John Z. Gyekenyesi

N&R Engineering and Management Services Corporation, Parma Heights, Ohio

Pappu L.N. Murthy

Glenn Research Center, Cleveland, Ohio

Subodh K. Mital

University of Toledo, Toledo, Ohio

Life Prediction for a CMC Component Using

the NASALIFE Computer Code

NASA/TM—2005-213887

September 2005
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John Z. Gyekenyesi

N&R Engineering and Management Services Corporation, Parma Heights, Ohio

Pappu L.N. Murthy

Glenn Research Center, Cleveland, Ohio

Subodh K. Mital

University of Toledo, Toledo, Ohio

Life Prediction for a CMC Component Using

the NASALIFE Computer Code

NASA/TM—2005-213887

September 2005

National Aeronautics and

Space Administration

Glenn Research Center
Acknowledgments

Thank you to D.N. Brewer, U.S. Army Aviation Systems Laboratory, Glenn Research Center, for providing the

support for this effort, and also to A.M. Calomino and M.J. Verrilli, NASA Glenn Research Center, for providing

empirical data.

Available from

NASA Center for Aerospace Information

7121 Standard Drive

Hanover, MD 21076

National Technical Information Service

5285 Port Royal Road

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Available electronically at http://gltrs.grc.nasa.gov
NASA/TM—2005-213887 1

Life Prediction for a CMC Component Using 

the NASALIFE Computer Code 

John Z. Gyekenyesi 

N&R Engineering and Management Services Corporation 

Parma Heights, Ohio 44130 

Pappu L.N. Murthy 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Subodh K. Mital 

The University of Toledo 

Toledo, Ohio 43606 

Abstract

The computer code, NASALIFE, was used to provide estimates for life of a SiC/SiC stator vane 

under varying thermomechanical loading conditions. The primary intention of this effort is to show how 

the computer code NASALIFE can be used to provide reasonable estimates of life for practical propulsion 

system components made of advanced ceramic matrix composites (CMC). Simple loading conditions 

provided readily observable and acceptable life predictions. Varying the loading conditions such that low 

cycle fatigue and creep were affected independently provided expected trends in the results for life due to 

varying loads and life due to creep. Analysis was based on idealized empirical data for the 9/99 Melt 

Infiltrated SiC fiber reinforced SiC. 

1. Introduction

Engine companies are constantly striving to improve the performance and life of their gas turbine 

engines. Materials are pushed to new limits as new materials and concepts are applied to gas turbines to 

increase operating temperatures, reduce weight, and improve aerodynamic efficiencies. Also, there is the 

need to reduce maintenance costs and down time making life prediction an ever more important part of 

the design process. The task of accurately determining the fatigue life of an engine component in service 

has become increasingly complex. This results, in part, from the complex missions which are now 

routinely considered during the design process. These missions include large variations of multiaxial 

stresses and temperatures experienced by critical engine parts. 

NASALIFE was written in an attempt to provide a convenient software package to the designer for 

determining the life of a component under cyclic or variable thermo-mechanical loading. The program 

was developed under the National Aeronautics and Space Administration’s (NASA) Enabling Propulsion 

Materials (EPM) and Ultra Efficient Engine Technology (UEET) programs. The EPM program and its 

objectives were summarized by Brewer (1999) and the UEET program overview was presented by Shaw 

and Peddie (2004). NASALIFE was covered in detail by Gyekenyesi, et al. (2005) with a brief overview 

provided earlier by Levine, et al. (2000). It should be noted that NASALIFE is not a final and all inclusive 

program for predicting component life, but is the part of the continuing process to improve life prediction 

techniques. Empirical data is used as a reference for predicting the fatigue and rupture life of a 

component. 
NASA/TM—2005-213887 2

In addition to the stress and lifing analysis, the high temperature ceramic matrix composite (CMC), 

silicon melt infiltrated (MI) silicon carbide fiber reinforced silicon carbide matrix, was developed under 

the aforementioned NASA programs. The MI SiC/SiC composite system was chosen for its low density, 

acceptable high temperature mechanical behavior, resistance to environmental attack within a gas turbine 

application, and potential for being processed in high volumes at an acceptable cost. The composite 

utilizes a silicon doped boron nitride fiber/matrix interface. The EPM program developed the composite 

for a gas turbine combustor application and part of the UEET program’s goal was to use the composite for 

a gas turbine stator vane. The programs helped advance the processing methods for the CMC and 

provided the support for generating mechanical properties data by thermomechanical testing of the 

material. 

Fatigue and creep at relatively high temperatures are two important material phenomena, among 

others, that have to be addressed when designing high temperature components for a gas turbine 

application. The work presented here, as one of the UEET program milestones, utilizes NASALIFE to 

predict an approximate number of missions a component is capable of performing prior to failure due to 

cyclic fatigue and/or creep rupture under a given mission loading condition. The high stress condition of a 

model of a thermomechanically loaded CMC stator vane, analyzed by N&R Engineering (2003), is used 

as the reference loading condition for the CMC. The CMC material of the vane and the corresponding 

data on which the analysis is based is designated as the 9/99 MI SiC/SiC composite. This work utilizes 

the limited empirical data generated for the 9/99 ceramic composite and the analysis capability of 

NASALIFE in an effort to verify the basic functionality of the life prediction program. In addition, the 

work provides a preliminary view to the behavior of a MI SiC/SiC stator vane with respect to life. Lastly, 

this work is the first step towards a sensitivity and probabilistic analysis for life prediction of the MI 

SiC/SiC stator vane. 

2. Model and Material 

The focus of this work is to approximate the life of a MI SiC/SiC stator vane. The vane is loaded at a 

high temperature. The highest stress at a given temperature is the reference load. 

The vane is fabricated from a MI SiC/SiC composite using a silicon doped boron nitride fiber/matrix 

interface and a surface environmental barrier coating (EBC) with a bond coat. The composite is laminated 

with each laminate having a five harness satin woven fiber architecture. Tow spacing is approximately 

20 ends per inch (EPI). The walls of the vane have six plies. More details of the ceramic composite are 

provided in the EPM Final Report (2000). A typical vane is illustrated in figure 1. The location of the 

highest stress is noted in the illustration also. 

