Jiangang G. Sun

Argonne National Laboratory, Argonne, Illinois

Michael J. Verrilli

Glenn Research Center, Cleveland, Ohio

Robert R. Stephan

Pratt & Whitney, East Hartford, Connecticut

Terry R. Barnett

Southern Research Institute, Birmingham, Alabama

Greg C. Ojard

United Technologies Research Center, East Hartford, Connecticut

Nondestructive Evaluation of Ceramic Matrix

Composite Combustor Components

NASA/TM—2003-212014

April 2003
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Jiangang G. Sun

Argonne National Laboratory, Argonne, Illinois

Michael J. Verrilli

Glenn Research Center, Cleveland, Ohio

Robert R. Stephan

Pratt & Whitney, East Hartford, Connecticut

Terry R. Barnett

Southern Research Institute, Birmingham, Alabama

Greg C. Ojard

United Technologies Research Center, East Hartford, Connecticut

Nondestructive Evaluation of Ceramic Matrix

Composite Combustor Components

NASA/TM—2003-212014

April 2003

National Aeronautics and

Space Administration

Glenn Research Center

Prepared for the

29th Annual Review of Progress in Quantitative Nondestructive Evaluation (QNDE)

sponsored by the Quantitative Nondestructive Evaluation Programs

Bellingham, Washington, July 14–19, 2002
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NASA/TM—2003-212014 1 

Nondestructive Evaluation of Ceramic Matrix Composite 

Combustor Components

Jiangang G. Sun 

Argonne National Laboratory 

Argonne, Illinois 60439 

Michael J. Verrilli 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Robert R. Stephan 

Pratt & Whitney 

East Hartford, Connecticut 06108 

Terry R. Barnett 

Southern Research Institute 

Birmingham, Alabama 35211 

Greg C. Ojard

United Technologies Research Center 

East Hartford, Connecticut 06108 

Summary 

 Combustor liners fabricated from a SiC/SiC composite (silicon carbide fibers in a silicon carbide 

matrix) were nondestructively interrogated before and after combustion rig testing by x-ray, ultrasonic, 

and thermographic techniques. In addition, mechanical test results were obtained from witness coupons, 

representing the as-manufactured liners, and from coupons machined from the components after 

combustion exposure. Thermography indications correlated with reduced material properties obtained 

after rig testing. The thermography indications in the SiC/SiC liners were delaminations and damaged 

fiber tows, as determined through microstructural examinations. 

Introduction 

 The pursuit of lower emissions and higher performance from gas turbine engines requires the 

development of innovative concepts, such as the use of advanced materials for key engine components. 

One such engine component is the combustor, where liners fabricated from a ceramic matrix composite, 

silicon carbide fibers in a silicon carbide matrix (SiC/SiC), are under evaluation. SiC/SiC composite 

technology has progressed from the fabrication of simple coupons to subelements and presently to actual 

components (refs. 1 to 4). This last area posed a significant challenge to nondestructive evaluation (NDE) 

techniques because of the complex geometries involved. 

 To evaluate SiC/SiC components in a combustion environment, the rich-burn, quick-quench, leanburn (RQL) sector rig was designed and fabricated. The purpose of the sector rig testing was to 

demonstrate the structural durability of the SiC/SiC liners in a combustion environment where stresses, 
NASA/TM—2003-212014 2 

temperatures, and pressures would as accurately as possible reflect the operating conditions found in a 

turbine engine. 

 X-ray, ultrasonic, and thermographic techniques were used to inspect the as-manufactured SiC/SiC 

components and the liners after combustion testing. The purpose of this study was to characterize the 

postexposure condition of one SiC/SiC combustor liner part through NDE evaluation via thermography 

and mechanical property measurement. 

Experimental 

Material and Component 

Combustor liners were manufactured by Honeywell Advanced Composites from a SiC/SiC composite 

developed under the NASA Enabling Propulsion Materials program. A chemical-vapor-infiltrated, slurrycast, and melt-infiltrated SiC matrix was reinforced with Sylramic SiC fibers. The fiber tows were woven 

into 5-harness satin weave cloth. More details on the SiC/SiC material can be found in reference 5. 

