.017 

0.008 

0.009 

0.023 

0.010 

0.008 

0.017 

0.011 

0.005 

0.021 

0.008 

0.013 

0.017 

0.009 

0.008 

0.003 

0.007 

0.009 

G1 ass 

Layer 

(ins. 1 

Minimum Average 

Value Thickness 

0.002 

0.001 

0,002 

0.002 

0.001 

0.001 

0 006 

0.002 

0.002 

0.008 

0.004 

0.002 

0.001 

0.001 

0.002 

0.002 

0.002 

0.001 

86 
TABLE E 

Chemical Analysis of Sic Coating on RCC-3 Substrate 

Batch Side Aluminum Boron Calcium - Iron Sodium Silicon 

0 Mol d 0.2 0.09 1.3 0.02 0.4 32.8 

-- 

0 Bag 0.2 0.09 0.04 0.02 0.3 33.0 

1 Mol d 0.1 0.1 0.04 0.02 0.3 35.3 

1 Bag 0.1 0.1 0.03 0.02 0.4 33.6 

2 Mold 0.1 0.07 0.03 0.01 0.2 34.2 

2 Bag 0.1 0.1 0.02 0.01 0.3 35.4 

TABLE F 

Chemical Analysis of C-C Substrate Ash After Coating 

Batch -- - Side Aluminum --- Boron Calcium Iron Sodium Silicon 

0 Mol d 0.9 0.7 0.2 0.2 2.0 71.7 

0 Bag 0.4 0.7 0.1 1.7 1.4 38.4 

1 Mold 0.6 0.4 0.2 0.2 1.0 30.5 

1 Bag 0.5 0.3 0.1 0.1 0.6 17.5 

2 Mold 0.2 0.4 0.1 0.1 1.1 25.8 

2 Ba 9 0.3 0.9 0.1 0.2 1.8 34.3 

87 
Peaks 

1 

2 

3 

4 

5 

6 

7 

8 

Major 

Phase 

Minor 

Phase 

TABLE G 

X-Kay Diffraction Analysis of Sic Coating 

B a tc h I' 0 I' Bat c h 'I 1 I' Batch "2" 

