NASA Technical Memorandum 103157

High Temperature Fatigue Behavior of a

SiC/Ti-24Al-11Nb Composite

P.A. Bartolotta and P.K. Brindley

Lewis Research Center

Cleveland, Ohio

Prepared for the

10th Symposium on Composite Materials: Testing and Design

sponsored by the American Society for Testing Materials

San Francisco, California, April 24-25, 1990

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HIGH TEMPERATURE FATIGUE BEHAVIOR OF A SIC/TI-24AL-11NB COMPOSITE

P.A. Bartolotta and P.K. Brindley

Nati,)nal Aeronautical and Space Administration

Cleveland, OH 44135

SUMMARY

A series of tension-tension strain- and load-controlled fatigue tests were conducted on

unidirectional SiC/Ti-24Al-!lNb (at %) composites at 425 and 815 oC. Several regimes of

damage were identified usillg Talreja's (ref. 1) concept of fatigue life diagrams. Issues of

test technique, test control mode, and definition of failure were also addressed.

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INTRODUCTION

The intermetallic composite SiC/Ti-24Al-llNb (at %) has generated considerable

interest for aerospace applications due to its high strength to density ratio. Recent investigations have examined various aspects of this composite such as fiber-matrix chemical interactions (refs. 2 to 7), fiber-matrix bond strength (ref. 8), tensile properties (refs. 2, 3, and

9), and thermal cycling (r_fs. 10 and 11). However, only one study to date has examined the fatigue behavior of thi._ composite (ref. 12). Further study of SiC/Ti-24Al-I1Nb is

necessary for complete understanding of fatigue behavior and damage mechanisms. The

current study examined the fatigue response of unidirectional SiC/Ti-24Al-11Nb isothermally

fatigued at 425 and 815 °C. The influence of test methodology (strain- versus loadcontrol) on failure mechanisms was also a major topic of this investigation. This study is

the first portion of an on-going examination of fatigue properties and damage mechanisms

in SiC/Ti-24Al-llNb.

EXPERIMENTAL PROCEDURE

The SiC/Ti-24Al-llNb composites were fabricated by the powder cloth technique

(ref. 13) followed by vacut_m hot pressing (VHP). The fiber used was the 140 _m

diameter, double-coated, SCS-6 SiC manufactured by Textron, Inc. The matrix material

was Ti-24Al-llNb and was obtained as prealloyed PREP powder. The composites were

produced with three plies of continuous fibers oriented in the 0 ° direction (parallel to the

loading axis during testing} with an approximate volume fraction of 27 percent. Composite

plates typically measured ]50 mm long x 50 mm wide x 0.7 to 0.8 mm thick. Consolidated plates were etched iJl nitric acid to remove the Mo cladding required during VHP.

Typical fiber distribution ia the fully consolidated composite is shown in figure l(a). As

shown in figure l(b), the composite microstructure was quite complex. The matrix was

comprised of equiaxed a 2 (Ti3Al) surrounded by disordered ft. A p-depleted zone was

present in the matrix adjacent to the fiber-matrix reaction zone which surrounded the

C-rich coating of the fiber A more extensive analysis of this as-fabricated composite

microstructure has been previously performed by scanning transmission electron microscopy

(ref.4).

Composite specimens were wire electrodischarge machined (EDM) from the plate as

shown in figures 2(a) and (b). The specimen design incorporated a reduced, parallel-sided

test section to facilitate cl,)se control of the temperature profile and a large fillet radius of

60 mm to minimize failures near the radius. After EDM, specimen surfaces were polished

with 180 grit SiC paper t,) remove a 10 ]_m cladding reaction with the Mo release sheets

used during fabrication and to remove damage due to wire EDM. Tabs, as shown in

figure 2(b), were applied with epoxy. This increased the gripped-section thickness to

approximately 6 mm, nearer to that of the grip opening, to help minimize possible bending

stresses due to gripping. Tabs were tapered to distribute throu.gh-thickness stresses

resulting from high clamping forces. The use of the above specimen and tab designs produced fracture within the gage section in 93 percent of the specimens.
The experiments were conducted under computer control on a closed-loop, servohydraulic system with a 222 kN maximum load capacity. A derated load capacity of 9 kN

full scale was used. Water-cooled, hydraulic grips were aligned independent of the specimen

to a tolerance of <3 percent bending during initial installation. The experimental equipment, data acquisition, and computer control system has been previously described (ref. 14).