Mechanical and other high temperature properties are taken from the EPM Final Report (2000). In 

addition, some material properties were provided by Calomino (2004) and Verrilli (2004). The reported 

values are the statistical average. With a maximum of three specimens per condition the data exhibited 

significant scatter also. The scatter was typical of cyclic fatigue tests for most materials, that is, quite 

significant. A range of an order of magnitude or greater in the life of the material is not unusual, 

especially with the limited number of specimens per condition. The range of data is still limited due to the 

fact the material system is relatively new and is still in a state of flux. As a result, for this study the range 

of data was expanded by extrapolation. Higher and lower stresses were linearly extrapolated using the log 

of life. Due to the limitations of the MI SiC/SiC data as described above the data should be viewed only 

for demonstrating the functionality of NASALIFE and not be used for design purposes. 

Pesting of the composite is an issue near a temperature of 1470 °F (800 °C) as noted by Ogbuji 

(2002). Ogbuji (2002) describes pesting as the oxidative degradation of CMCs in a service environment at 

intermediate temperatures. The experimental data illustrated this problem as the fatigue strength dropped 

significantly at 1500 °F (820 °C). For this study the data, where pesting is a concern, was not utilized. The 

phenomena of pesting may be incorporated in a future study. 
NASA/TM—2005-213887 3

 

Load VS. Time

0

26

0123

Time (h)

Stress

σmax

Figure 1.—MI SiC/SiC stator vane showing 

temperature distribution and the location 

of the maximum stress.

The low cycle fatigue (LCF) data was generated using a trapezoidal waveform loading condition on a 

uniaxial tensile specimen as described in the EPM Final Report (2000). The test specimen was loaded 

from approximately zero to the desired maximum tensile stress by ramping up to the desired load, holding 

for two hours, and unloading back to approximately zero. The rate of loading and unloading was such that 

the maximum stress or near zero stress was reached in 1 minute. Figure 2 illustrates the applied load 

cycle. The noted cyclic loading was repeated until complete fracture of the specimen. The complete 

fracture was used as the definition of failure for the cyclic fatigue and creep tests. The trapezoidal wave 

form load cycle was utilized to model the potential loading conditions a combustor may experience for an 

average duration flight. The resulting R-ratio, that is the minimum stress divided by the maximum stress, 

is approximately zero. Also, the resulting A-ratio, the stress amplitude divided by the mean stress, is 

approximately one. One should keep in mind that due to the uniaxial tensile test the coupons had a 

uniform stress throughout the cross section. Whereas, most modeled components will typically have some 

stress gradient through a cross section. 

Due to the extended hold time for each cycle of the LCF tests creep deformation is a significant part 

of the overall failure of the composite at higher temperatures. At the current time, this is the most 

complete matrix of data that is available for any type of high temperature LCF information for this 

composite system. As such, the available data was utilized for defining the behavior under LCF 

conditions. DiCarlo, et al. (2004) discussed some of the issues with creep with the MI SiC/SiC CMC 

system. At the time of this writing a single LCF data point for the 9/99 MI SiC/SiC generated with a 

triangular wave form load at a frequency of 0.33 Hz was presented by Kalluri, et al. (2005). That data was 

used to compare properties with a single data point for a slightly modified version of the ceramic 

composite designated as 1/01 MI SiC/SiC. Unfortunately, the one data point is insufficient information 

for this study. 

Figure 3 shows the LCF curves with the maximum stress, with an R-ratio of zero, as a function of 

cycles to failure. The curves utilize actual experimental data as the foundation but are highly massaged to 

provide smooth nonoverlapping curves for this study. Failure was defined as the complete fracture of the 

composite test specimen. The data has been modified for this work. Outlier points, due to significant 

scatter in the data, were deleted to provide smoother curves for this initial analysis. Points were linearly 

extrapolated to extend the range of the data. It is the idealized data that is illustrated in figure 3 and used 

for this preliminary analysis. As more tests are being conducted the developing database of the composite 

properties will provide a more sufficient base for design and analysis with the MI SiC/SiC composite 

system. It should be reiterated that available data is quite limited with significant scatter that typically 

comes with cyclic fatigue tests. The range in life at any one condition can be an order of magnitude or 

greater from the lowest to the highest number of cycles. The data presented with this work is for 

demonstrating NASALIFE functionality and should not be used for design purposes. 

Figure 2.—Illustration of load cycles used to 

generate the cyclic fatigue data for the 

9/99 MI SiC/SiC composite. 
NASA/TM—2005-213887 4

Low Cycle Fatigue

Maximum Stress VS. Life

0

5

10

15

20

25

30

35

1.E+00 1.E+01 1.E+02 1.E+03 1.E+04 1.E+05 1.E+06 1.E+07 1.E+08 1.E+09

Life (cycles)

σmax (ksi)

 70°F

400°F

800°F

1200°F

1600°F

1900°F

2200°F

R=0

A=1

Figure 3.—Idealized low cycle fatigue curve for MI SiC/SiC. Maximum

stress versus the life in cycles with an R-ratio of zero.

Creep Rupture Life

0

5

10

15

20

25

30

1.E+00 1.E+03 1.E+06 1.E+09 1.E+12 1.E+15 1.E+18 1.E+21 1.E+24 1.E+27 1.E+30 1.E+33

Life (h)

σ (ksi)

70°F

400°F

1200°F

2400°F

Figure 4.—Creep rupture stress as a function of time for the 9/99 

MI SiC/SiC ceramic matrix composite. Creep life at 400 °F and 

below was set at the maximum allowable value within 

NASALIFE essentially indicating no creep deformation. 

The MI SiC/SiC composite is being developed for high temperature applications. Another major 

concern with materials at high temperatures is creep deformation. Creep rupture is a significant issue 

along with the cyclic fatigue strength of the composite. Figure 4 presents creep rupture curves for the 9/99 

MI SiC/SiC composite. Creep at 400 °F (200 °C) and below was assumed to be negligible to nonexistent. 
NASA/TM—2005-213887 5

The life at 70 °F and 0.1 ksi (0.7 MPa) was set at the maximum value that NASALIFE accepts, that is,

1031 hours. The other life values at the low temperatures were reduced only slightly from the maximum 

acceptable life within NASALIFE. As a result, the low temperature damage due to creep is effectively 

zero relative to the damage from the cyclic loading. 

As with the LCF data the available test results for creep rupture is quite limited. The curves for the 

high temperature creep rupture data were extrapolated to lower stresses to extend the range of the 

available data. The definition of failure by creep is the same as for the LCF data, that is, the complete 

fracture of the specimen. 

