January 2011 

NASA/TM-2011-216878 

 

Fracture Mechanics Analyses of the Slip-Side 

Joggle Regions of Wing-Leading-Edge Panels 

Ivatury S. Raju/NESC 

Langley Research Center, Hampton, Virginia 

Norman F. Knight, Jr. 

General Dynamics Information Technology, Chantilly, Virginia 

Kyongchan Song 

ATK Space Division, Langley Research Center, Hampton, Virginia 

Dawn R. Phillips 

Marshall Space Flight Center, Huntsville, Alabama 
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January 2011 

NASA/TM-2011-216878 

 

Fracture Mechanics Analyses of the Slip-Side 

Joggle Regions of Wing-Leading-Edge Panels 

Ivatury S. Raju/NESC 

Langley Research Center, Hampton, Virginia 

Norman F. Knight, Jr. 

General Dynamics Information Technology, Chantilly, Virginia 

Kyongchan Song 

ATK Space Division, Langley Research Center, Hampton, Virginia 

Dawn R. Phillips 

Marshall Space Flight Center, Huntsville, Alabama 
Available from: 

NASA Center for AeroSpace Information 

7115 Standard Drive 

Hanover, MD 21076-1320 

443-757-5802 

The use of trademarks or names of manufacturers in the report is for accurate reporting and does not 

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National Aeronautics and Space Administration. 
iii 

Table of Contents 

1.0 Introduction .......................................................................................................................... 1

2.0 Wing-Leading-Edge Configuration ..................................................................................... 1

3.0 Loading Environments ......................................................................................................... 3

4.0 Building-Block Approach and Analysis Models ................................................................. 4

4.1 Analysis Models ............................................................................................................... 5

5.0 Fracture Mechanics Analyses .............................................................................................. 9

6.0 Results and Discussion ...................................................................................................... 14

6.1 Interface Defects – Effect of Defect Location ............................................................... 14

6.2 Substrate Defects – Effect of Defect Length .................................................................. 17

7.0 Effect of Various Variables on the Defect Driving Force ................................................. 18

8.0 Concluding Remarks .......................................................................................................... 19

9.0 References .......................................................................................................................... 20

List of Figures 

Figure 1. Space Shuttle WLE configuration and geometry. ......................................................... 2

Figure 2. RCC spallation history. ................................................................................................. 3

Figure 3. Typical peak entry temperature distribution for the Space Shuttle WLE. .................... 4

Figure 4. Building-block approach. .............................................................................................. 5

Figure 5. Analysis models. ........................................................................................................... 7

Figure 7. TTT stress distribution for peak-heating condition for Panel 10 model. ...................... 8

Figure 8. TTT stress distribution for peak-heating condition for plane-strain model. ................. 9

Figure 9. Typical interface and substrate defects. ........................................................................ 9

Figure 10. Plane-strain model for fracture mechanics analyses. .................................................. 10

Figure 11. Defect modeling in the finite element mesh. .............................................................. 10

Figure 12. Deflection of interface and substrate defects. ............................................................. 11

Figure 13. VCCT scheme for 4-node (linear) 2D elements. ........................................................ 12

Figure 14. Typical curve fit of mixed-mode fracture data. .......................................................... 13

Figure 15. Normalized GT as a function of interface defect location – Entry peak heating. ....... 15

Figure 16. Normalized GT as a function of interface defect location – On-orbit cold. ................ 16

Figure 17. Fracture mechanics response for the left substrate defect tip fixed at Location –2. ... 17

Figure 18. Effect of various variables on defect driving force. .................................................... 18
iv 

Nomenclature 

B-K Benzeggagh and Kenane 

IML Inner Mold Line 

max-Q Maximum Dynamic Pressure Condition 

OML Outer Mold Line 

OV Orbiter Vehicle 

RCC Reinforced Carbon-Carbon 

STS Space Transportation System 

TTT Through-the-Thickness 

VCCT Virtual Crack Closure Technique 

WLE Wing Leading Edge 
Abstract 

The Space Shuttle wing-leading edge consists of panels that are made 

of reinforced carbon-carbon. Coating spallation was observed near the 

slip-side region of the panels that experience extreme heating. To 

understand this phenomenon, a root-cause investigation was conducted. 

As part of that investigation, fracture mechanics analyses of the slip-side 

joggle regions of the hot panels were conducted. This paper presents an 

overview of the fracture mechanics analyses. 

