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INTEGRATION OF MSC.Nastran AND AFGROW TO DESIGN

REINFORCEMENT FOR FATIGUE LIFE EXTENSION

Lim Chi Keong1, Chow Wai Tuck1, T.E. Tay2

1Republic of Singapore Air Force, 2National University of Singapore

2001-109

ABSTRACT

The S211 fleet of jet trainers has been in service within the RSAF for the past

15 years. Throughout this period, local upgrades performed on the aircraft have

increased the weight significantly from that of the original design, hence reducing the

airframe life. The aim of the project is to restore the fatigue life of the airframe

through the use of reinforcement patches and doublers by reducing the stress level on

the fatigue control points identified in the full-scale fatigue test to meet the Damage

Tolerance Requirement, MIL-A-83444. To perform the assessment, computational

simulation based on finite element method (MSC.Nastran) is employed to evaluate the

applied loads on the control points. To model the airframe, the FE models of the

fuselage and wings were first generated. Wing components such as spars, ribs and

stiffeners were modeled in detail. Similarly, the fuselage with the empennages in

which the frames, the bulkheads and longerons were modeled. To validate the

simulation, the FE result is compared with the strain gage data from the static load test

result.

The upgrade designs on the control points were then simulated to evaluate the

applied loadings. Both the tensile and bearing loads can be studied by modeling the

fasteners and adjoining plates in detail. On the control points, reinforcement design of

doublers and composite patches were simulated and analyzed. From the studies

performed, the tensile and bearing loads could effectively be reduced based on the

proposed designs.

Based on the computed stress result of these fatigue control points, a crack

growth program, AFGROW, was used to compute the fatigue inspection interval to

satisfy the damage tolerance requirement, MIL-A-83444. These fatigue inspection

intervals would quantify the effectiveness of the simulated upgrade designs in terms

of extended flight hours. Based on the performed studies, the proposed upgrade

designs would restore the fatigue life of the airframe back to the original design life.

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1. Introduction

The project is developed in two phases; the first phase involves the generation of the

course grid wing and fuselage model. The second phase would involve refinement of

the coarse grid model on identified control points after which reinforcement designs

would be simulated to analyze its effectiveness. Stress results from the FE model

would then be used in the crack growth software AFGROW to access the fatigue life

of the aircraft based on the concept of Damage Tolerance Assessment (DTA).

2. Coarse Grid Model

Before the FE modeling begins, the airframe structure is examined carefully to

evaluate the criticality of each component and hence the level of detail required in the

modeling. The S211 airframe can be categorized into three components: fuselage,

wing and empennages. The aircraft is modeled according to the data specification

from the OEM which includes dimensions and material properties.

For the fuselage structure, only the main load carrying structures are modeled. These

include the major and minor longerons, the bulkheads and the skin. Secondary

structures like the radome, canopy and floorboards which do not carry much loads are

omitted.

The fuselage model and the wing model were first developed separately and assessed

for errors before being combined to perform a full analysis. For the fuselage model, 4

node plate elements were used for the skin as well as the web of the bulkheads. 2 node

beam elements were employed for longerons and stringers. For wing fitting

attachments, rod elements were used. For the fuselage structure, a total of 7555 nodes

and 12939 elements were used. The following figures shows the three main fuselage

structures of the aircraft.

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Figure 1. Cockpit and center fuselage mesh

Figure 2. Empennages

The wing structure has been modeled in detail, including both leading and trailing

edges. In addition, the holes in the skins, ribs and spars are also respectively modeled.

The wing model consists of 9892 nodes and 12814 elements.

Figure 3.Wing Structure

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3. Full Scale Fatigue Test Result

Based on the extended Full Scale Fatigue Test conducted by OEM, it was identified

that the immediate airframe life is determined by the fatigue critical locations (FCL)

at control points F2, F4, W4 and W5. F2 and F4 are located on the bulkhead near the

rear and front wing fitting station respectively (see Fig 4). W4 and W5 lie near the

wing fuel transfer hole on the stringer and the bottom skin respectively (see Fig 5)..