3. NASALIFE 

NASALIFE is a program for predicting the life, due to LCF and creep rupture, of a 

thermomechanically loaded component. The program was developed to model CMCs. A detailed 

description was provided by Gyekenyesi, et al. (2005). The program can take a multiaxial loading 

condition where the load and temperature may vary and calculate an approximate life based on cyclic 

fatigue and creep. The output includes the damage due to the creep, damage due to the variable load, and 

the combined damage. 

Empirical data from low cycle fatigue tests, creep rupture tests, and static tensile tests are used as the 

reference for predicting the number of missions a component can handle under a given thermomechanical 

loading condition. The creep rupture information can be entered in tabular form with stress versus life or 

with the use of Larson-Miller (1952) parameters. 

The program uses the Walker (1970) Mean Stress model to adjust life prediction for ranges in the 

R-ratio. Complex load cycles are reduced by the Rainflow Counting Method as presented by both, Endo 

et al. (1967) and Mitsunaga and Endo (1968). In addition, the Rainflow Counting Method is covered by 

the American Society for Testing and Materials (ASTM) (2004) standard E 1049–85. Minor’s (1945) 

Rule is utilized to combine the damage at different load levels. Finally, the program determines the total 

damage due to creep and combines it with the fatigue damage due to the cyclic loading and determines 

the approximate number of missions a component can handle up to failure. It should be noted that the 

definition of failure for NASALIFE is the same as the definition of failure used for the empirical data. 

The point of highest stress within the SiC/SiC vane, illustrated in figure 1, is the focus this study. At this 

time the changes in area due to crack propagation are ignored. 

4. Mission Load Cases for Analysis 

The mission loads were selected to test the output from NASALIFE as a verification process for its 

functionality and to provide an insight for a potential life of the modeled stator vane. The calculated 

maximum stress of 14.1 ksi (97 MPa) and a temperature of 2000 °F (1090 °C) within the stator vane, as 

determined by N&R Engineering (2003), were used as the primary reference load condition for this study. 

Also, the mission loads were set up like the original LCF loads with a ramp from approximately zero up 

to a load which was held for a duration followed by a ramp down to an approximately zero load. As noted 

earlier, figure 2 illustrates the load cycle. Other conditions included combinations of hypothetical 

situations by reducing the mean hold load to 7 ksi (48 MPa) at 2000 °F (1090 °C) and lastly, varying the 

14.1 ksi (97 MPa) stress at temperatures of 400 °F (200 °C) and 70 °F (20 °C). The combination of low 

temperatures and high stress is noted as a hypothetical situation since it is not practical for a gas turbine 

engine application but a good condition for checking NASALIFE’s functionality. Varying loads of 

different magnitudes were superimposed over the mean hold stress during the hold period. The 

superimposed loads over the main load of 14.1 ksi (97 MPa) may model what vibrations or localized 

changes in turbine gas flow may induce. The random number generator in Miscrosoft® Excel 2002 was 

utilized to generate the random loads for the runs that had a random load superimposed over the hold 
NASA/TM—2005-213887 6

Mission Loading

Applied Stress VS. Time

0

2

4

6

8

10

12

14

16

0 1000 2000 3000 4000 5000 6000 7000 8000 9000 10000

Time (s)

Stress (ksi)

ramp duration: 60 s

hold time

∆t=30 s between peaks

ramp duration: 60 s

Figure 5.—A sample mission loading with a mean stress of 14 ksi and superimposed 

random load of ±0 to 2.5 percent. The mission consists of the ramp up, hold, and 

ramp down. Time increment between load peaks are fixed at 30 seconds. 

stress. A typical mission with its load curve is illustrated in figure 5. The details of the individual 

missions will be presented in the following section along with the corresponding results for the life 

prediction. Appendix A illustrates one of the NASALIFE input files. The same material properties data 

were utilized for all the runs. The NASALIFE technical/user manual by Gyekenyesi, et al. (2005) 

describes all the entries of the input file. 

5. Results and Discussion 

The work presented here is the initial application of NASALIFE for predicting the life of a CMC 

component under variable thermomechanical loading conditions. As noted earlier, the runs were used to 

verify NASALIFE functionality and to provide an insight for a potential life of the modeled stator vane 

under various practical and hypothetical loading conditions. 

NASALIFE utilizes empirical data as a reference for cyclic fatigue strength and creep rupture 

strength. Gyekenyesi, et al. (2005) provided details of NASALIFE and the theories and models within the 

program used to predict the number of missions to failure. It was noted earlier that the MI SiC/SiC 

material properties data base for NASALIFE was not ideal. The data used for fatigue contained 

significant creep due to the long duration hold times with the trapezoidal loading wave form. Also, the 

stress range of the data had to be extended by extrapolation. As the CMC continues to be characterized 

and more data becomes available it will be incorporated for future applications of NASALIFE. 

The definition of failure is complete fracture for the test coupons used to generate the LCF and creep 

rupture data as per the EPM Final Report (2000). As a result, the same definition has to be applied to the 

stator vane. Also, one should keep in mind that each uniaxial tensile test coupon had a uniform stress 

through its cross section, whereas, the stator vane will have a stress gradient. 

Tables 1 thru 3 provide life prediction results along with the respective load conditions. Each table is 

at a single temperature as indicated in the table titles. The results are grouped as cases with one condition 

being varied within each case. The first condition in each case is used as the reference state and is 

indicated in bold print. The percentage drop in life, due to the joint effect of creep and LCF, is based on 

the reference condition within each case. Figure 5 illustrates a typical load curve that was utilized to 
NASA/TM—2005-213887 7

model a mission load cycle. Variants of the presented mission were used to generate the results with 

NASALIFE. 

The simple load cases that ramped up to a load, held the load for a duration, and unloaded can be 

compared directly with figure 3 for the life prediction due to cyclic fatigue. The predicted results 

compared well with the expected trends of the experimental data. 

The low temperature cases in table 1 at 70 °F (20 °C) and in table 2 at 400 °F (200 °C) were selected 

to effectively eliminate the effects of creep. Both tables of data indicate the missions to failure due to 

creep are large numbers showing creep damage to be essentially zero at these low temperatures. 

