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Failure analysis of the wing-fuselage connector of an agricultural aircraft
Engineering Failure Analysis 13 (2006) 572–581 www.elsevier.com/locate/engfailanal
Failure analysis of the wing-fuselage connector of an agricultural aircraft Lucjan Witek
*
Faculty of Mechanical Engineering and Aeronautics, Rzeszow University of Technology, 8 Powstan´co´w Warszawy Ave., 35-959 Rzeszo´w, Poland Received 10 December 2004; accepted 11 December 2004 Available online 31 March 2005
Abstract This paper presents an analysis of failure of the wing-fuselage connector for the agricultural aircraft. From the visual examination of the fractured surface, it was possible to observe beach marks, typical for fatigue failure. Careful observation showed that the crack origin surface was covered by corrosion products, therefore the failure had a character of fatigue combined with corrosion. A nonlinear finite element method was utilized to determine the stress state of the connector under operating condition. High stress zones were found at the region of the wing lug, where the failure occurred. Obtained results were next put into total fatigue life (S–N) and crack initiation (e–N) analyses performed for the load time history equivalent to 10-minute operating flight. In these analyses, the number of flight hours to the first fatigue crack and also to total damage of the wing lug were estimated. Moreover the influence of many factors such as: heat treatment, surface treatment, corrosion and overload on the fatigue life of the connector were analyzed. Ó 2005 Elsevier Ltd. All rights reserved. Keywords: Fatigue failure; Aircraft structures; Crack initiation; Fatigue life; FEM
1. Introduction The main wing-fuselage connector (Fig. 1) is one of the most critical parts of an aircraft. This connector joins the wing spar and the bulkhead of a fuselage. It transmits aerodynamic, inertial and gravity forces from the wing to the fuselage. Its life often limits the service life of the entire airframe due to large operational stresses. The agricultural operations often expose structural components to warm, humid and polluted air. The work environment of atomized chemicals has a negative influence on fatigue durability of *
Tel.: +48 17 865 1324; fax: +48 69 196 2047. E-mail address: [email protected].
1350-6307/$ - see front matter Ó 2005 Elsevier Ltd. All rights reserved. doi:10.1016/j.engfailanal.2004.12.029
L. Witek / Engineering Failure Analysis 13 (2006) 572–581
573
fuselage bulkhead upper wing lug
wing
upper fuselage lugs wing spar lower fuselage lugs lower wing lug (critical component) Fig. 1. Components of typical wing-fuselage connector of aircraft.
airframe components of agriculture aircraft. Corrosion can affect aircraft structural integrity since fatigue cracks can nucleate from corrosion pits and grow at an accelerated rate in the corrosive environment. It has been pointed out that fatigue combined with corrosion is one of the primary causes of aircraft fleet aging [1,2], thus encouraging studies related to this issue. Description and results of the investigations concerned with failure of aircraft components are presented in papers [3–13]. The problem of premature fracture failure of the wing-fuselage connector occurred in certain type of agricultural aircraft after about 5000–6000 h of operation. Failure of the connector occurred because of the growth to critical size of an undetected fatigue crack in the lower lug of wing-fuselage connector. In a few cases the failure of the connector was the reason for the crash. After these accidents, the producer of aircraft has given an order for detailed inspection of connectors in all planes which have more than 3000 flight hours operation. In a few cases both corrosion pits and small cracks were detected in the lugs of aircraft, which had a total time of 4000–5000 h of operation. Attention of this paper is mainly devoted to explain the reasons of low durability and premature failure of the wing-fuselage connector for the agriculture aircraft.
