Experimental Investigation on a Mechanically Fastened Splice Joint in a Composite Aircraft Wing D. Murali Krishna*, B. Ramanaiah, Polagangu James, Kotresh M. Gaddikeri, Byji Varughese, and M. Subba Raoa a
Advanced Composites Division, National Aerospace Laboratories, Bangalore, India ABSTRACT
Mechanically fastened splice joints like skin splice joints are essential in a large aircraft structure like wing to join inboard and outboard pieces of top and bottom skins. These joints transfer large direct loads and the joint behaviour is complex. The performance and behaviour of these joints need to be understood in the preliminary design stage of the wing itself to realize a safe structure. In the present work an experimental programme has been conducted on a representative specimen to study the behaviour and safety of a skin splice joint. The authors have discussed the features, testing and modeling of the test specimen in this paper. The behavior of the joint during the limit load and design ultimate load tests is discussed along with finite element analysis results. The comparison of results has showed a good correlation between test and FEM at most locations. The paper also discusses the possible causes of differences at a few locations between test and FE results and methods to qualitatively estimate the effects of these causes on the behavior of the specimen through FE modeling. Keywords: Composite wing, Skin splice joint
1. INTRODUCTION A large aircraft structure like wing is often made in multiple pieces like outboard and inboard or centre piece to achieve many advantages in fabrication and assembly to the fuselage structure. These pieces are joined through mechanically fastened splice joints on the primary load carrying members like top skin, bottom skin and spars. In a typical skin splice joint the top skin and the bottom skin of the inboard and outboard pieces are spliced at a rib location, where the rib flange acts as the splice plate. In a spar splice joint the inboard and outboard pieces of the spars are mechanically joined using splice plates. The skin splice joint predominantly transfers the bending stresses and the shear due to the load on the wing structure. The spar splice joint transfers the shear force across the out board and inboard pieces. Since the above spliced joints transfer often large loads and the joint behaviour is complex, safe design of them is very essential. The performance and behaviour of these joints need to be understood in the preliminary design. In the present work this is done through an experimental programme on test segments that represent the essential features of the skin splice joint. This work is a part of the ambitious composite wing development for the light transport aircraft programme at NAL. The testing programme fills up the requirement of element level testing in the building block approach. The building block approach is an important step toward the certification of composite aircraft structures. The finite element modeling and analysis of the joints is also considered as an essential part of the programme. It is aimed to establish reliable analysis techniques to simulate the joints and verification through a testing programme. In this paper, the authors share their experience at NAL on the modeling and testing of skin splice joint. The geometry of specimen, fabrication and test set up, modeling and analysis etc. are discussed in detail. Figure 1 shows the arrangement of outboard (LH and RH) pieces and centre piece of the wing for a typical medium sized transport aircraft. The fabrication of the centre and outboard pieces are done independently and the segments are joined at the splicing location as shown (station number 6) in the figure. *
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Figure 1. The three pieces (segments) – outboard and inboard(centre) – of a typical transport aircraft wing
2. GEOMETRY OF THE SKIN SPLICE TEST SPECIMEN The geometry of the skin splice test specimen was fixed such that the specimen represents the major features of the splice joint in the wing. In Fig. 2 the inboard and the outboard skin segments are shown forming a splice joint with the rib flange. The rib flange acts like a splice plate. The joint is eccentric like a single lap joint. A second splice plate can not be incorporated due to aerodynamic requirements on the wing external surface. At the joint the thickness of skin at the splice was increased locally to a value at which the joint fails only in bearing. Factor of safety in skin against bearing failure was kept the least of all failure modes like bearing, net tension, shear out etc. Thus, the thicknesses of the skin and rib flange were kept 8.64 mm from this consideration.
Figure 2. Geometry of the skin splice joint
The philosophy of designing the specimen was such that the test joint would see the same force that is seen at the splice location in real wing under the most critical load, i. e. the Vd (dive case). For this load the bottom skin of the wing is under tension. Moreover, there is a rise in tensile force in the bottom skin as there is a large cut out for landing gear close to the splice joint The width of the specimen was fixed as 202mm incorporating all design features of the joint such as thickness at the splice, diameter of fasteners and the pitch as in the wing structure. The nominal skin thickness close to the splice area in the skin was provided based on the first ply failure criterion as used in laminate analysis 1,2. Each skin segment in the specimen contains two stringers. The stringer ends are tapered (stringer run out) as in the figure. Mild steel lugs were placed at the ends of the specimen to be attached to the test rig and loading jack as shown in Fig.2. The ultimate load (Limit load x 1.5) on the specimen was 17658 Kg.
