A proposed fatigue test protocol for generic mechanical joints

A proposed fatigue test protocol for generic mechanical joints

Engineering Failure Analysis 13 (2006) 136–154 www.elsevier.com/locate/engfailanal A proposed fatigue test protocol for generic mechanical joints Chu...

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Engineering Failure Analysis 13 (2006) 136–154 www.elsevier.com/locate/engfailanal

A proposed fatigue test protocol for generic mechanical joints Chul Young Park b

a,*

, Alten F. Grandt Jr.

b

a Eclipse Aviation Corporation, 2503 Clark Carr Loop SE, Albuquerque, NM 87106, USA School of Aeronautics and Astronautics, Purdue University, 315 N. Grant St., West Lafayette, IN 47907, USA

Received 27 September 2004; accepted 16 October 2004 Available online 1 April 2005

Abstract This paper describes a generic test procedure to evaluate materials for aircraft joint applications. The life of mechanical joints is controlled by various parameters and it is difficult to reliably predict the influence of many parameters on life. Therefore, the designer often depends on extensive testing to evaluate the performance of various joint configurations. However, these extensive tests are sometimes impractical for initial material development and down-select decisions since new exploratory alloys are often produced in small quantities. This paper addresses that issue by proposing fatigue test methods that have the goal of discerning material performance in as simple a manner as possible. Generic specimen configurations and test protocols for low load transfer and high load transfer airframe joints are described herein. Ó 2005 Elsevier Ltd. All rights reserved. Keywords: Generic joint specimen; Mechanical joint; Load transfer; Fatigue testing

1. Introduction Although mechanical joints are some of the heaviest and most costly aircraft components, they remain one of the most common sources for structural cracking. The life of mechanical joints is controlled by complex stress states, several potential failure mechanisms (e.g., fatigue, fretting, corrosion, fracture), and is sensitive to a variety of manufacturing parameters (e.g., fastener size and material, interference level, hole spacing, surface finish, sheet thickness and material, etc.). Since it is difficult to reliably predict the influence of these parameters on life, the designer is often forced to depend on extensive testing to evaluate the performance of various joint configurations. Each of these factors can, in its own right, have a significant effect on the fatigue performance of a mechanically fastened joint. Therefore, given the attendant dependency on *

Corresponding author. Tel.: +1 505 724 1088; fax: +1 505 245 7888. E-mail addresses: [email protected] (C.Y. Park), [email protected] (A.F. Grandt Jr.).

1350-6307/$ - see front matter Ó 2005 Elsevier Ltd. All rights reserved. doi:10.1016/j.engfailanal.2004.10.013

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these additional parameters, relative improvements in a given alloyÕs fatigue performance determined by small coupon specimens, may not necessarily translate to an improvement in mechanically fastened joint configurations.

2. Objective The primary objective of this paper is to propose fatigue test protocol for mechanically fastened joint configurations. These protocols were developed to discern the joint fatigue performance of selected aluminum alloys. Since these specimens are designed to be representative of aircraft applications, load transfer, secondary bending, material surface preparation and fastener particulars are selected in the context of current aircraft fabrication techniques. An important consideration is to either standardize, or minimize, variability introduced by different joint configurations so that the effects of material performance can be evaluated.

3. Specimen design considerations This section describes several design parameters that have been considered for the proposed test procedure. 3.1. Load transfer (LT) The fact that fasteners transmit loads from one member to another, with various degrees of load transfer is an important specimen design parameter. Load transfer (LT) is defined here as the portion of the load transferred from one structural member to another by the mechanical fastener (Fig. 1), and can vary significantly for various types of joints. The portion of the load not transferred by the fastener is defined as the by-pass load. As demonstrated by Lee [1], who conducted fatigue tests with zero, low, medium and high load transfer specimens, the amount of load transfer has an influence on the fatigue life of structural joints. By increasing the load transfer, the fatigue crack initiation life and the total fatigue life were reduced. The LT, which was determined experimentally using strain gages, ranged from 0.057 to 0.062 in the low load transfer, 0.307 to 0.314 in the medium load transfer and 1.0 for the high load specimens [1]. Shown in the Table 1 is a list of typical aircraft joints with typical load transfer categories.