A standard 12.7 mm gage length extensometer was mounted on the edge of the sample in

order to avoid possible bending induced by mounting the extensometer on the width of the

specimen. The specimens were heated by a two-zone direct resistance furnace specifically

designed for the plate specimens (ref. 15). This furnace maintained a temperature profile

in the test section of _-1 percent of the nominal value. Beaded type K thermocouples were

tied onto the test section. Specimens were maintained at temperature for 1 hr prior to

testing to allow the equipment to stabilize. Note that 1 hr at the highest temperature of

815 °C causes negligible growth of the fiber-matrix reaction zone (ref. 3).

To address the issue of test control method, fatigue tests were conducted in both

strain-control and load-control at 425 and 815 °C. Strain-controlled tests were conducted

using a triangular control waveform which was limited by the required maximum strain

and zero load. The strain rate was lxl0 -3 sec -1. Load-controlled tests also used a

triangular control waveform in which load was limited by the maximum required load and

zero load. The loading rate and maximum load limits were selected to parallel the strain

rate and maximum loads achieved in the strain-controlled tests. For some of the higher

load limits, a hybrid program was substituted for the load-controlled tests. The hybrid

program was strain-controlled but had load limits at the maximum and minimum points of

the cycle instead of strain limits. This was necessary because some of the large loads

achieved in strain-controlled tests could not be achieved in load-control since the material

response caused an instability in the control loop.

RESULTS AND DISCUSSION

Fatigue Life Relationships

Fatigue life results for the SiC/Ti-24Al-llNb composite at 425 and 815 °C are

presented in figure 3 for both strain- and load-controlled test conditions. A fatigue life

diagram was created after Talreja (ref. 1) by plotting the data as maximum strain versus

number of cycles to failure. Fatigue lives of the strain- and load-controlled tests were

compared on a maximum strain basis since, in the absence of cracking or debonding, the

strain for both the fiber and the matrix are the same during a test but the stresses in the

composite constituents differ (ref. 16). The data displayed three distinct regions, identified

for this work as Regions I to III. Regions I and III did not vary with test temperature

while Region II was clearly temperature dependent. These three regions were similar in

shape to those of the fatigue life diagram for a polymer matrix composite at 23 °C

(ref. 1) as shown in figure 4. It should be noted that failure of the SiC/Ti-24Al-llNb

composite specimens in figure 3 was defined as fracture into two pieces except for the

strain-controlled tests in Region II where failure was defined as the first simultaneous

decrease in stress and stiffness of _>30 MPa and 10 GPa, respectively, as will be discussed

shortly.

For the SiC/Ti-24Al-11Nb composite, Region I of figure 3 was approximately

bounded by the tensile ductility ranges for this composite, which were independent of test

temperature at 425 and 815 °C, at values of 0.69 to 0.85 percent (ref. 9). It was previously suggested that the tensile properties of SiC/Ti-24Al-11Nb were dominated by the

SiC fiber over a temperature range of 23 to 815 °C (ref. 9). This was because the

composite failed at strains slightly less than or equal to the fiber strain-to-failure at all

temperatures tested while the total strain-to-failure of the matrix increased from 1.6 to

20.6 percent over the temperature range. Likewise, it is reasonable to suggest that fiber

strain-to-failure also largely affected the fatigue behavior of this composite at 425 and

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815 *C within Region I, especially since several tests failed upon load-up. Furthermore,

the probabilistic failure nature of this composite in tension was attributed to variations in

fiber strength (and thus, fiber strain-to-failure), both between fiber lots and within a fiber

lot (ref. 9). Fiber strength variability within a given specimen would be expected to

mimic the variability with!n a given fiber lot. Therefore, variations in fiber strengths and

thus, variations in fatigue lives, were also expected within this high strain Region I. In

fact, fatigue lives for the SiC//Ti-24Al-llNb composite ranged from 1/4 cycle to approximately 700 cycles at 815 _C and from 1/4 cycle to approximately 20 000 cycles at 425 *C

in this region. This apparent dominant effect of fiber failure does not exclude other

operative mechanisms.

