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Presented to the Israel Annual Conference on Aerospace Sciences, 2008
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THE DAMAGE-TOLERANCE BEHAVIOR OF INTEGRALLY STIFFENED
METALLIC STRUCTURES
A. Brot, Y. Peleg-Wolfin, I. Kressel and Z. Yosef
Engineering Division
Israel Aerospace Industries
e-mail: [email protected]
ABSTRACT
Israel Aerospace Industries (IAI) has studied the damage-tolerance behavior of integrally
stiffened metallic structures as part of an on-going international project called DaToN
(Innovative Fatigue and Damage Tolerance Methods for the Application of New Structural
Concepts), which is partially funded by the European Commission (EC). IAI has performed
both analytical and experimental studies of integrally stiffened metallic structures, in the
framework of this EC project, including crack growth testing of integrally stiffened panels
reinforced by composite materials. IAI also investigated and performed a crack growth
analysis for two configurations. Finite-element models were built using StressCheck
p-version software. StressCheck uses the J-integral method for the computation of mode 1
and mode 2 stress intensity factors using linear elastic fracture mechanics. From the
StressCheck stress-intensity results, a crack growth analysis was performed using NASGRO
ver. 5 software.
INTRODUCTION
Many aircraft structures that previously were assembled from several mechanically
fastened built-up parts are now manufactured as integral structures produced by high-speed
machining by laser-beam welding or by friction-stir welding. This change is being
implemented in order to reduce the manufacturing costs of aircraft. It also carries a bonus in
that many fastener holes, which are prone to fatigue cracking, have been eliminated.
On the other hand, integrally stiffened structures will not have the same crack-arrest
capability as built-up structures. On a typical built-up stiffened panel, cracks developing in
the skin will be contained in their bay, since the cracks cannot "jump" to the stiffeners. If the
design is changed to be integrally stiffened, a skin crack approaching the integral stiffener
can "climb" the stiffener and then proceed into the next bay and eventually resulting in a total
failure.
Totally new manufacturing techniques are generally tested in many ways before they are
applied. The following examples show that totally new designs bear a certain risk, which
must be considered.
In the 1940s, the US built a series of 2500 “Liberty Ships” using an innovative all-welded
design. 145 of these ships broke-in-two and 700 experienced serious failures [1]. Apparently,
the effect of cold sea temperatures on the weld integrity was not considered during the
design, and insufficient testing was performed. A photo of a Liberty Ship is shown in
Figure 1.
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Figure 1: All-Welded Liberty Ship
In the 1950s, the de Havilland Aircraft Company of the UK introduced the first modern
airliner, the Comet I, as is shown in Figure 2.
Figure 2: de Havilland Comet I Airliner
The Comet I was the first high-altitude, jet-propelled airliner and had many innovative
features. The design cabin pressure of the Comet I was approximately double that of previous
airliners. It entered airline service, after static and fatigue testing, in 1951.
In January 1954, a Comet disintegrated in flight near Elba. It had sustained only 1286
pressurized flights before failure.
In April 1954, another Comet disintegrated near Stromboli. This aircraft had experienced
only 903 pressurized flights. Fatigue testing confirmed that fatigue of the pressurized cabin
was the primary cause of failure [2].
What went wrong? The aircraft had, in fact, been designed and tested to the cabin pressure
requirements, and no problems were evident.
Several studies were performed on the Comet I accidents and two main deficiencies were
uncovered:
1. The fuselage skins, stringers and frames were designed to be similar to the design of
previous airliners. This means that the stringers were continuous, and the fuselage
frames had cutouts to accommodate the stringers. As a result, the structure was found
to be grossly deficient in its ability to arrest longitudinal cracks. This means that any
longitudinal crack that reached a critical size, would continue to grow unstably until
the entire fuselage had failed [2] [5]. (Modern airliners are designed with special
design features so that cracks of 1 or 2 bays will arrest at the adjacent fuselage
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frame.) The Comet I was not designed accordingly, so the fuselage disintegrated in
flight. Figure 3 shows the wreckage of the fuselage of the Comet. (The primary
cracks are shown propagating well beyond the fastener lines of several fuselage
frames, without arrest.)
Figure 3: Comet I Fuselage Wreckage, Showing the Lack of Crack Arrest
2. The second deficiency involved the static and fatigue testing that was performed on
the airframe. The first cracks, during the fatigue test, were detected after 16,000
pressure cycles [4]. Unfortunately, in order to save costs, the same fuselage that had
been tested until the ultimate design load, was later used for the fatigue test [3][4].
Residual compressive stresses, arising from the ultimate loading, retarded the
development of the fatigue cracks. As a result, there were no indications from the
fatigue test that the fuselage design was inadequate.