Load Transferred Bypass Load

Total Load

Fig. 1. Load transfer.

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Table 1 Aircraft joint configurations

Load transfer Type of joint Secondary bending Representative aircraft joint

LLT Joint detail

MLT Joint detail

HLT Joint detail

Low Single shear Low Wing or stabilizer span-wise skin-to-stiffener joint detail

Medium Single shear Minimized (low) Monocoque fuselage longitudinal lap-splice

High Double shear None Wing or stabilizer chord-wise double butt-splice detail

3.2. Secondary bending Secondary bending is a result of the inherent eccentricity in the loading of shear fasteners in structural joints where the applied load seldom coincides with the neutral axis of sheet. As defined in Fig. 2, secondary bending is expressed as the ratio of the axial and bending strains eA and eB, respectively, in the local vicinity of a fastener [2]. It is necessary to minimize the effects of secondary bending in the design of the joint details, since excessive secondary bending may mask potential improvements in the fatigue performance of the selected aluminum alloys. Moreover, a specimen with excessive bending would not be representative of typical aircraft joint details. Therefore, material thickness of single shear joint details was kept to a realistic minimum and a double shear scheme was used as appropriate. 3.3. Aircraft structural joint representation Although complex joints are more capable of accurately representing the lateral stress gradients and load transfer distributions between fastener rows in aircraft structures, the fatigue data acquisition would be complex and quite expensive. Alternatively, the use of simple and cheaper joints, which may only contain a small number of fasteners that nevertheless closely approximate aircraft structural characteristics may still be effective in ranking the fatigue performance of materials in joint applications. 3.4. Material selection The generic specimens proposed in this paper are intended to discriminate material performance in joint fatigue applications. Thus, selection of materials is left to the particular alloy development program under consideration. Secondary Bending Ratio =

A Remote strain

Fig. 2. Secondary bending.

Strain affected by secondary bending

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3.5. Protective systems and specimen treatment Another important parameter for aluminum alloy is surface treatment. This issue entails the condition of the mating or faying surfaces, as well as general surface treatments commonly applied to aluminum structures in the aviation industry. Since the joint specimens are intended to represent aircraft structures, the surface treatment needs to be representative of aircraft exterior surfaces [3]. In addition to resisting corrosion, the alodine, chromate or anodized coatings, and the priming layers also improve the fatigue resistance performance by providing additional wear resistant coatings that delay the onset of fretting, thereby augmenting the resistance provided by the faying surface sealant [4]. The aluminum plate and sheet were chosen here to have a chromate conversion coating applied to all surfaces, followed by priming with a zinc chromate primer. This is typical of the specimen treatment of other similar research reports and is representative of aluminum surface treatments in aircraft applications [4–6]. In addition, the majority of structural joints are assembled with an interfay sealant applied between mating surfaces. This sealant is very beneficial, for it has been proven to inhibit corrosion and fretting by delaying the onset of fatigue crack initiation at the faying surfaces. Faying surface sealants also improve the load transfer (via the sealant) and reduce the stress concentrations at the fastener holes [4]. Therefore, faying surface sealants were applied to all mating surfaces for all the joint details [6]. 3.6. Fastener and installation As a general rule, the total material thickness to fastener diameter ratio (t/D) for aluminum aircraft structures should be greater than 0.18 and less than 5.5 [3]. This minimizes transitional failures in joints, where transitional failures in this instance are defined as joint failures other than by normal failure modes such as shear-out, bearing failure, and so on. Another general rule is that for structural members with a thickness greater than 0.080 inch (2.03 mm), the use of alloy steel fasteners, as opposed to aluminum rivets, is preferred in a single shear joint [3]. Hi-Lok and Lockbolt fasteners are prevalent in wing and stabilizer joint details of commercial transport aircraft, and are therefore representative. Fastener HL19 of the HiLok family was chosen for baseline fastening [6,7]. The Hi-Lok fasteners are installed with collars having a frangible ÔhexÕ that Ôshears offÕ at a predetermined torque of 60–80 inch lbs (6.78–9.04 Nm). Therefore, the degree of clamping provided by the fastener and collar assembly is predetermined by this torque. Since the collar can only be removed destructively, HiLoks are permanent fasteners. In the aircraft industry, Hi-Loks are generally installed Ôwet,Õ that is, with faying surface sealant applied to the fastener shank and hole [3]. This procedure serves as additional protection against corrosion as wet fastener installation does assist in minimizing the effects of fretting fatigue in the bore of the hole [8]. The fasteners are installed in the ‘‘wet’’ condition for this proposed test method. 3.7. Interference-fit and hole preparation In the aviation industry, interference-fit fasteners are used to improve structural fatigue performance. The attendant hole expansion of an interference-fit results in stress fields that delay fatigue crack formation and growth [9]. Interference-fit fastener systems are relatively insensitive to effects of hole quality [3], where hole quality refers to the condition of the bore of the hole (e.g., degree of scoring on the bore, any elongation of the hole diameter, etc). Nonetheless, dimensioning to size to obtain a close tolerance fit results in holes with good fatigue quality [2], whereas inadequately prepared holes can, in fact, reduce the service life of a joint [10]. Other research literature [2], which tested specimens utilizing Hi-Loks in an interference fit, suggested that the interference level be at least 1% of the fastener diameter for optimal results. The specimens in this research use 0.25-inch (6.35 mm) diameter (nominal) fastener and the interference level of 64 ± 13 lm (0.0020 to 0.0030 inch, 0.8% to 1.2% of the nominal diameter) are proposed accordingly.