With one exception, failure within Region I occurred by complete fracture and was

defined as catastrophic or nonprogressive since the material response gave only minor

indications of the impending failure. The typical maximum and minimum stress response

as a function of cycles is shown in figure 5 for specimens at 425 and 815 *C. By test

design, the load-controlled tests showed a constant maximum stress until failure. A typical

specimen tested in strain-control at 425 °C displayed a relatively constant stress response

until the last few thousand cycles before failure in which stress slowly decreased by

40 MPa. At 815 *C, typical strain-controlled specimens displayed stress decreases which

totaled approximately 45 MPa in the first several cycles, as well as additional stress

decreases which totaled about 55 MPa in the last several hundred cycles before failure.

Figure 6 shows maximum and minimum strain as a function of cycles for samples in

Region I at 425 and 815 _C. Again, by design, strain-controlled tests exhibited a constant

maximum strain throughout the tests. The increases in minimum strain for the straincontrolled tests were artifacts of the control program which did not allow the composite to

go into compression and will be discussed shortly. Specimens tested in load-control

exhibited ratchetting up to 0.05 percent strain at 425 *C, which was approximately

10 percent of the total strain range, and up to 0.13 percent strain at 815 *C, which was

approximately 23 percent of the total strain range.

The elastic modulii (E) for both strain- and load-controlled specimens within

Region I are shown in figure 7. E remained constant, within experimental error, at both

425 and 815 *C until the last cycles of life during which it appeared to decrease. This

slight decrease in E wa_ most noticeable in the strain-controlled tests. An E of 168

GPa at 425 *C was obtaiaed in this study as compared with the E obtained in tension

of 180 GPa (ref. 9). E values of 151 to 157 GPa were obtained at 815 *C and were

slightly greater than the previous E values obtained in tension of 138 GPa (ref. 9).

One possibility for the differences in E is variation in fiber volume for samples tested in

tension and those of this ..study; volume fraction of fibers has been determined for each

individual tensile specimen (ref. 9) but is only approximately known for the samples of this

study.

Fatigue behavior w!thin Region II of figure 3 was defined as progressive because the

mechanical response of the strain-controlled tests at both 425 and 815 *C first displayed

gradual decreases in the maximum stress and then abrupt, simultaneous decreases in the

maximum stress (aMAX) _nd unloading modulus with continued cycling, as shown in

figures 8 and 9. Furthermore, the specimens typically remained intact after these abrupt

indications of damage. With these evidences of damage accumulation throughout the life of

the composite, it became necessary to define failure for the strain-controlled tests in Region

II as something other that_ final fracture, as has been suggested by Salkind (ref. 17).

Thus, failure for the strairL-controlled tests in Region II of this work was defined as the

cycle at which the first abrupt sign of damage was observed (e.g., the first abrupt,

simultaneous decrease in stress and modulus). In the lower strain portion of Region II,

this initial degradation in strength and modulus was approximately 30 MPa and 10 GPa,

respectively, and increased as the strain range was increased. However, the load-controlled

specimens did not display these stress drops, as shown in figure 8, but failed catastrophical-
ly by complete fracture of the specimen. It can be seen in figure 3 that the above

definition of failure in strain-control was a reasonable approach since, at a given temperature, failure in strain-control corresponded well with complete fracture in load-control. And

as previously mentioned, Region II of figure 3 displayed temperature dependent fatigue

behavior in which specimens tested at 425 °C had longer lives than those tested at 815 *C

at the same strain amplitude.

Figures 10 to 13 show typical stress-strain responses of the SiC//Ti-24Al-11Nb

composite within Region II of figure 3. Comparison of the responses from the strain- and

load-controlled tests under conditions with similar initial values of strain and stress at a

given temperature are presented for the 425 °C tests in figures 10 and 11 and for the

815 *C tests in figures 12 and 13. Both test methods at 425 *C resulted in approximately the same amount of inelastic strain (0.10 percent) after the first cycle as did both

test methods at 815 °C, which exhibited 0.04 percent inelastic strain after the first cycle.

