THE DAMAGE-TOLERANCE BEHAVIOR OF INTEGRALLY STIFFENED ...abebrot.com/papers/IACAS 2008.pdf ·...

Presented to the Israel Annual Conference on Aerospace Sciences, 2008

1

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


3

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.


4

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

50

100

150

200

250

300

350

400

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

20

40

60

80

100

120

140

160

180

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

20

40

60

80

100

120

140

160

180

200

0 10000 20000 30000 40000 50000 60000

Life [cycles]

Cra

ck S

ize [m

m]

Test Results

StressCheck/Nasgro Results


9

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

50

100

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%.


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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


12

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.

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