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The Hi-Lok is a flush head fastener that requires countersinking after the initial hole preparation. Countersunk holes have been responsible for many fatigue cracks in aircraft. In particular, Ôknife-edgeÕ conditions, where the entire thickness of a structural member is countersunk, result in relatively poor fatigue lives due to the high stress concentration factor. The countersink depth is, in general, required to be less than two thirds of the structural memberÕs thickness [3]. 3.8. Specimen geometry The standard edge margin is defined as the distance from the center of the fastener hole to the nearest edge of the structural member, in the direction perpendicular to the primary load direction (see Fig. 3(a)). For fasteners installed in aluminum structures, this dimension is nominally required to be at least twice the nominal fastener diameter (D) [3,12]. In general, the fastener or rivet pitch is recommended to be between 4.0 and 8.0D; where D represents the nominal fastener diameter (see Fig. 3(b)). However, more stringent requirements on the pitch are commonly placed due to the tighter spacing required for inter-rivet buckling or fuel tank sealing considerations. Hence, in these instances, the fastener pitch is recommended to be between 4.0 and 6.0D [3]. 3.9. Specimen length and grain direction In order to simplify testing and minimize fabrication costs, the specimen dimensions were designed to be relatively small. Another important requirement was that the grain or rolling direction of the material utilized in the joint details be representative of aircraft applications. Hence, for the representative wing and

Edge Margin ≥ 2.0D

Loading direction

(a)

Fastener pitch along a row

(b)

Fastener pitch between rows

Fig. 3. (a) Edge margin. (b) Fastener pitch.

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Fig. 4. High humidity test environment.

stabilizer applications (low and high load transfer joint details), the rolling direction of the material was selected to be along the specimenÕs longitudinal axis. 3.10. Loading conditions The loading type and stress ratio (R) was chosen here to be similar to many of the tests described in the literature, which predominantly used constant amplitude loading at R = 0.1. Examples are given in Refs. [1,2,4,5,9,13,14] although tests were conducted at different frequencies. Although the exact reasons for selecting this value of R cannot be traced, Ford [13] suggested it may be that the early fatigue test machine operated best at R = 0.1. Since the mean flight stresses of an upper wing surface are compressive, negative R might be more relevant for the loading of the low load transfer configuration. However, negative R can possibly require additional testing complexities such as anti-buckling restraints. In order to minimize potential changes in fatigue behavior associated with changes in the laboratory environment, it is suggested that the tests be conducted within a controlled, high humidity (greater than 90% relative humidity) environment. This high humidity environment is readily obtained by sealing the test specimen in a plastic bag along with a wet cloth (see Fig. 4).