What these two examples have in common is the principle that innovation often leads to
new problems. In the case of using innovative techniques to manufacture integral metallic
aircraft structures, we must not lose sight that the degree of innovation can lead to specific
problems that were not as acute while using conventional manufacturing processes. The
problem of crack-arrest, that was the primary cause of the Comet disaster, can once more
become a problem when dealing with one-piece (integral) structures. Therefore, it is
imperative to perform sufficient research, by analysis and testing, in order to be certain where
the structural integrity boundaries exist. And it is important to learn how to optimize
structural configurations in order to provide sufficient margins.
Israel Aerospace Industries (IAI) has studied the damage-tolerance behavior of integrally
stiffened metallic structures as part of an on-going international project called DaToN
(Innovative Fatigue and Damage Tolerance Methods for the Application of New Structural
Concepts), which is partially funded by the European Commission (EC). IAI has performed
both analytical and experimental studies of integrally stiffened metallic structures, in the
framework of this EC project.
Presented to the Israel Annual Conference on Aerospace Sciences, 2008
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TESTING OF INTEGRALLY STIFFENED PANELS
As a part of the DaToN project, IAI performed crack growth testing of integrally stiffened
panels. Three panels, all of a two-stringer configuration, were crack growth tested, as is
shown in Figure 4. Table 1 summarizes the results, while the detailed results are shown in the
next section together with the analytical predictions. An additional panel was tested with
carbon-epoxy reinforcements, as is described later in this paper.
Figure 4: A Two-Stringer Panel in the Test Rig and a Panel after Failure
Table 1: Summary of Results of Two-Stringer Panel Testing
CRACK GROWTH ANALYSIS OF INTEGRALLY STIFFENED PANELS
As a part of DaToN project, IAI investigated and performed a crack growth analysis for
two configurations:
1. A 7475-T7451 seven-stringer aluminum panel (Figure 5) with an initially broken
stringer and a skin crack of ±15 mm, under a skin stress of 95 MPa and R=0.1. (This
configuration was not tested by IAI.)
Specimen
No
Initial
Crack [mm]
Max Stress
[Mpa]R
Life
[cycles]
1 15 80 0.1 79450
2 13.9 110 0.5 99250
3 33.5 70 0.1 53500
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2. A 2024-T351 two-stringer aluminium panel (Figure 6), whose testing was described
in the previous section.
The analytical results were compared to the measured test results.
A finite-element model (FEM) was built using StressCheck p-version software [6].
StressCheck uses the J-integral method for the computation of mode 1 and mode 2 stress
intensity factors (SIF) using linear elastic fracture mechanics. The method is super-
convergent. This mean that the error in the SIF converges to zero much faster than the error
in the energy norm, as the number of degrees-of-freedom is increased [6].
In the vicinity of the crack tip, the solution to the linear problem is singular and the stress
values are infinite. Whether or not a crack will propagate, and at what rate, depends on the
energy available to drive crack extension. The available energy is characterized by the SIF
and the J-integral, which depend on the geometry of the body, the configuration of the crack,
the boundary conditions and the loading.
In order to reduce the number of elements, a symmetric model was meshed. The
calibration of each model was based on strain gage readings obtained during testing. The SIF
was calculated at the crack tip for various crack lengths using this method, as is shown in
Figure 7.
Two phases of the crack growth were analyzed: The first phase consisted of crack growth
on the skin before reaching the integral stiffener. The second phase included two crack path
one on the skin and other on the stiffener. For this phase, a matrix of combinations of skin
crack size and stringer crack size was built. From the StressCheck stress-intensity results, a
crack growth analysis was performed using NASGRO ver. 5 software [7].
Figure 5: Seven-Stringer Panel, StressCheck Model
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Figure 6: Two-Stringer Panel, StressCheck Model
The StressCheck (p-version) FEM of the seven-stringer panel was developed, as is
described in Figure 5, using approximately 270 three-dimensional elements (half-symmetric
model). A 7th order polynomial was used to describe the stress distribution of the elements
and stress-intensities at the crack-tips were computed.
Figure 7: Extraction of Stress-Intensities at the Crack Tips Using the J-Integral Method
The test results were compared to StressCheck results using NASGRO 5.01 for crack
growth analysis (using data table models DT01 & DT03). The results, for the seven-stringer
panel, are shown in Figure 8. It should be noted, that the initial conditions for the seven-
stringer test were a totally broken center stringer and a skin crack of ±15 mm.
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Figure 8: Comparison of the Analysis and Test Results for the Seven-Stringer Panel
For the two-stringer panels, a 3D StressCheck model was built using approximately 100
p-version elements in a quarter-symmetric model, as is shown in Figure 6. An 8th order
polynomial was used to describe the stress distribution of the elements. A geometric
nonlinear analysis was performed and the model loading distribution and crack growth rate
were adjusted, based on the test calibration results of the first test panel.
