Skin Splices Research Articles

The Effects of Fatigue Cracks on Fastener Loads during Cyclic Loading and on the Stresses Used for Crack Growth Analysis in Classical Linear Elastic Fracture Mechanics Approaches

High strength threaded fasteners are widely used in the aircraft industry, and service experience shows that for structures where shear loading of the joints is significant, like skin splices, fuselage joints or spar caps-web attachments, more cracks are initiated and grow from the edges of the fastener holes than from features like fillets radii and corners or from large access holes. The main causes of this cracking are the stress concentrations introduced by the fastener holes and by the threaded fasteners themselves, with the most common damage site being at the edge of the fastener holes. Intuitively, it is easy to visualize that after the crack initiation, during the growth stages, some of the load transferred initially by the fastener at the cracked hole will decrease, and it will be shed to the adjacent fasteners that will carry higher loads than in uncracked condition. Using currently available computer software, the method presented in this paper provides a relatively quick and quantitatively defined solution to account for the effects of crack length on the fastener loads transfer, and on the far field and bypass loads at each fastener adjacent to the crack. At each location, these variations are determined from the 3-dimensional distribution of stresses in the joint, and accounting for secondary bending effects and fastener tilt. Two cases of a typical skins lap splice with eight fasteners in a two rows configuration loaded in tension are presented and discussed, one representative for wing or fuselage skins configurations, and the second case representative for cost effective laboratory testing. Each case presents five cracking scenarios, with the cracks growing from approx. 0.03 inch to either the free edge, next hole or both simultaneously.

Realisation of Shear Flow at Crucial Spar Splice Joints of Composite Wing in Idealised Wing Test Box

CSIR -National Aerospace Laboratories (CSIR-NAL) has developed a patented processing technique called “Vacuum Enabled Resin Infusion Technology (VERITy” for the manufacturing of composite wing for its transport aircraft programme. The building block approach was adopted during the design stage of the composite wing to understand the structural response in increasingly complex structural levels. In the subcomponent level, a wing test box was designed, fabricated and was subjected to structural static testing for critical loads of the wing. The wing test box which was idealized into a rectangular box included the most critical joints such as top & bottom skin splice joints and front & rear spar splice joints. The stacking sequence and thickness of laminates used in the respective positions in wing test box replicated the wing structure. The challenge was to maintain the shear flow despite differences in the geometry of wing and test box without compromising on the magnitude of shear force at the critical spar splice joint. This paper discusses the methodology adopted for transforming shear flow from wing structure on to the idealized wing test box

The torsion of the multicell sections

The paper is focused on the stress analysis of thin-walled multicell sections subjected to pure torsion. The shear flow and stiffness characteristics of the cross section for torsion are given. Example: aircraft wing section. The theory for thin-walled closed sections used in this paper was developed by Bredt (3). The shear flows obtained are used in the design of skins and interior webs, ribs and fasteners at skin splices, skin web junctions and the joints where the ribs meet the skin or webs.

Equivalent initial flaw size testing and analysis of transport aircraft skin splices

ABSTRACTThe equivalent initial flaw size (EIFS) concept was developed nearly 30 years ago in an attempt to account for the initial quality, both manufacturing and material properties, of a structural detail prone to fatigue cracking. Widespread use of this concept has been limited due to the large amount of test data required to develop a reliable EIFS distribution. In this effort, an EIFS distribution was determined for four types of flat, production like transport aircraft fuselage skin joints loaded by remote tension. Two crack growth prediction codes, AFGROW and FASTRAN, were used to not only develop the EIFS but also to compare the crack growth algorithms in each code. The EIFS calculations are prone to compounding errors in the crack growth analysis due to the changing stress intensity factor solutions and stress fields as the crack gets longer. Thus, only including EIFS calculations for mechanically small cracks, crack lengths less than 1.27 mm, results in a mean EIFS of 18.0 μm with a standard deviation of 3.78 μm.