1.0 Introduction 

Each Space Shuttle Orbiter wing is comprised of 22 leading edge panels that are made of 

reinforced carbon-carbon (RCC). These panels are part of the thermal protection system that 

protects the wings from extreme heating that takes place during entry into Earth’s atmosphere 

[1]. On some of the panels that experience extreme heating, spallation of RCC coating was 

observed in the slip-side regions of the panels. To understand the reasons for this coating 

spallation anomaly, a root-cause investigation was conducted. The root-cause investigation team 

consisted of researchers from various disciplines such as structures, materials, non-destructive 

evaluation, aerothermal, flight hardware, arc-jet testing, coupon testing, vibro-acoustic testing, 

aeroloading testing, photo-micrographic investigation, etc. This paper describes an overview of 

the structural and fracture mechanics analyses that were conducted. 

The paper is organized as follows: First, the Shuttle wing-leading-edge (WLE) configuration is 

presented, and the loading environments to which the Shuttle is subjected are described. Next, 

the building-block approach and analysis models that were used for the stress and fracture 

mechanics analyses are described. Then, the fracture mechanics analyses and methods used to 

characterize defects in the panels are presented. Finally, analysis results and findings are 

discussed. 

2.0 Wing-Leading-Edge Configuration 

The Space Shuttle WLE configuration and geometry are presented in Figure 1. Between 

adjacent panels, T-seals are installed to close or cover the gap. Figure 1(a) shows the Space 

Shuttle WLE, and Figures 1(b) and 1(c) show close-up and cross-sectional views, respectively, 

of Panels 9 and 10 and T-seal 10. The regions of the panels that have S-shaped curvature are 

called “joggles” (see Figure 1(c)). When T-seal 10 is installed, it is locked to a joggle on Panel 

9; that side of the panel is called the “lock side” and is shown in Figure 1(c). In addition, T-seal 

10 floats over a joggle on Panel 10; that side of the panel is called the “slip side” and is also 

shown in Figure 1(c). The region between the lock side and slip side of each panel is called the 

acreage region. The outer mold line (OML) and inner mold line (IML) are the configuration 

boundaries on the outer and inner surfaces of the WLE, respectively. The chord direction runs 

around the WLE and is defined in Figure 1(b), while the span direction runs across the WLE 

panels and T-seals and is defined in Figure 1(c). The RCC material contains a substrate region 

and two conversion coating regions containing craze cracks, as shown in Figure 1(d). Note that 

these craze cracks form in the coating during the cool-down phase of the conversion process of 
2 

fabrication. Then a sealant is applied to both the OML and IML surfaces, filling the craze 

cracks. The conversion-coating regions are commonly referred to as coating layers. 

 a) Space Shuttle WLE. b) WLE Panels 9 and 10 and T-seal 10. 

 c) Cross-sectional view of Panels 9 and 10 d) Cross-sectional views of slip-side joggle. 

 and T-seal 10. 

Figure 1. Space Shuttle WLE configuration and geometry. 

After two different Space Shuttle landings (Missions STS-102 and STS-103) at the NASA 

Kennedy Space Center, small areas of the RCC coating were observed to be missing from the 

WLE apex near the top of the joggle in the slip-side region, as shown in Figure 2. The two 

spallation events occurred (a) after space transportation system (STS)-103 on orbiter vehicle 

(OV)-103 Panel 8L and (b) after STS-102 on OV-103 Panel 10L. After the Shuttle landed 

following the return-to-flight mission (STS-114), an infrared thermography indication (from 

non-destructive evaluation (NDE)) was observed on OV-103 Panel 8R (Figure 2(c)), indicating a 

loss of coating strength. In addition, the spallation event on the nose cap on OV-105 was 

observed during refurbishment (Figure 2(d)). The root-cause investigation was subsequently 

conducted to determine what factors and scenarios contribute to these spallation events (e.g., see 

References 1–6). 
3 

a) Post STS-103: OV-103 Panel 8L. b) Post STS-102: OV-103 Panel 10L. 

 c) Post STS-114 themography: d) Observed during refurbishment: 

 OV-103 Panel 8R. OV-105 nose cap. 

Figure 2. RCC spallation history. 

3.0 Loading Environments 

During a mission, the Space Shuttle is subjected to four distinct loading environments [1]. 

During lift-off and ascent, aerodynamic loads act on the Shuttle; while the Shuttle is in orbit, it is 

subjected to extreme cold temperatures; during entry, the Shuttle experiences peak heating; and 

during descent and landing, different aerodynamic loads act on the Shuttle. 

Loads analyses showed that the lift-off and ascent loading bounds the loading on descent and 

landing. Thus, fracture mechanics analyses were performed using a bounding pressure load over 

an assumed defect for the lift-off and ascent loading condition. The bounding pressure load 

occurs at the maximum dynamic pressure condition (max-Q) about 80 seconds after lift-off. The 

defect driving forces were calculated for the bounding pressure load, and they were nearly zero. 