 The sector rig liner set consisted of 6 component geometries and a total of 28 liners (fig. 1). For this 

study, one of the six components, the lean-zone inside-diameter liner (LZID) was examined. The LZIDs 

are curved plates, each covering about a 30° arc of the total 60° sector, as shown in figure 1. A twelveply-thick leading edge reinforces the attachment hole region. The rest of the LZID is six plies thick. 

LZIDs containing 34- and 40-percent fiber volume were tested. 

 Flexure coupons were machined from the as-manufactured and exposed sector parts to obtain 

mechanical properties. The as-manufactured parts were supplied with extra material that was machined 

into test coupons. The exposed parts were machined to obtain test coupons of the same geometry as that 

used for the witness testing. Three-point flexure testing was conducted at 815 °C. Reference 6 gives more 

details on the mechanical test procedures. 

Combustion Testing 

 The Preheated Combustor and Materials Test Facility (PCMTF) at NASA Glenn Research Center was 

used to conduct the RQL sector rig testing. Air temperature (T3), pressure (P3), and flow rate (W3) were 

varied to approximate the anticipated service cycle of the High Speed Civil Transport (HSCT) engine 

combustor (ref. 7). All aspects of an HSCT flight operation, such as takeoff, climb, cruise, and decent, are 

included in the 230-min test cycle. 

 The combustion testing was periodically interrupted to remove some of the SiC/SiC liners for 

postexposure analysis. After 260 hr of operation, the rig was disassembled and all SiC/SiC combustor 

liners were removed. 

NDE Techniques 

 As-manufactured and combustion-exposed SiC/SiC components were inspected using x-ray film 

radiographic, ultrasonic, and thermographic techniques. Radiography is sensitive to volumetric features, 

such as large voids and gross porosity or thickness variations, but cannot detect planar (laminar) features, 

such as delaminations in ceramic matrix composites. Ultrasonic techniques have a high sensitivity for the 

detection of delaminations, but a quantitative interpretation of the data is difficult as they are affected by 

many experimental setup parameters. Thermography techniques are also very sensitive to delaminations 

and when used in a through-transmission setup to determine thermal diffusivity, which is a material 
NASA/TM—2003-212014 3 

property, the data can be quantitatively analyzed to relate with the detailed structures of defects (ref. 8). In 

this study, only the thermography results will be discussed. 

 A transient through-transmission (through-thickness) thermography method was used on the tested 

liners. Two high-energy flash lamps generate a heat pulse that travels through the thickness of the parts 

(fig. 2). The raw data consist of multiple-layered images acquired at 120 Hz. A total of 40 frames was 

typically suitable to capture the temperature rise. The analysis technique was based on that given by 

Parker (ref. 9). The front surface temperature rise is given by 

T Lt Q

CL

n

L

t n

n

b g b g bg , exp = +− −

F

H

G I

K

J

L

N

M

M

O

Q

P

P =

∞

ρ ∑ π 12 1 1 α

2 2

2

1

where T is the temperature, L is the liner thickness, t is the time elapsed, Q is the amount of heat absorbed 

by the liner from the flash lamps, α is the thermal diffusivity, ρ is the material density, and C is the 

specific heat of the material. Using the time to reach half the maximum detected temperature (t1/2), the 

thermal diffusivity at each pixel can be calculated by 

α

π = 137 2

2

2

1 2

.

/

L

t

b g

 At regions of planar defects, such as delaminations, the material surface temperature is lower than 

that of the undamaged composite and will take a longer t1/2 to reach the equilibrium part temperature (note 

that the part will eventually reach an equilibrium constant temperature if sufficient time is given). 

Therefore, from equation (2), the regions with damaged material will have lower thermal diffusivities. 