All four FCLs have to be remedied to restore the fatigue life of the aircraft to

8,000FH.

Figure 4. Crack initiation on FCL F2 and FCL F4

Figure 5. Crack initiation on FCL W4 and FCL W5

4. Mesh Refinement on Fatigue Critical Locations

The mesh on the identified critical location has to be further refined and higher detail

element formulation to allow a more accurate stress analysis. As the nature of the

problem on FCL F2 is due to a crack occurring from the riveted lap joints, the effect

of bearing stress induced by the rivet has to be considered. To simulate the joints, the

F-4

W4

W5

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fasteners are modeled with DOF spring elements (Figure 6) between the flanges and

deck-skin through which the load transfer through the fasteners may be determined.

The fastener spring constant can be converted to equivalent structural member using

the following hypothetical formula [10]:

DOF Spring constant for aluminum plate and fasteners:

}0.1]2)(12.2[2)(13.0{8

++=d

tavd

tavtavE

C

Where spring constant K = C1 and

2ps

av

ttt

+=

The flanges itself was changed from beam elements to fine plate element so that the

stresses may be evaluated more accurately

Figure 6 DOF spring elements

The problem on FCL F4 is similar in nature to that of FCL F2. Thus the modeling

technique is similar. However, as F4 is a less critical region, the element mesh is not

as fine as that for F2. The significant difference is that stringers run beneath the

flanges on F4 (Figure 7), which has to be properly modeled, and their offsets taken

into consideration.

Figure 7 Stringers offset beneath flanges on FCL F4

On FCL W4, a crack occurs on a stringer but on FCL W5, the crack appears on the

butt-strap near the fuel transfer hole. For W4 it is impractical to remodel the stringers

and wing skin to simulate the rivets. To find the bearing loads, a hypothetical

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approach is used. Therefore, the elements in the region were only re-meshed to finer

sizes. On W5, the butt-strap and the fuel transfer attachment has to be modeled,

together with the surrounding rivets. Figure 8 shows the butt-strap model with riveted

joints.

Figure 8 Butt-Strap and Fuel transfer attachment (inset: modeled fastener)

Validation of Finite Element Model

After refinement, the model was analyzed using MSC.Nastran. The fatigue critical

locations from the OEM report were compared to that of the FEM to validate the

model for use as a platform for modification. Figure 9 and Figure 10 shows the

analysis model of the front and rear fuselage bulkhead under static load.

Figure 9 Major Principle Stress Distribution on rear bulkhead FCL F2

Figure 10 Major Principle Stress Distribution on front bulkhead FCL F4

On FCL F2, where the mesh size is finer, it is possible to visualize the expected

higher stress on one side of the fastener due to the bearing load transfer. On FCL F4,

F2

Uneven stress distribution due to bearing load

F4

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high stress concentration is located at the rivet hole where the crack initiated. For the

fuel hole region (Figure 11), the model also displays high stress concentration on both

the stringer at W4 and on the bottom skin at the rivet hole position, W5. The model’s

critical locations coincide with the OEM’s inspection location on W5. Thus the FE

model on all three sections simulates the characteristics of the actual airframe.

Figure 11 Major Principle Stress distribution on bottom skin

Correlation of Finite Element Model

SM’s static test results provided strain gage results on the main spar and stringers

along the wing. Therefore, a comparison was made between the FE model and the

static tests’ strain gage readings to correlate the results. Most values are close to the

strain gage data with variance of less than 20%. Table 2 displays the comparison

results on the control points.

FCL STRAIN GAGE NO.

STATIC TEST RESULT (KG/MM2)

FE RESULT

(KG/MM2)

CORRELATION FACTOR (%)

F2 007 14.9 16.8 -12.9 F4 004 13.6 14.6 -7.3 W4 001 12.0 11.8 1.6 W5 001 / 002 13.1 11.9 9.2

The percentile differences from the strain gage readings were used as correlation

factors for the corresponding FCL. Interpolation is applied on W5 as it lies between

the two strain gages. All FEM stress values used henceforth would be corrected with

the corresponding correlation factor. With a correlation between the finite element

model and the static test results, upgrade design can then be performed.