Hence, both the lift-off and ascent conditions and descent and landing conditions were dismissed 

as being benign contributors to the spallation root cause [6]. 

During entry, the Shuttle WLE panels experience entry heating that depends in part on the entry 

trajectory. For a representative entry trajectory from the International Space Station, the peakheating entry temperature along the slip-side joggle regions for each of the 22 WLE panels is 

indicated in Figure 3. Peak heating reaches temperatures up to about 1650°C (3000°F) and 

occurs in the hot panels, Panels 8, 9, and 10. For the on-orbit heating/cooling thermal cycle, the 

temperature range is –130°C ≤ T ≤ 90°C (±200°F), with –130°C (–200°F) being the cold 
4 

condition. In this paper, fracture mechanics analyses are presented for these two bounding 

thermal loading conditions (i.e., on-orbit cold and entry peak heating) that were performed in 

support of the root-cause investigation. 

a) Definitions of zones. 

b) RCC WLE slip-side inboard joggle non-catalytic peak temperatures. 

Figure 3. Typical peak entry temperature distribution for the Space Shuttle WLE. 

4.0 Building-Block Approach and Analysis Models 

A building-block analysis approach begins with basic elements and builds in complexity in a 

systematic and progressive manner [2]. Such an approach permits each step in the process to be 

verified and its influence on the overall response determined. The hierarchy of the buildingblock approach is presented in Figure 4. First, the problem, objectives, products, resources, 

schedule, and stakeholders are defined. Second, the structural configuration, loads, materials, 

boundary/interface conditions, and tools to be used to solve to problem are identified. Next, 

analysis models, such as finite-element models of the individual components, are created. The 

material modeling procedure, boundary conditions, and other approximations are assigned to the 

individual component models, and the models are solved and verified for accuracy. The process 

at the individual component level is an iterative process that is repeated until the results can be 

verified by comparison to reference solutions or test data. Then, the individual component 

models are assembled, incorporating component interface conditions, different design 

configurations, and various load cases. The assembled models are solved, and the results are 

assessed. The process at the assembly level is also an iterative process that is repeated until 

confidence in the results can be demonstrated and advocated by the analysis team itself. Finally, 
5 

the results are presented to external reviewers and the stakeholders, engineering reports are 

written, and the entire effort is subjected to rigorous peer review. If necessary, the process may 

be repeated from any of the building-block steps. 

Problemdefinition, objectives, products,resources,schedule,stakeholders

Configuration, loads,materials, boundary/interface conditions,tools

Individual componentmodeling andmeshing

Materialpropertiestraceability and material modeling

Assemble multiple components&solve

Componentinterface conditions

Results presentation

Boundary conditions&approximations

Peerreview

Verification of individualmodels

Engineering report

Design configuration&cases

Assessmentofresults

Iterate

Figure 4. Building-block approach. 

This building-block analysis approach provides different entry points for analysts, stakeholders, 

and external reviewers. Analysis assumptions can be challenged and revised, analysis findings 

can be compared to observed behavior, and bounding responses can be determined; and the 

analysis end products have technical rigor and their applicability limits defined. This type of 

building-block approach was applied to the global structural and fracture mechanics analysis 

effort as part of the RCC spallation root-cause investigation. All finite element analyses reported 

in this paper were performed using the ABAQUSTM1

 commercial software [7]. 

4.1 Analysis Models 

The finite element models used for the analysis effort are presented in Figure 5. First, an 

integrated model was considered (see Figure 5(a)) including Panels 9 and 10 (i.e., two of the hot 

panels), their adjoining T-seal, and attachment hardware. The peak-heating temperature 

distribution was applied to this integrated model. Negligible interaction was observed between 

 

1

 ABAQUSTM is a registered trademark of Dassault Systèmes. 
6 

Panel 10 and T-seal 10 for the peak-heating conditions. These results demonstrate that a single 

panel can be analyzed alone for the entry thermal conditions. 

Next, the Panel 10 model shown in Figure 5(b) was considered because its slip-side joggle region 

experienced the highest entry temperature of the integrated model. The model had 3D elements 

all along the chord in the joggle regions and shell elements in the acreage regions. The peak 

entry temperature distribution shown in Figure 6 was applied to the model. The resulting 

through-the-thickness (TTT) stress distribution in the substrate is presented in Figure 7. (To 

show the TTT stresses clearly, the coating elements are removed in Figure 7.) As shown in 

Figure 7(c), an elevated stress value is observed all along the slip-side joggle in the local region 

near the panel apex (i.e., the hot region) and is nearly constant all along the chord direction (i.e., 

the gradient of stress in the chord direction is nearly zero). These results demonstrate that any 

slice perpendicular to the chord direction could be taken and used in a plane-strain analysis of the 

slip-side joggle [5]. 

