Static structures (e.g. – buildings) – Imperforate panel with integral reinforcing
Reexamination Certificate
2001-02-22
2004-02-03
Friedman, Carl D. (Department: 3635)
Static structures (e.g., buildings)
Imperforate panel with integral reinforcing
C244S119000
Reexamination Certificate
active
06684593
ABSTRACT:
FIELD OF THE INVENTION
The invention relates to a structural panel component, and particularly a curved structural shell component that is especially for an aircraft fuselage, and that includes at least one skin sheet as well as stiffening elements thereon. The invention further relates to a method of manufacturing such a structural component.
BACKGROUND INFORMATION
The present day conventional fabrication of large format structural components, and particularly the fuselage shell components for aircraft, typically uses skin sheets having dimensions of approximately 2.5 m by 10 m. The maximum largest possible size of the skin sheets is typically used for fabricating the fuselage shell, in order to minimize the number of the longitudinal and circumferential or transverse joints of the finished fuselage, and thereby reduce the weight and the fabrication effort and expense of the aircraft fuselage. Minimizing the structural weight of the fuselage is a very important factor in the manufacture of aircraft, with regard to the ever present effort to reduce the fuel consumption and therefore improve the economy of operating the aircraft.
Such a weight minimization of the fuselage requires a structure-mechanical optimization of the fuselage shell structure, so that, for example, the skin sheets must have different thicknesses at different areas depending on the respective loads that will prevail at each respective area. This aims to avoid the weight penalty of otherwise having an unnecessary excess thickness at any given area of the fuselage shell that will be subjected to below-average loading. Similarly, areas of the fuselage shell that will be subjected to above-average loading can be provided with a thicker fuselage skin (and/or stronger stringers and frames), without unnecessarily thickening other areas of the fuselage shell. For example, it may be necessary to provide a thickening of the fuselage skin metal sheet in the area of each stringer joint for proper load introduction.
The various different thickness areas of the skin sheet are typically conventionally achieved by riveting or adhesively bonding a doubling or thickening sheet onto the base skin sheet. Alternatively, a reduction of thickness of the fuselage skin can be achieved by mechanical milling or by chemical milling and material removal at these areas. This process of chemical milling and material removal of a skin sheet is carried out by masking the skin sheet, cutting and partially removing portions of the mask at areas that are to be etched, and then applying an etching chemical to remove material from the exposed surface areas, followed by neutralizing, cleaning, etc.
In order to strengthen and support the skin sheet and thereby fabricate the structural shell component, longitudinally extending stringers are typically riveted or adhesively bonded onto the fuselage skin. Then, crosswise or circumferentially extending frames are joined onto the structure being fabricated, by first riveting angle elements, i.e. so-called clips, onto the skin sheet and the stringers. Then the frames are joined onto the clips.
The above described conventional fabrication process for manufacturing a fuselage shell structure is rather complicated, time consuming, and costly. Such a process is described in greater detail in an article by Peter Heider entitled “Lasergerechte Konstruktion und lasergerechte Fertigungsmittel zum Schweissen grossformatiger Aluminium-Strukturbauteile” (“Laser Compatible Construction and Laser Compatible Manufacturing Means for Welding Large-Format Aluminum Structural Components”), published by the VDI Verlag Publishers in VDI Fortschritt-Berichte (“Progress Reports”), Series 2: Fertigungstechnik (“Manufacturing Technology”), No. 326, Dissertation July/1994, especially at pages 3 to 5.
German Patent DE 198 44 035 discloses a method of fabricating large format structural components, as well as different possible manners of construction of a structural component, which use a laser beam welding process for joining a stiffening structure onto the fuselage skin. Particularly, the stringers running in the longitudinal direction of the aircraft are laser welded onto the skin sheet. From this prior art reference it is also known to embody the frames or frame elements that run in the crosswise or circumferential direction in a weldable manner. Thus, the overall structural component is realized predominantly by welding together numerous individual parts.
FIG. 3
of the present application shows a representative example of a conventionally fabricated aircraft fuselage shell
10
, including a skin sheet
11
with stringers
12
arranged and joined thereon. Particularly, the stringers
12
may be riveted or adhesively bonded onto the skin sheet
11
. Before carrying out such an adhesive bonding process, all of the individual parts must be subjected to a special pretreatment process, including degreasing, cleaning, pickling or etching, anodizing in a chromic acid solution, and finally being coated with a primer, before the actual adhesive application can take place. Then, a suitable adhesive is applied to the components, which are joined and clamped or held together by appropriate jigs, and then the adhesive is allowed to cure at an elevated temperature and pressure in an autoclave for a sufficient amount of time to ensure adequate and proper bonding of the stringers
12
onto the skin sheet
11
. Thereafter, the bonded components must be cleaned, and excess adhesive must be removed. Further, the adhesive joints must be protected against corrosion by applying a bead of an appropriate sealant, and then an additional protective coating to protect against attack by aggressive media.
On the other hand, the above mentioned riveting process for joining the stringers
12
onto the skin sheet
11
is similarly complicated, time-consuming, and costly. Namely, the components must be subjected to a rather complicated preparation process, especially for achieving adequate surface protection. Namely, the skin sheet
11
and the stringers
12
must be anodized, coated with a primer, cleaned with an activator along the joint surfaces, and provided with a sealant along the joint surfaces. Also, a complicated rivet hole boring, cleaning and preparation process must be carried out. Only thereafter can the actual riveting process be carried out. Then the rivets
15
are inserted into the mating holes, and upset or riveted to secure the stringers
12
onto the skin sheet
11
.
Thereafter, angle bracket elements, particularly so-called clips
13
, are riveted onto the stringers
12
and onto the skin sheet
11
. This riveting process also involves the complicated and time consuming steps that were described above. In a further assembly process, the actual crosswise or circumferential frames
14
are joined onto the fuselage shell
10
being fabricated. Particularly, the frame
14
is riveted onto the respective angle bracket elements or clips
13
. During this process, the above described complicated, time consuming and costly preparation measures and application of a suitable sealant are necessary at least on certain areas of the fuselage shell or skin.
SUMMARY OF THE INVENTION
In view of the above it is an object of the invention to provide a structural component and especially a structural panel or shell component that has a simplified structure and construction, so that it can be fabricated with a simpler, less costly and less time consuming fabrication process, whereby the fabrication time and fabrication costs of the component can be reduced. It is a further object of the invention to provide such a suitable simplified and less costly fabrication process. The invention more particularly aims to avoid complicated riveting and adhesive bonding process and the associated preparatory steps, and to achieve varying dimensions or thicknesses of the structural component at different areas in a simple and economical manner. The invention further aims to avoid or overcome the disadvantages of the prior art, and to achieve additional advantages, as ap
Brenneis Hartmut
Zink Walter
Airbus Deutschland GmbH
Fasse W. F.
Fasse W. G.
Friedman Carl D.
McDermott Kevin
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