Note: Descriptions are shown in the official language in which they were submitted.
CA 02317366 2000-09-OS
FIELD OF THE INVENTION
The invention relates to a large format or large surface area
structural component for an aircraft, which includes at least one
skin field or skin panel, and further relates to a method of
s manufacturing such a large surface area structural component.
BACKGROUND INFORMATION
In the present day manufacturing of large format or large surface
area structural components, and especially fuselage shell compo-
nents for an aircraft, the skin panels, also called skin fields,
~o used for manufacturing such components typically have dimensions
of about 2.5m x lOm. Particularly, the largest possible size is
used for each individual skin sheet that makes up the skin panel
for manufacturing the fuselage shell, in order to minimize the
number of lengthwise and crosswise seams between adjacent skin
sheets or panels, and thereby to minimize the overall weight of
the aircraft fuselage as well as the required amount of assembly
work. Minimizing the structural weight is an especially impor
tant consideration in the construction of aircraft, in view of
the energy consumption and therewith the economical operation of
zo the resulting aircraft.
A minimization of the total weight of the fuselage requires a
mechanical structural optimization of the fuselage shell at each
location to meet the locally effective requirements, e.g. differ-
ent strength requirements at different locations . Thus, the skin
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CA 02317366 2000-09-OS
panels are individually embodied to have different or varying
thicknesses, depending on the particular local loading that will
be applied to the skin panel at different locations. For exam-
ple, it may be necessary to provide thickenings on the skin
s panels in the area of the stringer connections in order to pro-
vide more support for the connected stringer.
Such areas of differing thickness of a particular skin panel are
typically produced by present day techniques using sheetmetal-to-
sheetmetal riveting, or sheet to sheet adhesive bonding or glu-
~o ing, or mechanical milling, or chemical etching or other material
removal techniques. The chemical etching or material removal of
a skin panel is carried out by masking the panel, cutting and
partially removing areas of the mask to form a prescribed pat-
tern, and then removing material by chemical etching or mechani-
cal material removal from the surface area or areas that have
been exposed by the patterning and partial removal of the mask.
Next, the skin panels are joined together to form so-called half
shells whereby riveting as well as adhesive bonding are used as
joining techniques. Before carrying out the riveting process,
2o the sheets and stringers are anodized to provide surface protec-
tion, primed, cleaned with an activator along the joint surfaces,
and then provided with a surface seal. Then the stringers are
fitted in position and fixed by means of tacking rivets. The
riveting is carried out automatically, whereby sealant is
z5 squeezed or exuded out of the joint locations . This squeezed-out
sealant forms a so-called sealant rope or bead that must be
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CA 02317366 2000-09-OS
smoothed out by hand and then protected by a protective coating
against aggressive media.
In the context of adhesive bonding or gluing, the skin sheets and
doubling members or reinforcing members are produced by contour
s milling of semi-finished plates. The stringers are sheet metal
profiles or extruded profiles, which are cut to proper length and
them subsequently deformed into the required shape. Before the
adhesive bonding process, the individual parts are subjected to
a specialized adhesive pretreatment. The parts are degreased,
1o cleaned, pickled, anodized in chromic acid, and finally primed.
In order to carry out the adhesive bonding, the adhesive joint
surfaces are provided with an adhesive film, the individual parts
are positioned and pressed together and then held together using
fixing screws. Then, the adhesive is cured and hardened under
~S the effect of increased temperature, pressure and time in an
autoclave. Thereafter, the component must be cleaned, and then
provided with a sealant rope or bead and once again with a pro-
tective layer to provide protection against corrosion.
The above described rather complicated manufacturing processes
zo for fabricating a fuselage shell structure are described in more
detail in the vDI progress reports ("Fortschrittsberichten")
Series 2: Fabrication Technology (Fertigungstechnik), No. 326,
Dissertation 0794 by Dipl.-Ing. Peter Heider "Lasergerechte
Konstruktion and lasergerechte Fertigungsmittel zum Schweissen
zs grossformatiger Aluminium-Strukturbauteile" ("Laser Compatible
Construction and Laser Compatible Production Means for Welding
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CA 02317366 2000-09-OS
Large Format Aluminum Structural Components"), pages 3 to 5.