2. Visual examination of failed connector The fractured lug of the connector was first subjected to visual examination. The connector failure location is presented in Fig. 2. As seen from this figure, the wing lug (main component of connector) was fractured in two zones. The crack was initiated in the bottom zone of lug (region I), where the typical fatigue beach marks are observed (Fig. 3). After fatigue fracture of the bottom lug area occurred plastic bending and rupture of the top lug zone (region II). Fig. 3(a) presents the cracked lug surfaces with top and bottom fracture areas distinguished. Fig. 3(b) is a magnified view of the lower fracture surface. The curvature of the beach marks indicated that incremental
574
L. Witek / Engineering Failure Analysis 13 (2006) 572–581
Fig. 2. View of lower wing lug after damage.
Fig. 3. (a) Cracked lug surfaces and (b) magnified view of lower fracture surface.
crack growth had taken place from the corner of the lug fracture region (zone A in Fig. 3(b)). Careful observation showed that the crack origin surface was covered by corrosion products, thus the crack initiation and crack growth processes were accelerated by corrosion. The fractured surface showed a clear difference in colour in three different regions. The crack origin zone, the fatigue fracture area with presence of beach marks, and the ruptured zone are marked A, B and C respectively. The region, which finally ruptured, is dark grey (C).
3. Finite element stress analysis of the connector Using the Patran 2000r2 program [14], the geometry and mesh for all components of the model was created. The model consists of the following parts (Fig. 4):
L. Witek / Engineering Failure Analysis 13 (2006) 572–581
575
Fig. 4. Components of the connector before assembling (region I from Fig. 2 is in this model located in the top part of lug).
Lug connected to the wing spar (in which the fatigue fracture has appeared) (1); Double lugs connected to bulkhead of the fuselage (2); Expanding pin (3); Screw with conical head (4); Conical sleeve and the nut (5).
FE model of the connector consists of 12,460 first-order, 8-node brick elements. All components of the connector before assembling are shown in Fig. 4. To model interaction between adjacent surfaces of solids, 3D contact interface elements with the friction coefficient of 0.05 were used [14]. This value of the coefficient represents interaction of the components when they are lubricated. The lugs and the pin of the connector are made out of high tensile steel 30 HGSA (0.32C; 0.9Mn; 1.1Si; 0.9Cr; 0.3Ni according to Polish Standard PN 89/M-84030/04) with the following properties after heat treatment:
Ultimate tensile strength (UTS) – 1200 MPa, Yield stress – 1050 MPa, Young modulus – 210 GPa, Poisson ratio – 0.3.
The fuselage lugs (2) were fully restrained in the vertical cut-off plane, while a force of 126 kN was applied to the vertical cut-off section of the wing lug (1). Finite element code AFEA [14] was used for stress analysis of the connector. In the calculations, reported here, the nonlinear (incremental), Newton–Raphson method, with initial time step of 0.1 was applied. In all presented results, Megapascal (MPa) units were used to describe the stress fields.
576
L. Witek / Engineering Failure Analysis 13 (2006) 572–581
Fig. 5. Von Mises (a), and circumferential (b) stress distribution in the lug.
Fig. 5 shows the von Mises (a) and the circumferential (b) stress distributions in the lug. The zone of maximum von Mises stress (1144 MPa) was located in the left-top part of the hole. The area of the maximum circumferential stress was situated at the top part of the hole, where the tensile stress is equal to 739 MPa. The second result (Fig. 5(b)) is particularly interesting from the point of view of the fatigue strength because just the tensile circumferential stresses contribute most to the appearing of fatigue cracks and hence to damaging the connector. Fig. 6 shows the von Mises stress distribution in the horizontal cross-section of connector. As seen from this figure, the sleeve-screw assembly is bent. The bending effect causes additional stress concentrations on the edges of the wing lug (marked by circles).
Fig. 6. Von Mises stress distribution in the axle-section of connector. The stress field was displayed on the deformed model with displacement visualization of 20:1.
L. Witek / Engineering Failure Analysis 13 (2006) 572–581
577
Fig. 7. Von Mises stress on the edge (curve A–B) and in the internal zone of the wing lug (curve C–D).