For both test methods at both temperatures, the hysteresis did not close (i.e., did not

become a loop) upon the first reverse cycle since neither control program allowed the

composite to experience compressive stress. However, in all subsequent cycles, complete

hysteresis loops were evident. For the 425 °C strain-controlled test, the inelastic strain

decreased to zero before the first abrupt stress drop and increased again thereafter. The

425 °C load-controlled test and the 815 *C strain-and load-controlled tests all exhibited

approximately constant inelastic strain (measured from hysteresis widths) from the second

through final cycles at 0.005, 0.02, and 0.04 percent, respectively. Both the 425 and

815 *C strain-controlled tests exhibited modulus degradation simultaneously with the abrupt

stress decreases as previously discussed in figures 8 and 9. In contrast, the load-controlled

tests at 425 and 815 °C exhibited approximately constant modulii throughout the tests

with no stiffness degradation before failure. In this discussion of fatigue response, it is

important to note that the usual meaning of inelastic strain and hysteresis for monolithic

materials may be different than its definition for composites. Inelastic deformation for

composites may be due to one or more of events such as matrix deformation, matrix

cracking, fiber/matrix debonding, or fiber fracture. It is not possible to attribute particular

types of damage to inelastic behavior until metallographic work is completed. However, the

abrupt decreases in strength and modulus in the strain-controlled tests can certainly be

presumed to have involved at least fiber fracture in order to have obtained such large

quantities of degradation. This reasoning relies on the facts that the fibers provide the

majority of the strength and stiffness in this composite (ref. 9) and that the decreases in

strength and stiffness occurred simultaneously as would be expected with fiber fracture.

Two other general observations can be made from the stress-strain responses of

figures 10 to 13 which result from the type of test method used. In the strain-controlled

tests of figures 10 and 12, the maximum stress of the composite decreased slowly in the

cycles before the first abrupt stress drop. These stress decreases could be attributed to one

or more of the following: softening, fatigue damage, or effects of the test control program.

Since the test control program did not allow compressive stresses to be induced, the straincontrolled specimens actually experienced smaller strain amplitudes with each cycle, as can

be seen in figures 10 and 12 and by examining the continuously increasing minimum strain

in figure 9. Thus, it was possible that the test control mode was influential in continuously reducing the stress at the maximum control strain as it induced the continuously

smaller strain amplitudes throughout the test. The possible contributions of material

softening and decreased stress due to control mode could not be determined from these

tests; increasing the strain amplitude after particular cycles would be necessary to characterize this behavior. In the load-controlled tests, strain ratchetting was observed up to

0.08 percent at 425 °C, which was 18 percent of the total strain, and up to 0.12 percent,

which was 25 percent of the total strain, was observed at 815 °C. For both Regions I

and II, strain ratchetting at 425 °C was less than at 815 °C, presumably due to the

difference in creep rate of the matrix at the two temperatures (ref. 18).
Region III of figure 3 incorporates the high cycle fatigue lives of the SiC/Ti-24A1-

11Nb composite with live.,_ >105 cycles. The maximum stresses and strains were so low in

this regime that the specimens were typically run-outs. No indications of damage, such as

stress drops for the strain-controlled sample, were observed. Therefore, tests were discontinued after 105 cycles. The strain- and load-controlled samples at 425 °C and the straincontrolled sample at 815 °C appeared to establish an endurance limit at approximately

0.3 percent strain for this SiC/Ti-24Al-11Nb composite.

As previously men_,ioned, Talreja's (ref. 1) concept of the fatigue life diagram,

figure 4, was developed for polymer matrix composites at 23 °C. Three distinct regions

were evident, each with a different dominant fatigue mechanism. For polymer matrix

composites, Region I repn_sented the region where fiber fracture and interracial debonding

were the dominant mechanisms of damage. This region's width represented the scatter in

the monotonic fracture strain of the composite. The scatter in this band depicted the

fiber's probabilistic failure nature and was associated with low cycle fatigue lives. In

Region II, matrix cracking and interracial shear failure were the dominant mechanisms.

Region III began at the fatigue limit of the matrix and the dominant damage mechanism

for polymer matrix composites in this region was matrix microcrack nucleation. This

region was associated with high cycle fatigue lives and usually resulted in test runouts.