TABLE 1.—LIFE PREDICTIONS FOR A CMC STATOR VANE WITH 

GIVEN MISSION LOADING CONDITIONS AT 70 °F 

Mission % drop

Load Conditions Due to

Creep

Due to

Cyclic 

Fatigue

Joint

Effect

relative to

base 

condition

Vary frequency of superimposed load

60 s ramp, 7200 s hold at 14.1 ksi, 60 s ramp 4.89E+30 47433 47433

60 s ramp, 7200 s hold at 14.1 ksi with ±0.5% and ∆t=2400 s, 60 s ramp 4.89E+30 46000 46000 3.0

60 s ramp, 7200 s hold at 14.1 ksi with ±0.5% and ∆t=1200 s, 60 s ramp 4.89E+30 45982 45982 3.1

60 s ramp, 7200 s hold at 14.1 ksi with ±0.5% and ∆t=600 s, 60 s ramp 4.89E+30 45928 45928 3.2

Vary magnitude of superimposed load

60 s ramp, 7200 s hold at 14.1 ksi, 60 s ramp 4.89E+30 47433 47433

60 s ramp, 7200 s hold at 14.1 ksi with + 0-0.5% random and ∆t=30 s, 60 s ramp 4.89E+30 44831 44831 5.5

60 s ramp, 7200 s hold at 14.1 ksi with + 0-1.5% random and ∆t=30 s, 60 s ramp 4.89E+30 42155 42155 11.1

60 s ramp, 7200 s hold at 14.1 ksi with + 0-2.5% random and ∆t=30 s, 60 s ramp 4.89E+30 39701 39701 16.3

bold font - reference or base condition within each case

∆t - time between sequential load peaks

Missions to Failure

TABLE 2.—LIFE PREDICTIONS FOR A CMC STATOR VANCE WITH 

GIVEN MISSION LOADING CONDITIONS AT 400 °F 

Mission % drop

Load Conditions Due to

Creep

Due to

Cyclic 

Fatigue

Joint Effect

relative to

base 

condition

Vary frequency of superimposed load

60 s ramp, 7200 s hold at 14.1 ksi, 60 s ramp 4.81E+30 13301 13301

60 s ramp, 7200 s hold at 14.1 ksi with ±0.5% and ∆t=2400 s, 60 s ramp 4.81E+30 12943 12943 2.7

60 s ramp, 7200 s hold at 14.1 ksi with ±0.5% and ∆t=1200 s, 60 s ramp 4.81E+30 12928 12928 2.8

60 s ramp, 7200 s hold at 14.1 ksi with ±0.5% and ∆t=600 s, 60 s ramp 4.81E+30 12884 12884 3.1

Vary magnitude of superimposed load

60 s ramp, 7200 s hold at 14.1 ksi, 60 s ramp 4.81E+30 13301 13301

60 s ramp, 7200 s hold at 14.1 ksi with + 0-0.5% random and ∆t=30 s, 60 s ramp 4.81E+30 12027 12027 9.6

60 s ramp, 7200 s hold at 14.1 ksi with + 0-1.5% random and ∆t=30 s, 60 s ramp 4.81E+30 11368 11368 14.5

60 s ramp, 7200 s hold at 14.1 ksi with + 0-2.5% random and ∆t=30 s, 60 s ramp 4.81E+30 10773 10773 19.0

bold font - reference or base condition within each case

∆t - time between sequential load peaks

Missions to Failure

In reference to table 1 for the first group at 70 °F (20 °C) the frequency of the superimposed stress 

was varied. The hold conditions were 14.1 ksi (97.2 MPa) for 2 hours. The variable load was ±0.5 percent 

of 14.1 ksi (97.2 MPa). The peak to peak duration was held constant within each case. The peak to peak 

durations were varied between cases from 600 to 2400 seconds giving a range of three load peaks to 

twelve load peaks per mission. The superimposed load being a low magnitude of ±0.5 percent of the 

14.1 ksi (97.2 MPa) or ±0.071 ksi (0.49 MPa) only had a minor effect on the number of missions the 

component can provide up to failure. The number of missions with a simple loading with a ramp from 

approximately zero to 14.1 ksi (97.2 MPa), hold at 14.1 ksi (97.2 MPa), and unload was 47433. The 
NASA/TM—2005-213887 8

number of missions was reduced by 3.2 percent by superimposing a varying load of ±0.071 ksi 

(0.49 MPa) over the hold stress of 14.1 ksi (97.2 MPa) with 600 seconds between load peaks. 

The next case, at 70 °F (20 °C) had a random superimposed load increase in magnitude. Runs were 

compared with no superimposed load going up to a varying load of 0 to +2.5 percent over the 14.1 ksi 

(97.2 MPa). The duration between load peaks was held at 30 seconds resulting in 240 load peaks in the 

2 hours. The number of missions went from 47433 to 39701 or a 16 percent reduction in life by going 

from a fixed hold stress to a 0 to +2.5 percent superimposed random load with 30 seconds between peaks. 

The load cases at 70 °F (20 °C) were repeated at 400 °F (200 °C). The results are presented in 

table 2. As with the previous case creep is insignificant. The basic load cycle of ramping from 

approximately zero to 14.1 ksi (97.2 MPa), holding at 14.1 ksi (97.2 MPa) for 2 hours, and ramping down 

was maintained. Superimposing a load of ±0.5 percent over the 14.1 ksi (97.2 MPa) load with increasing 

frequency reduced the life by 3 percent within the tested range. 

Increasing the magnitude of the random load from no superimposed load to a varying load of 0 to 

+2.5 percent at 400 °F (200 °C) reduced the missions to failure by a significant 19 percent. The load peak 

to peak duration was held constant at 30 seconds with the superimposed variable load. The resulting 

number of load peaks was 240 within the 2 hour hold period. The remaining results are shown in table 3 

for a temperature of 2000 °F (1090 °C). At this temperature creep becomes a significant part of the failure 

process when considering the life of a thermomechanically loaded component. Significant creep 

deformation can promote matrix cracking which exposes the constituents to the harsh environment in 

which a CMC component would be utilized. Upon cracking of the matrix oxidation of the constituents 

becomes a major concern. Lewinsohn, et al. (2000) discussed time dependent failure of nonoxide CMCs 

such as the SiC/SiC system. The EPM final report (2000) discusses the oxidation process and its 

detrimental effects on the mechanical properties of the SiC/SiC composite. The creep and oxidation 

effects can be observed in table 3 with the significant drop in life as loading conditions and durations 

under load increase. 

