Pressurized Aircraft Fuselage Research Articles

Effect of hexagonal stringer design on bulging factor and stress intensity factor of cracks in the skin of pressurized aircraft fuselage

Aircraft structural fatigue is a serious issue. Untreated, it could lead to failure. Several aircraft accidents were caused by widespread fatigue damage. For cracks in the fuselage, the stress state is due to the bulging around the crack caused by applied internal pressure. This condition is characterized by a parameter called the bulging factor; it compares the stress intensity factor of a crack in curved shell to its counterpart in a plate. Bulging factor expressions are available for longitudinal and circumferential cracks but less for slanted cracks. The underlying structure that stiffens the skin of the fuselage has a direct impact on the stress intensity factor and the bulging factor of the crack. A hexagonal grid stiffening pattern has been shown to provide sufficient stiffening to the fuselage skin while using a lesser amount of material compared to the traditional orthogonal grid design. However, the response of this grid pattern in the presence of a cracked fuselage has not been studied. The current paper aims to estimate the effect of the hexagonal grid pattern on the bulging factor and stress intensity factor of cracks of various lengths and orientations in the skin of the fuselage. Results obtained are compared to the conventional orthogonal grid stiffening pattern of a fuselage structure. Several patterns were considered. Results show that cracks of different lengths and orientations in hexagonal grid stiffened panels had stress intensity factors and bulging factors that are comparable to the base case within a margin of 2%–8%.

Mechanical response of a panel section with a hexagonally tessellated stiffener grid

Stringers are stiffening members of pressurized aircraft fuselage. They provide support to the fuselage’s skin. A new stringer grid concept is proposed for conventional aircraft fuselage. Optimization is used to find the hexagonal grid that best replaces the original while keeping the same total stringer length. A finite element model is built to analyze the optimal hexagonal grid stiffened structure and compare it with the original orthogonally stiffened structure in terms of eigenfrequencies and static response to external loading. The finite element model is validated through Flugge’s analytical expressions for stiffened shells. Results show that the hexagonal grid stiffened structure yields higher eigenfrequencies with stresses and displacements comparable with that of the original structure.

Blast Response of a Pressurized Aircraft Fuselage Measured Using High-Speed Image Correlation

As part of a project examining the blast response of a pressurized aircraft fuselage, high-speed digital image correlation was used to measure the change in surface strain and out-of-plane displacement during an explosive event. Two explosive charges were used for this test, one located in the aft cargo compartment and one located in the forward cargo compartment. Digital image correlation was used to monitor the charge in the aft compartment with the second charge in the forward compartment being detonated 100 ms after the first. Digital image correlation was able to monitor the change in out-of-plane displacement as a result of the high-temperature gas generated during the explosion. Strain measurements made during the explosion were used to calculate the Von Mises stresses and confirmed that the aluminum skin did not plastically deform. The second explosion in the forward baggage compartment did result in fracture of the fuselage skin and post-test analysis of the fracture surfaces showed that tensile stresses in the rivets likely led to failure.

Rapid decompression of pressurized aircraft fuselages

A methodology for determining compartment pressures at any point during a rapid decompression of a pressurized aircraft fuselage is presented. The approach takes into consideration the mass and mass moment of inertia of panels and doors through which pressurized air must pass. Failure to consider vent panel mass is shown to severely underestimate the pressure differential across partitions separating compartments during a rapid-decompression event.

Discrete crack growth analysis methodology for through cracks in pressurized fuselage structures

AbstractA methodology for simulating the growth of long through cracks in the skin of pressurized aircraft fuselage structures is described. Crack trajectories are allowed to be arbitrary and are computed as part of the simulation. The interaction between the mechanical loads acting on the superstructure and the local structural response near the crack tips is accounted for by employing a hierarchical modelling strategy. The structural response for each cracked configuration is obtained using a geometrically non‐linear shell finite element analysis procedure. Four stress intensity factors, two for membrane behaviour and two for bending using Kirchhoff plate theory, are computed using an extension of the modified crack closure integral method. Crack trajectories are determined by applying the maximum tangential stress criterion. Crack growth results in localized mesh deletion, and the deletion regions are remeshed automatically using a newly developed all‐quadrilateral meshing algorithm. The effectiveness of the methodology, and its applicability to performing practical analyses of realistic structures, is demonstrated by simulating curvilinear crack growth in a fuselage panel that is representative of a typical narrow‐body aircraft. The predicted crack trajectory and fatigue life compare well with measurements of these same quantities from a full‐scale pressurized panel test.

Growth of multiple cracks and their linkup in a fuselage lap joint

An issue of concern in aging aircraft is the growth of multiple cracks emanating from a row of fastener holes, typically in a pressurized aircraft fuselage lap splice. This multisite damage (MSB), or widespread fatigue damage, if allowed to progress, can suddenly become catastrophic. The understanding of the failure behavior dictates the level of compromise between safety and economy. The complexity of the structure due to various stiffening elements makes it unamenable to a simple direct analysis. A two-step elastic finite element fatigue analysis combining a conventional finite element method and the Schwartz-Neumann alternating method with analytical solutions is developed to understand fatigue growth of multiple cracks and to obtain a first estimate of the residual life of a stiffened fuselage shell structure with MSD in the riveted lap joint. The analysis procedure is validated by simulating a laboratory fatigue test on a lap joint in a flat coupon. Both the coupon and the shell panel are found to have fatigue lives only up to the first linkup of neighboring crack tips.

Fatigue growth of multiple-cracks near a row of fastener-holes in a fuselage lap-joint

The fatigue growth of multiple cracks, of arbitrary lengths, emanating from a row of fastener holes in a bonded, riveted, lap joint in a pressurized aircraft fuselage is studied. The effects of residual stresses due to a rivet misfit, and of plastic deformation near the hole, are included. A Schwartz-Neumann alternating method which uses the analytical solution for a row of multiple colinear cracks in an infinite sheet (the crack-faces being subject to arbitrary tractions), is developed to analyze this MSD problem on a personal computer. It is found that for a range of crack lengths, a phenomena wherein the shorter cracks may grow faster than longer cracks may exist.

Practical use of the ‘equivalent’ measured stress intensity factor to control fatigue crack propagation rates in aircraft full-scale fatigue tests—First assessment of the method in testing of a pressurized aircraft fuselage

In the case of circumferential cracks in a cylindrical fuselage, the comparison of some analysis and test results shows that the theoretical stress intensity factor is a suitable correlation parameter of fatigue crack propagation rates, both in aircraft fuselages and in plane panels. Values of the ‘equivalent’ stress intensity factor, computed by applying the Barrois-Bhandari method to slot-opening measurements performed under decreasing loading levels, agree well with the values computed from two dimensional Theory of Elasticity, using the method of finite elements. In the case of longitudinal cracks, the experimental values of the ‘equivalent’ stress intensity factor, i.e. the stress intensity factor of the infinite plane sheet containing a centre crack with the same elastic strain and stress distributions near the boundary of the plastically strained region around the crack tip, yield a good correlation of fatigue crack propagation rates of the cracked fuselage and of cracked plane structures. The values of the ‘equivalent’ stress intensity factor are lower than those of the theoretical stress intensity factor, the interest of which disappears, but are also far higher than the bidimensionally computed values, which are no longer to be considered. Some meant of safety provided to limit crack openings will make it possible, in the near future, to investigate test conditions reaching the ultimate residual static strength of cracked structures, while avoiding, however, catastrophic failures.