This publication similarly describes the laser beam welding
process for the production of large format structural components,
whereby possible manners of construction of a structural compo-
s nent are exemplified at pages 42 to 47 of this report. In all
of these manners of construction however, the starting point is
always an available skin panel, which must be thickened and~or
have material removed at corresponding appropriate locations, as
has been discussed above, and which shall always have the largest
~o possible dimensions in order to minimize the necessary lengthwise
and crosswise seams or joints and thereby minimize the total
resulting weight of the fuselage.
U. S. Patent 3, 023, 860 (Ellzey) discloses a body construction and
a corresponding method in which sheet metal parts are joined
together to form a body construction or structural component.
However, this body construction involves a corrugated sheet with
stiffening elements and a flat sheet that are arranged adjacent
one another and joined together, in order to form the structural
component. A selection of different semi-finished parts is not
zo provided for, in order to achieve an adaptation to different
requirements and an optimization of the total resulting weight
of the structural component.
SUMMARY OF THE INVENTION
In view of the above, it is an object of the present invention
z5 to provide a large format or large surface area structural compo-
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CA 02317366 2000-09-OS
nent that may be produced by a simplified and less costly fabri-
cation process and that minimizes or avoids the necessity of
introducing additional reinforcements onto the skin sheet by
applying additional sheet metal layers by means of riveting or
s adhesive bonding, or by the mechanical or chemical material
removal of a partial thickness of the metal sheet at other areas .
It is a further object of the invention to provide such a struc-
tural component that has an optimized weight in consideration of
meeting the structural requirements at any location. The inven-
~o tion also aims to allow such a structural component to be fabri-
Gated with a customized or specialized structure that is particu-
larly adapted to the structural requirements in each particular
application. The invention further aims to avoid or overcome the
disadvantages of the prior art, and to achieve additional advan-
~s tages, as are apparent from the present specification.
The above objects have been achieved according to the invention
in a large surface area or large format structural component for
an aircraft that comprises at least one skin panel, wherein the
skin panel is formed of a plurality of individual structural
2o elements, which have each respectively been selected in terms of
the material and the thickness in view of the locally effective
requirements and loads that will be applied to a particular
portion or area of the structural component in its finished end
use. After being so selected and arranged, the structural ele-
2s ments are joined to each other by a welding process.
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CA 02317366 2000-09-OS
Particularly, the skin panel is formed of a plurality of struc-
tural elements, and especially separate skin sheets, that respec-
tively have different thicknesses and that are respectively made
of different materials. Thus, the resulting skin panel has
s different thicknesses and different material compositions at
different areas, thereby providing different strengths and other
structural characteristics at different locations so as to meet
the different requirements at these different locations. Also,
this structural component has been fabricated without requiring
~o strengthening members to be added onto the skin panel, and with-
out requiring material removal from the skin panel.
Also, it is particularly advantageous that the individual skin
sheets making up the overall structural component are no longer
required to have the largest possible dimensions as is the case
in the prior art. Instead, smaller individual skin sheets are
joined together, with stringers and frame or rib segments to form
a large structural component and particularly a fuselage shell
structure. The maximum resul ti nn ai ~A of tho ~"~o, ~,..~ ..~,.., , ...
not limited by, but rather is independent of, the size of the
zo individual skin sheets from which the shell is fabricated. Thus,
it is possible to manufacture fuselage shells having a size that
cannot be achieved with presently existing fabrication processes
and equipment. This is especially true since the smaller indi-
vidual skin sheets that are used according to the invention are
2s easier to handle and manipulate during the fabrication. This is
true for the fabrication equipment and processes, as well as for
the fabrication workers. The concept of using individually
CA 02317366 2000-09-OS
specialized skin sheets having smaller dimensions to build larger
resulting structural components is directly contrary to the prior
art teachings that call for using the largest possible dimensions
of a skin sheet to minimize the number of seams and thereby
s minimize the joining effort and the resulting weight. The pres-
ent invention overcomes these problems and achieves additional
advantages as described herein.
The above objects have further been achieved according to the
invention in a method of producing a large surface or large
~o format structural component, involving the following steps.
First, individual skin sheets are cut to length or otherwise cut
to the proper dimensions, from various different available semi-
finished materials that respectively are made of different mate-
rial compositions and have different material thicknesses. The
material and the thickness for a respective skin sheet to be used
at a particular location of the structural component is selected
dependent on the respective requirements that will apply to that
location of the structural component in its end use application.