3.MATERIAL AND FABRICATION PROCESS The laminate for the test specimen was fabricated with the VERITy (Vacuum Enhanced Resin Infusion Technology) process. HS CARBON UD FABRIC G0827-B1040-HP03-1F was used as reinforcement and RTM120/HY2954 as the resin, both manufactured by M/s Hexcel composites. Design allowables and elastic properties of the material were evaluated through standard tests on coupons in room temperature and hot wet conditions. The base line data for lamina are given in Table 1. Thickness of cured ply is 0.16 mm. Table 1. Elastic and strength properties of lamina
Elastic Properties
Strength Properties
Modulus along Fibre Dir. E L =130 GPa
Tensile strength along Fibre Dir. = 585 MPa
Modulus along Transverse Dir. E T = 8GPa
Compressive strength along Fibre Dir. = 494 MPa
Shear Modulus, G LT = 3GPa
In plane Shear strength = 46 MPa
Poisson’s ratio, υ LT = 0.32 All metal components which are part of the test specimen were fabricated using mild steel with yield stress of 250 MPa.
4.EXPERIMENTAL SETUP
a
b
Figure 3 a. Schematic of test rig set up. b. Close up view of specimen on test rig
A test rig was set up with standard steel sections so as to accommodate the test specimen and the jack in the vertical position as shown in Fig. 3a. The test rig was supported on the ground through appropriately designed bracket. The
outboard end of the specimen with its metallic fittings was attached to the loading jack. The inboard end of the specimen was attached to the horizontal beam on the top of the test rig. The top attachment point on the beam and the bottom supporting point on the jack were nearly brought in the same vertical line during the test rig set up. Pin connections were provided at the ends of the specimen where it is attached to test rig and loading jack. The entire set up allowed vertical self alignment under loaded condition and hence, the possibility of bending of the specimen due to misalignment was minimized. The rib web part of the specimen was held laterally on to the vertical member of the test rig as in Fig. 3b. This was to closely simulate the connection of the rib web to the top skin in the wing structure. This rib support was provided in such a way that it allowed free axial movement (along load direction) and arrested lateral movement (normal to load) by providing sliding connection between skeleton structure and rib The test specimen was adequately strain gauged in the vicinity of the splice joint and away from the joint to capture the behaviour of the splice joint under the load. Since the specimen was subjected to axial tensile load, linear strain gauges were placed along the loading direction. Fig. 4 shows the locations of strain gauges on the skin part of the specimen. Strain gauges were also placed across the width (symmetrically about the center line) of the specimen to understand the strain variation across the width. Gauges were placed on either side of the skin as it was expected that the specimen would locally bend due to the eccentricity of the load path in skin and the rib flange at the splice location. A few gauges were placed at the stringer run out zone close to the joint. Deflection of the specimen was measured through a few dial gauges placed at the required locations. S1, S3, S5…., S17, S55, S56, S1
S3
S19 S20 S56 S21
S2
S25 S26 S58 S27
S4 S22 S23 S55 S24
S5 S6
S28 S29 S57 S30
S57 & S58 – Strain Gauges S7,S8
S13,S14
outside the skin S2, S4, S6..., S18- Strain
DG-1
DG-2
S9,S10
S15,S16
Gauges inside the Skin DG-3
S19 to S30
Strain Gauges
inside the skin at stringer run S11,S12
S17,S18
out area DG-1, DG-2, DG-3 - Dial Gauges
Figure 4. Locations of strain gauge and dial gauges.
5.FINITE ELEMENT MODELLING
Max. Displacement =2.86mm
a
b
Figure 5a. Isometric view of FE model of specimen. b. Deformed view (exaggerated) of specimen under load
The pre and post processing of FE model were carried out in HyperMesh7.0 and solving in NASTRAN. The finite element meshing was carried out using QUAD4 and TRIA3 elements in NASTRAN3 with 2D orthotropic layered shell properties for all composite parts and with isotropic shell properties for all metallic parts. Various parts of the test specimen (like skin, ribs and stringers) were modeled with 2D shell element. Since the eccentricity in load transfer between skin and rib flange at single lap joint varied for every step of loading, geometric nonlinearity was considered in analysis. The fasteners used at skin splice location were assumed as elastically rigid and modeled with Multi Point Constraint (MPC) equations. This simulated rigid behavior of fastener in extension, shear and bending under loading. The fasteners provided at remaining locations were modeled with BAR elements of NASTRAN with isotropic material (mild steel) property. Skelton structure supporting rib laterally was modeled with BAR elements. MPC equations were used to simulate sliding support condition between test rig and specimen. The FE model had a total of 32,580 degrees of freedom. The isometric view of FE model of specimen is shown in Fig. 5a.