Residual strength analysis using CTOA criteria for fuselage structures containing multiple site damage

An extensive experimental program was conducted by the Boeing Company under the funding of the Federal Aviation Administration (FAA), National Aeronautics and Space Administration (NASA), and the United States Air Force Research Laboratory (USAF/RL) to investigate the effects of multiple-site damage (MSD) on the residual strength of typical fuselage splice joints. The experimental results were used to validate the analytical prediction using various methodologies, including STAGS (a generalized shell finite element code) with the crack-tip-opening angle and T * fracture criteria. The test specimens consisted of large flat panels, curved panels, and an aft pressure bulkhead. The flat panel specimens included three types of longitudinal splice joints and one type of circumferential splice joint. For each type, one panel contained only a lead crack and the other two panels contained MSD 1.3 and 2.5 mm in size, respectively, at the fastener holes ahead of the lead crack. The curved panels were tested under simulated loads of combined cabin pressure and fuselage down bending. Two skin splice types were tested. For each splice type, one panel contained a lead crack only and the other had a lead crack with various sizes of MSD. A section of an aft fuselage containing a large lead crack and MSD in the pressure dome was also tested to demonstrate the capabilities of the methodologies in analyzing actual aircraft structures. This paper presents the analytical approaches and the comparison of predictions with the experimental results in terms of crack linkup stress and residual strength.

An Engineer'S Viewpoint on Design and Analysis of Aircraft Structural Joints

The design of structurally efficient joints in aircraft fuselage structures and wing skin splices is addressed. It is contended that the joints should be designed first and the gaps in between filled in afterwards, taking pains not to optimize the basic structure first and then discover that it either cannot be assembled or that, when it is assembled, it is full of weak-link fuses. Both adhesively bonded and mechanically fastened joints are covered. Analogies are drawn between the characteristics of both classes of joints. The aspects of static joint strength and fatigue lives are included. The work is applicable to metallic as well as composite structures, and covers both high-load wing joints which have already been tested and new ideas for fuselage splices which have not. The effects of flaws and defects are associated with the need for damage tolerance, particularly in fuselage structures.

Buckling and vibration of externally pressurized conical shells withcontinuous and discontinuous rings

External pressure buckling, free vibration, and external pressure-vibration interaction data were obtained on a sixth scale Lexan model of the Titan IV (prototype) nose cone fairing. The objectives of these experiments were to validate analytical models used in the prototype design and to provide experimental data on the effects of discontinuous circumferential rings, flexible ring-to-shell attachment, separation rails, and skin splices. In addition, a nondestructive buckling load predictive test technique based on natural frequencies was evaluated using the Lexan model. The advantage of a Lexan model is that it can be buckled without sustaining damage, thus allowing repeated buckling experiments to be performed. Correlation studies were carried out using NASTRAN finite element and BOSOR finite difference models. Buckling and frequency data showed good agreement between analysis and experiment. Pressure-vibration interaction data were obtained to investigate an experimental procedure for extrapolating the interaction curve to predict the critical buckling load. While it was found that extrapolation of these data always over predicted the buckling pressure, the data were used to develop a procedure that was successfully applied as stop-test criteria in the prototype test.

Assurance of Continued Safe Operation for Ageing Structure

When an operated period of a structure is close to or over the original design life, it is necessary to reassess the regular inspection program applied to secure the structural integrity. Because, the program is effective only to detect damages in the structural details which occur within the design life. Therefore, another inspection program, that is, a supplemental inspection program which is intended to detect significant damages occurring after the design life with high detection capability, is required for the sake of continuing safe operation of ageing structure.As to aircraft structures, FAA and ICAO have recently imposed to develop a supplemental inspection document of older aircrafts on manufacturers. In response to the requirements, the Boeing company has released the supplemental inspection documents for Model 707/720 and 727.The purposes of the present study are ; 1) to define deterministic and probabilistic factors which are necessary in a reliability analysis on the effects of supplemental inspection, 2) to examine results obtained by the Boeing method, and 3) to propose a method to estimate possible extension of service life resulting from a supplemental inspection.A numerical example of typical aircraft structural significant detail, which is a wing spanwise skin splice consisting of skin and stringer, is given in order to evaluate the validity of the present study.