Finally, the plane-strain model of the Panel 10 slip-side joggle region (shown in Figure 5(c)) was 

considered. For the stress analysis, the finite element model did not include craze cracks or 

defects. Because the temperatures in the local region near the panel apex are nearly constant (see 

Figure 6(c)), the maximum peak entry temperature was applied to the entire plane-strain model. 

The resulting TTT stress distribution in the substrate (the coating elements have been removed) 

is presented in Figure 8. 

A failure criterion (or criteria) that is (or are) based on linear-elastic stress analysis results alone 

would suggest that wide-spread spallation would occur on many WLE panels on every mission, 

while the flight history shows that wide-spread spallation did not occur. Note that the observed 

spallation events shown in Figure 2 from flight history were limited to a single event at a single 

location on a single panel on each occurrence. However, the stress analysis results are used to 

indicate locations where potential subsurface defects may contribute to a spallation anomaly and 

prompted the following fracture mechanics analyses. 
7 

a) Integrated model. 

b) Panel 10 model. 

c) Plane-strain model. 

Figure 5. Analysis models. 
8 

 a) Panel 10 model. b) Slip-side joggle region. c) Local region near apex. 

Figure 6. Peak entry temperature distribution for Panel 10. 

 a) Panel 10 model. b) Slip-side joggle region. c) Local region near apex. 

Figure 7. TTT stress distribution for peak-heating condition for Panel 10 model. 
9 

11.5 MPa (1664 psi)

‐2.7 MPa (‐387 psi)

TTT Stress

Figure 8. TTT stress distribution for peak-heating condition for plane-strain model. 

5.0 Fracture Mechanics Analyses 

The objective of the fracture mechanics analyses is to evaluate the defect driving forces, which 

are characterized by the strain energy release rates, and determine if defects can become unstable 

for each of the loading conditions. 

The fracture mechanics analyses were performed using the plane-strain model. For the fracture 

mechanics analyses, subsurface defects were introduced at the maximum stress location (see 

Figure 8) and perpendicular to the maximum tensile TTT stress. Three types of defects were 

considered: interface defects (those along the coating-substrate interface), substrate defects 

(those completely within the substrate), and combined interface-substrate defects (those in both 

regions). Typical interface and substrate defects are presented schematically in Figure 9. In this 

paper, only interface and substrate defects are discussed. 

a) Interface defect. 

b) Substrate defect. 

Figure 9. Typical interface and substrate defects. 

The finite element model and terminology are presented in Figure 10 (mesh is excluded for 

clarity). The panel acreage is to the left of the model. (Note that the orientation of Figure 10 is 

opposite to Figure 8. The orientation of Figure 10 is used in the remainder of the paper.) The 
10 

substrate and coating materials are shown in yellow and blue, respectively. The defects are 

inserted by introducing coincident nodes, as demonstrated in Figure 11. To insert a subsurface 

defect between elements A-D and E-H, nodes 2–4 are duplicated to create new nodes 2'–4'. The 

connectivity of elements A-D is left unaltered, while the connectivity of elements E-H is 

modified to use the new nodes 2'–4'. These new nodes are initially coincident with the original 

nodes; i.e., both sets of nodes have identical coordinates. Nodes 1 and 5 define the defect tips. 

Similarly, the craze cracks are modeled explicitly by using coincident nodes in the finite element 

mesh. The craze cracks are oriented normal to the joggle surface, as shown in Figure 10. The 

defects are placed symmetrically and directly underneath a craze crack as specified by the Defect 

Location Index. Location 0 corresponds to the craze crack at the joggle shoulder (see Figure 

5(c)). The craze cracks are numbered consecutively from Location 0. Negative-numbered craze 

cracks are consecutive towards the panel acreage, and positive-numbered craze cracks are 

consecutive down the joggle. The left and right tips of the defect are towards the acreage and 

towards the joggle, respectively. 

‐1 0

+1

+2

+3

Interface Defect

Defect Location

Index

Craze Cracks

OML

IML

Substrate

Coating

Left Tip

(towards acreage)

Right Tip

(towardsjoggle)

‐2

Figure 10. Plane-strain model for fracture mechanics analyses. 

1 2 3 4 5

2' 3' 4'

A B C D

Initially coincident nodes

E F G

H

Initially coincident

defectsurfaces

Figure 11. Defect modeling in the finite element mesh. 