6.RESULTS AND DISCUSSION The test was conducted in two phases. In the first phase the specimen was loaded up to limit load and then unloaded. Subsequently the specimen was subjected to non-destructive testing (NDT) based on an A-scan. The fastener holes were inspected for any deformation during the limit load test. It was found from the NDT that there was no delamination or any bearing failure at the fastener locations. The strains extracted from the FE analysis have been compared with strains obtained from the test. In general there is a good correlation between the FE and test strains at most locations. The trend observed in both results, both at high strain areas and low strain areas, is nearly same. This is true for gauges that are on either side (outside and inside) of the specimen. In order to understand the trend in the strains from FE analysis and test, the strains from typical gauges are presented in Table 2. At the limit load the maximum strains observed in the specimen is at S10 (2580microstrain). The comparison of strains in the test and FEM at the highest strain (S10 & S8) locations is quite good. Gauge locations away from the splice joint like S1, S2, S5 and S6 show low strains (800 to 900 microstrains) and show good comparison with FE results. This is true for gauges like S13 to S18 (1500 to 1800 microstrains) which are on the other side(outboard) of the splice. The strains on the outside and inside of the skin are nearly same. This shows that bending of skin at these locations is negligible. The FE model captured this effect. Table 2. Comparison of strain (microstrains) from Test and FEM at Limit Load
Strain Gauge S1 S5 S2 S6 S3 S4 S7 S11 S8 S12 S9 S10 S13 S17 S14 S18 S15
Strain, (test) 846 905 840 841 1113 757 851 894 2496 2145 783 2580 1631 1526 1625 1566 1630
Strain, (FEM) 863 863 828 828 929 888 838 838 2312 2312 735 2530 1550 1550 1568 1568 1685
% change w.r.t. test -2.01 4.64 1.43 1.55 16.53 -17.31 1.53 6.26 7.37 -7.79 6.13 1.94 4.97 -1.57 3.51 -0.13 -3.37
Strain Gauge S16 S19 S24 S20 S23 S21 S22 S25 S30 S26 S29 S27 S28 S55 S56 S57 S58
Strain, (test) 1764 555 556 954 870 785 555 821 703 1036 1049 768 869 911 851 1017 1038
Strain, (FEM) 1643 765 765 858 858 732 732 582 582 832 832 529 529 950 950 1090 1090
% change w.r.t. test 6.86 -37.84 -37.59 10.06 1.38 6.75 -31.89 29.11 17.21 19.69 20.69 31.12 39.13 -4.28 -11.63 -7.18 -5.01
Closer to the splice joint skin is subjected to local bending due to thickness variation in skin from joint location to nominal skin zone and eccentricity due to single lap joint. This effect is revealed in the strains at S7(outside) & S8 (inside) and S11(outside) & S12(inside). The outside strain is nearly three times more than the inside strain due to the
additional tensile strain developed due to local bending. This effect is well captured by FEM analysis. The local bending of the FE model under load can be seen in the exaggerated deformed view in Fig. 5b. Strain gauges, S19 to S30 are located on the skin at the stringer run out areas, close to the splice joint. The strain values observed in these locations are low. In majority of the locations there is a good correlation between the FEM and the test results. The trend observed is same. In the case of a few gauges there is considerable difference between the test and FEM strains (10% to 40%) and also there is difference between test results observed on symmetrically placed gauges (Examples S7 and S11, S8 and S12, S13 and S17, and S14 and S18) about load axis and structural symmetric axis. It is also noted that symmetric strain gauges S21 and S22 are showing different strain values (785 and 555 micro strains). S22 is showing less strain value than S21. Further, from test and analysis results, it has been noted that strain values at gauges provided on both outside and inside near stringer run out areas (S20&S55, S23&S56, S26&S57 and S29&S58) that skin is bending more in FEM than in test. During any real time structural testing, it is not always possible to simulate ideal conditions assumed in the theoretical simulations. And it is also not possible to simulate exact test conditions and uncertainties during test in theoretical models due to limitations coming from assumptions made during analysis. After careful analysis of results, the following possible causes have been identified for the mismatch of analysis and test results at some strain gauge locations. •
Small misalignment of load from symmetric axis of the test specimen
•
Minor errors in orientation and placement of strain gauges