For the first high temperature case the hold duration at 14.1 ksi (97.2 MPa) was varied from 

5 seconds to 4 hours. As expected the creep damage increased as the duration increased. The number of 

missions relative to creep damage ranged from 22532 to 42 or a 99.8 percent reduction in number of 

missions to failure within the range of the study. The life relative to the joint effect of creep and LCF was 

reduced by 94.9 percent. The damage due to cyclic fatigue did not change as it was the duration and not 

the number of cycles that changed. No load frequency effect is anticipated within the range of this study. 

As with the low temperature cases the frequency of the superimposed load was varied at 2000 °F 

(1090 °C). The holding stress was 14.1 ksi (97.2 MPa) with a ±0.5 percent load superimposed over it. 

Load peak to peak durations were varied from 2 hours with no superimposed load to 600 seconds with the 

±0.5 percent superimposed load. Damage due to creep remained constant with 84 missions to failure. At 

the same time the missions to failure with respect to the cyclic loading went from 811 to 387. That is a 

52 percent reduction in life with respect to the cyclic loading. The combined missions to failure ranged 

from 76 to 69 for a 9.2 percent reduction in the number of missions to failure. 

For the next case the magnitude of the maximum load of an applied random load was increased. The 

range included the simple load with no superimposed load variations to a superimposed load of 0 to 

+2.5 percent. The load peak to peak duration was held constant at 30 seconds. The number of missions 

relative to creep was 84 and 85 which was relatively constant. The number of missions with respect to the 

cyclic loading went from 811 for the simple loading case down to 42 missions with the superimposed 0 to 

+2.5 percent random load. Although, just the application of a 0 to +0.5 percent random load reduced the 

missions to failure due to cyclic fatigue to 49. 

The next trial maintained the holding stress of 14.1 ksi (97.2 MPa) for 2 hours with a 0 to 

+0.5 percent random load superimposed over it. The duration between load peaks was varied from 

30 seconds down to 1 second. As a result, the number of load peaks ranged from 240 up to 7200 in the 

2 hour hold period. The number of missions to failure due to creep held fairly constant at 84 and 85. 
NASA/TM—2005-213887 9

TABLE 3.—LIFE PREDICTIONS FOR A CMC STATOR VANE 

WITH GIVEN MISSION LOADING CONDITIONS AT 2000 °F 

Mission % drop

Load Conditions Due to

Creep

Due to

Cyclic 

Fatigue

Joint Effect

relative to

base

condition

Vary hold load duration

60 s ramp, 5 s hold at 14.1 ksi, 60 s ramp 22532 811 782

60 s ramp, 30 s hold at 14.1 ksi, 60 s ramp 11715 811 758 3.1

60 s ramp, 60 s hold at 14.1 ksi, 60 s ramp 7433 811 731 6.5

60 s ramp, 7200 s hold at 14.1 ksi, 60 s ramp 84 811 76 90.3

60 s ramp, 14400 s hold at 14.1 ksi, 60 s ramp 42 811 40 94.9

Vary frequency of superimposed load

60 s ramp, 7200 s hold at 14.1 ksi, 60 s ramp 84 811 76

60 s ramp, 7200 s hold at 14.1 ksi with ±0.5% and ∆t=2400 s, 60 s ramp 84 668 75 1.3

60 s ramp, 7200 s hold at 14.1 ksi with ±0.5% and ∆t=1200 s, 60 s ramp 84 562 73 3.9

60 s ramp, 7200 s hold at 14.1 ksi with ±0.5% and ∆t=600 s, 60 s ramp 84 387 69 9.2

Vary maximum magnitude of superimposed load

60 s ramp, 7200 s hold at 14.1 ksi, 60 s ramp 84 811 76

60 s ramp, 7200 s hold at 14.1 ksi with + 0-0.5% random and ∆t=30 s, 60 s ramp 84 49 31 59.2

60 s ramp, 7200 s hold at 14.1 ksi with + 0-1.5% random and ∆t=30 s, 60 s ramp 85 45 29 61.8

60 s ramp, 7200 s hold at 14.1 ksi with + 0-2.5% random and ∆t=30 s, 60 s ramp 84 42 28 63.2

Vary frequency of random magnitude superimposed load

60 s ramp, 7200 s hold at 14.1 ksi with + 0-0.5% random and ∆t=30 s, 60 s ramp 84 49 31

60 s ramp, 7200 s hold at 14.1 ksi with + 0-0.5% random and ∆t=10 s, 60 s ramp 85 17 14 54.8

60 s ramp, 7200 s hold at 14.1 ksi with + 0-0.5% random and ∆t=5 s, 60 s ramp 84 9 8 74.2

60 s ramp, 7200 s hold at 14.1 ksi with + 0-0.5% random and ∆t=1 s, 60 s ramp 84 2 2 93.5

Vary maximum magnitude of superimposed load

60 s ramp, 14400 s hold at 14.1 ksi, 60 s ramp 42 811 40

60 s ramp, 14400 s hold at 14.1 ksi with + 0-0.5% random and ∆t=30 s, 60 s ramp 42 26 16 60.0

60 s ramp, 14400 s hold at 14.1 ksi with + 0-1.5% random and ∆t=30 s, 60 s ramp 42 22 15 62.5

60 s ramp, 14400 s hold at 14.1 ksi with + 0-2.5% random and ∆t=30 s, 60 s ramp 42 22 14 65.0

Vary duration of hold load

60 s ramp, 5 s hold at 7.0 ksi, 60 s ramp 225591 1756 1742

60 s ramp, 30 s hold at 7.0 ksi, 60 s ramp 143545 1756 1734 0.5

60 s ramp, 60 s hold at 7.0 ksi, 60 s ramp 99931 1756 1725 1.0

60 s ramp, 7200 s hold at 7.0 ksi, 60 s ramp 1363 1756 767 56.0

Vary maximum magnitude of superimposed load

60 s ramp, 7200 s hold at 7.0 ksi, 60 s ramp 1363 1756 767

60 s ramp, 7200 s hold at 7.0 ksi with + 0-0.5% random and ∆t=30 s, 60 s ramp 1364 59 56 92.7

60 s ramp, 7200 s hold at 7.0 ksi with + 0-1.5% random and ∆t=30 s, 60 s ramp 1364 52 50 93.5