Interface and substrate defects interact with the craze cracks differently for elevated temperature 

conditions. As shown in Figure 9(a), the craze-crack surfaces above an interface defect are 

completely disconnected from the substrate material. As heating occurs, the two craze-crack 

edges may come into contact and try to slide past each other, as shown in Figure 12(a). For this 
11 

reason, friction is included in the finite element model to account for the craze-crack edge 

interaction. In contrast, as shown in Figure 9(b), the craze-crack surfaces above a substrate 

defect are connected to substrate material. When heating occurs, the two craze-crack edges may 

come into contact, but they are restrained from sliding past each other, as shown in Figure 12(b). 

a) Interface defect with craze-crack edge interaction. 

b) Substrate defect. 

Figure 12. Deflection of interface and substrate defects. 
12 

The capability of a structure with an embedded defect can be described using the total defect 

driving force GT calculated using fracture mechanics analyses; i.e., the problem is no longer a 

stress analysis problem. Once the defect driving force is computed, it can be compared to 

material toughness values for the different modes of fracture (Mode I, II, and/or their mixity for 

plane strain). 

In this paper, the G calculation for the defects is performed via the Virtual Crack Closure 

Technique (VCCT). In finite element analysis, the VCCT uses the nodal forces at the node at the 

defect tip and the displacements at the nodes behind the defect tip [8–10]. In the analyses in this 

paper, the finite element models were developed such that the element edge length (a) of the 

elements ahead and behind the defect tip was of equal size. Figure 13 shows local modeling at 

the defect tip in a 2D analysis. The defect tip is represented by node i. Elements I and J

contribute to the internal forces, Zi and Xi, at node i. The internal forces at node i and the relative 

displacements between nodes j and k, wj,k = wj – wk and uj,k = uj – uk, are used to evaluate the 

individual G component values as 

a

wZ

G kji



  2

,

I

, a

uX

G kji



  2

,

II

 (1) 

In Equation 1, GI and GII are the Mode I and Mode II strain energy release rates, the G values. 

The total strain energy release rate GT can be evaluated using 

  GGG IIIT (2) 

X, u

Z, w

a

J

J-1 I-1

j

k i I

a

Figure 13. VCCT scheme for 4-node (linear) 2D elements. 
13 

In this paper, a somewhat conservative comparison is made by comparing the total strain energy 

release rate GT to the Mode I fracture toughness GIc, which is the smallest of the fracture 

toughness values for the different modes. Unstable defect growth (i.e., defect is likely to grow in 

a catastrophic manner) may occur when GT is greater than or equal to GIc; i.e., GT/GIc ≥ 1. 

Alternately, a mixed-mode fracture criterion is used to determine whether the defect is stable or 

unstable. In this paper, the 2D plane-strain criterion given by Benzeggagh and Kenane (B-K) 

[11] is used as the mixed-mode fracture criterion. This B-K criterion is expressed as 

1 T 

R

G

 (3) 

where GT represents the total defect driving force, given in Equation 2, and R represents the 

resistance to the total defect driving force. The resistance R is also referred to as Gc, the critical 

fracture toughness based on test data. The value for the resistance R is given by 

    IIII

T

II

IIII m ccc

m

ccc GGG

G

G GGGR  

















 (4) 

where GIc and GIIc represent the Mode I and Mode II material toughness values, respectively,  = 

GII/GT represents the mode-mixity ratio, and m represents the exponent of  used to fit the 

mixed-mode test data as shown systematically in Figure 14. When the ratio of GT to R is less 

than unity, the defect is said to be stable, and defect growth is unlikely. When the ratio equals or 

exceeds unity, the defect is said to be unstable, and catastrophic defect growth is likely. 

GII/GT

0 1

Gc

GIc

GIIc

 ~0.2

Mode I dominated

Linear m=1 curve fit

Mixed

mode

dominated

Determined as a curve fit 

to mixed-mode data, m>1

Figure 14. Typical curve fit of mixed-mode fracture data. 
14 

For a mixed-mode response, the B-K resistance R given in Equation 4 is dependent on the value 

of the exponent m as indicated in Figure 14. This exponent is typically determined using a curve 

fit for mixed-mode fracture data obtained from material testing. For a value of m = 1, a linear 

response like the green dashed line in Figure 14 is obtained, which is a non-conservative 

criterion. Mixed-mode fracture tests of polymeric matrix composites show that the B-K criterion 

using m = 2 or 3 appears to fit mixed-mode fracture data accurately. A typical mixed-mode 

fracture criterion curve is shown as red in Figure 14. For mode-mixity ratios less than , the 

response is dominated by Mode I behavior. 