The von Mises stress values as a function of the radial node position of lug, presented in Fig. 7 show that on the vertical cross section of the wing lug a non-even stress distribution occurred. For the internal zone of the lug the stress range is 100–570 MPa (curve C–D). The bending effect of the expanding pin caused the stress on the edge of the lug to increase to value of 1030 MPa (point B).
4. Crack initiation and fatigue life analysis To estimate the fatigue durability of the connector, the program MSC Fatigue ver. 9.0 was used. This program enables one to perform two main kinds of analyses: the total fatigue life (S–N) and the crack initiation (e–N), for non-limited geometry defined by user [15]. The load time history for the wing-fuselage connector, presented in Fig. 8 was defined on the base is of the simplified spectrum, which is equivalent to 10-minute operating flight, typical for the agricultural aircraft. This history consists of: take-off, a few working maneuvers over the field area and the landing. During the analysis, program Fatigue uses the procedures of ‘‘rain flow counting’’ and ‘‘linear damage summation’’
300
Load [KN]
250 200 150 100 50 0 1
4
7
10
13
16
19
22
25
28
31
Number of cycle
Fig. 8. Simplified load spectrum of connector for 10-minute operating flight of the agricultural airplane.
578
L. Witek / Engineering Failure Analysis 13 (2006) 572–581
[15–17] to transpose the non-symmetric time history with different levels of loads on the results of experimental standard tests performed for constant amplitude of load. In the S–N analysis presented here, the correction of mean stress according to Goodman theory additionally was applied. The result of the S–N analysis is presented in Fig. 9. The practical sense of this result is the minimum estimated life of 103.7032 h (5012 flight hours). The critical failure zone estimated here overlaps the area, where the maximum value of circumferential stress occurred (Fig. 5(b)). The e–N analysis, implemented into the MSC Fatigue program, known as ‘‘crack initiation analysis’’ [15] gives much more information than the S–N analysis. The influence of many factors such as: heat treatment, surface treatment, corrosion and overload on the time to crack initiation can be estimated using the e–N analysis. This kind of computation is based on results of e–N tests performed for standard specimens, which were made by different technological processes and tested in various environments. In the e–N calculations reported here, the correction of mean stress according to Smith–Watson–Topper and NeuberÕs elastic–plastic correction of strain were used. The result of e–N analysis for ground surface treatment of the lug is presented in Fig. 10. The estimated number of cycles to first fatigue crack (for crack depth about 1.5 mm [15]) is equal 103.6024 = 4003 flight hours. The results of e–N analysis for different surface treatments of lug and for its work in a corrosive environment are presented in Table 1. As seen from this table, polishing and the shot peening can increase the fatigue life by over 50–100%. Moreover the corrosion environment (i.e. chemical agents atomized by the agricultural aircraft and mixed with the rain water) has a huge influence on the decreasing of life. Sometimes, in the manufacturing process certain parameters such as temperature or speed of cooling for the heat treatment of the component are not observed. In these cases, material with different mechanical properties can be obtained and in consequence, the fatigue life is often lower. For example, if the UTS (ultimate tensile strength) of the 30 HGSA steel is only 1100 MPa (this is 100 MPa less than the value of UTS recommended by the designer of this aircraft), the fatigue life can decrease about 25%. Results of e–N analysis for 30 HGSA steel with the lover value of yield point and for the similar steels classified according to SAE standards are presented in Table 2. The specific type of duty of the agricultural aircraft is often overloaded. This often causes magnified spectrum of load. In this study, the influence of the under- and over-loading on the fatigue durability of the connector was also investigated.
Fig. 9. Result of the S–N analysis (zones of the FE model with determination of total fatigue life).
L. Witek / Engineering Failure Analysis 13 (2006) 572–581
579
Fig. 10. Results of e–N analysis – number of flight hours to first crack initiation and location of the critical fatigue region.