While the fatigue lives of the SiC/Ti-24Al-llNb composite at 425 and 815 °C

showed similar shapes in the three regions as compared to the polymer matrix composites,

much of the damage mechanisms for the isothermally fatigued SiC/Ti-24AI-11Nb composite

remains to be determined. Fatigue tests of the Ti-24AI-11Nb matrix are also necessary to

estimate the contribution of matrix inelastic deformation in the composite in each of the

three regions. Matrix tests are also necessary to determine whether Region III began at

the fatigue limit of the matrix. However, as discussed previously, differences in fatigue

response and fatigue failure imply certain mechanisms within each region. It appeared that

Region I was dominated by catastrophic fiber fracture, particularly at low numbers of

cycles-to-failure, though ol_her mechanisms may also have been operative. In Region II, it

appeared likely that the abrupt stress decreases observed in strain-control were associated,

at least in part, by local fiber fracture(s) since these stress decreases were accompanied by

large modulus decreases. In Region III, fatigue life of the SiC/Ti-24AI-11Nb composite was

similar to polymer matrix composites (ref. 1) in that only test runouts were observed in

this region.

TEST CONTROL MODE

Since the development of the Manson-Coffin (refs. 19 and 20) relationship, straincontrolled tests have been prevalent in the fatigue testing of monolithic, polycrystalline

materials. Mitchell (ref. 21) has presented several fundamental reasons for preferring

strain-control over load-control. Two of the most important were that strain-control

provides a kinematic constraint which does not allow ratchetting to take place and that

straln-control provides a _cenario most often like an actual component; i.e., stress-strain

gradients exist in many practical applications causing material at critical locations to be

constrained to some extent.

Recent studies investigating high temperature fatigue of metal matrix continuous

fiber composites (refs. 22 to 24) used loaxl-control for other practical reasons. Chief among

them was the fact that the materials were available only in the form of thin plates.

Thus, there was the very real possibility of producing significant compressive stresses and

buckling of the composite specimens due to geometric restraints if tested in strain-control.

By conducting these tests in load-control, specimen buckling was alleviated but in exchange,

ratchetting was produced. In the present study, load-controlled testing also produced

ratchetting of up to 19 percent of the total strain range at 815 °C in Region II.
In addition to the present study, one previous study on SiC/Ti-24Al-llNb has used

strain-control without specimen buckling (ref. 12). In the present study, cyclic tests were

conducted with computer control in a manner which was strain controlled and load monitored. In these tests, the specimen was loaded to a maximum strain and then unloaded to

zero stress. This method did not allow the specimen to experience compressive stresses and

by design, did not allow ratchetting to occur. However, as previously discussed and as

shown in figures 10 and 12 for SiC/Ti-24Al-11Nb, the control program imposed successively

smaller strain amplitudes on the composite as it controlled to zero stress on the reverse

cycle. Thus, the tension-tension strain-control mode provided less and less severity as the

test progressed. However, it is believed that this type of test is justified since thin plates

of continuous fiber composite are not suitable for applications involving fully reversed

cycling with compressive loads in the fiber direction.

In addition to the successively smaller strain ranges and the ratchetting discussed

above, the type of test control mode (tension-tension strain-control versus tension-tension

load-control) used in the fatigue of this SiC/Ti-24AI-11Nb composite was seen to produce

other instances of different material response. The most striking differences in behavior

occurred in Region II and will be summarized here along with the similarities.

The primary similarity in the fatigue response of the SiC/Ti-24AI-11Nb composite in

strain- and load-controlled testing occurred on the first cycle, where both types of tests

produced a large quantity of inelastic strain. This inelastic strain was equal in both types

of tests at equivalent strain amplitude and at a given temperature. Also, the inelastic

strain of the first cycle was greater than that produced on any subsequent cycle in both

types of tests. Therefore, metallographic examination of interrupted tests from either test

control mode would be expected to result in the determination of equivalent mechanisms

after the first cycle of either test. Furthermore, if one were primarily concerned with

determining life of the composite, either strain- or load-controlled testing would generate

this information, provided an appropriate definition for failure in strain-controlled tests is

used.

The primary differences between test control modes obtained with SiC/Ti-24Al-11Nb

occurred late in the test, at the first abrupt stress drop and beyond. Complete fracture in

the load-controlled tests occurred catastrophically and coincided with the first abrupt stress

drop observed in the strain-controlled tests, as shown in figure 8. Thus, an advantage of

strain-controlled testing is that it allows easier identification of the failure mechanism by

interruption and examination of a specimen after the stress drop has occurred. The other

difference in test control mode involved the definition of fatigue life in strain-control. Even

after the first abrupt stress drop, defined as failure in this study, the composite continued

to have appreciable load-carrying capacity for many thousands of cycles. If an application

were most similar to the tension-tension strain-controlled tests, the SiC/Ti-24Al-11Nb

composite may be considered to have a longer useful life than that indicated by the loadcontrolled tests.