Mold 

2.53 

1.54 

1.32 

2.64 

2.37 

2.59 

2.67 

1.42 

$Sic 

aSiC 

- Bag 

2.52 

1.54 

1.32 

2.63 

2.36 

2.58 

1.25 

1.42 

BSiC 

aSiC 

Mold 

2.51 

1.54 

1.31 

2.34 

2.62 

2.57 

2.17 

1.42 

- 

$Sic 

aSiC 

Bag 

2.54 

1.55 

1.32 

2.66 

2.38 

2.60 

1.43 

2.20 

$Sic 

aSiC 

Mol d 

2.54 

1.54 

1.32 

2.38 

2.65 

2.60 

1.42 

2.20 

- 

BSiC 

aSiC 

Bag 

2.54 

1.55 

1.32 

2.65 

2.38 

1.43 

2.61 

2.69 

BSiC 

aSiC 

TABLE H 

Room Temperature F1 exural Strengths of Uncoated Specimens 

Specimen Panel 

Number Number 

19 

46 

69 

89 

135 

159 

184 

i56 

2 

3 

5 

6 

8 

9 

10 

11 

Average 

Standard Deviation 

Fl exural 

Strength 

ksi MPa 

13.7 94 

13.6 94 

13.7 94 

13.6 94 

13.1 50 

14.1 97 

14.0 97 

12.8 88 

13.6 94 

0.4 3 

- - 

88 

F1 exural 

Modulus 

Msi GPa 

2.3 16 

2.7 19 

2.4 17 

2.3 16 

2.2 15 

2.4 17 

2.2 15 

2.0 14 

2.3 15 

0.2 1 

- - 

Failure 

Type 

Partial 

Complete 

Partial 

Compl ete 

Part i a1 

Partial 

Part i a1 

Partial 
TABLE I 

Room Temperature F1 exural strengths of Coated Specimens 

Specimen Panel F1 exural F1 exural Failure 

Number Number Strength Modulus Type 

- ksi - MPa - - Msi GPa 

0-17 2 12.7 88 3.1 21 Partial 

0-18 2 13.0 90 2.7 19 Partial 

1-17 6 12.9 89 3.0 21 Partial 

1-18 6 12.3 85 2.9 20 Complete 

2-26 9 12.1 83 3.0 21 Part i a1 

2-27 9 13.5 93 3.4 23 Partial 

Average 12.8 88 3.0 21 

Standard Deviation 0.5 3 0.2 1 

TABLE J 

Room Temperature Shear Strengths of Uncoated Specimens 

Shear F1 exural 

Specimen Panel Stress Stress 

Number Number at Failure at Failure 

ksi - MPa - ksi - MPa 

45 3 2.6 18 12.9 89 

112 7 2.4 17 12.0 83 

183 10 2.5 17 12.8 88 

Average 2.5 17 12.6 87 

89 
TABLE K 

Room Temperature Shear Strengths of Coated Specimens 

Shear F1 exural 

Specimen Panel Stress Stress 

Number Number at Failure at Failure 

- ksi - MPa - - ksi MPa 

0-37 3 2.9 20 14.6 101 

1-35 7 2.4 17 12.4 86 

2-45 10 2.7 18 13.9 96 

Average 2.7 18 13.6 94 

TABLE L 

1000°F F1 exural Strengths 

Propor- Flexure tional F1 exure Defl . Spec - Strength Limit Modulus Deflection Failure Curve 

imen Panel ksi MPa ksi MPa Msi GPa in. cm. Type Type 

0-04 2 12.7 88 4.1 28 5.5 38 0.14 0.35 Partial I 

0-32 3 12.2 84 3.9 27 6.4 44 0.13 0.33 Partial I 

1-12 6 12.4 86 4.7 32 5.4 37 0.15 0.37 Complete I 

2-04 8 12.2 84 3.8 26 5.8 40 0.13 0.34 Complete I 

2-32 10 12.3 85 3.6 25 6.9 48 0.16 0.40 Partial II 

-- 

A vera ge 12.4 85 4.0 28 6.0 41 0.14 0.36 

Standard 0.2 1 0.4 3 0.6 4 0.01 0.03 

Deviation 

90 
r 

F1 e 

TABLE M 

1500°F F1 exural Strengths 

D 

Propor- ire tional F1 exure fl. 

Spec - Strength Limit Modulus Deflection Failure Curve 

-- imen Panel ksi MPa ksi MPa Msi GPa in. cm. Type Type 

0-12 2 12.0 83 4.8 33 7.0 49 0.13 0.34 Partial I11 

0-42 5 12.8 88 4.4 30 6.7 46 0.14 0.35 Partial I11 

1-28 7 12.6 87 4.5 31 7.1 49 0.12 0.32 Partial I11 

2-40 10 12.6 87 4.6 32 6.5 45 0.13 0.34 Partial I11 

2-50 9 12.1 83 4.7 32 7.7 53 0.12 0.30 Partial I11 

Average 12.4 86 4.6 32 7.0 48 0.13 0.33 

Standard 0.3 2 0.2 1 0.4 3 0.01 0.02 

Deviation 

TABLE N 

2000°F F1 exural Strengths 

Propor- F1 exure t ional F1 exure Defl . Spec - Strength Limit Modulus Deflection Failure Curve 

imen Panel ksi MPa ksi MPa Msi GPa in. cm. Type Type 

0-19 3 12.8 88 6.2 42 7.3 50 0.11 0.29 Partial I 

0-49 5 13.8 95 5.9 41 6.9 48 0.14 0.35 Partial I 

1-19 7 13.8 95 6.7 46 6.9 48 0.13 0.33 Partial I 

1-31 7 12.8 88 6.4 44 6.5 45 0.13 0.32 Partial I 

2.17 9 13.5 93 6.3 43 6.6 46 0.15 0.38 Partial I 

-- 

Average 13.3 92 6.3 44 6.8 47 0.13 0.33 

Standard 0.5 3 0.3 2 0.3 2 0.01 0.03 

Deviation 

91 
TABLE 0 

2400°F F1 exural Strengths 

Propor- F1 exure tional F1 exure Defl . Spec - Strength Limit Modul us Deflection Fail ure Curve 

