fretting fatigue in mechanical se-161 11 bromma, email … · hi-lok t! al rivet h!-lok steel...

FRETTING FATIGUE IN MECHANICAL

JOINTS:

Critical Experiments and Numerical Analyses

G. Segerfrqjd*t, G.-S. Wang* and A.F. Blom*t

* The Aeronautical Research Institute of Sweden, P.O. Box 11021,SE-161 11 Bromma, email [email protected], [email protected] [email protected].

f The Royal Institute of Technology, Department of Aeronautics,SE-100 44 Stockholm, Sweden

ABSTRACT

Two different fatigue initiation mechanisms have been observed in 1 1/2-dogbone joints. Somecracks initiated at the edge of the fastener hole at the location of the maximum stressconcentration, as would normally be expected. However, fatigue cracks frequently initiatedseveral mm:s away from the edge of the fastener hole. This anomalous behavior is due tofretting and is particularly prevalent in joints assemled with certain fastener systems. Thispaper summarizes a part of a larger experimental programme and shows which designparameters, e.g. secondary bending, fastener stiffness, etc., control the two different initiationmechanisms. Also included are results from detailed numerical modeling that shows that theobserved behavior can actually be predicted and the numerical results are shown to correlate wellwith experimental observations.

INTRODUCTION

An aircraft structure is most often assembled of several mechanically joinedpieces, which often causes the design to be heavier than desired. The stressconcentration induced by the fasteners in a shear loaded joint is often theoriginating point of fatigue failure.

Due to the large number of parameters involved in the fatigue behavior ofmechanical joints and the difficulties in measuring several of these parameters,the fatigue design and analyses of such joints is usually cumbersome. Extensiveexperimental investigations regarding fatigue behaviour of various mechanicaljoints are presently performed at the authors' institutes. Among the experimentalinvestigations, a type of 1 1/2-dogbone joint specimen showed that, dependingon fastener system, fatigue cracks often initiated due to fretting in front of thehole. However, in some cases the initiation took place at the edge of the hole atthe maximum stress concentration site, as would normally be expected.

Transactions on Engineering Sciences vol 14, © 1997 WIT Press, www.witpress.com, ISSN 1743-3533

Contact Mechanics HI

PURPOSE

The main purposes of this investigation are to perform detailed numericalanalyses of the mentioned joint type and to compare results of the analyses toexperimental results. The specific objective is to identify the mechanism thatgoverns the fatigue initiation process in the 1 1/2-dogbone type of joint.

EXPERIMENTAL DETAILS

Material and Test SpecimensThe so called 1 1/2 dogbone joint is used in this investigation. It simulates theload transfer (LT) and the secondary bending (SB) characteristics of the stiffenerrunouts attached to the outer skin of an aircraft. The test specimen designobjectives were to achieve a load transfer of approximately 40 % and asecondary bending ratio of 0.50. However, in these type of joints, LT as well asSB are critically dependent on fastener system and installation parameters, e.g.the type of fastener, clamping force, fastener fit etc., [1-2].

All specimens were cut out with the rolling direction parallel to the loadingdirection, i.e. in the L-direction. Additionally, all specimens were subjected tosurface treatment, i.e. they were anodized with phosphoric acid according toSaab standard STD 1991 and primer treated (9344-08) according to Saabstandard STD 2403. No wet assembling or interfay sealant were applied.

Totally, 100 test specimens were manufactured from aluminum alloy A A2024-T3, with sheet thickness t=2.5mm. However, only 30 specimens havebeen tested at our institutes [3], since the other 70 specimens were tested byHuck in a joint programme between Saab Military Aircraft and Huck, [4]. Thepresent paper is restricted to the study of nine selected test specimens. The staticmechanical properties in the rolling direction (L-direction), according to SaabMaterial Standard, are: cr =360MPa, cr =485 MPa, elongation =15.5%.These are guaranteed minimum values. Hence, they are lower than actual valuesmeasured on each batch of material.