4. Low load transfer (LLT) specimen 4.1. General The LLT joint detail employed here is the reverse double dogbone (RDD) configuration shown schematically in Fig. 5 and detailed in Fig. 6. This type of specimen is commonly used for test applications requiring fasteners with a LLT such as that experienced in wing or stabilizer spanwise skin-to-stiffener joint details [1,2,11]. The proposed specimen is 14 inches (0.36 m) in length, and is comprised of two asymmetric 0.250 inch (6.35 mm) thick dogbone pieces fastened in reversed order (see Fig. 6). Two fasteners are employed in the test section of the detail, with the location of the manufacturerÕs head of the fastener reversed to achieve symmetry. The RDD configuration has been shown to be particularly sensitive to clamping considerations [2]. When testing the RDD, Van der Linden [2] reported that the clamping procedure is crucial when using wedge type grips. Exploratory tests at NLR, for example, showed that load transfer might even reverse

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Rolling direction

Fig. 5. Low load transfer specimen. 14.000" 7.000" 0.625"

0.500"

0.500"

1.250" 0. 625"

2.500"

2.500"

Rolling Direction Symm Centerline 0.625"

1.250"

3.000"

0. 625"

Locating Hole See Instructions for Hole Diameter

Locating Hole See Instructions for Hole Diameter

R3.0"

R3.0" 1.000"

0.250"

2.000"

Countersunk Near-Side See Instructions for Hole Diameter and Countersink Depth

14.

No Countersink See Instructions for Hole Diameter

Fig. 6. Low load transfer specimen dogbone plate. (Dimensions are in inch.)

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Table 2 Low load transfer joint detail specification parameters Configuration

Reversed double dogbone

Representative joint detail

Upper wing or lower stabilizer span-wise skin-to-stiffener/skin-to-spar joint details

Material Baseline New alloy

To be determined To be determined

Material thickness

Each dogbone is 0.250 inch (6.35 mm) thick

Specimen treatment Aluminum surface Faying surface

Chromate conversion followed by zinc chromate primer (no top coat applied) Pro-Seal 870 applied to all mating surfaces

Fastening system Fastener type Part number Nominal diameter (D) Head configuration Material Shear strength Collar part number Collar material Hole preparation Fastener fit Fastener loading

Hi-Lok HL19PB8-9 0.2485–0.2495 inch (6.32–6.34 mm) Shear flush Alloy steel 95 ksi (655 MPa) HL70-8 2024-T6 Aluminum 15/64 inch pilot-drill/0.2465 inch ream Interference (0.0020 to –0.0030 inch, 64 ± 13 lm) Single shear

Load transfer

Approx. 5% for each fastener

Secondary bending

Approx. SBR between 0.1 and 0.25

Geometric parameters Edge margin (e) Fastener pitch (p) Gross cross-sectional area (AG)

2.5D = 0.625 inch (15.88 mm) 5.0D = 1.25 inch (31.75 mm) 0.625 in2 (403 mm2)

Specimen loading direction

Material rolling direction

if no special precautions were taken. Thus, it is essential to prevent any relative motion of the two dogbone plates at the ends of the specimen. To ensure that this did not occur during testing, alloy steel dowel pins were placed in a press fit into the two locating holes at the ends of the specimen. The effectiveness of the clamping provided by the test machine wedge grips could be further enhanced by application of an epoxy adhesive to the mating surfaces of the specimenÕs ends, which would also assist in preventing relative movements [6]. A summary of the LLT specimen parameters is given in Table 2. 4.2. Load transfer Assuming that the lower and upper plates in the center section of the LLT specimen carry half of the total load applied, and using the displacement compatibility of the lower and upper plate, the load carried by each plate can be approximated. According to the current LLT specimen geometry, L1 = 1.875 inch (47.63 mm), L2 = 2.0 inch (50.80 mm), l1 = 1.0 inch (25.4 mm), l2 = 1.875 inch (47.63 mm) and l3=1.0 inch (25.4 mm) (refer to the Fig. 7). All the loads are assumed to be transferred by the fastener only and not by friction. The resultant displacements of the upper and lower plates due to the load carried by each should be the same in order to meet the displacement compatibility (Eq. (1)).