Table 1 Influence of the surface treatment and work in corrosion environment on the time to first crack initiation of the wing lug Surface treatment and environment of work
Number of flight hours
Nitrided Shot pened
25300 8122
Polished Ground Well machined Machined Badly machined Forged Cast
6248 4003 2365 1703 1251 245 199
Water corrosion Sea water corrosion
168 88
Table 2 Critical time of crack initiation for wing lug manufactured out of different sorts of steel 30 HGSA UTS = 1100 MPa
30 HGSA UTS = 1200 MPa
SAE 4130
SAE 4340
2896 h
4003 h
5299 h
9327 h
Fig. 11 shows results of the e–N analysis performed for the wing lug for different factors of overload. As seen from this plot, the wing-fuselage connector is very sensitive to exceeding the maximum take-off weight. Only 10% overloading decreases the fatigue life by about 40%. The high level of the operational stress in the lug (close to the yield point of material) is why the decrease in fatigue life at overload is so significant.
L. Witek / Engineering Failure Analysis 13 (2006) 572–581 Number of flight hours
580
18000 16000 14000 12000 10000 8000 6000 4000 2000 0 0.8
0.9
1
1.1
1.2
1.3
1.4
Factor of overload
Fig. 11. Influence of overloading of the connector on the time to crack initiation of lug.
5. Conclusions In this study the complex failure analysis of the wing-fuselage connector for the agricultural aircraft was carried out. The visual examination shows that failure was a typical fatigue fracture with beach marks. The crack origin surface was covered by corrosion products, thus the premature crack initiation process and crack growth were in this case accelerated by corrosion. It was evident, that the excessive stresses in the critical lug area were the main reason for premature fatigue failure. Moreover the corrosive environment has accelerated the cracking process. The main remark can be formulated based on the results of visual inspection and numerical investigation carried out through the work: the wing-fuselage connector should be re-designed for better fatigue durability.
6. Recomendations (a) (b) (c) (d) (e)
Increase the load-carrying areas of lug about 10–15%. Increase the bending stiffness of the expanded sleeve. Change the rotational position between the expanding pin and the wing lug. Introduce a frequent obligatory inspection of connector. Improve the heat treatment procedure in order to increase the toughness properties of 30HGSA steel or use a different sort of steel. (f) Introduce an anti-corrosion layer for lugs and pin surface. (g) Introduce an intermediate sleeve to reduce the local stress concentration in the wing lug.
Fig. 12. Wing-lug designed with fatigue resistance philosophy.
L. Witek / Engineering Failure Analysis 13 (2006) 572–581
581
An example of a fatigue-resistant wing lug, with special sleeve to reduce the excessive stress concentration in the critical lug area is presented in Fig. 12 [18]. The solution for the wing-fuselage connector was also described in [19].
References [1] Simpson DL, Hoffman PC. Corrosion and fatigue research – structural issues and relevance to aviation. Int J Fatigue 2001;23:1–10. [2] Simpson DL, Brooks CE. Tailoring the structural integrity process to meet the challenges of aging aircraft. Int J Fatigue 1999;21:1–14. [3] Withey PA. Fatigue failure of the De Haviland Comet I. Eng Fail Anal 1997;4:147–54. [4] Chen ChS, Wawrzynek PA, Ingraffea AR. Prediction of residual strength and curvilinear crack growth in aircraft fuselages. AIAA J 2002;40:1644–51. [5] Glodez S, Sraml M, Kramberger J. A computational model for determination of service life of gears. Int J Fat 2002;24:1013–20. [6] Vogwell J. Failure in lug joints and plates with holes. Eng Fail Anal 1995;2(2):129–35. [7] Lee HC, Hwang YH, Kim T-G. Failure analysis of nose landing gear assembly. Eng Fail Anal 