-- imen Panel ksi MPa ksi MPa Msi GPa in. cm. Type Type 

0-27 3 14.2 98 7.0 48 6.2 43 0.14 0.35 Partial I 

1-03 6 14.7 101 7.6 53 7.2 50 0.14 0.35 Partial I 

1-41 8 14.2 98 7.1 49 6.0 41 0.14 0.36 Partial I 

2-25 9 14.6 101 7.3 50 6.4 44 0.16 0.40 Partial I 

2-49 11 15.2 105 7.5 52 6.2 43 0.16 0.40 Partial I 

Average 14.6 101 7.3 51 6.4 44 0.15 0.37 

Devi at ion 

Standard 0.4 3 0.3 2 0.4 3 0.01 0.02 

TABLE P 

1000°F Stress Rupture Sumnary 

Number 

0-31 

2-30 

2-33 

2-31 

1-06 

0-36 

1-13 

0-05 

0-30 

2-07 

1-11 

1-07 

2-01 

1-05 

2-05 

0-02 

2-34 

1-21 

0-35 

2-02 

1-09 

0-33 

2-28 

Panel 

3 

10 

10 

10 

6 

3 

6 

2 

3 

8 

6 

6 

8 

6 

8 

2 

10 

7 

3 

6 

6 

3 

10 

Stress 

MPa ksi 

21 3.0 

21 3.0 

21 3.1 

21 3.1 

22 3.2 

22 3.2 

23 3.3 

23 3.4 

23 3.4 

25 3.6 

28 4.1 

34 5.0 

34 5.0 

45 6.5 

45 6.5 

45 6.5 

45 6.5 

55 8.0 

55 8.0 

55 8.0 

65 9.5 

65 9.5 

65 9.5 

Time of 

Test 

hours 

130 

130 

114 

75 

460 

41 

30 

63 

20 

23 

18 

13 

10.9 

10.6 

7.1 

6.4 

5.1 

5 

4.7 

4.3 

3.2 

3.1 

2.8 

Failure 

Type 

* * * 

Singl e 

Spl it 

Singl e 

Singl e 

Spl it 

Wide 

Spl it 

Shear 

Shear 

Shear 

Shear 

Shear 

Singl e 

Shear 

Shear 

Shear 

Singl e 

Shear 

Wide 

* 

Number 

Optical 

Cracks 

0 

0 

2 

0 

2 

3 

1 

3 

5 

4 

11 

9 

7 

8 

7 

4 

5 

12 

6 

5 

8 

9 

- 

Initial 

Strain 

Rate 

-- 

-- 

-- 

-- 

-- 

0.15 

1.02 

0.18 

0.10 

0.14 

0.40 

0.60 

0.62 

0.70 

0.57 

0.62 

0.56 

0.63 

0.51 

0.81 

0.66 

-- 

* indicates no failure 

92 
J 

TABLE Q 

1500°F Stress Rupture Sumnary 

Time of 

Stress Test 

Number Panel MPa ksi hours 

2-11 

1-27 

1-23 

0-09 

2-12 

1-30 

0-13 

2-41 

0-41 

2-39 

0-11 

0-40 

2-37 

0-10 

1-25 

2-38 

9 32 4.6 

7 43 6.3 

7 44 6.4 

2 44 6.4 

9 44 6.4 

7 45 6.5 

2 45 6.5 

10 45 6.5 

5 45 6.6 

10 47 6.8 

2 48 7.0 

5 55 8.0 

10 55 8.0 

2 65 9.5 

7 65 9.5 

10 65 9.5 

228 

122 

27 

100 

165 

8 

58 

160 

10 

12 

6.9 

3.9 

5 

1.4 

1.6 

2.2 

Number 

Failure Optical 

Type Cracks 

* * 

Singl e 

Spl it 

Spl it 

Singl e 

Spl it 

Spl it 

Spl it 

Singl e 

Spl it 

Singl e 

Singl e 

* * 

* 

* indicates no failure 

TABLE R 

2000°F Stress Rupture Summary 

- 

1 

2 

1 

5 

6 

1 

4 

1 

7 

10 

6 

14 

16 

- 

- 

Time of Number 

Stress Test Failure Optical Number Panel MPa ksi hours Type Cracks 

0-20 

2-20 

0-22 

1-16 

2-16 

1-32 

0-47 

2-42 

0-48 

1-33 

2- 19 

0-21 

3 46 6.7 

9 53 7.7 

3 54 7.9 

6 54 7.9 

9 55 8 

7 55 8 

5 55 8 

10 55 8 

5 56 8.2 

7 65 9.5 

9 65 9.5 

3 65 9.5 

235 

137 

100 

126 

3.2 

3.7 

5 

300 

4 

1.5 

1.6 

3.3 

* * * * 

Singl e 

Singl e 

Singl e 

Singl e 

Singl e 

Spl it 

Spl it 

* 

- 

4 

6 

7 

5 

14 

6 

1 

7 

4 

- 

- 

Initial 

Strain 

Rate 

-- 

0.15 

0.23 

0.16 

0.55 

0.72 

0.39 

0.46 

0.44 

0.23 

0.19 

0.44 

0.45 

0.55 

0.44 

0.78 

Initial 

Strain 

Rate 

0.74 

0.18 

-- 

-- 

0.98 

0.43 

0.74 

0.62 

0.75 

0.41 

-- 

-- 

* indicates no failure 

93 
TABLE S 

2400°F Stress Rupture Summary 

Time of Number Initial 

Stress Test Failure Optical Strain 

Number Panel MPa ksi hours Type Cracks Rate 

2-24 

1-40 

0-38 

2-23 

0-29 

1-02 

0-24 

0-26 

2-47 

1-43 

9 

8 

5 

9 

3 

6 

3 

3 

11 

8 

55 8.0 

63 9.0 

63 9.2 

63 9.3 

64 9.3 

65 9.4 

65 9.4 

65 9.5 

65 9.5 

65 9.5 

270 

117 

100 

152 

1.4 

115 

2.2 

1.7 

1.5 

0.9 

* * * * 

Singl e 

Singl e 

Singl e 

Singl e 

Singl e 

* 

. -- 

7 -- 

9 0.30 

9 -- 

10 0.78 

11 -- 

10 0.60 

9 0.78 

1 0.61 

9 0.90 

* indicates no failure 

94 
I- 

* 

Appendix C 

Strain Measurement 

A four point bend test configuration is employed in the stress 

rupture test and the strain on a specimen is measured with an 

extensometry system. A specimen in this test deforms with a curvature 

and has compressive and tensile strains distributed through the 

thickness. The radius of curvature is geometrically proportional to 