Fastener SystemsThree fastener systems are considered in this investigation, a titanium Hi-Loksystem, a steel Hi-Lok system and a solid aluminum rivet system. All arecountersunk (CSK) fasteners. The nominal diameter of the fastener holes, aswell as CSK details are dependent on the applied fastener system, see Table 1below for details.

Fastener System Designation

Hi-Lok TiHi-Lok SteelAl Rivet

HL11VAP-6-3HL19PB-6-3AF139DD-6-5

Hole Diameter[mm]

Ream 4.72-4.79Ream 4.72-4.79

Drill 4.9

Countersunk[0mm]

100°, 7.5-7.8100°, 7.5-7.8100°, 8.6-8.8

Saab StandardHole/CSK

STD1957/2018STD1957/2018FiV432:066

Table 1. Fastener System Compilation and Hole Manufacturing Details


Contact Mechanics HI 795

Schematic blueprints of the fastener systems are shown below.

Hi-Lok T! AL Rivet

H!-Lok Steel

Figure 1. Hi-Lok and solid aluminum rivet systems

The fastener fit is dependent on the chosen fastener system. The Hi-Lok systemshave interference fitted fasteners with approximately 50 |uim interference. Thesolid aluminum rivet system is a hole filling system.

The manufacturer of Hi-Lok fasteners guarantees a minimum fastenerclamping force (CF), i.e. an axial fastener pre-load after installation. Theguaranteed minimum magnitude depends on fastener diameter. However, theguaranteed magnitude is not a representative value of fastener CF. Evaluationtests indicate a fastener CF in the order of 50%-80% of ultimate fastener fractureload. Thus, the tested Hi-Lok systems are characterized by a CF ofapproximately 5-7 kN. Solid rivet systems are characterized by almost zero CF,i.e. CF=0.1-1 kN. However, in design and fatigue life estimations the CF inrivet systems is usually neglected.

Flight Simulation Load HistoryThe specimens in this investigation are subjected to spectrum loading with theMiniTWIST load history at a maximum gross stress level of 250 MPa.

The MiniTWIST [5] load history is a shortened version of TWIST [6], thestandardized load sequence for flight simulation tests on Transport Aircraft WingStructures. This spectrum is representative for the load history of the wing rootof a transport aircraft. The TWIST spectrum consists of 10 different flight types(A-J) with an average flight length of 90 cycles. In the shortened version, i.e. inMiniTWIST, low-amplitude cycles were omitted in order to decrease testing timeand thus, the average flight length decreased from 90 to 15 cycles, although thenumber of flights in one block remained the same.


196 Contact Mechanics HI

EXPERIMENTAL RESULTS

In Table 2 below, fatigue test results are presented in terms of the number offlights to failure for each specimen and with mean value and standard deviationgiven for each type of fastener system.

Spec* No.

456789131415

Type of FasSystem

Hi-Lok TiHi-Lok TiHi-Lok TiHi-Lok SteelHi-Lok SteelHi-Lok SteelAL Rivet typeAL Rivet typeAL Rivet type

tener

III

Fatigue Life[F]104200108492109217394643997649579107211146011706

Meanm

107303

43006

11296

Std Dev.[F]

2712

5698

513

Std Dev.f%1

2.5

13.2

4.5

Table 2. Fatigue test results of three different fastener systems assembled in a 1 1/2dogbone joint specimen

These results indicate that the Hi-Lok fastener systems show a significantlyhigher fatigue resistance than the solid aluminum rivet system. In the Hi-Loksystems, titanium fasteners show more than twice the fatigue resistance of steelfasteners. This may partly be explained by a higher flexibility of the titaniumfastener. Thus, less secondary bending (SB) is imposed during fatigue loading,which will result in decreased peak contact pressure between the dogbone andthe splice plate. Hence, the fretting area for the most critical stress amplitudes isdisplaced away from the fastener hole, i.e. the maximum stress concentrationsite.