Alternatively or additionally, extruded panels are produced by
zo an extrusion process to meet the prescribed requirements at any
particular location. Then, the metal skin sheets and/or the
extruded panels are joined together to form a skin panel. The
thus-formed skin panel is then further processed, and particu-
larly is deformed into a cylindrically or spherically curved
z5 structural component by means of a deforming process, such as
rolling or stretch forming, i.e. drawing.
_ g _
CA 02317366 2000-09-OS
In a further embodiment of the invention, the above objects have
been achieved in a method for producing profile sectional parts
involving the following steps . First, blank parts are cut to the
proper length and/or other dimensions and possibly heat-treated
s or deformed as needed, whereby each respective blank part is made
from a respective material having a respective thickness depend-
ent on and responsive to the prescribed profile geometry and
requirements of the profile sectional part to be manufactured.
The several blank parts are joined together by means of a welding
~o process to form respective profile sectional parts or structural
elements. Then the profile sectional parts or structural ele-
ments are after-machined, and particularly contour machined or
processed.
BRIEF DESCRIPTION OF THE DRAWINGS
15 In order that the invention may be clearly understood, it will
now be described in connection with example embodiments, with
reference to the accompanying drawings, wherein:
Fig. 1 is a schematic perspective diagram of the fabrication
sequence for manufacturing a skin panel from semi-
2o finished parts according to a first embodiment;
Fig. lA is a schematic sectional diagram illustrating the
cause of a bending moment that arises in joined skin
panels using a conventional overlapping joint;
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CA 02317366 2000-09-OS
Fig. 2 is a schematic perspective view of the fabrication
sequence for manufacturing profile sectional parts
such as stringers, clips and frames;
Fig. 3 is a schematic perspective view of a fuselage shell
s structure according to the invention;
Fig. 4 is a schematic perspective view of an extruded panel
as a semi-finished part for the further manufacturing
of a skin panel;
Fig. 5 is a schematic perspective view of several extruded
panels joined to each other along respective joints;
and
Fig. 6 is a schematic perspective view of respective sheet
metal parts joined onto an extruded panel along re-
spective joints.
~5 DETAILED DESCRIPTION OF PREFERRED EXAMPLE EMBODIMENTS AND OF THE
BEST MODE OF THE INVENTION
Fig. 1 schematically shows the fabrication sequence for manufac-
turfing a large surface area or large format structural component
1 according to the invention. The structural component 1 is
zo especially an aircraft fuselage shell component, which is formed
from a plurality of semi-finished materials or parts 2, 3, 4 and
that are respectively made from different materials and respec-
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CA 02317366 2000-09-OS
tively have different thicknesses, and that are joined to each
other to form the fuselage shell component 1. The different
semi-finished parts 2, 3, 4 and 5 are schematically indicated in
the drawing figures through the use of different lining or hatch-
s ing, which is intended to indicate the different thicknesses or
material compositions thereof. The individual semi-finished
parts 2, 3, 4 and 5, which are generally regarded as metal sheets
2, 3, 4 and 5, can be provided, for example, in the form of
substantially continuous long rolls of the semi-finished mate-
~o rial, from which the individual parts are cut out as needed. In
the selection of the material for each respective semi-finished
part, it must be take into account that a good weldability of the
material relative to the materials of the adjacent parts must be
achieved, which, for example, may be achieved using aluminum
~s alloys such as a AlMgSiCu, AlMgLi, or AlMgSc as the materials of
the respective semi-finished parts 2, 3, 4 and 5.
According to the invention, the individual semi-finished parts
2, 3, 4 and 5 are each respectively cut only to that particular
size, i.e. the particular dimensions,, as required for that par-
zo ticular material thickness and material composition at the given
location of the finished structural component 1. As an example,
when a skin area having a particular thickness and a particular
set of material characteristics is needed at an area having
certain dimensions in the finished structural component 1, a
zs correspondingly sized skin sheet 21 is cut to have the proper
length from a roll of sheet metal material 2 having the proper
width, thickness and material to meet the requirements at the
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CA 02317366 2000-09-OS
given location. A somewhat narrower skin sheet 22 is cut from
the same semi-finished part 2, but is cut to the required smaller
dimensions as needed for a different area of the finished struc-
tural component 1. In a similar manner, various sheet metal
s blanks are cut from the other semi-finished parts or material
rolls 3, 4 and 5, so as to provide the skin sheets 31, 32, 33,
34, 35, 36 and 37 from the semi-finished part 3, the skin sheets
41, 42, 43, 44 and 45 from the semi-finished part 4, and the skin
sheets 51, 52 and 53 from the semi-finished part 5.
