1

THE EFFECT OF DENTS IN FUSELAGE STRUCTURES ON FATIGUE AND STATIC STABILITY

Cornelis Guijt, Daniel Hill, Justin Rausch, Scott Fawaz

Center for Aircraft Structural Life Extension, Department of Engineering Mechanics, United States Air Force Academy

Key Words: Dents, Fuselage, Fatigue, Shear, Compression, Stability Abstract: A significant workload for large transport aircraft is related to fuselage dent repair. This research program addresses both the effect of dents and reformed dents on durability (fatigue) and static stability. In fatigue, tension-tension loading is used, and both compression and shear loading are considered for static stability. Different dent shapes and depths in unstiffened flat sheets were subjected to fatigue loading. Reformed dents were fatigue tested as well. The current status is that although dents do decrease fatigue life, and reforming dents does somewhat restore the loss in fatigue life; none of the dents tested would shorten the fatigue life of the skin below the required life for the airframe. For fatigue loaded structures with dents, this opens the possibility to greatly reduce the workload and prevent unnecessary repairs. For static stability, representative flat stiffened aircraft panels were loaded in compression and shear. Based on the effective width of the skin in compression, which is small for the configuration tested, the dents have no effect on structural stability and do not need to be repaired. The outcome of this program is that a significant amount of time and money can be saved by avoiding unnecessary structural repairs for the typical fuselage configuration tested. These results can be applicable to similar fuselage structures of other large transports.

2

INTRODUCTION Dents in fuselages are a widespread maintenance concern. A field survey performed by Boeing1 revealed significant amounts of damage as shown in Figure 1.

Figure 1: Overall results of Boeing field survey on 28 aircraft

Currently the Technical Order (T.O.) for this aircraft has strict limits on structurally allowable dents and dent locations. Typically in a skin bay (stiffener-frame) single dents up to 6.35 mm are allowed; however, if the dent is within 75 mm of a stiffener or frame the allowable depth is only 0.76 mm. With the ever increasing maintenance workload, the T.O. guidelines need to be reassessed regarding structural needs for dent removal. Initially, a literature survey revealed no applicable data on dent limits used for the T.O.. Further research indicated that Airbus has done some work regarding dents in fatigue, and addresses dents in fatigue as stress concentration factors2. Both the effect of dents and reformed dents on durability and static stability is investigated. In fatigue, tension-tension loading is used, and both compression and shear loading are considered for static stability. Since a reformed dent brings the skin back to the moldline of the structure, static stability should be restored; thus, reformed dents were omitted from the static stability portion of the program. A special impact swing hammer was developed to provide consistent dents. The impact energy can be varied by changing the drop height and/or impact weight. The energy level used for impact did not cause any cracks near the dent. Within the dent definition used for this aircraft, depth and diameter, there are many possibilities of differently shaped dents, some examples are given in Figure 2

3

Figure 2: Dent shapes

The dents for this program were formed artificially. In order to investigate the sensitivity of the dent shape, several forming techniques and dent shapes were investigated. Due to the sharp transition between the dent and the skin, the spherical dent was the worst performing dent in fatigue, but obviously this shape was not very realistic. In order to control the size and depth, the natural partly supported dent was used for the remaining part of the program. A ring support at the bottom of the skin controlled the diameter of the dent. The depth of the dent was controlled using the energy level of the impact3 . The majority of testing for this program was performed using 1 mm 2024-T3 clad skins but some additional testing was done using 1.5 mm skins and bare material. These materials cover typical skin configurations of interest.

FATIGUE TESTING IN TENSION Dented panels were subjected to tension-tension fatigue, 120 MPa, R=0.05. Figure 3 shows some typical results for flat unstiffened skin panels with a dent diameter of 50.8 mm.

Cycles to Failure vs Dent Depth120 Mpa, R=0.05

0

200,000

400,000

600,000

800,000

1,000,000

1,200,000

1,400,000

1,600,000

1,800,000

0 1 2 3 4 5 6 7 8 9

Depth (mm)

Cyc

les

to fa

ilure

Single dent allowableStiffener proximity dent allowable

Estimated service life

Reformed

Figure 3: Fatigue life of natural partly supported dents

4

From these results, it is clear that dents do affect fatigue life. The actual dent depth has only a limited effect. Some dents were hand reformed using a hammer, the roughness or dent dept after reforming was close to 1 mm depth. The reforming did actually improve the fatigue life due to a reduction in secondary bending. However, the load cycle, which represents a high stress for a typical pressurization cycle, none of the dents affect the service-life requirement which will be less than 100,000 cycles (flights). As mentioned earlier the shape of the dent is very significant. For a spherical dent, see Figure 2,of about 8 mm deep, the fatigue life was only on the order of 80,000 cycles vs. 200,000 cycles for a more naturally shaped dent. The radius, or lack of thereof, between the dent and the skin which is the critical location for fatigue, see Figure 4, is extremely sharp for the spherical dent. This caused the shorter fatigue life.