Fatigue Crack Initiation SitesA fatigue crack initiation process may start at different locations due to differentdriving forces. Two types of crack initiation mechanisms were experimentallyidentified [7]:

• Type I, which is characterized by a fatigue crack initiation in theminimum net section (rivet row) at the edge of the fastener hole. Thislocation of crack initiation is named a Stress Concentration InitiatedCrack, (SCIC).

• Type II, which is characterized by a fatigue crack initiationapproximately one radius ahead of the fastener hole center. Thislocation of crack initiation is typical for systems with a high clampingforce, i.e. a higher percentage of load is transferred by friction, andthus, the most critical stress location at the minimum net section ispartly relieved. The higher clamping force also initiates smalldisplacement movements ahead of the fastener hole, i.e. frettingcorrosion, that subsequently initiate the crack. In the following, thistype of crack initiation is named a Fretting Initiated Crack, (FIC).

In general, remains of oxide debris at the fracture initiation sites were found forboth type I and type II crack initiation mechanisms.


Contact Mechanics III 797

For the Hi-Lok systems, fracture in all specimens initiated due to fretting, i.e.a type II initiation. Specimens with aluminum rivets seem to have been subjectedto extensive fretting. However, they initiated at the fastener hole stressconcentration site, i.e. a type I crack initiation.

A scanning electron microscope (SEM) was used to locate the fatigue crackorigins. Positions of the crack initiation sites are surveyed below for the two Hi-Lok systems and the aluminum rivet system.

O OA B

HOLE A HOLEB

Ti HI-Lok Steel Hi-Lok Al-Rivet

Figure 2. Experimentally observed fatigue crack origins

A majority of fatigue crack initiation is located to the free edge side, i.e. at theleft side of fastener hole A or at the right side of fastener hole B [3].

CONTACT ANALYSIS WITH FE-FORMULATION

The fretting contact is usually divided into three different regimes; stick, mixedstick-slip and gross-slip regimes. The mixed stick-slip regime is characterized bylow amplitudes and as the initiation site of fretting fatigue cracks. These cracksare usually observed at the boundary between stick and slip, or sometimes at theboundary of the contact zone.

A series of finite element models have been created systematically, with animprovement in refinement between each model in order to evaluate the modelrefinement needed to study stress distributions accurately [1].

FE Contact Analysis: DetailsIn this investigation an FE-model with a total of 7024 elements and 19026DOF:s is used. The hole in the plate and the solid fastener with a nut and acountersunk head are modeled. A detailed mesh around the fastener hole is used.The fastener fit is taken into account in this model as well as the clamping force.The fastener fit is derived by an expansion of the fastener diameter, implying atotal interference fit of 50 |Lim. The clamping force is obtained by prescribing afastener deformation of-0.051 mm in the length of the fastener. This resulted ina clamping force of about 7 kN which is comparable to the experimentalmeasurements mentioned above.


198 Contact Mechanics III

The contact and friction conditions are modeled by using ABAQUSisCONTACT PAIR/CONTACT SURFACE approach [8]. A friction coefficient of|1=0.65 is used on fretting contact surfaces. On all other contact surfaces afriction coefficient of |J=0.25 is used.

The analysis is performed under load control condition with the far-field loadapplied in the longitudinal direction of the specimen. The load is applied as aconcentrated load at all nodes in the dogbone top section. One full load cycle ismodeled, i.e. max 250 MPa, min -58 MPa and max 250 MPa corresponding tothe extreme values of MiniTWIST load spectrum. A ramp function is used toapply the load in every step.