Single cantilever beam and end-notch flexure test configurations were used in RCC fracture 

testing to determine GIc and GIIc. In the absence of further test data, a value of two was selected 

as the value of the exponent m. In addition, the B-basis fracture toughness values are used in 

evaluating Equation 4 as they provide a smaller value for R, which is conservative. 

6.0 Results and Discussion 

In this section, only representative results in terms of defect driving forces, the GT values, for 

entry peak heating and the on-orbit cold condition are presented. Both interface and substrate 

defects are considered. More detailed analyses and results can be found in References 1–6. 

6.1 Interface Defects – Effect of Defect Location 

Interface defects were studied by assuming a constant-length defect at different craze-crack 

locations. Both entry peak heating and the on-orbit cold condition were considered. The results 

for the normalized GT values are presented as the ratio GT/GIc. Values of GT/GIc less than unity 

indicate that the defect is stable and unlikely to grow catastrophically. Values greater than unity 

require further study. Such values could mean the defect is unstable and likely to grow 

catastrophically, or they could mean the mode of fracture is not Mode-I dominated. Hence, the 

total and individual component G values need to be calculated and assessed. 

Entry Peak Heating. For entry peak heating, the craze-crack surfaces come into contact and slide 

past each other, as shown in Figure 15(a), and friction along the craze-crack edges influences the 

defect deformations. Figures 15(b) and 15(c) present the normalized GT values for two values of 

the coefficient of friction  for various locations of the defect along the joggle. For each 

location, the interface defect is centered beneath the craze crack. Both the left and right tip 

normalized GT values are plotted as a function of the defect location. For  = 0 (Figure 15(b)), 

the left tip values are considerably smaller than the right tip values for all defect locations 

considered. For both  = 0 and  = 0.4, the maximum normalized GT occurs at the shoulder of 

the joggle, Location 0. However, the normalized GT for  = 0 is considerably higher than the 

normalized GT for  = 0.4 indicating that the craze-crack edge interaction is an important 

variable to the defect driving force. 

The individual mode G values were also examined. For both  = 0 and  = 0.4, the GI values are 

larger than the GII values. This result suggests that the defect opening for entry peak heating is 

dominated by Mode I. 
15 

a) Interface defect at craze-crack Location +1 and deformed configurations for interface defects 

at varying craze-crack locations (deformations scaled by a factor of 20). 

b) Normalized GT for μ = 0. 

c) Normalized GT for μ = 0.4. 

Figure 15. Normalized GT as a function of interface defect location – Entry peak heating. 
16 

On-Orbit Cold. For the on-orbit cold condition, the craze-crack surfaces displace away from 

each other, as shown in Figure 16(a), and the craze-crack edges do not interact and do not 

influence the defect deformations. Figure 16(b) presents the normalized GT values as a function 

of the interface defect location. The maximum normalized GT occurs in the acreage area (for the 

left tip at Locations –1 and 0). As the defect is moved into the joggle region, the normalized GT

values decrease and reach a plateau. Both the left and right tips have nearly the same normalized 

GT values except at Location 0, where the left tip (towards the acreage) has the higher value. 

This difference in behavior at Location 0 is due to the geometry; at Location 0, half of the defect 

is in the acreage, and half is down the joggle curve. 

The individual mode G values were also examined. The GII values are larger than the GI values, 

suggesting that the defect opening for on-orbit cold is dominated by Mode II and indicating that 

the comparison of GT to GIc in Figure 16(b) may be overly conservative. 

a) Interface defect at craze-crack Location +1 and deformed configurations for interface defects 

at varying craze-crack locations (deformations scaled by a factor of 20). 

b) Normalized GT. 

Figure 16. Normalized GT as a function of interface defect location – On-orbit cold. 
17 

6.2 Substrate Defects – Effect of Defect Length 

The fracture mechanics response for substrate defects is examined by fixing the left tip of the 

defect at different acreage locations and then extending the defect along the joggle (i.e., changing 

the defect size or length). The notation ‘A→B’ is used in this paper to indicate that a subsurface 

defect exists between Location A and Location B. The left and right defect tips are located at 

Locations A and B, respectively, as illustrated in Figure 17(a). 

Entry Peak Heating. To study “very long” substrate defects for entry peak heating, the left defect 

tip is placed at Location –2, and the location of the right defect tip is varied along the joggle up 

to Location +4. The fracture response is given in Figure 17(b) for the left tip and in Figure 17(c) 

for the right tip. The blue curves represent the total defect driving force, or GT, normalized by 

GIc, and the green dashed curves represent the resistance to the defect driving force, or R, 

evaluated using Equation 4 and normalized by GIc. For the initial case (i.e., case –2→0), both 

defect tips are stable. For longer defects (i.e., cases –2→+1 and –2→+2), the left tip is stable, 

but the right tip is unstable. For the case –2→+3, the normalized R at the right tip becomes 

larger than the normalized GT suggesting a return to a stable region. For the “very long” defect 

(i.e., case –2→+4), the left tip normalized GT exceeds the normalized R and hence the left tip 

becomes unstable. 

a) Various defects with the left tip fixed at Location –2; case –2→+4 shown. 

 b) Left tip. c) Right tip. 