The test results were compared to StressCheck results using NASGRO 5.01 for crack
growth analysis (again using data table models DT01 & DT03). The results, for the two-
stringer panels, are shown in Figures 9 −11.
Figure 9: Comparison of the Analysis and Test Results for the 1st Two-Stringer Panel
Two-Stringer, First Specimen 80 MPa (R=0.1)
0
20
40
60
80
100
120
140
160
180
200
0 10000 20000 30000 40000 50000 60000 70000 80000Life [cycles]
Cra
ck S
ize [m
m]
Test Results
StressCheck/ NASGRO
Results
Seven Stringer, Test Vs. Analysis
0
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0 2000 4000 6000 8000 10000 12000 14000 16000 18000 20000 22000 24000 26000
Cycles
Mean C
rack S
ize [m
m]
Seven StringerTest - Crack on Skin
Seven Stringe Test - Crack on Stringer
Analysis - Crack on Skin
Analysis -Crack on Stringer
Stringers 3&5
Stringers 2&6
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Figure 10: Comparison of the Analysis and Test Results for the 2nd Two-Stringer Panel
Figure 11: Comparison of the Analysis and Test Results for the 3rd Two-Stringer Panel
From the seven-stringer panel test results, it is clear that there was a marked slow-down in
measured crack growth rate as the cracks encountered stringers 3 and 5. On the other hand,
for the two-stringer panel tests, this phenomenon was virtually absent. This difference can be
Two-Stringer, Second Specimen 110MPa (R=0.5)
0
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0 10000 20000 30000 40000 50000 60000 70000 80000 90000 100000
Life [cycles]
Cra
ck S
ize [m
m]
Test Results
StressCheck/ Nasgro Results
Two-Stringer, Third Specimen 70MPa (R=0.1)
0
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0 10000 20000 30000 40000 50000 60000
Life [cycles]
Cra
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StressCheck/Nasgro Results
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attributed to the heavy stringers used for the seven-stringer panel and the relatively light
stringers for the two-stringer panels. In any case, the effect of the stringers in increasing the
crack growth life of the panels was quite small, and probably did not exceed 20%. This
confirmed the expected belief that integrally stiffened panels do not exhibit good damage-
tolerance performance.
From the seven-stringer test results, it is clear that the analysis did not account sufficiently
well for the slow-down that occurred when the crack approached the junction between the
skin and the stringer. One of the reasons of this behavior can be the lack of analytical results
within the transition area between the skin and the stringer. The stress-intensity analysis
considered the case where the skin crack just reached the stringer. The next case calculated
the stress-intensities for a skin crack just beyond the stringer, with a crack in the stringer.
Linear interpolation was used in the NASGRO runs for the transition area between these two
cases. This matter needs to be re-evaluated in order to improve the analytical results.
EFFECTIVENESS OF THE INTEGRALLY STIFFENED PANEL
The seven-stringer had relatively heavy stiffeners whose total cross-section area was 54%
of the sheet area. A comparative analysis was performed between the crack growth life of the
stiffened panel and that of an unstiffened panel, both under the identical stress level.
Figure 12 shows the result of this study. To our great surprise, the analysis showed that the
stiffened panel is expected to have only 62% of the life of the unstiffened panel. The reason
for this surprising result is that every time a stiffener breaks, large forces are transferred to
the sheet which accelerates the crack growth in the panel. Since the test began with the
middle stringer broken, the crack in the stiffened panel grew faster than the unstiffened panel
from the start of the test.
Crack Growth of the Seven-Stringer Panel vs. an Unstiffened
Panel
0
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150
200
250
300
350
400
0 5000 10000 15000 20000 25000 30000
life [cycles]
cra
ck s
ize [m
m]
Stiffened Panel
Unstiffened Panel
Figure 12: Comparison of the Expected Crack Growth of a Stiffened and Unstiffened
Panel
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REINFORCING THE INTEGRAL PANEL USING CARBON-EPOXY STRIPS
In recent years, there has been much discussion of the advantages of a "hybrid" stiffened
panel which has composite materials bonded to the aluminium [8] [9] [10]. The composite
material reinforces the aluminium panel and serves to bridge any cracks that may develop in
the aluminium panel. This bridging effect was proven during the last 30 years in many
composite bounded repairs of aging aircraft [11].
In order to improve the performance of the two-stringer integral panel, two 35mm wide
strips of AS4/3502 carbon-epoxy were bonded to the panels, as shown in Figure 13. Each
strip consisted of three layers of carbon-epoxy material. The purpose of the strips is to reduce
the stress-intensity of a crack that grows under it, thereby increasing the crack growth life of
the panel. On the other hand, detrimental residual thermal stresses will exist, induced by the
thermal expansion coefficient mismatch between the composite material and the metal
substrate. These residual stresses may be significant because of the difference between the
curing temperature 120°C and the operating temperature at altitude. Therefore, a crack
growth test was needed to determine the net gain in crack growth life.