FE Contact Analysis: ResultsThe contact and friction behavior at the dogbone and splice plate interface in thevicinity of the fastener holes are evaluated during the analyzed load cycle. Inorder to determine the fretting contact regimes at the joint interface, the stick,mixed stick-slip and gross-slip regimes are visualized at maximum and minimumload levels for the three fastener systems. A Coulomb friction model is used, i.e.the model assumes that no relative motion occurs if the equivalent frictionalstress

T, =V< + (1)

is less than the critical stress, T^, which is proportional to the contact pressure,/?, in the form:

T.=WP (2)

where ]H is the friction coefficient and r, and r are perpendicular shear forceson the contact surface. If the equivalent stress equals the critical stress( T^ = T ), then slip may occur. As the friction is assumed isotropic, thedirection of the slip and the frictional stress coincides, which is expressed in theform

1 7- = -4 (3)T 7«9 / *,

where 7. is the slip rate in direction i and 7 is the equivalent slip rate

(4)

The stick area increases when comparing results at maximum load for the firstand second load cycle. Therefore, results of the maximum load in Figure 3below correspond to the maximum load reached during the second load cycle.


Contact Mechanics HI 199

Hole A Hole A

Maximum Load250MP&

Minimum Load-58MPa

Figure 3. Contact interface response: steel Hi-Lok system at maximum and minimum load.

Hole A Hole A

Sliparea

Stickarea

TIHI-Lok

Maximum Load250MPa

MWmum Load-58MPa

Figure 4. Contact interface response: Ti Hi-Lok system at maximum and minimum load.

Hole A Hole A

Maximum Load350MPa -58MPa

Load

Figure 5. Contact interface response: Solid aluminum rivet system at maximum andminimum load.



According to the results in Figures 3-5, there seems to be a distinct correlationbetween analysis and experimental results. The Hi-Lok systems are characterizedby mixed stick-slip regimes at the top of the fastener holes. The steel Hi-Lokfastener system shows a mixed stick-slip regime at approximately 3 mm ahead ofthe fastener hole center for maximum and minimum load. The titanium Hi-Loksystem shows a mixed stick-slip regime closer to the minimum net section. Thisis located on the edge of the hole, approximately 1.7 mm from the center line ofthe fastener hole. The solid aluminum rivet system shows a negligible contact formaximum load and no contact at all in the fretting area for minimum load. This iswell in accordance with experimental crack initiation sites, see Figure 2.

However, the specimens are subjected to spectrum loading and most cycleshave less range than maximum and minimum load. Based on the exceedancedistribution for the actual load spectrum, see Figure 6 below, the most frequentrange is from 136 MPa to 64 MPa.

Mini-Twist Spectrum

100

Excedances

10000

Figure 6. Exceedance distribution for MiniTWIST load spectrum

In order to evaluate the contact response at corresponding load levels, a secondanalysis with the titanium Hi-Lok fastener system was performed. Forsimplicity, the mid value, 100 MPa, was chosen. The analysis was run in aseven step procedure, see Figure 7 below. Contact results were extracted at eachpass of load level 100 MPa, both during loading and unloading.

300 +Analysis load history

Time

Figure 7. Analysis load history, contact results extracted at 100 MPa, i.e. in steps 1, 3,5and?.



The fretting contact regimes for points 1, 3, 5, and 7 are depicted below inFigure 8.

Hole A Hole A

Step 1lOOMPoi

St*p 3iOOMPa

Hole A Hole A

Step 5lOOMPa

Step 7lOOMPa

Figure 8. Contact interface response for titanium Hi-Lok system at the most frequent loadlevel according to spectrum exceedance distribution, 100 MPa. Analysis steps according toFigure 6 above.

Results shown in Figure 8 indicate that the fretting contact differs in loading andunloading. In contrast to the first loading, the 100 MPa load level of the secondloading exhibits a complete stick area without slip. However, during unloadingsome slip takes place in the typical 1 1/2 dogbone critical fretting area. Thecontact behavior seems to stabilize after the first loading.


202 Contact Mechanics III

FATIGUE CRACK GROWTH ANALYSES

The fracture mechanics method is used to analyze both the crack initiation andsubcritical propagation due to fretting based on the finite element stress analysisand stress intensity factor solutions at the actual crack location. The analysis isintended to quantify which parameters are of major importance for crackinitiation respective propagation when the fretting mechanism is involved.