2003;10:77–84. [8] Carboni M, Beretta S, Finzi A. Defects and in-service fatigue life of truck wheels. Eng Fail Anal 2003;10:45–57. [9] Giglio M. FEM submodelling fatigue analysis of complex helicopter component. Int J Fat 1999;21:445–55. [10] Bellinger NC, Komorowski JP, Benak TJ. Residual life predictions of corroded fuselage joints. Int J Fat 2001;23:349–56. [11] Urban MR. Analysis of the fatigue life of riveted sheet metal helicopter airframe joints. Int J Fat 2004;25:1013–26. [12] Kumar KS. Elastic plastic contact stress analysis of joints subjected to cyclic loading. Comput Struct 1996;60:1067–77. [13] Glodez S, Ren Z, Flasker J. Surface fatigue of gear teeth flanks. Comput Struct 1999;73:475–83. [14] MSC-Patran/AFEA. UserÕs manual – Modeling and analysis guidelines 2000. [15] MSC Fatigue 9.0. Quick Start Guide, Los Angeles, 2000. [16] Kocan´da S, Szala J. Fundamentals of fatigue calculation. Warszawa: PWN; 1985. [17] Anderson TL. Fracture mechanics. Fundamentals and applications. Boca Raton, FL: CRC Press; 1991. [18] Schijve J. Fatigue of structures and materials. Dordrhert: Kluwer Academic Publishers; 2001. [19] Witek L. Numerical stress and fatigue life analysis of the wing-fuselage connector for the agricultural aircraft. Technical Report (non-published), Rzeszow University of Technology, Poland, 2001.
Failure analysis of the wing-fuselage connector of an agricultural aircraft Lucjan Witek
*
Faculty of Mechanical Engineering and Aeronautics, Rzeszow University of Technology, 8 Powstan´co´w Warszawy Ave., 35-959 Rzeszo´w, Poland Received 10 December 2004; accepted 11 December 2004 Available online 31 March 2005
Abstract This paper presents an analysis of failure of the wing-fuselage connector for the agricultural aircraft. From the visual examination of the fractured surface, it was possible to observe beach marks, typical for fatigue failure. Careful observation showed that the crack origin surface was covered by corrosion products, therefore the failure had a character of fatigue combined with corrosion. A nonlinear finite element method was utilized to determine the stress state of the connector under operating condition. High stress zones were found at the region of the wing lug, where the failure occurred. Obtained results were next put into total fatigue life (S–N) and crack initiation (e–N) analyses performed for the load time history equivalent to 10-minute operating flight. In these analyses, the number of flight hours to the first fatigue crack and also to total damage of the wing lug were estimated. Moreover the influence of many factors such as: heat treatment, surface treatment, corrosion and overload on the fatigue life of the connector were analyzed. Ó 2005 Elsevier Ltd. All rights reserved. Keywords: Fatigue failure; Aircraft structures; Crack initiation; Fatigue life; FEM
1. Introduction The main wing-fuselage connector (Fig. 1) is one of the most critical parts of an aircraft. This connector joins the wing spar and the bulkhead of a fuselage. It transmits aerodynamic, inertial and gravity forces from the wing to the fuselage. Its life often limits the service life of the entire airframe due to large operational stresses. The agricultural operations often expose structural components to warm, humid and polluted air. The work environment of atomized chemicals has a negative influence on fatigue durability of *
Tel.: +48 17 865 1324; fax: +48 69 196 2047. E-mail address: [email protected].
1350-6307/$ - see front matter Ó 2005 Elsevier Ltd. All rights reserved. doi:10.1016/j.engfailanal.2004.12.029
L. Witek / Engineering Failure Analysis 13 (2006) 572–581
573
fuselage bulkhead upper wing lug
wing
upper fuselage lugs wing spar lower fuselage lugs lower wing lug (critical component) Fig. 1. Components of typical wing-fuselage connector of aircraft.