60 s ramp, 7200 s hold at 7.0 ksi with + 0-2.5% random and ∆t=30 s, 60 s ramp 1360 48 46 94.0

Apply increased magnitude superimposed load

60 s ramp, 7200 s hold at 7.0 ksi

 with ± 0-0.5% of 14.1 ksi random load and ∆t=30 s, 60 s ramp 1366 47 45

Vary frequency of superimposed load

60 s ramp, 7200 s hold at 7.0 ksi with + 0-0.5% random and ∆t=30 s, 60 s ramp 1364 59 56

60 s ramp, 7200 s hold at 7.0 ksi with + 0-0.5% random and ∆t=10 s, 60 s ramp 1364 19 19 66.1

60 s ramp, 7200 s hold at 7.0 ksi with + 0-0.5% random and ∆t=5 s, 60 s ramp 1363 10 10 82.1

60 s ramp, 7200 s hold at 7.0 ksi with + 0-0.5% random and ∆t=1 s, 60 s ramp 1363 2 2 96.4

Vary magnitude of hold load

60 s ramp, 7200 s hold at 6.5 ksi, 60 s ramp 1686 1905 894

60 s ramp, 7200 s hold at 7.0 ksi, 60 s ramp 1363 1756 767 14.2

60 s ramp, 7200 s hold at 7.5 ksi, 60 s ramp 1102 1627 657 26.5

60 s ramp, 7200 s hold at 13.6 ksi, 60 s ramp 98 844 88 90.2

60 s ramp, 7200 s hold at 14.1 ksi, 60 s ramp 84 811 76 91.5

60 s ramp, 7200 s hold at 14.6 ksi, 60 s ramp 72 780 66 92.6

bold font - reference or base condition within each case

∆t - time between sequential load peaks

Missions to Failure
NASA/TM—2005-213887 10

On the other hand, missions to failure with respect to cyclic loading dropped significantly from 49 to 

2 or by 96 percent. The missions to failure from the combined effect of creep and cyclic loading dropped 

from 31 to 2 or 94 percent. 

Another case extended the hold time to 4 hours. The set compared the simple load condition with a 

hold stress of 14.1 ksi (97.2 MPa) with runs having a superimposed random load of 0 up to +2.5 percent. 

The time between load peaks was held constant at 30 seconds. The missions to failure due to creep held 

constant at 42. The missions to failure due to cyclic fatigue dropped from 811 for a simple load to 22 for a 

0 to +2.5 percent superimposed random load. That is a 97 percent decrease due to the cyclic loading. The 

majority of the life reduction is just from the application of the random load. Superimposing a random 

load of 0 to +0.5 percent resulted in 26 missions to failure due to the cyclic loading. The combined effect 

of the creep and cyclic loading reduced the number of missions from 40 for a simple load to 14 with the 

superimposed 0 to 2.5 percent random load. 

For the next few cases from table 3 the holding stress was reduced to 7.0 ksi (48 MPa). The remaining 

variables were adjusted in the same manner as the previous cases. 

The first 7.0 ksi (48 MPa) set had the duration of the hold stress adjusted. The time ranged from 

5 seconds to 2 hours. The resulting missions to failure due to creep ranged from 2.256105

 to 1363. The 

missions to failure due to the cyclic stress held constant at 1756. The combined effect missions to failure 

ranged from 1742 to 767. 

The next set compares a simple load cycle with various magnitudes of superimposed positive random 

loads. The maximum stress of the random loads ranged from 0.5 to 2.5 percent of 7.0 ksi (48 MPa). The 

hold stress duration was held constant at 2 hours as was the time between load peaks at 30 seconds. 

Missions to failure due to creep showed a negligible change from 1364 to 1360. The missions to failure 

due to the cyclic loading dropped from 1756 with a simple load curve with constant 7.0 ksi (48 MPa) to 

59 with the addition of a 0 to 0.5 percent random load. From there the missions to failure dropped more 

gracefully to 48 as the peak load was increased to 2.5 percent. The combined effect missions to failure 

ranged from 767 to 46. 

One case was run with a 2 hour hold stress of 7.0 ksi (48 MPa) with a superimposed positive random 

load ranging from 0 to 0.5 percent of 14.1 ksi. The intent is to superimpose the same magnitude stress 

amplitude as used with the 14.1 ksi load. Missions to failure due to creep was 1366 and missions to 

failure due to cyclic loading was 47. The combined effect missions to failure was 45. 

The frequency of the superimposed 0 to +0.5 percent random load was varied over the 7.0 ksi 

(48 MPa) hold stress also. The time between load peaks was varied from 30 seconds down to 1 second. 

The missions to failure due to the creep remained at 1364 and 1363. Whereas, the missions to failure due 

to the cyclic loading exhibited a significant decrease from 59 to 2. The missions to failure due to the 

combined effect from creep and the cyclic loading decreased from 56 to 2 or by 96 percent. 

Lastly, a simple loading situation was applied by ramping to a fixed stress, holding for 2 hours, and 

ramping back to a no load condition. The hold stress was varied from 6.5 ksi (45 MPa) to 14.6 ksi 

(101 MPa). Damage from creep and cyclic loading were significant. Missions to failure due to creep 

ranged from 1686 to 72 and missions to failure due to cyclic loading ranged from 1905 to 780. The 

combined effect of creep and cyclic loading provided missions to failure of 894 at 6.5 ksi (45 MPa) 

maximum stress to 66 at 14.6 ksi (101) maximum stress. 

The life prediction results from NASALIFE show trends as one would expect. The missions that 

consist of a simple ramp up, hold stress and ramp down provide results for the missions to failure with 

respect to the cyclic loading that can be compared directly to the chart shown in figure 3. 

The varying of the hold duration but maintaining the maximum stress did provide the expected results 

of a constant missions to failure due to cyclic fatigue but a reduction in missions to failure due to creep as 

the hold duration increased. 

On the other hand, superimposing a low magnitude variable load over the hold stress and altering the 

frequency of the variable load had minimal effect on creep but increased the damage due to the cyclic 

loading. 
NASA/TM—2005-213887 11

6. Summary and Conclusions 

Accurate life prediction under conditions of LCF and creep is a difficult process. The analysis 

requires a complete database of material behavior in fatigue and creep under different temperatures and 

loading conditions. 