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Fig. 7. LLT Freebody diagram.

dU ¼

R h¼0:6567

dL ¼

R l1

rad

0

0

RL dx w1 tE

þ

RU R cos h dh ðw1 2Rð1cos hÞÞtE

R h¼0:6567 0

rad

þ

R L2 0

RU dx ; w2 tE

RL R cos h dh ðw1 2Rð1cos hÞÞtE

þ

R l3 0

RL dx ; w2 tE

ð1Þ

dU ¼ dL : Furthermore, based on the free body diagram of Fig. 7, the following relations can be established (Eq. (2)). P trans þ RU  P2 ¼ 0; P trans  P2 þ RL ¼ 0; RL þ RU ¼ P ; P ¼ 1:19RU þ RU ¼ 2:19RU ;

ð2Þ

P P trans ¼ P2  2:19 :

As a result of Eqs. (1) and (2), we get RL = 1.19RU, from which we can conclude that approximately 4.3% of total load P is transferred through the fastener in the LLT specimen. This result does not consider the

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countersunk fastener hole nor the three dimensional geometry. It is suggested that strain gage measurements be employed if a more precise measurement of the actual load transfer is desired for a particular application.

5. High load transfer (HLT) specimen 5.1. General The HLT joint detail is the symmetric double butt-splice configuration shown schematically in Fig. 8 and detailed in Fig. 9. This type of specimen utilizes a double row of fasteners in a very simplistic double-splice configuration. This configuration is quite commonly used for test applications requiring fasteners with a HLT such as those encountered in wing or stabilizer chord-wise double butt-splice details. Examples of similar configurations can be found in Refs. [2,11]. The total length of the proposed specimen is 14.6 inches (0.37 m), and is comprised of two center plates and two splice plates as shown in Fig. 9. The fastener pattern employs two fasteners on either side of the spliced gap, and is loaded in double shear. The location of the manufacturerÕs head of the fasteners is reversed either side of the splice to achieve symmetry. A summary of the HLT specimen parameters is given in Table 3. 5.2. Load transfer Applied load is transferred into the splice plates and fasteners are loaded in double shear. It is assumed that 50% of the loads are transferred through each fastener. 5.3. Secondary bending Symmetric double shear joints have effectively no secondary bending [2]. In this instance, the center of gravity of the splice plates coincides with the center of gravity of the center base plates, implying zero gross eccentricity and therefore no major secondary bending affects.

Fig. 8. High load transfer specimen.

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5. 25"

1. 00"

ø 0. 2465

5. 25" 1. 25"

0. 20"

1. 00"

(a)

1. 00"

R 3.00"

3. 13"

2.38" 1. 00"

2. 00"

ø 0. 25

1. 00"

0. 50"

0. 50"

7. 16"

(b) Fig. 9. (a) Splice plate drawing. (b) Center plate drawings. (Dimensions are in inch.)

6. Experimental results 6.1. Material comparison Generic LLT and HLT specimens were assembled with steel Hi-Lok fasteners and cycled to failure under constant amplitude loading on a 20 kip (89 kN) closed-loop servo-hydraulic MTS testing machine and used to

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Table 3 High load transfer joint detail specification parameters Configuration

Symmetric double butt-splice

Representative joint detail

Wing or stabilizer chord-wise double-butt joint detail

Material Baseline New Alloy

To be specified To be specified

Material thickness

Each center plate is 0.5 inch (12.70 mm) Each splice plate is 0.2 inch (5.08 mm)

Specimen treatment Aluminum surface Faying surface

Chromate conversion followed by zinc chromate primer (no top coat applied) Pro-Seal 1422 applied to all mating surfaces

Fastening system Fastener type Part number Diameter (D) Head configuration Material Shear strength Collar part number Collar material Hole preparation Fastener fit Fastener loading

Hi-Lock HL19PB8-15 0.2485–0.2495 inch (6.32–6.34 mm) Shear flush Alloy steel 95 ksi (655 MPa) HL70–8 7075-T73 Aluminum 15/64 inch pilot-drill/0.2465 inch ream Interference (0.0020 to 0.0030 inch, 64 ± 13 lm) Double shear

Load transfer

50% each row

Secondary bending

Effectively zero

Geometric parameters Edge margin (e) Fastener pitch (p) Gross cross-sectional area (AG)