 The thermal diffusivity calculated from equation (2) is proportional to the thickness squared. Because 

it is impossible to provide thickness values for all pixels of an imaged part, it is common to use a 

representative thickness (typically the averaged thickness) for the calculation of the diffusivity image for 

the entire part. This method has some shortcomings when thickness variation is large. For example, in the 

case of a uniform material of constant diffusivity, equation (2) dictates that for a pixel with a thickness 

twice that of a second pixel, the half-rise time t1/2 at the first pixel will be four times longer than that of 

the second. However, if we choose to use the thickness of the second pixel for calculating diffusivities at 

all pixels, the calculated diffusivity at the first pixel would appear to be four times lower than its real 

diffusivity. Similarly, the diffusivity calculated from equation (2) appears to be higher for a thinner pixel. 

Therefore, care should be taken when interpreting measured diffusivity images. Note that this equation is 

exact when a test part has a uniform thickness. 

Results 

 An example of thermography results obtained for three LZIDs after 115 hr of sector rig testing is 

shown in figure 3. Part 216−1 contains 34-percent fiber volume fraction, and the other two contain 

40 percent. The thickness of the middle sections of the liners was used to calculate the diffusivity values 

in the image. The apparent diffusivities are lower at the strips with fastener holes where the material is 

thicker and are higher at the left edges where the material is thinner. In addition to these diffusivity 

variations due to a thickness change, regions with considerably lower diffusivities can be observed in 

all liners, especially for 216−1. 

 To investigate the origin of local lower diffusivity values, a LZID was sectioned and polished. 

Figure 4 shows the microstructure of a cross section taken from 216−1. The line on the thermography 

image in figure 3 shows the location of the cross section. Damaged SiC fiber tows and delamination 
NASA/TM—2003-212014 4 

cracks can be seen. The diffusivity reduction shown in figure 3 is related to the damage (i.e., 

delamination) of the ceramic matrix composite material due to the sector testing. For a single 

delamination, Sun (ref. 8) showed that the diffusivity reduction could be used to determine the (air gap) 

thickness of the delamination. 

 Thermal diffusivity data for another LZID with 34-percent fiber volume, which was exposed in the 

combustion environment for 145 hr, is shown in figure 5. Thermal diffusivity values are plotted in 

figure 5(a) along seven horizontal lines passing through several interesting regions in the thermal 

diffusivity image of 216−2. Although there is a thickness variation near both vertical edges of the liner, 

the thickness is uniform in the middle section and this value is used in the diffusivity calculation. In this 

middle section, the measured diffusivity is uniform in undamaged regions (e.g., line 7). Reduced 

diffusivity representing damage of varying severity is shown in other regions (lines 1 to 6). Figure 5(b) 

plots the width-averaged diffusivity along the length for the regions where four coupons were machined 

from the part to obtain flexure strength. It is apparent that coupons 2 to 4 contained large damaged areas 

with significant thermal diffusivity reductions and that coupon 1 had only slight damage near the region 

of thickness transition at the right. It is interesting to note that fracture occurred near the locations where 

the lowest diffusivities were observed for coupons 2 to 4. 

 Flexure strengths obtained from these coupons are compared with the thermal diffusivity in table I. 

The postexposure strength obtained from these coupons decreased with the decreasing ratio of the 

coupon-averaged thermal diffusivity to the mean thermal diffusivity (α/αm), revealing the correlation 

between the thermal diffusivity data and strength. 

 The normalized flexure strength versus combustion exposure time for coupons machined from all 

LZIDs tested in the sector rig is shown in figure 6. The normalized flexure strength was calculated as the 

ratio of the strength of a coupon machined from a tested LZID to the average as-fabricated strength 

obtained from the LZID witness coupons for the same part. The strength decreases with increasing 

exposure time for the parts with 34-percent fiber volume, whereas the strength of the 40-percent fiber 

volume decreased after 115 hr of exposure. Also, coupons machined from component regions with 

thermography indications had lower strengths than coupons from regions with no NDE indications. As a 

function of exposure time, the strength data sets for coupons with and without NDE indications follow 

the same trends.

Summary 

 1. Combustor liners fabricated from SiC/SiC were tested in an aircraft engine environment using the 

rich-burn, quick-quench, lean-burn (RQL) sector rig for up to 260 hr. SiC/SiC liners were periodically 

removed during the rig operation to conduct postexposure analyses. 

