Aeronautics and astronautics – Aircraft structure – Fuselage and body construction
Reexamination Certificate
2000-11-30
2003-07-22
Eldred, J. Woodrow (Department: 3644)
Aeronautics and astronautics
Aircraft structure
Fuselage and body construction
C244S132000
Reexamination Certificate
active
06595467
ABSTRACT:
FIELD OF THE INVENTION
The invention relates to a structural component and particularly a fuselage shell component for an aircraft, including a skin sheet and several stiffening profile members connected to the skin sheet at least partially by welding.
BACKGROUND INFORMATION
In the construction of aircraft fuselages, it has become known to connect the stiffening profile members, such as frame members and stringers, to the outer fuselage skin by means of welding, at least partially or at certain locations of the aircraft. For example, see German Patent Publication DE 196 39 667 and corresponding U.S. Pat. No. 5,841,098, or German Patent Publication DE 198 44 035. Particularly, the stringers and frame members are welded onto large format skin sheets by means of laser beam welding, so as to fabricate structural components in the form of fuselage shell components that are assembled together to form the fuselage of the aircraft.
Such fuselage shell components must have a sufficient strength and stiffness to support the ordinary operating loads applied to the aircraft fuselage, as well as extreme loads applied under unusual conditions, and a further safety margin or safety factor beyond such loads. Particularly in the future, fuselage shell components will have to satisfy a so-called “two bay crack” criteria. Namely, the fuselage shell structure will have to be able to withstand a longitudinally or circumferentially extending crack that spans or extends over two frame sections or two stringer sections (i.e. the crack extends into two bays), without resulting in a failure of the overall shell structure. In this context it is further to be assumed that the stiffening profile member at the middle of the crack is broken. Thus, the remaining structure of the fuselage shell must be able to withstand the requisite loads, without failing.
In the previously typical construction, the stiffening profile members, such as frame members and stringers, were connected to the skin sheets by riveting or adhesive bonding. Such a joining method of the stiffening profile members onto the skin sheets is disadvantageous in comparison to laser welding, because the riveting and adhesive bonding result in a greater total weight, and involve greater costs and efforts in the fabrication procedures. On the other hand, the structure resulting from such rivet connections or adhesive bonding of the stiffening profile members onto the skin sheets provides a greater residual strength and a better crack stopping characteristic (i.e. resistance to crack propagation) than a corresponding shell structure in which the stiffening profile members have been laser welded onto the skin sheets.
Particularly, with a riveted or adhesively bonded junction between the skin sheets and the stiffening profile members, a crack that initiates in the skin sheet and progresses to a location of a stiffening profile member will generally not propagate into the stiffening profile member itself, because the rivets or adhesives do not provide the necessary degree of local force coupling to transmit the crack into the stringer or frame member. Thus, while the crack in the skin sheet might propagate past the location of a stringer or frame member, it does not directly damage the associated stringer or frame member. Therefore, the respective stringer or frame member maintains its original strength and holds together the skin sheet through the rivets or adhesive on opposite sides of the crack, thereby inhibiting the propagation of the crack.
The respective stiffening profile member is able to maintain this condition for a certain number of load alternations, until the extra loading transmitted from the skin into the stiffening profile member eventually fatigues and overloads the profile member, leading to a failure of the respective stiffening profile member. At that point, the fuselage skin and the affected stiffening profile member will fail, typically in a sudden rupturing manner, which leads to a failure of the fuselage shell structure. However, the fact that the stiffening profile member maintains its integrity and load-carrying ability even after a crack has formed in the adjoining skin sheet, generally allows the aircraft to fly safely to a landing, whereupon the crack defect in the skin sheet can be detected and repaired.
The above described advantageous property of crack propagation resistance or inhibition is not generally achieved by fuselage shell structures in which the stiffening profile members are welded onto the skin sheets. Namely, with such a welded junction, any crack that forms in a skin sheet and propagates to the junction of a stiffening profile member will directly propagate through the welded joint into the stiffening profile member, where the crack will then propagate further into or even entirely through the stiffening profile member. Since there is no effective interruption between the skin sheet and the stiffening profile member, there is no “crack stopping” effect which would prevent the crack from propagating into the respective stringer or frame member. As a result, any crack in the skin sheet will readily propagate through the stringers and frame members as well, which leads to a significantly lower residual or remaining strength of the overall fuselage shell structure upon the occurrence of such a crack. Namely, once such crack forms, it will readily propagate through both the skin and the stiffening profile members, and there is no structural component remaining to hold together the fuselage shell at the location of the crack, thus leading to a failure of the overall shell structure.
In view of the above, the shell structure components would have to be thickened and thereby strengthened in areas of the aircraft fuselage in which the post-crack residual strength is the predominant design criterium, in order to achieve an adequate residual strength in such areas. These areas especially include the sides and the upper or top portion of the fuselage, since these areas are especially subjected to tension loads during operation, with a consequent tendency toward crack opening and propagation. Such thickening of the fuselage shell in these areas would, however, lead to an unacceptable increase in the overall weight of the fuselage. For these reasons, prior aircraft fuselages have not used welded stringers in these areas at the sides and top of the fuselage, but instead have only used welded stringers, for example, in the bottom or belly of the fuselage, while using riveted or adhesively bonded stringers on the sides and top of the fuselage.
German Patent DE 199 24 909 has further disclosed a fuselage shell component in which each stiffening profile member includes an integral thickening at a location adjacent to the base or root of the profile member at which the profile member is welded onto a skin sheet. The ratio of the thickness of this thickening or protruding portion of the profile member relative to the thickness of the root of the profile member that is welded onto the skin sheet is at least two to one. The protruding portion or thickening is an integrally formed portion of the same material as the rest of the profile member. The object of this thickened portion or protrusion is to stop the propagation of any crack that might progress from the skin sheet through the welded junction into the base or root of the profile member. Thus, even if the crack propagates into the base or root of the profile member, it shall not propagate further beyond the thickened protrusion into the rest of the profile member. This provides a crack propagation stopping characteristic as well as an improved residual strength of the fuselage shell structure after a crack has formed in the skin. While such an integral protrusion or thickening of the stiffening profile member aims to provide a certain degree of crack stopping performance, it has been found that further improvements are possible.
SUMMARY OF THE INVENTION
In view of the above, it is an object of the invention to provide a shell structural component and particularly an aircr
Airbus Deutschland GmbH
Eldred J. Woodrow
Fasse W. F.
Fasse W. G.
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