airframe components of agriculture aircraft. Corrosion can affect aircraft structural integrity since fatigue cracks can nucleate from corrosion pits and grow at an accelerated rate in the corrosive environment. It has been pointed out that fatigue combined with corrosion is one of the primary causes of aircraft fleet aging [1,2], thus encouraging studies related to this issue. Description and results of the investigations concerned with failure of aircraft components are presented in papers [3–13]. The problem of premature fracture failure of the wing-fuselage connector occurred in certain type of agricultural aircraft after about 5000–6000 h of operation. Failure of the connector occurred because of the growth to critical size of an undetected fatigue crack in the lower lug of wing-fuselage connector. In a few cases the failure of the connector was the reason for the crash. After these accidents, the producer of aircraft has given an order for detailed inspection of connectors in all planes which have more than 3000 flight hours operation. In a few cases both corrosion pits and small cracks were detected in the lugs of aircraft, which had a total time of 4000–5000 h of operation. Attention of this paper is mainly devoted to explain the reasons of low durability and premature failure of the wing-fuselage connector for the agriculture aircraft.
2. Visual examination of failed connector The fractured lug of the connector was first subjected to visual examination. The connector failure location is presented in Fig. 2. As seen from this figure, the wing lug (main component of connector) was fractured in two zones. The crack was initiated in the bottom zone of lug (region I), where the typical fatigue beach marks are observed (Fig. 3). After fatigue fracture of the bottom lug area occurred plastic bending and rupture of the top lug zone (region II). Fig. 3(a) presents the cracked lug surfaces with top and bottom fracture areas distinguished. Fig. 3(b) is a magnified view of the lower fracture surface. The curvature of the beach marks indicated that incremental
574
L. Witek / Engineering Failure Analysis 13 (2006) 572–581
Fig. 2. View of lower wing lug after damage.
Fig. 3. (a) Cracked lug surfaces and (b) magnified view of lower fracture surface.
crack growth had taken place from the corner of the lug fracture region (zone A in Fig. 3(b)). Careful observation showed that the crack origin surface was covered by corrosion products, thus the crack initiation and crack growth processes were accelerated by corrosion. The fractured surface showed a clear difference in colour in three different regions. The crack origin zone, the fatigue fracture area with presence of beach marks, and the ruptured zone are marked A, B and C respectively. The region, which finally ruptured, is dark grey (C).
3. Finite element stress analysis of the connector Using the Patran 2000r2 program [14], the geometry and mesh for all components of the model was created. The model consists of the following parts (Fig. 4):
L. Witek / Engineering Failure Analysis 13 (2006) 572–581
575
Fig. 4. Components of the connector before assembling (region I from Fig. 2 is in this model located in the top part of lug).
Lug connected to the wing spar (in which the fatigue fracture has appeared) (1); Double lugs connected to bulkhead of the fuselage (2); Expanding pin (3); Screw with conical head (4); Conical sleeve and the nut (5).
FE model of the connector consists of 12,460 first-order, 8-node brick elements. All components of the connector before assembling are shown in Fig. 4. To model interaction between adjacent surfaces of solids, 3D contact interface elements with the friction coefficient of 0.05 were used [14]. This value of the coefficient represents interaction of the components when they are lubricated. The lugs and the pin of the connector are made out of high tensile steel 30 HGSA (0.32C; 0.9Mn; 1.1Si; 0.9Cr; 0.3Ni according to Polish Standard PN 89/M-84030/04) with the following properties after heat treatment:
Ultimate tensile strength (UTS) – 1200 MPa, Yield stress – 1050 MPa, Young modulus – 210 GPa, Poisson ratio – 0.3.
The fuselage lugs (2) were fully restrained in the vertical cut-off plane, while a force of 126 kN was applied to the vertical cut-off section of the wing lug (1). Finite element code AFEA [14] was used for stress analysis of the connector. In the calculations, reported here, the nonlinear (incremental), Newton–Raphson method, with initial time step of 0.1 was applied. In all presented results, Megapascal (MPa) units were used to describe the stress fields.