Figure 13: Test Panel with Carbon-Epoxy Reinforcing Strips, before and after Failure
The first hybrid panel was tested at room temperature, under a 7% higher loading than
what was used for the first unreinforced panel. The purpose of the 7% increase was to
compensate for the additional EA cross-section contribution of the reinforcing strip.
Figure 14 shows the crack growth test results of the hybrid panel. The results clearly show
that the hybrid panel had a significantly slower crack growth rate. Figure 14 shows that the
crack growth life was doubled from the unreinforced panel to the reinforced panel, in spite of
the increase in the loading by 7%.
Presented to the Israel Annual Conference on Aerospace Sciences, 2008
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The crack propagation rate of the reinforced panel seems to be constant, almost up to
failure. This phenomenon is with good agreement with the Rose Model [11] that predicts a
constant stress-intensity factor under a bonded composite patch.
It should be noted that no debonds between the composite strips and the metal substrate,
or delaminations between the layers, were observed up to failure.
Two additional panels will be tested at -30°C and -60°C, in order to simulate aircraft
operation at altitude. The detrimental thermal stresses in the aluminium panel can be expected
to be larger than those encountered during the room temperature test.
Figure 14: Comparison of the Crack Growth of the Carbon-Epoxy Reinforced Panel to
the First Unreinforced Two-Stringer Panel. (The loading on the reinforced panel was
increased by 7% relative to the unreinforced panel.)
FUTURE ACTIVITIES TO BE PERFORMED
An analytical study will be performed to determine the optimum ratio of stiffener to plate
cross-section area which will yield the most favorable damage-tolerance performance. Two
additional two-stringer reinforced panels will be tested at -30°C and -60°C, in order to
simulate aircraft operation at altitude.
PRINCIPAL CONCLUSIONS REACHED
1. Analysis and testing has shown that integrally stiffened panels do not generally have
good damage-tolerance performance. If the integral stiffeners are "light", they offer
almost no resistance to the crack growth. If the integral stiffeners are "very heavy",
the failure of any stiffener greatly accelerates the crack growth in the panel.
2. Introducing a hybrid design, with composite material strips bonded to the panel, has
the potential of significantly enhancing the damage-tolerance performance of the
integrally stiffened panel.
3. The optimum ratio of stiffener to plate cross-section area needs to be determined for
optimizing the damage-tolerance performance of a stiffened panel.
0
20
40
60
80
100
120
140
0 20000 40000 60000 80000 100000 120000 140000 160000Life [cycles]
Cra
ck S
ize [m
m]
Unreinforced Panel
Reinforced Panel
Beginning of Reinforcment
Stiffener Center-line
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REFERENCES
[1] Broek, D., Elementary Engineering Fracture Mechanics, Martinus Nijhoff Publishers,
Third edition, 1983.
[2] Swift, T., “Damage Tolerance in Pressurized Fuselages”, Plantema Memorial Lecture,
14th Symposium of the International Committee on Aeronautical Fatigue (ICAF), Ottawa,
Canada, 1987.
[3] Schijve, J. “Fatigue of Aircraft Materials and Materials”, International Journal of
Fatigue, Vol. 15, No. 1, 1994.
[4] Schijve, J. “Case Histories in Fatigue”, USAF Aircraft Structural Integrity Program
(ASIP) Conference, San Antonio, USA, 1999.
[5] Blom, A. “Fatigue Science and Engineering – Achievements and Challenges”, Plantema
Memorial Lecture, 21st Symposium of the International Committee on Aeronautical
Fatigue (ICAF), Toulouse, France, 2001.
[6] StressCheck ver. 7 Master Guide, ESRD Inc., 2005.
[7] NASGRO ver. 5.01 Reference Manual, Southwest Research Institute, February 2007.
[8] Zhang, X. et al, "Improving Fail-Safety of Aircraft Integral Structures through the Use of
Bonded Crack Retarders", Proceedings of the 24th ICAF Symposium, Naples, May 2007.
[9] Heinimann, M., "Validation of Advanced Metallic Hybrid Concept with Improved
Damage Tolerance Capabilities for Next Generation Lower Wing and Fuselage
Applications", Proceedings of the 24th ICAF Symposium, Naples, May 2007.
[10] Baker, A. A., "Repair of Cracked or Defective Metallic Aircraft Components with
Advanced Fibre Composites – an Overview of Australian Work", J. of Composite
Structures, vol. 2, 1984.
[11] Baker, A., Rose, F. and Jones, R. "Advances in the Bonded Composite Repair of Metallic
Aircraft Structure", Elsevier, 2002.