Stress Analysis ResultsBased on the numerically determined fretting area, critical stresses are evaluatedat the mixed stick-slip boundary for the steel and titanium Hi-Lok systems.Stresses for the solid aluminum rivet system are evaluated at the stress,concentration site in the minimum net section. The positions of stress evaluationare depicted in Figure 9 below.

SCIC— FIC 3nm--- FIC 2mm

Hole A ' Hole BDogbone

Plate

Figure 9. Stress evaluation along depicted lines. Steel Hi-Lok (FIC 2mm), titanium Hi-Lok(FIC 3mm) and solid aluminum rivet (SCIC).

The (7, stress is evaluated at load steps 1-3 for the Hi-Lok systems according to

Figure 10 below. The a stress distribution is almost identical for maximumapplied far-field stress in the first and second load cycle, i.e. points 1 and 3 inFigure 10.

500-, 0=250 MPa

-500-J

X-axis, mm

1 Steel A 1 Ti

2 Steel ffi 2 Ti

a 3 Steel 3Ti

Figure 10. Stress distributions for the Hi-Lok system at the critical fretting area



In the fatigue crack growth analyses, stress amplitudes are of interest: At thecritical fretting area, the stress amplitude for the steel Hi-Lok specimens is

(5)

For titanium Hi-Lok specimens, the stress amplitude is curve fitted as

with the origins of x and z at the crack center and

The stress distribution for a crack that initiates at the edge of the hole is for thesteel Hi-Lok much less than the gross stress amplitude, namely

Acr . / a = 0.5577 - 0.4701a + 0.2782a' (8)

and for the titanium Hi-Lok, the peak level is close to the gross stress amplitude:

AO^./CJ = 1.0188 -1.0946a + 0.7082a* (9)

Fatigue Crack Growth CalculationsThe stress intensity factors are computed based on the weight function technique[9] and the crack growth analysis is based on the strip yield crack growth model[10-11].

Since the stress amplitude at the edge of the hole is less than the gross stress,only the crack growth at the fretting location is analyzed for the Hi-Lokspecimens. Regarding the aluminum rivet specimen, the maximum stressconcentration site at the edge of the fastener hole is analyzed. The results arebased on an initial crack assumed to be in the order of the maximum detectedmetallurgical defects, i.e. an elliptical inclusion with half axes; x=15 |Lim, y=7.5|im and z=7.5 |im. The x-y plane of this coordinate system is in the contactplane with the y-direction in the longitudinal direction of the specimen.

The calculated crack growth versus the number of flights is depicted in Figure1 1 below.



lE-2-g

1E-3-

sof

1E-4-

1E-5-

2024 T3, (a/c)-=0.5, Mini-TWIST

O tf

1000 10000 100000

FlightsFigure 11. Calculated crack growth at the fretting area for Hi-Lok specimens.

Maximum crack length in calculations, approximately 2.5 mm, corresponds tothe specimen thickness meaning that the crack developed into a through-the-thickness crack. The subsequent crack growth is not taken into account. Acomparison of numerical calculations to experimental results is shown in Table 3below.

Fastener System Fatigue Failure(Experiments)

Hi-Lok TiHi-Lok SteelAl Rivet

107 3034300611 296

RatioH-Lok/Al

9.53.81.0

Through-the-Thicknessc Crack

45000174002800

RatioHi~Lok/Ai

166.21.0

Table 3. Comparison of experimental and analysis results

The predictions are qualitatively in good agreement with experimental results,and the different fastener systems are ranked in the correct way. However, thenumerical results are very sensitive to assumptions of initial crack size.

The number of flights as function of initial flaw size is depicted in Figure 12below.



100000 -

10000-

1000

00g 100-y

Mini-Twist Spectrum

I I M.001 .010 .100 1.000

a0, mm10.000

Figure 12. Number of flights as function of initial flaw size.