~o If it is necessary for the requirements of the particular appli-
cation, the thus-cut sheet metal blanks may now be chemically or
mechanically pretreated or machined, or provided with a surface
protective coating or the like, as individually needed. Thus,
it becomes possible to carry out a chemical material removal or
etching from particular or limited ones of the pre-cut sheet
metal blanks, without requiring a masking process or any other
special measures for limiting the chemical treatment to a partic-
ular area. Such a chemical treatment may be necessary, for
example, if a special or particular surface contour is to be
zo achieved on a given skin sheet, or if a very localized reinforce
ment is necessary for load introduction paints or the like.
It is advantageous in this fabrication process, that not the
entire structural component, but rather only individual ones of
the precut blank skin sheets are treated or machined in the above
2s described manner, as needed. Thereby, the various apparatus
necessary for transporting the skin sheets as well as the chemi-
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CA 02317366 2000-09-OS
cal bath for carrying out a chemical treatment dv not need to be
sized for the maximum dimension of the fuselage shell, but rather
only must have the dimensions of the largest individual skin
sheet, which is significantly smaller than the overall maximum
s size of the fuselage shell. This especially makes it possible
to achieve a very economical chemical bath apparatus and treat-
ment, and also achieves an improved utilization factor for such
a smaller chemical bath apparatus.
As a further process step, after the skin sheets 21, 22, 31, 32,
~0 33, 34, 35, 36, 37, 41, 42, 43, 44, 45, 51, 52 and 53 have been
precut to size, and possibly subjected to a chemical or mechani-
cal pretreatment as mentioned above, the plurality of individual
sheet metal parts are then welded together to form a large skin
panel lA. The respective joining of the individual skin sheets
to each other is preferably achieved by means of laser beam
welding to form butt welds of the individual parts. This process
achieves a high welding process speed (for example in the range
of approximately 10 to 15 m/min using presently available equip-
ment and techniques), and is also rather low in the rate of
zo defects and deformation of the workpiece, while also being eco-
nomically practical. Another possible joining method for fabri-
cating the butt joints is, for example, a frictional contact
welding or so-called friction stir welding process, such as
described in the published international patent application WO
2s 93/10935 and which is known to persons of skill in this art.
Note that the drawing of panel lA still shows all individual skin
sheets with seams therebetween (e. g. before completed welding),
while the drawing of the finished component 1 omits seams between
same-material skin sheets and shows the finished result in the
3o major plane of the component 1.
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CA 02317366 2000-09-OS
The butt weld joints are preferably provided at those locations
where step-wise variations in the thickness are necessary due to
the layout of the fuselage shell. In other words, at locations
at which such step-wise variations in thickness are necessary
s anyway, it is sensible to provide a seam or joint at such a
location between two respective suitable skin sheets having the
two different thicknesses. Thus, the skin panel lA can be opti-
mally adapted to the requirements and particularly the loading
conditions of the finished fuselage shell or the finished fuse-
~o lage, by proper selection or adaptation of the thickness of the
individual skin sheets 21, 22, 31, 32, 33, 34, 35, 36, 37, 41,
42, 43, 44, 45, 51, 52 and 53. In this case, the additional need
for reinforcements or thickening members, as well as chemical or
mechanical removal of material from the skin panel lA, are mini-
mized or completely eliminated.
The weld seam geometry and particularly the path of the respec-
tive weld seams can be laid out on the skin panel lA or on the
structural component 1 in such a manner. so that the weld seams
or joints provide further functions, for example, the effect of
2o acting as a rip-stop seam, also known as a dummy weld seam. By
appropriately designing the weld seams in such a manner as rip-
stoppers, the tension level is thereby reduced and the continuous
progression of a rip or crack in the material is hindered or
prevented. Namely, when a rip or crack is initiated, and extends
z5 to one of the rip-stop weld seams, the rip or crack is intended
to be terminated at the seam so that it does not progress any
further.