W4

W5

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4. Reinforcement Designs

After validating the FE model, reinforcement designs could be simulated to evaluate

their effectiveness. On FCL F4, W4 and W5, typical repair like installing a metallic

doubler to reinforce the region could be employed. Figure 18 shows the FE design of

the upgrade on FCL F4.

Figure 12 Initial and proposed design of doubler on FCL F4

Figure 13 Finite Element mesh of upgrade doubler on FCL F4

Using FE simulation, the 2mm thick doubler was designed to bypass the 1st rivet

position next to the wing fitting. The analysis shows that this design would effectively

redistribute tensile loads from the 1st rivet position without causing an increase in the

bearing load. The simulation uses 2D plate elements for the doublers which are

connected to the flanges via DOF spring elements which act as the rivets.

On FCL W4 and FCL W5, FE analysis was performed to evaluate a butt-strap

designed by the OEM to reinforce the bottom-skin at the fuel transfer hole. Figure 13

displays the designs of the original buttstrap and that of the modified design. As the

new butt-strap design is comparatively large, the offset between the butt strap and the

bottom-skin is ignored and they are modeled together as plate elements in which the

fasteners are ignored.

Thick fitting block

Thin Flange

1st set of fasteners

Avoid 1st set of fasteners

Reinforcement doubler

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Figure 13 Original and modified buttstrap design

For FCL F2, due to geometrical constraint on the rear bulkhead of the airframe, a

typical metallic repair is inapplicable. Therefore, the problem is approached with a

less conventional method of using composite materials to reinforce the bulkhead. The

salient idea is to bond a cured boron epoxy patch onto the bulkhead with FM73

adhesive using a heat blanket cure. FE simulation would then be used extensively to

determine the physical parameters of the patch required to reduce the tensile loads on

FCL F2 effectively.

The composite patch is first modeled using laminate elements on the web of the

bulkhead of which the shape and size is determined by physical constraints.

Figure 14 Simulated composite patch on F2

In optimizing the patch design, analysis was performed to find the ideal orientation of

the uni-directional boron epoxy patch. As shown in Fig 15, the patch is most effective

in reducing tensile loads when the fiber direction is parallel to the flange of the

bulkheads.

θ

Z

X

Y

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Figure 15 The effect of ply orientation on major principal stress

The thickness of the composite patch is largely limited by the ability of the heat

blanket to produce a good cure. Therefore, the number of boron ply used has to be

determined by compromising between obtaining a good cure with a thin patch and

effective load reduction using a thick one. The use of FE simulation thus allows the

evaluation to be performed efficiently, after which it was determined that a [0]20 ply

configuration would reduce the tensile loads sufficiently.

Other then reducing the tensile load, simulation was also done to investigate the effect

on the bearing load due to varying doubler thickness. Fig 17 shows the reduction of

the bearing load on the 1st rivet position due to various upgrade configurations.

Figure 16 Plot of Bearing Load from various configurations

20.0

21.0

22.0

23.0

24.0

25.0

26.0

27.0

28.0

29.0

0 15 30 45 60 75 90 105 120 135 150 165 180 195

Ply Orientation (θ)

Maj

or P

rinci

ple

Stre

ss

(kg/

mm

2 )

4.44mm 0.81mm

2mm doubler

1.6mm doubler

0.81mm skin

2mm flange

1st rivet

2nd rivet

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Figure 17 Plot of Bearing Load from various configurations

From the analysis performed, it can be seen that reaming away the doublers above the

flange can significantly reduce the bearing load. With the FE model, it was also

possible to ensure that by reducing the bearing load on the 1st rivet position, the 2nd

rivet position would not be aggravated.

3. Stress Analysis Results

After simulation of the reinforcements on the various control points, the effectiveness

of the designs can be judged by comparing the stress results before and after the

upgrade. The relevant stress values obtained from the FE analysis are the remote and

bypass stresses and the bearing load on the riveted joints.

The stress analysis result from the upgrade on FCL F2 is shown in Figure 18.