CONCLUSIONS

1. A fatigue life diagram was beneficial in identifying three distinct regions of

behavior in the SiC/Ti-24Al-11Nb composite. Regions I and III did not vary with test

temperature and Region II was temperature dependent. Furthermore, distinct damage

mechanisms are also expected to be defined in these regions.

2. SiC/Ti-24AI-11Nb fatigue failure within Region I was defined as catastrophic and

appeared to be fiber dominated although other mechanisms may be operative. Fatigue

behavior in Region II was defined as progressive since the strain-controlled tests exhibited

first gradual decreases in stress and then abrupt, simultaneous decreases in strength and

modulus during the life of the composite. In the high cycle fatigue regime of Region III,

an apparent endurance limit at 0.3 percent strain was observed.

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3. The definition of failure used for the strain-controlled tests in Region II (the first

abrupt, simultaneous decrea.._e in strength and modulus) was a suitable approach since it

produced fatigue lives in close agreement with lives from load-controlled tests which were

determined by complete frat:ture.

4. The advantages of using the tension-tension strain-controlled test method were

that specimen buckling was prevented while still providing a strain-control type of test and

stable fatigue cycling continued after large, abrupt decreases in load-carrying capacity. This

allows easy interruption and inspection of fatigue damage. The disadvantage of this test

method was the application of a continuously decreasing strain amplitude caused a continuous decrease in the severity of the test.

5. The advantage of using load-controlled testing was that failure was unambiguously

defined. The disadvantages were that ratchetting was induced and stable fatigue cycling

was not possible after a lalge reduction in the composite's load-carrying capability.

6. Since each type of test control mode has unique features and results in unique

fatigue response both strain- and load-controlled tests are important to the complete

understanding of SiC//Ti-24Al-11Nb composite fatigue behavior.

ACKNOWLEDGEMENT

The authors gratefully acknowledge the diligence and expertise of R.M. Shinn of

Sverdrup Technology, Inc. in the Fatigue Laboratory.

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ORI_':TNAE PAGE I' ;.

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Figure 1, - As-fabricated SiC/Ti-24AI-11Nb composite.

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.:_a..;,,a.........-=-,_W-_...._--

.or- Z-Tensil e iB_ t

.5 - ductility "1:_, -"__

"_ ,4" range I_ZL _%

F_gion I! J

.3-

E Region Ill -_

:_ .2 -- • 815 °C Strain controlled

£] 815 °C Load controlled

• 425 °C Strain controlled

O 425 °C Load controlled

,f I I I I I

100 10 t 10 2 103 104 105

N f (Cycle)

Figure 3. - Fatigue life diagram of SiCKi-24AI-11Nb at

elevated temperatures.

I

106
Catastrophic failure: fibre breakage,

,. interfacial debonding

I-f-J-rI-F-I I-I-J-

_"_,Region I[_,,_ _ Progressive failure: ," matrix cracking,

_ ."_ interfadal failure shear

Failure limit of matrix _ _

log N

Figure 4. - Talreja's fatigue life diagram for unidirectional polymer matrix composites at room temperature.

13.

v

(D

1000

8001

600

40O

200

o!

-200

100

-- I Maximum l

• 815 °C Strain controlled

_ [] 815 °C Load controlled

• 425 °C Strain controlled

O 425 °C Load controlled

m

I I I I

101 102 103 104

Cycle

Figure 5. - Variation of maximum and minimum stress

for region [ tests.

.6

.5

._ .4

u')

.3

m m _

• 81socStraincontrolled

[] 815 °C Load controlled

• 425 °C Strain controlled

O 425 °C Load controlled

2--m _ ,,,Am, a--- I-_-,, ----

o I I I I

10 o 101 102 103 104

Cycle

Figure 6. - Variation of maximum and minimum strain

tor region ! tests.

10

ORIGINAL PAGE IS

OF POOR QUALITY
• 815 °C Strain controlled

[] 815 °C Load controlled

• 425 °C Strain controlled

0 425 °C Load controlled

1BO

-- 170 __i____

"_%` En 160 _-== o(.9

1_°_'_l"--"_"i I _ I

uJ 140

100 101 102 103 104

Cycle

Figure 7. - Variation of elastic modulus for region I

tests.