Figure 4: Critical fatigue location for dents

The critical fatigue locations around the dent show that the dent is behaving much like an open hole with load-bypass causing the fatigue cracking. The cracks start at the three and nine o’clock position and grow slowly into the dent (dent crack) but grow faster out of the dent (skin crack).

STATIC STABILITY OF DENTED PANELS Besides fatigue considerations, the effect of dents on static stability was unknown. Two load cases were selected; compression and shear. The panels represent a typical fuselage configuration, 1 mm skin, 1000 countersunk rivets on 25.4 mm spacing, hat stiffeners on 203 mm spacing, and a frame spacing of 508 mm.

5

Compression testing Flat stiffened panels, see Figure 5, with multiple small dents and large dents were tested. The purpose of these tests was to provide experimental data to validate analytical modeling of more complex dent configurations consisting of multiple small dents, related proximity effects (i.e. dents to dents and dents to stringers), and small dent / large dent equivalency. To have the same support conditions in the test as an actual fuselage frame provides can be complicated. The frame does prevent out of plane movements but it does allow some rotation. Pouring a resin support at the end of the panel allows a smooth load introduction but almost completely prevents any rotation of the skin. Reference [4] showed that the failure load of panels is relatively insensitive to the exact boundary conditions used in testing4. Due to the fact that in this configuration the stiffeners are a significant load carrying component, great care must be taken to symmetrically load the stiffeners. This was verified by using strain gauges on the two stiffeners. Shims were used on the stiffeners to apply symmetric loading.

Figure 5: Compression panel

For the panel tested with a 203 mm stiffener spacing, using Euler-Johnson5 the predicted effective skin width for this panel was only 45 mm. This meant that the largest part of the skin was not effective in compression. The strain gages, see Figure 6, confirmed that the skin was not effective in a large area between the stiffeners.

6

Figure 6: Compression panel strain gage locations

Figure 7 shows the skin strains in an area where the skin buckles and does not pick up any load in compression; average strains at gages 10 and 12 are nearly zero.

50mm out from stiffener, gage 10 & (12)Non-effective skin in compression

-1500

-1000

-500

0

500

1000

1500

0 10 20 30 40 50 60

Load (KN)

mic

ro s

train

gage 10gage 12avg 10&12

Figure 7: Strain in not-effective skin

7

Closer to the stiffener, at 25 mm, the skin becomes effective as predicted using Euler-Johnson, see Figure 8. The average strain between the front and backside of the skin now does pick up a compressive load.

25mm out from stiffener, gage 9 & (11)Effective skin in compression

-3500

-3000

-2500

-2000

-1500

-1000

-500

0

500

1000

0 10 20 30 40 50 60

Load (KN)

mic

ro s

train

gage 9gage 11avg 9&11

Figure 8: Strain in effective skin area

All this information from the compression panels indicates that damage in a large non-effective skin area will not affect the performance of the panel in compression. In order to verify this, the largest dent that would fit inside the stiffeners (125 mm diameter and 7.7 mm deep) was added to the compression panel and tested. A configuration with 5 dents inside the 203*203 mm test section was also tested see Figure 9.

Figure 9: 5-dent panel under load

8

These five dents were 50.8 mm diameter and 5.8 mm deep. Two pristine, two large dent panels, and two panels with the five dents were tested. Table 1 shows the failure loads of the different panels.

Panel Dent diameter depth Failure Load

# mm mm KN pristine 0 0 0 84.0 pristine 0 0 0 81.1

single dent 1 125 7.7 79.8 single dent 1 125 7.7 82.4

5 dent 5 50.8 5.8 86.9 5 dent 5 50.8 5.8 92.5

Table 1: Failure loads of compression panels with and without dents The failure loads are all very close; however it is remarkable that the 5-dent panel with the highest amount of damage actually seemed slightly stronger than the pristine panels. All panels failed due to local stiffener crippling and inter-rivet buckling of the skin. The other component that could be affected by the dents is the portion of the load that the stiffeners carry. The strains in the stiffeners are very similar but slightly lower in the stiffeners for the dented panel. This could be caused by the fact that the skin with the dents actually is slightly more effective due to the dents stiffening the skin.

Shear testing A small-unstiffened shear configuration (203 x 203 mm, representing the stiffener spacing) was tested, and a larger stiffened configuration (508 x 508 mm, representing the frame spacing) with two stiffeners at 203 mm spacing was tested, see Figure 10.

. Figure 10: Shear panel

Initial shear tests were performed using the 203 mm test panel to verify proper shear loading. The numbering used for the strain gages is given in Figure 11, with gage 4 on the back side of gage 1, gage 5 opposite of gage 2, and gage 6 opposite of gage 3. A double sided rosette was used to determine shear and bending conditions in the panel.