Figure 17. Fracture mechanics response for the left substrate defect tip fixed at Location –2. 
18 

7.0 Effect of Various Variables on the Defect Driving Force 

As mentioned previously, there are many variables that contribute to the defect driving force, the 

GT value, for this application. Some of these variables are identified in Figure 18, and the effect 

of each of these variables is discussed qualitatively. 

Defect

initiation site Coating/substrate

transition zone

Craze crack

edge interaction

Stress‐free

temperature

Extent of

fiber bridging

Extent of

ply convolutions

Defect

growth model

3D effects

GT

GT

GT

Nonlinear

‐ response

GT

GT GT

Defectsize

GT

Craze crack

orientation

GT

Figure 18. Effect of various variables on defect driving force. 

The effect of the variable in the green cloud (i.e., defect initiation site) was determined by 

studying the photomicrographs like the one shown in the center of this figure. 

The effect of the variables in the orange clouds was determined [1 – 6] through finite element 

and fracture mechanics analyses like those presented in this paper. For the coating/substrate 

transition zone, the initial analyses assumed a sharp coating/substrate interface and yielded 

higher GT values. Because a sharp interface does not exist in reality, analyses were conducted 

using different size zones where the material properties were transitioned from substrate to 

coating, yielding lower GT values. For defect size, larger defects yield higher GT values. For 3D 

effects, part-through defects yield lower GT values than those computed using plane-strain 

analysis (i.e., simulating a through defect). These results demonstrate that plane strain is the 

bounding case. For the extent of fiber bridging, including fiber bridging yields lower GT values. 

For the extent of ply convolutions, including knuckles, voids, etc. yields higher GT values. For 

stress-free temperature, using a stress-free temperature that increases the T yields higher GT

values. For craze-crack edge interaction, including friction in the analysis affects the results; 

high values for the coefficient of friction yield lower GT values, while simulating a smooth edge 
19 

using zero friction yields higher GT values. For craze-crack orientation, craze-cracks oriented 

normal to the joggle yield the lowest GT values. 

The effect of the variables in the gray clouds (i.e., defect growth model and nonlinear stressstrain response) could not be quantified because enough RCC material was not available to 

perform comprehensive testing, in addition, testing at elevated temperatures is much more 

complex than at room temperature. 

8.0 Concluding Remarks 

The Space Shuttle wing-leading edge consists of panels that are made of reinforced carboncarbon. On some panels that experience extreme heating, spallation of coating was observed in 

the slip-side joggle regions of the panels. Global structural and local fracture mechanics 

analyses were performed on these panels as a part of the root cause investigation of this coating 

spallation anomaly. The global structural analyses showed minimal interaction between adjacent 

panels and T-seals that bridge the gaps between the panels for entry thermal conditions. A 

bounding temperature distribution was applied to a representative panel, and the resulting stress 

distribution was examined. For this loading condition, the through-the-thickness normal stresses 

showed negligible variation in the chord direction and increased values in the vicinity of the slipside joggle shoulder. As such, a representative span-wise slice on the panel was taken, and the 

cross section was analyzed using plane-strain analysis. In the plane-strain models, both interface 

and substrate defects were introduced. Various size defects were considered. Plane-strain finite 

element analyses were conducted for entry peak heating and on-orbit cold conditions. Defect 

driving forces in terms of the strain energy release rates were used to characterize the defects. 

This paper presents some of the fracture mechanics analyses and results. 

Various parameters that affect the driving force of defects present in the Space Shuttle wingleading-edge joggle regions have been identified. In this paper, the effects of defect initiation 

site and defect size were briefly presented. For the fracture mechanics analyses, interface, 

substrate, and combined defects were introduced into the 2D plane-strain finite element models 

of the slip-side joggle region. 3D analyses showed that the plane-strain case is the bounding 

case. The defect driving forces were computed using the Virtual Crack Closure Technique, and 

the Benzeggagh-Kenane mixed-mode fracture criterion was considered. Parameters that affect 

the defect driving force include effects of the coating/substrate transition region, fiber bridging, 

ply convolutions, stress-free temperature, and craze-crack orientation and edge interaction. The 

fracture mechanics analysis results were ultimately used to define tests and test methods and 

helped in understanding possible factors and scenarios that contribute to the RCC spallation 

anomaly. 
20 

9.0 References 

1. Knight, N. F., Jr.; Raju, I. S.; Song, K.; and Phillips, D. R.: “Global and Fracture Mechanics 

Analyses for the Space Shuttle Orbiter Wing-Leading-Edge Panels.” Proceedings of the 

International Conference on Computational and Experimental Engineering and Sciences ’10 

Conference, March 29–April 1, 2010, Las Vegas, Nevada, Paper No. ICCES1020091224163. 