Figure 18 Locations of elements in obtaining various stress values

Units of σbypass ,σremote in kgmm-2 and ∆P in kg

FCL Original Upgraded Stress Reduction σbypass σremote ∆P σbypass σremote ∆P σbypass σremote ∆P

F2 15.2 25.5 299.6 13.2 17.9 167.8 13.2% 29.8% 44.0%

4.44mm (Original)

4.44mm (Patched)

2.44mm (Patched)

0.81mm (Patched)175

200

225

250

275

300

325

350

375

1 2 3 4 5 6

Bea

ring

Load

(kg)

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The stress analysis result from the upgrade on FCL F4 is shown in Figure 19.

Figure 19 Stress contour after upgrade on FCL F4

Units of σbypass ,σremote in kgmm-2 and ∆P in kg

FCL Original Upgraded Stress Reduction σbypass σremote ∆P σbypass σremote ∆P σbypass σremote ∆P

F4 16.6 27.0 149.5 14.4 23.8 124.4 13.3% 11.9% 16.8%

Figure 20 displays the stress contour on the bottom-skin after the upgrade.

Figure 20 Stress distribution on bottomskin after upgrade

The rivets were not modeled on FCL W4 due to impracticality. To account for the

bearing load, a conservative approximation based on fastener load distribution studies

has been used. Based on the geometric parameters, the bearing stress to remote stress

ratio on the evenly spaced rivets is approximately 0.24.

Units of σbypass ,σremote in kgmm-2 and ∆P in kg

FCL Original Upgraded Stress Reduction σbypass σremote ∆P σbypass σremote ∆P σbypass σremote ∆P

W4 - 13.9 - - 11.4 - - 18.0% - W5 15.8 17.5 70.0 8.1 8.9 23.3 43.7% 49.1% 66.7%

Therefore, using finite element analysis, the upgrade designs on the S211 airframe

could be simulated and the respective stress results obtained from the computational

W

W5

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analysis. With knowledge of the applied stress on the various control points, damage

tolerance analysis could then be performed to predict their respective fatigue life.

5. Damage Tolerance Analysis using AFGROW

Damage Tolerance of airframes is based upon the supposition of pre-existing

flaws in the structure with the initial flaw size specified in MIL-A-38444. Crack

growth rate are then predicted using fracture mechanics method.

The main elements for DTA of a component would include the applied loadings, the

initial flaw size, the load spectrum, crack growth model and the component geometry.

The initial flaw size was referenced from MIL-A-38444. For interference fit

fasteners, often associated with cold-worked holes, the initial flaw size is 0.005”. For

standard holes, the initial flaw size is based on slow crack growth structure of 0.05”.

For field inspection during the depot servicing, an even larger initial flaw size has to

be assumed in the analysis. It is based on the minimum detectable flaw (using NDI)

without the removal of the fastener which is assessed to be 0.1”.

The load spectrum was generated using flight data obtained from fatigue

meters on the aircraft and counted using the Rainflow method to identify hysterisis

loops within the load cycles.

The choice of the Nasgro crack growth equation is chosen because of it

common use on aerospace application and the wide range of material data available

within AFGROW ‘s library.

From the FE stress analysis results, the applied loadings on the component are

determined. The DTA results for the four control points is as shown:

FCL F2 FATIGUE INSPECTION INTERVAL Flaw Size Original Upgrade w/o patch Upgrade with patch

0.005” 1,141FH 2,151FH 3,591FH 0.05” 800FH 783FH 1,325FH

The inspection interval required for FCL F2 has extended from 800FH to 3,591FH

after the upgrade which also requires the cold-working of the rivet hole. When the

upgrade is performed at the 3rd Major Structural Inspection after 4,500FH, the fatigue

life on F2 can be restored from 5,300FH to 8,091FH. Thus, the objective of restoring

the fatigue life back to 8,000FH is achieved.