9

Figure 11: Shear panel rosette gage

For static stability, the compression component (gage 3 and 6) of shear loading will be critical. For perfect shear conditions, in absence of bending, ε1 = -ε3, and ε4 = -ε6. Due to the fact that the relative thin skin does bend, the double sided gages can be used to calculate the bending stresses. Initial test results did not show good shear loading of the skin. With a 1.5 mm skin panel, typical strains for the skin are given in Figure 12.

1.5 mm shear panel

-3000-2500-2000-1500-1000

-5000

500100015002000

0 5 10 15 20

load (KN)

mic

ro s

trai

n 1436avg 3 & 6

Figure 12: Shear strains 1.5 mm panel

For clarity only gage 1 (and 4), and 3 (and 6) are plotted. As mentioned earlier, for perfect shear the principle stresses should be the same magnitude with opposite signs. The strain data does not show this behavior. In the tension direction, gage 1 and 4 behave very similar, no bending in the tension direction. In the compression direction, the front (3) and back gage (6) behave opposite of each other due to bending. Closer investigation of gage 3 and 6 showed that the skin immediately starts bending and buckles in the compression direction. The 1.5 mm skin panel does not show proper shear loading, and proper loading using this shear frame cannot be verified due to the early onset of buckling in the skin. The fact that gage 3 and 6 almost have the opposite strain shows that the skin in the compression direction hardly carries any load, see avg 3 & 6 in Figure 12.

10

In order to verify that the shear frame properly applied shear loading, a much thicker 6.35 mm sheet was used in the shear frame. With the bending absent in the compression direction, this panel showed perfect shear conditions, see Figure 13. Gage 1 and 4 have the same strain but opposite in sign of gage 3 and 6. At about 35 kN the compression gages (3 and 6) start to show the onset of bending in this panel.

1.5 mm shear panel

-400

-300

-200

-100

0

100

200

300

400

0 10 20 30 40 50

load (KN)

mic

ro s

trai

n 3164

Figure 13: Shear strains 6.35 mm panel

With the proper loading conditions using the shear frame known, and the fact that the thin skins immediately buckle in compression, the behavior of the shear panels with and without dents was expected to be similar to the behavior in compression. Figure 14 shows the behavior of dented and pristine large shear panels, see Figure 10. The dent in this panel was the large single dent with a 125 mm diameter and 7.7 mm deep. The failure load of the shear panels was nearly identical, around 155 kN, despite the large dent in the skin. Just like the compression panels, the dent did not have a negative effect on the static stability of the shear panel. There is also an indication that the dented panels are slightly stronger.

11

Shear Panels

0

20

40

60

80

100

120

140

160

0 2 4 6 8 10 12

Displacement (mm)

Loa

d (K

N) Pristine-2

Dented-3Pristine-4Dented-5

Figure 14: Shear panel behavior

CONCLUSIONS Any damage in the non-effective skin does not affect panel strength for compression and shear loading. The large dent in the skin did not affect the static stability of the stiffened panel. The multiple dents in a skin bay did not affect panel strength negatively either. Dents might actually stiffen the skin region and make it slightly more effective, thus increasing panel strength. Only damage inside the effective width of the skin might affect panel strength. The effective width for the panels tested is so small that a dent in this region will cause stiffener damage. Obviously a damaged stiffener must be repaired and was not part of this investigation. Since static stability in shear is only affected by the compression component, shear data indicates that the skin is also not effective in the compression direction. For the configuration tested, a very favorable conclusion for dent related damage can be made: dents outside the effective width of a panel are not be an issue for static stability. This conclusion might be applicable for more types of structures but needs to be investigated further. Since the dents also do not critically affect the required fatigue life, fatigue performance degradation is of no concern regarding dents in the structure either, but dents might need additional inspections if fatigue is more critical for the application. With this new information available, aircraft availability will not be affected by dent repair actions and significant savings can be achieved by preventing not critical maintenance actions.

ACKNOWLEDGEMENTS This project is financially supported by the Oklahoma Air Logistics Center at Tinker Air Force Base; Mr. Jim Harper, Mike Sneed and Bob Nelson. Mr. Joe Luzar of The Boeing Company has provided technical supported for this project.

12

REFERENCES: 1 Fuselage Dent Evaluation Phase I, Final Technical Interchange Meeting EA 04-006-135OTH, USAF Academy, June 2004 2 B. Schmidt-Brandecker, Dent assessment guideline: A method to assess dents in fuselage skins regarding durability, International Committee on Aeronautical Fatigue 1999, p 5-19 3 M. Hagenbeek, Impact properties, Fibre Metal Laminates an introduction, p 409, Kluwer Academic Publishers ISBN 1-4020-0038-3 4 Marshall Holt, The effect of testing on the ultimate loads supported by stiffened flat sheet panels under edge compression, NACA Tech Notes 811 5 Airframe stress analysis and sizing; Niu M.C.Y.; ISBN 962-7128-08-2, January 1999