2.0D = 0.5 inch (12.70 mm) 4.0D = 1.0 inch (25.40 mm) 0.4 in2 (258 mm2)

Specimen loading direction

Material rolling direction

6.20

Log fatigue life

6.00 5.80

Mean: 5.74

5.60 5.40 Mean: 5.32 5.20 5.00 4.80 HL19 Fastener Alloy #3

HL19 Fastener Alloy #4

Fig. 10. Fatigue performance comparisons of two experimental aluminum alloys in LLT joint configuration (rmax = +186 MPa, rmin = 19 MPa).

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rank two proprietary aluminum alloys developed by Alcoa. A total of four different materials were used (Two 7XXX alloys for the LLT specimens and two 2X24 alloys for the HLT specimens). Alignment and calibration of the testing machine were properly checked beforehand. The test environment was 95% humidity room temperature, the stress ratio (R = rmin/rmax) was 0.1, and the cyclic frequency was 11 Hz for all testing. Fig. 10 5.40

Log fatigue life

5.30 5.20 5.10

Mean: 5.107

Mean: 5.110

5.00 4.90 4.80 HL19 Fastener Alloy #1

HL19 Fastener Alloy #2

Fig. 11. Fatigue performance comparisons of two experimental aluminum alloys in HLT joint configuration (rmax = +138 MPa, rmin = +14 MPa).

Fig. 12. Fracture surface of LLT specimen A3S5. (Note the two arrows pointing crack initiation areas.)

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compares fatigue performances of two materials, alloy #3 and alloy #4 in the LLT joint configuration. Maximum gross stress of 27 ksi (186 MPa) was used for all LLT tests and it is seen the alloy #4 has better fatigue performance than alloy #3 based on the mean fatigue life comparison. However, larger scatter is observed in the alloy #4 test results. A similar trend is observed in the HLT test results during which the maximum gross stress was 20 ksi (138 MPa). Fig. 11 shows that alloy #2 performs equally or slightly better than the alloy #1 based on the mean fatigue life, but larger scatter was seen in alloy #2 test results. Therefore, the superiority of one material over the other in the joint fatigue performance was not pronounced in these comparisons. 6.2. Fractographic investigation Fig. 12 shows a typical fracture surface for one of the LLT specimens (ID: A3S5). The specimen failed at 154,353 cycles under constant amplitude loading of 27 ksi (186 MPa) with R = 0.1. Although it is difficult to locate the initial flaw locations with an optical microscope, it is obvious that symmetric cracks developed and grew to failure. A SEM was used to locate fatigue striations and traced back to approximate the initial cracking site located along the bore of the fastener hole. The fracture surface of another LLT specimen (ID: T3) suggests that cracking initiated around the sharp edge of countersunk hole (Fig. 13). All HLT specimens tested developed edge corner cracks at the faying surface, as shown in Fig. 14. The edge corner crack is thought to be a typical initial flaw type for HLT in the countersink geometry. Fig. 15

Fig. 13. Fracture surface of LLT specimen T3.

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Fig. 14. Fracture surface of HLT specimen A2S8.

shows the failure modes of LLT and HLT specimen observed during this research. While all HLT specimens had failures at the joints, some of the LLT specimens failed away from the joint. 6.3. Material and fastener comparison test by Alcoa Technical Center This section is devoted to joint fatigue tests conducted by the Alcoa Technical Center in Pittsburgh, PA with the proposed protocol. The LLT and HLT joint specimens, designed by Purdue University, were fatigue tested under constant amplitude loading for the purpose of comparing materials and fastening systems. All the specimen details remained the same except that different fastener types were used as necessary. Fig. 16 is reproduced from Ref. [15] and compares the S–N fatigue performance of the standard version and improved versions of 7055-T7751 alloy in open hole configuration (R = 0.1). The LLT joint test results shown in Fig. 17, also reproduced from Ref. [15], further demonstrates and verifies the superior fatigue performance of the ÔHigh FatigueÕ version of 7055-T7751 over the standard version. Here the same negative stress ratio (rmin = 215 MPa and rmax = +95 MPa) was used for all LLT joint tests. Although the proposed test procedure employs a standard fastener and installation procedure, and was originally developed to rank materials, the joint specimens could also be used to evaluate the performance of various fastener systems for a given material application. The LLT specimen, for example, was used by Alcoa to compare the fatigue performance of different fasteners in aluminum alloy 7055-T7751 [15]. Here

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(a)

Mode 1

(b)

Mode 2 Fig. 15. Failure modes defined for: (a) LLT specimen, (b) HLT specimen.