The fatigue inspection interval required after the upgrade for the other three critical

locations is as shown:

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FCL FATIGUE INSPECTION INTERVAL Original Upgrade

F4 1,711FH 4,212FH W4 5,370FH 9,666FH W5 166FH 14,804FH

Similarly on FCL F4, the inspection interval after upgrade and cold-working has

increased from 1,711FH to 4,212FH. Therefore, the fatigue life of F4 would be

restored to 8,712FH from 6,211FH when the upgrade is performed at the 3rd MSI.

From the application of the doubler, the inspection interval on W4 has increased from

5,370FH to 9,666FH. From 3rd MSI, the fatigue life of the stringer on W4 would be

9,870FH. This would meet the 8,000FH requirement even without the upgrade. This

is not surprising as it is the least critical location among the four control points. Still,

the upgrade of the butt-strap is required for FCL W5.

From the results of W5, it can be seen that there is a large difference between the

flight hours prior to and after the upgrade. This can be explained from the approach

taken in the analysis. The metal adhesive between the butt-strap and bottom-skin were

not accounted for and causes large load transfer onto the bottom-skin. This is justified

by the relatively narrow rim in the original configuration compared to the modified

design. Nonetheless, the results shows that with the OEM’s butt-strap design, the

fatigue life on FCL W5 would meet our requirements. When performed at the 3rd

MSI, the fatigue life would extend from 4,666FH to 19,304FH.

6. Conclusion

Computer aided design using the finite element method (MSC.Nastran) has

been employed for designing the upgrade. The existing S211 finite element model

was developed and validated with the OEM’s static test results. The comparison

results are encouraging and the correlations required on the critical locations are all

within 13%. The finite element model has also successfully located the stress

concentration region as identified within the OEM report.

With the validated finite element model as platform, upgrade designs on the

critical locations were then carried out. On FCL F4, W4 and W5, mechanical doublers

were designed to re-distribute loads away from the critical locations. On FCL F2,

however, due to physical constraint, a composite patch has to be employed instead.

With the finite element model, the analyses for various configuration of the composite

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patch can be performed efficiently. Variables such as patch orientation, ply thickness

and lay up could be simulated in the program and corresponding results evaluated.

The model was also used to study the effectiveness of bearing load reduction by

reducing doubler thickness on the rivet positions. From the analyzed FE model, the

reduction in stress on the control points after the upgrade has been found.

To quantify the effectiveness of the upgrade design, Damage Tolerance

Analysis was performed on the control points using the fatigue crack growth program,

AFGROW. Together with the finite element stress results, inspection interval before

and after the upgrade was found for the four control points.

From the Damage Tolerance Analysis results, it was assessed that by carrying out the

upgrade at 3rd MSI (4,500FH), the fatigue life of the S211 can be restored to 8,091FH.

REFERENCES

[1] S211 RSAF – HUMS. Rep. No. 211-90-50-01, Release date: 15 June 1996

[2] S211 RSAF – HUMS: Analytical Conditional Inspection Assessment, Rep.

No. 560-211-149, Release date: 21 April 1997

[3] SIAI Marchetti Direzione Technica N.C. 003

[4] ATS Paper “Computational Structural Analysis of S211 Airframe” 1999.

[5] Command Reference, User Guide, MSC.Nastran for Windows Version 4.0

[6] Prof. G.Glinka, et al, AFGROW Version 3.9846.11.8

[7] A.C. Urugal, Mechanics of Materials. McGraw-Hill,Inc. 1991

[8] Michael C.Y. Niu, Composite Airframe Structures. Conmilit Press Ltd. 1988

[9] Michael C.Y. Niu, Airframe Structural Design. Conmilit Press Ltd. 1988

[10] USAF specification for damage tolerance requirements, (MIL-A-83444)

[11] Fatigue assessment of RSAF S211,2000

[12] David Broek, The Practical Use of Fracture Mechanics. Kluwer Academic

Publishers, 1989

[13] Julie A. Bannantine, et al, Fundamentals of Metal Fatigue Analysis. Prentice

Hall,1990.

[14] Forman, et al, NASGRO 3.0 Reference Manual, 2000.

[15] ATS Paper “Design of composite patch for fatigue life extension” 2000.