576
L. Witek / Engineering Failure Analysis 13 (2006) 572–581
Fig. 5. Von Mises (a), and circumferential (b) stress distribution in the lug.
Fig. 5 shows the von Mises (a) and the circumferential (b) stress distributions in the lug. The zone of maximum von Mises stress (1144 MPa) was located in the left-top part of the hole. The area of the maximum circumferential stress was situated at the top part of the hole, where the tensile stress is equal to 739 MPa. The second result (Fig. 5(b)) is particularly interesting from the point of view of the fatigue strength because just the tensile circumferential stresses contribute most to the appearing of fatigue cracks and hence to damaging the connector. Fig. 6 shows the von Mises stress distribution in the horizontal cross-section of connector. As seen from this figure, the sleeve-screw assembly is bent. The bending effect causes additional stress concentrations on the edges of the wing lug (marked by circles).
Fig. 6. Von Mises stress distribution in the axle-section of connector. The stress field was displayed on the deformed model with displacement visualization of 20:1.
L. Witek / Engineering Failure Analysis 13 (2006) 572–581
577
Fig. 7. Von Mises stress on the edge (curve A–B) and in the internal zone of the wing lug (curve C–D).
The von Mises stress values as a function of the radial node position of lug, presented in Fig. 7 show that on the vertical cross section of the wing lug a non-even stress distribution occurred. For the internal zone of the lug the stress range is 100–570 MPa (curve C–D). The bending effect of the expanding pin caused the stress on the edge of the lug to increase to value of 1030 MPa (point B).
4. Crack initiation and fatigue life analysis To estimate the fatigue durability of the connector, the program MSC Fatigue ver. 9.0 was used. This program enables one to perform two main kinds of analyses: the total fatigue life (S–N) and the crack initiation (e–N), for non-limited geometry defined by user [15]. The load time history for the wing-fuselage connector, presented in Fig. 8 was defined on the base is of the simplified spectrum, which is equivalent to 10-minute operating flight, typical for the agricultural aircraft. This history consists of: take-off, a few working maneuvers over the field area and the landing. During the analysis, program Fatigue uses the procedures of ‘‘rain flow counting’’ and ‘‘linear damage summation’’
300
Load [KN]
250 200 150 100 50 0 1
4
7
10
13
16
19
22
25
28
31
Number of cycle
Fig. 8. Simplified load spectrum of connector for 10-minute operating flight of the agricultural airplane.
578
L. Witek / Engineering Failure Analysis 13 (2006) 572–581
[15–17] to transpose the non-symmetric time history with different levels of loads on the results of experimental standard tests performed for constant amplitude of load. In the S–N analysis presented here, the correction of mean stress according to Goodman theory additionally was applied. The result of the S–N analysis is presented in Fig. 9. The practical sense of this result is the minimum estimated life of 103.7032 h (5012 flight hours). The critical failure zone estimated here overlaps the area, where the maximum value of circumferential stress occurred (Fig. 5(b)). The e–N analysis, implemented into the MSC Fatigue program, known as ‘‘crack initiation analysis’’ [15] gives much more information than the S–N analysis. The influence of many factors such as: heat treatment, surface treatment, corrosion and overload on the time to crack initiation can be estimated using the e–N analysis. This kind of computation is based on results of e–N tests performed for standard specimens, which were made by different technological processes and tested in various environments. In the e–N calculations reported here, the correction of mean stress according to Smith–Watson–Topper and NeuberÕs elastic–plastic correction of strain were used. The result of e–N analysis for ground surface treatment of the lug is presented in Fig. 10. The estimated number of cycles to first fatigue crack (for crack depth about 1.5 mm [15]) is equal 103.6024 = 4003 flight hours. The results of e–N analysis for different surface treatments of lug and for its work in a corrosive environment are presented in Table 1. As seen from this table, polishing and the shot peening can increase the fatigue life by over 50–100%. Moreover the corrosion environment (i.e. chemical agents atomized by the agricultural aircraft and mixed with the rain water) has a huge influence on the decreasing of life. Sometimes, in the manufacturing process certain parameters such as temperature or speed of cooling for the heat treatment of the component are not observed. In these cases, material with different mechanical properties can be obtained and in consequence, the fatigue life is often lower. For example, if the UTS (ultimate tensile strength) of the 30 HGSA steel is only 1100 MPa (this is 100 MPa less than the value of UTS recommended by the designer of this aircraft), the fatigue life can decrease about 25%. Results of e–N analysis for 30 HGSA steel with the lover value of yield point and for the similar steels classified according to SAE standards are presented in Table 2. The specific type of duty of the agricultural aircraft is often overloaded. This often causes magnified spectrum of load. In this study, the influence of the under- and over-loading on the fatigue durability of the connector was also investigated.