2. Knight, N. F., Jr.; Raju, I. S.; Song, K.; and Phillips, D. R.: “Building-Block Analysis 

Strategy Applied to Space Shuttle Orbiter Wing-Leading-Edge Panels.” Proceedings of the 

International Conference on Computational and Experimental Engineering and Sciences ’10 

Conference, March 29–April 1, 2010, Las Vegas, Nevada, Paper No. ICCES1020091224164. 

3. Phillips, D. R.; Raju, I. S.; Knight, N. F., Jr.; and Song, K.: “Influence of Craze Cracks on 

Reinforced Carbon-Carbon Material Modeling.” Proceedings of the International 

Conference on Computational and Experimental Engineering and Sciences ’10 Conference, 

March 29–April 1, 2010, Las Vegas, Nevada, Paper No. ICCES1020091224165. 

4. Knight, N. F., Jr.; Song, K.; Raju, I. S.; and Phillips, D. R.: “Fracture Response of a 

Composite Joggle Structure Including Ply Convolutions.” Proceedings of the International 

Conference on Computational and Experimental Engineering and Sciences ’10 Conference, 

March 29–April 1, 2010, Las Vegas, Nevada, Paper No. ICCES1020091224166. 

5. Knight, N. F.; Jr., Song, K.; and Raju, I. S.: “Space Shuttle Orbiter Wing-Leading-Edge 

Panel Thermo-Mechanical Analysis for Entry Conditions.” Proceedings of the 

AIAA/ASME/ASCE/AHS/ASC 51st Structures, Structural Dynamics, and Materials 

Conference, April 12–15, 2010, Orlando, Florida, Paper No. AIAA-2010-2688. 

6. Raju, I. S.; Phillips, D. R.; Knight, N. F., Jr.; and Song, K.: “Fracture Mechanics Analyses of 

Reinforced Carbon-Carbon Wing-Leading Edge Panels.” Proceedings of the 

AIAA/ASME/ASCE/AHS/ASC 51st Structures, Structural Dynamics, and Materials 

Conference, April 12–15, 2010, Orlando, Florida, Paper No. AIAA-2010-2689. 

7. Anon.: ABAQUS Analysis User’s Manual: Volumes I – VI, Version 6.7. Dassault Systèmes 

Simulia Corp., Providence, Rhode Island, 2007. 

8. Rybicki, E. F.; and Kanninen, M. F.: “A Finite Element Calculation of Stress Intensity 

Factors by a Modified Crack Closure Integral.” Engineering Fracture Mechanics, Vol. 9, No. 

4, pp. 931–938, 1977. 

9. Raju, I. S.: “Calculation of Strain-Energy Release Rates with Higher Order and Singular 

Finite Elements,.” Engineering Fracture Mechanics, Vol. 28, pp. 251–274, 1987. 

10. Krueger, R.: “The Virtual Crack Closure Technique: History, Approach and Applications.” 

Applied Mechanics Reviews, Vol. 57, No. 2, pp. 109–143, 2004. 

11. Benzeggagh, M. L.; and Kenane, M.: “Measurement of Mixed-Mode Delamination Fracture 

Toughness of Unidirectional Glass/Epoxy Composites with Mixed-Mode Bending 

Apparatus.” Composites Science and Technology, Vol. 56, No. 4, pp. 439-449, 1996. 
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Fracture Mechanics Analyses of the Slip-Side Joggle Regions of 

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Raju, Ivatury S.; Knight, Norman F., Jr.; Song, Kyongchan; 

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14. ABSTRACT

The Space Shuttle wing-leading edge consists of panels that are made of reinforced carbon-carbon. Coating spallation was 

observed near the slip-side region of the panels that experience extreme heating. To understand this phenomenon, a root-cause 

investigation was conducted. As part of that investigation, fracture mechanics analyses of the slip-side joggle regions of the

hot panels were conducted. This paper presents an overview of the fracture mechanics analyses.

15. SUBJECT TERMS

Reinforced carbon-carbon, fracture mechanics; stress-free temperature; wing leading edge; joggle region; thermal loading; 

subsurface defects; finite element analysis; strain energy release rates

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