CONCLUDING DISCUSSION

It has been shown that two different fatigue initiation mechanisms have beenobserved in 1 1/2-dogbone joints. The reasons for this behavior was explainedand the parameters controlling the observed behavior have been clarified.

It appears that numerical modeling is well capable of predicting when frettinginitiation may occur, and also at which location this can be expected.

What still remains is to better quantify the numerical predictions of total fatiguelife behaviour. Here, we have outlined one approach, based on fracturemechanics modeling of growth starting from an existing defect assumed tocorrelate to the largest inclusions in the material itself. The predictions arequalitatively in good agreement with experimental results, and the differentfastener systems utilized are ranked in the correct way. However, the numericalresults are very sensitive to assumptions of initial crack size. In the future weplan to concentrate on this early phase of cracking by excluding the crack growthfrom transition of a part through crack to a through crack and growth to finalfailure. This can be done by fractographic observations from the brokenspecimens. It is presently unclear if the present approach of starting with anassumed initial flaw is too conservative. There might exist a traditional initiationperiod for the original flaws to become crack like. If so, there should be anobserved effect on fatigue life not only from the here studied parameters, butalso from material strength. Such studies will be undertaken in the future.

KEYWORDS

Mechanical joint, contact analysis, fretting fatigue, crack growth, finite elementmodel



REFERENCES

[1] Segerfrojd, G. and Giovannelli, G., "Fatigue Behavior of MechanicalJoints - Detailed Finite Element Modeling and Stress Analysisof a 1 1/2 Dogbone Joint," FFA TN 1996-35, The AeronauticalResearch Institute of Sweden, Stockholm, December, 1996.

[2] Linden van der, H. H., Lazzeri, L. and Lanciotti, A., "Fatigue RatedFastener Systems in 1 1/2 Dogbone Specimens," NLR TR 86082U, National Aerospace Laboratory, NLR, Amsterdam, August, 1986.

[3] Segerfrojd, G., Zuccherini, S., Giovannelli, G.and Magnusson, L.,"Fatigue Behavior of Mechanical Joints - An ExperimentalEvaluation of Ten Different Fastener Systems and theirInfluence on Fatigue Life," FFA TN 1996-63, The AeronauticalResearch Institute of Sweden, Stockholm, January, 1997.

[4] Magnusson, L., "Fatigue of joint specimen type 1 1/2 dogbonewith blind fasteners or rivets as reference fastener", TUDL R-4000, Saab Military Aircraft, Linkoping, Sweden, 1994.

[5] Lowak, H., De Jonge, J. B., Franz, J. and Schutz, D., "MiniTWIST - AShortened Version of TWIST", NLR MP 79018 U, Amsterdam, May1979.

[6] De Jonge, J. B., Schutz, D., Lowak, H. and Schijve, J., " AStandardized Load Sequence for Flight Simulation Tests onTransport Aircraft Wing Structures," NLR TR73029, Amsterdam,1973.

[7] Segerfrojd, G., Zuccherini, S., Giovannelli, G. and Wang, G-S., " FrettingFatigue in Mechanical Joints - Detailed Numerical Analyses,"FFA TN 1997-11, The Aeronautical Research Institute of Sweden,Stockholm, March, 1997.

[8] Hibbit, Karlsson & Sorensen, Inc, "ABAQUS/Standard User'sManual" Volume II, Version 5.4, 1994.

[9] Wang, G-S., "A generalised WF solution for mode I 2D part-elliptical cracks", Engng. Frac. Mech. Vol.45, No.2, pp. 177-208, 1993.

[10] Wang, G-S., "The plasticity aspect of fatigue crack growth",Engng. Fracture Mech. Vol.46, No.6, pp.909-930, 1993.

[11] Wang, G-S. and A.F.Blom, "A strip model for fatigue crack growthpredictions under general load conditions", Engng. Fracture Mech.Vol.40, No.3, pp.507-533, 1991.


fretting fatigue in mechanical se-161 11 bromma, email … · hi-lok t! al rivet h!-lok steel...

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