 2. Degradation of the SiC/SiC material occurred and was in the form of delaminations and damaged 

fiber tows for the combustor liner parts examined, namely, the lean-zone inside-diameter liners (LZIDs). 

 3. Thermography was effective in imaging this damage in the SiC/SiC liners, whereas radiography 

and ultrasonic inspections were less sensitive. 

 4. Strength was measured by testing coupons machined from combustion-exposed parts and 

compared with witness (as-manufactured) data. For parts with 40-percent fiber volume, the strength 

decreased after 115 hr. However, exposure of only 10 hr decreased the strength of parts with 34-percent 

fiber volume. 

 5. For the exposed parts, coupons machined from regions that had thermography indications had 

lower strengths than coupons machined from regions that had no indications. 

 6. Local thermal diffusivity correlated with measured coupon strength for exposed SiC/SiC 

combustor liners. 
NASA/TM—2003-212014 5 

References 

1. Brewer, D.; Ojard, G.; and Gibler, M.: Ceramic Matrix Composite Combustor Liner Rig Test. ASME 

Paper TE00CER03−03, 2000. 

2. Miriyala, Narendernath; Fahme, Anthony; and van Roode, Mark: Ceramic Stationary Gas Turbine 

ProgramCombustor Liner Development Summary. ASME 2001−GT−0512, 2001. 

3. Corman, Gregory, et al.: Rig and Gas Turbine Engine Testing of MI−CMC Combustor and Shroud 

Components. ASME 2001−GT−0593, 2001. 

4. Igashira, Ken-Ichiroh, et al.: Development of the Advanced Combustor Liner Composed of 

CMC/GMC Hybrid Composite Material. ASME 2001−GT−0511, 2001. 

5. Brewer, David: HSR/EPM Combustor Materials Development Program. Mater. Sci. Eng., vol. A261, 

1999, pp. 284−291. 

6. Verrilli, Michael J., et al.: Evaluation of Post-Exposure Properties of SiC/SiC Combustor Liners 

Tested in the RQL Sector Rig. NASA TM2002-211380, 2002. 

http://gltrs.grc.nasa.gov/cgi-bin/GLTRS/browse.pl?2002/TM-2002-211380.html 

Accessed Jan. 6, 2003. 

7. Shaw, Robert J.; Koops, Leigh; and Hines, Richard: Progress Toward Meeting the Propulsion 

Technology Challenges for a 21st Century High-Speed Civil Transport. NASA TM−113161 

(ISABE−97−7045), 1997. http://gltrs.grc.nasa.gov/cgi-bin/GLTRS/browse.pl?1997/TM-113161.html

Accessed Jan. 7, 2003. 

8. Sun, J.G.: Analysis of Quantitative Measurement of Defect by Pulsed Thermal Imaging. Paper 

presented at the Review of Progress in Quantitative NDE, Brunswick, ME, 2001. 

9. Parker, W.J., et al.: Flash Method of Determining Thermal Diffusivity, Heat Capacity, and Thermal 

Conductivity. J. Appl. Phys., vol. 32, no. 9, 1961, pp. 1679−1684. 

TABLE I.COMPARISON OF THERMAL DIFFUSIVITY AND COUPON STRENGTH 

FOR LEAN-ZONE INSIDE-DIAMETER LINER (LZID) 216−2 AFTER 

145 HR OF TESTING IN COMBUSTION ENVIRONMENT 

Coupon Ratio of coupon-averaged thermal 

diffusivity to mean thermal diffusivity, 

α/αm

Posttest 815 °C 

flexure strength, 

MPa 

1 1.0 262 

2 .6 229 

3 .56 207 

4 .47 131 
NASA/TM—2003-212014 6

Figure 1.—Rich-burn, quick-quench, lean-burn (RQL) sector rig and images of all 

 SiC/SiC parts.

Rich-zone liner

Lean-zone bulkhead

heatshield

Lean-zone 

inside-diameter liners

Lean-zone 

outside-diameter 

liners

Lean-transitionzone liners

Lean-zone 

rig sidewall

Temperature

Lamps Time

Figure 2.—Through-transmission (through-thickness) thermography

 test configuration. 