Fig. 9. Result of the S–N analysis (zones of the FE model with determination of total fatigue life).
L. Witek / Engineering Failure Analysis 13 (2006) 572–581
579
Fig. 10. Results of e–N analysis – number of flight hours to first crack initiation and location of the critical fatigue region.
Table 1 Influence of the surface treatment and work in corrosion environment on the time to first crack initiation of the wing lug Surface treatment and environment of work
Number of flight hours
Nitrided Shot pened
25300 8122
Polished Ground Well machined Machined Badly machined Forged Cast
6248 4003 2365 1703 1251 245 199
Water corrosion Sea water corrosion
168 88
Table 2 Critical time of crack initiation for wing lug manufactured out of different sorts of steel 30 HGSA UTS = 1100 MPa
30 HGSA UTS = 1200 MPa
SAE 4130
SAE 4340
2896 h
4003 h
5299 h
9327 h
Fig. 11 shows results of the e–N analysis performed for the wing lug for different factors of overload. As seen from this plot, the wing-fuselage connector is very sensitive to exceeding the maximum take-off weight. Only 10% overloading decreases the fatigue life by about 40%. The high level of the operational stress in the lug (close to the yield point of material) is why the decrease in fatigue life at overload is so significant.
L. Witek / Engineering Failure Analysis 13 (2006) 572–581 Number of flight hours
580
18000 16000 14000 12000 10000 8000 6000 4000 2000 0 0.8
0.9
1
1.1
1.2
1.3
1.4
Factor of overload
Fig. 11. Influence of overloading of the connector on the time to crack initiation of lug.
5. Conclusions In this study the complex failure analysis of the wing-fuselage connector for the agricultural aircraft was carried out. The visual examination shows that failure was a typical fatigue fracture with beach marks. The crack origin surface was covered by corrosion products, thus the premature crack initiation process and crack growth were in this case accelerated by corrosion. It was evident, that the excessive stresses in the critical lug area were the main reason for premature fatigue failure. Moreover the corrosive environment has accelerated the cracking process. The main remark can be formulated based on the results of visual inspection and numerical investigation carried out through the work: the wing-fuselage connector should be re-designed for better fatigue durability.
6. Recomendations (a) (b) (c) (d) (e)
Increase the load-carrying areas of lug about 10–15%. Increase the bending stiffness of the expanded sleeve. Change the rotational position between the expanding pin and the wing lug. Introduce a frequent obligatory inspection of connector. Improve the heat treatment procedure in order to increase the toughness properties of 30HGSA steel or use a different sort of steel. (f) Introduce an anti-corrosion layer for lugs and pin surface. (g) Introduce an intermediate sleeve to reduce the local stress concentration in the wing lug.
Fig. 12. Wing-lug designed with fatigue resistance philosophy.
L. Witek / Engineering Failure Analysis 13 (2006) 572–581
581
An example of a fatigue-resistant wing lug, with special sleeve to reduce the excessive stress concentration in the critical lug area is presented in Fig. 12 [18]. The solution for the wing-fuselage connector was also described in [19].
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