%`

13_

o

O9

1000

800

6O0

400

2O0

_:Jo-<x:oea_

_33-43_

• 815 °C Strain controlled

[] 815 °C Load controlled _=

• 425 °C Strain controlled II.

-- 0 425 °C Load controlled !1

Minimun_

,t- _t __ I I I

10 0 101 10 2 103 104 10 5

Cycle

Figure 8. - Variation of maximum and minimum stress

for region TI tests.

180

160

[]

0

6O

10 o

815 °C Strain controlled

815 °C Load controlled

425 °C Strain controlled

425 °C Load controlled

I I I I I

101 102 103 104 105

Cycle

Figure 9. - Variation of elastic modulus for region I[

tests.

11

ORfGINAE PAGE IS

OF POOR QUALITY
1000

8OO

_" 600

40O

2OO

-- Cycle

_._ _i2)- looO

_JY'7 11,000 Before

__ //././././././././._z/ first

/ /_/ / 12,000 After

-- / ///'/ ./" several ///// , e,soroos

/ /Ae'/ / j 16,OOOAfter

-- / /J/ .// _ several

_-/ S/J N1 =11063

V I///_// I I I I I I

.1 .2 .3 .4 .5 .6 .7 .8

Strain (%)

Figure 10. - Typical stress-strain response of SiC/Ti-24AI-11Nb conducted at 425 °C with _;max= 0.67% under strain control (region I1).

1000 I'-" .,--1-2

| ,,;..-loo-looo

I .1.." .... sooo

800 "" -I0000

_" 600

400

(/3

2OO

ov J'f//,/I I I I I I I

0 .1 .2 .3 .4 .5 .6 .7 .8

Strain (%)

Figure 11. - Typical stress-strain repsponse of SiC/Ti-24AI-11Nb conducted at 425 °C with Omax = 830 MPa under load control (region [I).

12
n

O9

1000

800

600

400

200

1-2

-- 100-6000

6500 After 1st stress drop

10,000 After several

stress drops

N1= 6486

0 .1 .2 .3 .4 .5 .6 .7

Strain (%)

Figure 12. - Typical stress-strain response of SiC/Ti-24AI- 11Nb conducted at 815 °C with Emax = 0.45% under strain control (region If).

I

.8

IO(K) I

8c_ cycle

1-2 100 2000

_" 600

_o 400

o L//J" L/_" I I i I I

0 .1 .2 .3 .4 .5 .6

Strain (%)

Figure 13. - Typical stress-strain response of SiC/'I'i-24AI-11Nb

conducted at 815 °C with _max = 710 MPa under load control

(region [I).

13
National Aeronautics and

Space Adminpstratlon

1. Report No. 2. Government Accession No.

NASA TM-103157

4. Title and Subtitle

High Temperature Fatigue Behavior of a SiC/Ti-24AI-1 l Nb Composite

3. Recipient's Catalog No.

5. Report Date

Report Documentation Page

7. Author(s)

P.A. Bartolotta and P.K. Brindley

9. Performing Organization Name and Address

National Aeronautics and Space Administration

Lewis Research Center

Cleveland, Ohio 44135-3191

12. Sponsoring Agency Name and Address

National Aeronautics and Space Administration

Washington, D.C. 20546-0001

6. Performing Organization Code

8. Performing Organization Report No.

E-5524

10. Work Unit No.

510-01-0A

11. Contract or Grant No.

13. Type of Report and Period Covered

Technical Memorandum

14. Sponsoring Agency Code

15. Supplementary Notes

Prepared for the 10th Symposium on Composite Materials: Testing and Design sponsored by the American

Society for Testing Materials, San Francisco, California, April 24-25, 1990.

16. Abstract

A series of tension-tension strain- and load-controlled fatigue tests were conducted on unidirectional

SiC/Ti-24AI-IINb (at %) composites at 425 and 815 "C. Several regimes of damage were identified using

Talrega's (mlLe_4) concept of fatigue life diagrams. Issues of test technique, test control mode, and definition of

failure were also addressed.

17. Key Words (Suggested by Author(s))

Fatigue; Composites; Titanium aluminide; Silicon carbide;

Test methods

18. Distribution Statement

Unclassified- Unlimited

Subject Category 39

19. Security Classif. (of this report) 20. Security Classif. (of this page) 21. No. of pages 22. Price*

Unclassified Unclassified 14 A03

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