the strains in the specimen. The extensometry of the stress rupture 

test measures the relative displacement of a loaded specimen. The 

displacement is a factor in a calculation for the radius of curvature 

and hence the strain. 

specimen into a arc in the flexure test. The specimen has compressive 

strains at the surface in the concave arc between the load points 

while the opposite surface is convex and has tensile strains. Between 

the two exterior surfaces of the specimen, a surface exists without 

strain and is called the neutral surface. The neutral surface is 

illustrated by a dashed line in Figure 1C. 

Figure 1C displays the deformation of a 

The longitudinal strains, E, on a bend specimen are directly 

proportional to the radius of curvature, R, at the neutral surface and 

the distance, y, from the neutral surface by: 

Eq. 1C 

Equation 1C is derived solely from the geometry of the test and is 

independent of the properties of the specimen*. The maximum 

Timoshenko. S. P., and Gere J. M., Mechanics of Materials, D. Van 

Nostrand Co., Cincinnati, 1972 pp 113-115. 

95 
/ 

strain on the specimen occurs at the exterior surfaces because this is 

the greatest distance from the neutral surface. For all homogeneous 

material, this distance is one half of the thickness of the specimen. 

The maximum strain equation is: 

t 

max = BT E Eq. 2C 

where t is the thickness of the specimen. 

The extensometry of the stress rupture test system is designed to 

determine the radius of curvature of the loaded specimen. The 

extensometry probes contact the specimen on the convex side and senses 

the relative displacement, d, as in Figure IC. 

curvature on the convex surface is first calculated by an equation 

The radius of 

derived from the Pythagorian relationship for 

=L 2+(Rm-dm) 2 

Rm 

where Rm and L are the radius of curvature at 

a right triangle: 

Eq. 3C 

the convex surface and 

the distance between the probes respectively and as illustrated in 

Figure 1C. The value for L is one half inch. Solving Equation 3C for 

the radius with this value for L yields: 

2 

R,,, = 1/4 + dm 

2d m 

Eq. 4C 

The radius of curvature at the neutral surface can be calculated 

by subtracting the distance of the neutral to the convex surface from 

the radius of curvature at the convex surface. The first 

96 
ipproximation of this distance is one half of the thickness of the 

specimen and this yields: 

t R=R,-x L Eq. 5C 

Combining equations 5C, 4C, and 2C yield the strain equation: 

t /2 E.' 

7- 

1/4 + dm2 t/2 

m 

t 

- +dm-t 

4dm 

E= 

Equation 6C is employed in the data acquisition system to 

calculate the strain. 

Eq. 6C 

Figure 1C. Bending and Extensometry Geometry. 

97 
1 Report No 

NASA CR-180863 

Gary A. Rozak and John F. Wallace 

2 Government Accession No 3 Recipient's Catalog No 

( E-3708 1 

4 Title and Subtitle 

Stress Rupture Behavior of Si 1 Icon Carbide Coated, 

Low Modulus CarbonICarbon Composites 

7 Author@) 

5 Report Date 

January 1988 

6 Performing organization Code 

8 Performing Organization Report No 

12. Sponsoring Agency Name and Address 

National Aeronautics and Space Administration 

Lewis Research Center 

Cleveland, Ohio 44135-3191 

9 Performing Organization Name and Address 

Case Western Reserve University 

Cleveland, Ohio 44106 

Dept. of Material Science and Engineering 

15. Supplementary Notes 

505-63-0 1 

11. Contract or Grant No. 

NAG 3- 4 6 4 

13. Type of Report and Period Covered 

Project Manager, Carl Lowell, Materials Division, NASA Lewis Research Center. 

This report was a thesis by Gary A. Rozak submitted in partial fulfillment of 

the requirements for the degree of Master of Science to Case Western Reserve 

University, January 12, 1988. 

___-_ -- 

I7 Key Words (Suggested by Author@)) 

Carbonlcarbon composites; Stress 

rupture; Flexure; Elevated temperatures; Coating; Silicon carbide 

Testing 

18. Distribution Statement 

Unclassified - Unlimited 

Subject Category 24 

--_- -_- 

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