•
Fastener hole clearances which allow local load redistribution at the joints4
•
Eccentricity values assumed in FE modeling
To have a qualitative feel of effects of the above causes, assumed quantities of above uncertainties have been incorporated in the FE model to estimate change in behaviour of the specimen and its effect on strain values. It has been observed from the study that each of these quantities either increases or reduces the difference between the test and FE results at those gauge locations. This study has helped to assess qualitatively the effects of the above quantities. It can be concluded that these effects would have affected the behavior of the specimen either independently or in combination. 4200
Test FEM
3500
strain, microstrains
strain, microstrains
4200
Test FEM
3500 2800 2100 1400 700
2800 2100 1400 700 0
0 0%
40% 80% 120% % of limit load
160%
a
0%
40%
80% 120% % of limit load
160%
b
Figure 6. Comparison of strains a. at S8 b. at S10 In the second phase of the test the specimen was loaded up to design ultimate load. The test specimen withstood the design ultimate load successfully without any major failure The NDT carried out after the test showed that a few fastener holes were elongated. Due to this the dial gauge that measured axial deformation showed some residual deformation even after the unloading. The bearing failure at the hole locations was quite local and was not of any major concern at the ultimate load. The strains were plotted at all gauge locations and most of them were linear up to the DUL. Two typical strain plots obtained from the ultimate load test, are shown in Figs. 6a and 6b. The maximum strain observed at the DUL is 3919 microstrains at S10. This is well below the allowable tensile strain of 4500 microstrains of the material. The linearity of strain gauges at most of locations shows that the bearing failure at holes is local in nature and does not effect the overall structural response of the specimen. This proved the integrity of the joint to transfer the load for which it was designed.
7. SUMMARY AND CONCLUSION Mechanically fastened splice joints skin splice joints are essential in a large aircraft structure like wing to join inboard and outboard pieces of top and bottom skins. These joints transfer large direct and shear loads and the joint behaviour is complex. The performance and behaviour of these joints need to be understood in the preliminary design stage to realize a safe structure. In the present work an experimental programme has been conducted on a representative specimen to study the behaviour and safety of a skin splice joint. Tests were conducted up to limit load and ultimate load done in two phases. From the NDT conducted after the limit load test it was found that there was no delamination or any bearing failure at the fastener locations. The strains extracted from the FE analysis have been compared with strains obtained from the test. In general there is a good correlation between FE and test strains at most locations. The specimen also withstood the design ultimate load and this showed the integrity of the joint to transfer the load. The NDT after the test showed that the specimen didn’t have any delamination or any serious damage. At the fastener locations the holes were elongated due to local bearing failure. The strain values from the test and FE analysis have showed good correlation at most locations. The results have showed that the FE model has captured the joint behavior quite well. For a complex joint like the skin splice joint it is quite difficult to exactly model the behaviour in the FE analysis. The study has helped to identify a few possible causes that affect the close correlation of behavior in test and FE analysis.
ACKNOWLEDGMENT The authors wish to thank the Director, NAL for his valuable support in carrying out this work. Authors thank Mr. H.V. Ramachandra and Mr. Sanjeev Kumar S. for their support in conducting the test and the NDT.
REFERENCES 1. R. M. Jones, Mechanics of Composite Materials, Taylor and Francis, Inc., USA, Second Edition, 1999. 2. M.C.Y. Niu, Composite Airframe Structures, Conmilit Press Ltd, Hong Kong, 1992. 3. Robert S. Lahey, Mark p. Miller and Michael Reymond, MSC/NASTRAN Version 68 Reference Manual, The MacNeal-Schwendler Corporation , 1995.. 4. D. Murali Krishna, Kotresh M. Gaddikeri, Byji Varughese and M. Subba Rao, Effect of Bolt Hole Clearance on the Load Redistribution in Multi-Fastener Composite Joints, INCCOM-5, November, 24-25, 2006.