Infrared

camera

Test 

sample
NASA/TM—2003-212014 7

Delamination

Figure 4.—Microstructure of LZID 216–1 after 115 hr of combustion

 exposure (black line in fig. 3 shows location of this section). Note

 the arrows that indicate damaged tows.

Damaged fiber bundles

Figure 3.—Thermography of three SiC/SiC liners after 115 hr

 of sector rig testing. Part 216–1 contains 34-percent fiber 

 volume fraction; parts 208–2 and 218–2 contain 40-percent

 fiber volume fraction. 
NASA/TM—2003-212014 8

40

60

80

100

120

140

160

180

200

1

2

3

4

5

6

7

0 10 20 30 40 50 60 70 80

Thermal diffusivity, arbitrary units

Distance from left edge, pixels

Thicker

region 

Reference

level

Thinner 

 region

1

2

3

4

5

6

7

(a)

Line

Line

Thicker

region 

40

60

80

100

120

140

160

180

200

4

3

2

1

0 10 20 30 40 50 60 70 80

Distance from left edge, pixels

Location of lowest

mean diffusivity

Thermal diffusivity, arbitrary units

4

3

2

1

Location of

fracture

Figure 5.—Thermal diffusivity data for LZID 216–2 after 145 hr of combustion exposure. (a) Line 

 plots at seven part locations. Note: line 7 is reference level. (b) Average values for regions of 

 coupons machined for strength measurement. 

(b)

Coupon Coupon
NASA/TM—2003-212014 9

Figure 6.—Normalized flexure strength data obtained at 815 °C as function

 of exposure time for coupons machined from LZIDs. (a) Data for coupons

 from 34-percent fiber volume parts. (b) Data for coupons from 40-percent

 fiber volume parts. 

0.0

0.2

0.4

0.6

0.8

1.0

1.2

1.4

0 20 40 60 80 100 120 140 160

Exposure time in combustion environment, hr

Exposure time in combustion environment, hr

Normalized flexure strength

(as-fabricated strength/posttest strength)

0.0

1.0

Normalized flexure strength

(as-fabricated strength/posttest strength)

0.2

0.4

0.6

0.8

1.2

1.4

1.6

0 50 100 150 200 250 300

(b)

(a) Coupons from regions with no NDE indications

Coupons from regions with NDE indications

Coupons from regions with no NDE indications

Coupons from regions with NDE indications
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E–13702

WBS–22–714–04–15

15

Nondestructive Evaluation of Ceramic Matrix Composite Combustor Components

Jiangang G. Sun, Michael J. Verrilli, Robert R. Stephan, Terry R. Barnett, and

Greg C. Ojard

Nondestructive tests; Ceramic matrix composites; Combustion; Silicon carbide;

Composite structures; Mechanical properties

Unclassified -Unlimited

Subject Categories: 24 and 35 Distribution: Nonstandard

Prepared for the 29th Annual Review of Progress in Quantitative Nondestructive Evaluation (QNDE) sponsored by the Quantitative

Nondestructive Evaluation Programs, Bellingham, Washington, July 14–19, 2002. Jiangang G. Sun, Argonne National Laboratory,

Argonne, Illinois 60439; Michael J. Verrilli, NASA Glenn Research Center; Robert R. Stephan, Pratt & Whitney, East Hartford,

Connecticut 06108; Terry R. Barnett, Southern Research Institute, Birmingham, Alabama 35211; Greg C. Ojard, United Technologies

Research Center, East Hartford, Connecticut 06108. Responsible person, Michael J. Verrilli, organization code 5160, 216–433–3337.

Combustor liners fabricated from a SiC/SiC composite (silicon carbide fibers in a silicon carbide matrix) were

nondestructively interrogated before and after combustion rig testing by x-ray, ultrasonic, and thermographic techniques. In addition, mechanical test results were obtained from witness coupons, representing the as-manufactured

liners, and from coupons machined from the components after combustion exposure. Thermography indications

correlated with reduced material properties obtained after rig testing. The thermography indications in the SiC/SiC

liners were delaminations and damaged fiber tows, as determined through microstructural examinations.