With a limited database NASALIFE is shown to provide a good estimate for life of a 

thermomechanically loaded structural component made of the NASA 9/99 MI SiC/SiC ceramic composite 

system. The output for life compares well with the LCF data when the loading condition is the same as 

the loading conditions used to generate the LCF data. The trends shown by NASALIFE are as expected 

for the loading conditions that were used for this study. 

A full range of material properties data is required to make any kind of life prediction. In particular, 

the data must cover the operating conditions for which the analysis is to be made. 

As part of the development program for the SiC/SiC composite system a full range of LCF and creep 

data needs to be generated. The LCF data should have a high enough frequency such that the effects of 

creep are significantly reduced relative to the currently available data. Also, the definition of failure may 

have to be revised to reflect the typical design failure that may be used for a component in a gas turbine 

engine. 

Future work may include LCF tests with variable loading at high temperatures to verify the output of 

NASALIFE. Issues, such as pesting, will have to be studied and incorporated to increase life prediction 

accuracy, especially if there is significant operating time in the vicinity of the pesting temperature. 

Although, as environmental barrier coatings improve the problem of pesting may be reduced or 

eliminated altogether. 

Also for future work it was noted earlier that this effort is the first step towards utilizing NASALIFE 

for sensitivity and probabilistic analysis of high temperature structural components made with the 

advanced SiC/SiC ceramic composite system. The program may be used in conjunction with a finite 

element analysis package as part of the sensitivity and probabilistic analysis. 

References 

1. American Society for Testing and Materials, “E 1049–85 (Reapproved 1997) Standard Practice for 

Cycle Counting in Fatigue Analysis.” ASTM International, West Conchohocken, PA, 2004. 

2. Brewer, D.; “HSR/EPM Combustor Materials Development Program.” Materials Science and 

Engineering, A261, 1999. pp. 284–291. 

3. Calomino, A.M.; NASA Glenn Research Center, Cleveland, OH, personal communication, 2004. 

4. DiCarlo, J.A.; Yun, H.–M.; Morscher, G.N.; and Bhatt, R.T., “SiC/SiC Composites for 1200 °C and 

Above.” NASA Technical Memorandum, NASA/TM—2004-213048, November, 2004. 

5. Enabling Propulsion Materials Program–Final Technical Report, NASA Contract NAS3 26385, 

May, 2000. 

6. Endo., T.; Mitsunaga, K.; and Nakagawa, H.; “Fatigue of Metals Subjected to Varying Stress–

Prediction of Fatigue Lives.” Preliminary Proceedings of the Chugoku-Shikoku District Meeting, The 

Japan Society of Mechanical Engineers, Nov., 1967, pp. 41–44 and The Rainflow Method in Fatigue. 

Ed. By Murakami, Y. Butterworth-Heinemann Ltd, Boston, ISBN 0 7506 0504 9. 1992. p. vii. 

7. Endo., T.; Mitsunaga, K.; Nakagawa, H.; and Ikeda, K.; “Fatigue of Metals Subjected to Varying 

Stress–Low Cycle, Middle Cycle Fatigue.” Preliminary Proceedings of the Chugoku-Shikoku District 

Meeting, The Japan Society of Mechanical Engineers, Nov., 1967, pp. 45–48 and The Rainflow 

Method in Fatigue. Ed. By Murakami, Y. Butterworth-Heinemann Ltd, Boston, ISBN 0 7506 0504 9. 

1992. p. xi. 

8. Gyekenyesi, J.Z., Murthy, P.L.N., Mital, S.K.; “NASALIFE -Material Life Prediction Program.” 

NASA Technical Memorandum TM-213886, 2005. 
NASA/TM—2005-213887 12

9. Kalluri, S.; Calomino, A.M.; and Brewer, D.N.; “Comparison of Elevated Temperature Tensile 

Properties and Fatigue Behavior of Two Variants of a Woven SiC/SiC Composite.” Presented at the 

American Ceramic Society 29th International Conference & Exposition On Advanced Ceramics & 

Composites, Cocoa Beach, FL, Jan. 23–28, 2005. 

10. Larson, F.R. and Miller, J., “A Time-Temperature Relationship for Rupture and Creep Stresses.” 

Trans. ASME, vol. 74, 1952, p. 765. 

11. Levine, S.R.; Calomino, A.M.; Ellis, J.R.; Halbig, M.C.; Mital, S.K.; Murthy, P.L.; Opila, E.J.; 

Thomas, D.J.; Ogbiji, L.U.; Virrilli, M.J.; “Ceramic Matrix Composites (CMC) Life Prediction 

Method Development.” NASA Technical Memorandum, NASA/TM—2000-210052, May, 2000. 

12. Lewinsohn, C.A.; Henager Jr., C.H.; Simonen, C.F.; Windisch Jr., C.F.; and Jones, R.H.: “The 

Relationship Between Interphase Oxidation and Time-Dependent Failure In SiCf/SiCm Composites.” 

Mechanical, Thermal and Environmental Testing and Performance of Ceramic Composites and 

Components, ASTM STP 1392, M.G. Jenkins, E. Lara-Curzio, and S.T. Gonczy Eds., American 

Society for Testing and Materials, West Conshohocken, Penn., 2000, pp. 229–241. 

13. Minor, M.A.; “Cumulative Damage in Fatigue.” Journal of Applied Mechanics, vol. 12, no. 3, 1945, 

pp. A159–A164. 

14. Mitsunaga, K. and Endo., T.; “Fatigue of Metals Subjected to Varying Stress–Fatigue Lives Under 

Random Loading.” Preliminary Proceedings of the Kyushu District Meeting, The Japan Society of 

Mechanical Engineers, Mar., 1968, pp. 37–40 and The Rainflow Method in Fatigue. Ed. By 

Murakami, Y. Butterworth-Heinemann Ltd, Boston, ISBN 0 7506 0504 9. 1992. p. xv. 

15. N&R Engineering, “Probabilistic Analysis of SiC/SiC CMC Vane In Support of Vane Sub-Element 

Fabrication and Testing: Final Report.” NASA Glenn Research Center, Jan. 2003. 

16. Ogbuji, L.U.J.T.; “Recent Developments in the Environmental Durability of SiC/SiC Composites.” 

NASA Contractor Report NASA/CR—2002-211687, Jul. 2002. 