34 Standard version

Net stress [ksi]

32

High Fatigue version

30 28 26 24 22 20 1.E+04

1.E+05

1.E+06

1.E+07

Cycles to failure

Fig. 16. Fatigue performance comparisons of two aluminum versions of 7055-T7751 in open hole configuration (R = 0.1) [15].

the proposed LLT specimens were manufactured with either the Huck Lockbolt (LGPL4SC-VO8-09AD) or the AeroLite (AL755-08-09) fasteners. The Fig. 18 shows the influence of different fasteners on joint fatigue test results for the same material using the LLT specimen under constant amplitude loading with

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Log fatigue life

5.70 5.60 Mean: 5.53

5.50 5.40 5.30

Mean: 5.31

5.20 5.10 5.00 4.90 7055-T7751 Standard Version

7055-T7751 High Fatigue Version

Fig. 17. Fatigue performance comparisons of two 7055-T7751 versions in LLT joint configuration at same constant amplitude cyclic stress (rmax = +95 MPa, rmin = 215 MPa) [15].

5.80 5.70

Log fatigue life

5.60

Mean: 5.59

5.50 5.40 5.30

Mean: 5.31

5.20 5.10 5.00 AeroLite 7055-T7751 STD

Lockbolt 7055-T7751 STD

Fig. 18. Fatigue performance comparisons of two fastener types in LLT joint configuration at same constant amplitude cyclic stress (rmax = +95 MPa, rmin = 215 MPa) [15].

Log fatigue life

5.30 5.20 Mean: 5.16 5.10 5.00

Mean: 5.09 Mean: 5.01

4.90

Mean: 4.87

4.80 4.70 C433-T39 C433-T39 AeroLite Lockbolt

C433-T351 AeroLite

C433-T351 Lockbolt

Fig. 19. Fatigue performance comparisons of two fastener types in HLT joint configuration at same constant amplitude cyclic stress (rmax = +200 MPa, rmin = 40 MPa) [15].

a negative stress ratio (rmin = 215 MPa and r max = +95 MPa) [15]. Other results shown in Fig. 19 compares the performances of the Lockbolt with the AeroLite fastening product using the HLT specimen (rmin = 40 MPa and rmax = +200 MPa), with two different experimental plate materials (C433-T39 and

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C433-T351). For the particular test conditions, it is shown that the Huck Lockbolt delivered better fatigue performance than the AeroLite fastener in both low and high load transfer joint configurations. This superiority was observed for three different aluminum alloys.

7. Summary and conclusions During this research, the design and the test protocol for two types of generic joint specimens, low load transfer (LLT) and high load transfer (HLT) specimen, were developed. The specimens were designed to be representative of aircraft joints and fabricated to the industrial practices. The fabricated generic joint specimens were then tested at Purdue University and at the Alcoa Technical Center for the purpose of evaluating the effect of different alloys and fastening systems on the fatigue characteristics of mechanical joints. The design and the allowable stress of a joint structure are governed by several complex parameters. Therefore, in order to fully investigate and demonstrate the fatigue performance of certain material and design detail, the joint test protocol rather than simple fatigue testing using coupon-type specimen, are required. The generic mechanical joint specimen and its test protocol presented in this paper successfully address this issue.

Acknowledgments Portions of this work were supported by the Alcoa Technical Center through a research project with Purdue University entitled ‘‘Durability of Aircraft Joints’’. The authors greatly appreciate the test data and support provided by Dr. G.H. Bray, Dr. R.J. Bucci, Dr. J.R. Yeh and Mr. H.R. Zonker from the Alcoa Technical Center. The early work on this project by Purdue graduate student D.J. Hahn is also much acknowledged.

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