17. Shaw, J. and Peddie, C.; “NASA Ultra Efficient Engine Technology Project Overview.” NASA 

Conference Proceedings 2004–212963, vol. 1, Sep. 2004. 

18. Verrilli, M.J.; NASA Glenn Research Center, Cleveland, OH, personal communication, 2004. 

19. Walker, K.; "The Effect of Stress Ratio during Crack Propagation and Fatigue for 2024-T3 and 7075-

T6 Aluminum," Effects of Environment and Complex Load History on Fatigue Life. ASTM STP 462, 

American Society for Testing and Materials, Conshohocken, PA, 1970, pp. 1–14. 
NASA/TM—2005-213887 13

Appendix A 

One of the NASALIFE input files is presented here. The material properties data were used for all the 

runs. The mission loads are at the end of the input deck. The loading conditions were varied as indicated 

in the tables of the results section of this work. The variables in the lines were defined in the NASALIFE 

technical and user manual by Gyekenyesi, et al. (2005). 

PRIN 

3 1 1 1 1 2 2 2 2 1 

LCF 

TITLE 

Example XLIFE lcf file 

ARAT 

 1 

TEMP FLIF SMAX FLAG 

70 1.0 32.2 000 

70 3.8 30. 000 

70 10.0 28.4 000 

70 1.0E+08 2. 000 

70 3.4E+08 0.002 000 

400 2.9 30. 000 

400 10.0 27.7 000 

400 1E+07 2. 000 

400 2.927E+07 0.002 000 

800 10.0 26.9 000 

800 1500000.0 2. 000 

1200 10.0 26. 000 

1200 1000000.0 2. 000 

1600 10.0 25.2 000 

1600 100000.0 2. 000 

1900 1.8 24.6 000 

1900 482.8 12. 000 

1900 6192.0 2. 000 

1900 1.005E+04 0.10 000 

1900 1.031E+04 0.002 000 

2200 3.0 22. 000 

2200 26.7 16. 000 

2200 110.4 13.4 000 

2200 439.8 9.4 000 

2200 900.0 2. 000 

2200 1.082E+03 0.10 000 

2200 1.092E+03 0.002 000 

EOF 

RUPT 

TITLE 
NASA/TM—2005-213887 14

Example XLIFE rupture table 

IRUP 

1 

TEMP RSTR RLIF FLAG 

70 28.4 9.90E+30 100 

70 1.0 1.00E+31 100 

400 27.7 9.66E+30 100 

400 1.0 9.90E+30 100 

1200 26.0 23.4 100 

1200 18.0 150. 100 

1200 14.0 400. 100 

1200 12.0 600. 100 

1200 1.0 640144. 100 

2400 23.5 4.7 100 

2400 15.0 90.8 100 

2400 12.5 211.5 100 

2400 10.0 515.2 100 

2400 1.0 11647. 100 

EOF 

MATL 

IOP1 IOP4 

 1 1 

TEMP M FLAG 

 60 .5 020 

1300 .5 000 

2200 .5 010 

TEMP E K N V FLAG 

 70 30400 28.4 0 .13 20 

 100 30400 28.4 0 .13 0 

 200 30400 28.1 0 .13 0 

 300 30400 27.9 0 .14 0 

 400 30400 27.7 0 .14 0 

 500 30400 27.5 0 .15 0 

 600 30400 27.3 0 .15 0 

 700 30400 27.1 0 .15 0 

 800 30400 26.9 0 .16 0 

 900 30400 26.7 0 .16 0 

 1000 30400 26.5 0 .16 0 

 1100 30400 26.3 0 .17 0 

 1200 30400 26.0 0 .17 0 

 1300 30400 25.8 0 .18 0 

 1400 30400 25.6 0 .18 0 

 1500 30400 25.4 0 .18 0 

 1600 30400 25.2 0 .19 0 

 1700 30400 25.0 0 .19 0 

 1800 30400 24.8 0 .19 0 

 1900 30400 24.6 0 .20 0 

 2000 29200 24.4 0 .20 0 

 2100 28000 24.2 0 .21 0 

 2200 26800 23.9 0 .21 0 
NASA/TM—2005-213887 15

 2300 25500 23.7 0 .21 0 

 2400 24300 23.5 0 .22 0 

 2500 23100 23.3 0 .22 10 

EOF 

TIME TEMP S11 

0 2000 0.00 

60 2000 14.1 

1260 2000 14.1705 

2460 2000 14.0295 

3660 2000 14.1705 

4860 2000 14.0295 

6060 2000 14.1705 

7260 2000 14.1 

7320 2000 0.00 
This publication is available from the NASA Center for AeroSpace Information, 301–621–0390.

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Technical Memorandum

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NASA TM—2005-213887

E–15258

WBS–22–714–30–20

21

Life Prediction for a CMC Component Using the NASALIFE Computer Code

NASALIFE; Life; Life prediction; Fatigue; Low cycle fatigue; LCF; Creep rupture; Ceramics; Ceramic matrix

composites; CMC; Silicon carbide; SiC/SiC; MI SiC/SiC; Probabilistic analysis; CMC vane; Cumulative

distribution function; Probability density function; Scatter; Weibull distribution; Strength; Proportional limit;

Design requirements

Unclassified -Unlimited

Subject Categories: 24 and 39

John Z. Gyekenyesi, Pappu L.N. Murthy, and Subodh K. Mital

The computer code, NASALIFE, was used to provide estimates for life of an SiC/SiC stator vane under varying

thermomechanical loading conditions. The primary intention of this effort is to show how the computer code

NASALIFE can be used to provide reasonable estimates of life for practical propulsion system components made of

advanced ceramic matrix composites (CMC). Simple loading conditions provided readily observable and acceptable

life predictions. Varying the loading conditions such that low cycle fatigue and creep were affected independently

provided expected trends in the results for life due to varying loads and life due to creep. Analysis was based on

idealized empirical data for the 9/99 Melt Infiltrated SiC fiber reinforced SiC.

John Z. Gyekenyesi, N&R Engineering and Management Services Corporation, 6659 Pearl Road, Suite 400,

Parma Heights, Ohio 44130; Pappu L.N. Murthy, NASA Glenn Research Center; and Subodh K. Mital, University

of Toledo, 2801 W. Bancroft Street, Toledo, Ohio 43606. Responsible person, Pappu L.N. Murthy, organization code

RSL, 216–433–3332.