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CA 02317366 2000-09-OS
Another advantageous effect achieved by the use of butt welds
according to the present invention, is that the tension level of
the skin panel joints can be positively influenced due to the
avoidance of a bending moment that typically arises when the
s lengthwise seams or joints are formed as overlapped joints. To
provide a further understanding in this context, Fig. lA shows
how an additional bending moment is generated in the overlap
areas of conventional overlapped joints. Due to the prevailing
internal pressure P within the fuselage, this pressure P acts on
~o the skin sheets 200 and 300 of the fuselage shell, which in turn
gives rise to a circumferential force F that acts on the skin
sheets 200 and 300. Due to the overlap of respective portions
of the two skin sheets 200 and 300, there exists a skin sheet
offset S in the radial direction between these two skin sheets
~s 200 and 300. The above mentioned circumferential force F thus
acts on the respective skin sheets 200 and 300 respectively
offset by the skin sheet offset distance S, which acts as a lever
arm and induces a corresponding bending moment, generally repre-
sented by F x S. This bending moment has a negative influence
zo on the tension level. By entirely avoiding overlapping joints,
according to the invention, the additional bending moment and its
negative effects will also be avoided.
After the skin panel lA has been fabricated, especially in a
substantially flat planar configuration, by respective butt weld
as joining of the individual skin sheets 21, 22, 31, 32, 33, 34, 35,
36, 37, 41, 42, 43, 44, 45, 51, 52 and 53, the resulting skin
panel lA is subjected to a further manufacturing step involving
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CA 02317366 2000-09-OS
the bending or deforming of the skin panel lA into a cylindri-
cally or spherically curved structural component 1. This bending
or deforming process is carried out with any known forming meth-
ods and forming equipment, for example by means of drawing and/or
s rolling, dependent on and adapted to the curvature of the struc-
tural component 1 that is to be achieved.
Fig. 2 schematically shows an example of a fabrication sequence
for manufacturing profile sectional parts according to the inven-
tion. Similarly, as the skin panel lA discussed above, the
~o present example profile sectional parts 10, 11, 12, 13, 14 and
15 are each individually formed by joining proper components from
different semi-finished materials or products 6, 7, 8 and 9.
These semi-finished products 6, 7, 8 and 9 preferably are made
of different materials and/or have different thicknesses, and may
15 be cut to different sizes as required, and then joined together
to form the resulting profile sectional parts having different
strength characteristics and the like in different areas or
portions thereof. For example, a profile sectional part 10 is
formed by precut blank parts 61 and 91 that have been cut from
zo the semi-finished products 6 and 9 respectively, and then welded
to each other. A further profile sectional part 11 is formed by
joining together precut blanks 62 and 81 that have been cut from
the semi-finished products 6 and 8 respectively.
Before the welding process, it is possible to carry out any
z5 necessary steps such as deformation steps or mechanical machining
steps on respective individual ones of the precut parts 61, 91,
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CA 02317366 2000-09-OS
62 and 81. Thus, it is possible to manufacture very individual-
ized profile sectional parts according to the demands or require-
ments at hand, without giving rise to high storage or warehousing
costs for keeping a corresponding assorted selection of the
s finished parts at hand. In other words, the individual profile
sectional parts can be fabricated to meet individual requirements
at the time they are needed, thus avoiding the need for keeping
different specialized profile sectional parts on hand. Prefera-
bly, the fabrication is carried out in a computer controlled or
~o computer aided manner, so that the respective required profile
geometry can simply be called up as needed, and then automati-
cally manufactured under the control of corresponding data sets
that define the profile geometries and required characteristics
for the respective required profile sectional parts such as
15 frames, clips, stringers, and fixtures as required for an indi-
vidual aircraft or an entire aircraft type or family. The fin-
fished profile sectional parts 12, 13, 14 and 15 demonstrate that
different profile shapes can be provided in the finished parts,
simply by appropriately pre-cutting the semi-finished product
zo materials from which the profile sectional parts are fabricated.
After the various profile sectional parts, which are particularly
embodied as frames, clips, stringers, etc. , have been fabricated,
these are installed in or joined onto the structural component
1 for manufacturing a fuselage shell of an aircraft, whereby the
2s profile sectional parts act as reinforcements or the like for the
structural component 1. A portion of a finished fuselage shell
100 is shown in Fig. 3. In this example, the fuselage shell 100
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CA 02317366 2000-09-OS
comprises a skin panel 1C, which has been manufactured or assem-
bled, for example, from a plurality of individual skin sheets
according to the fabrication process described above in connec-
tion with Fig. 1. Exemplary butt weld joints 101 and 102 are
shown in Fig. 3. Further lengthwise butt weld seams have been
hidden or covered by the arrangement of stringers on the skin
panel 1C.
The stringers 103, 104, 105 and 106 are arranged on the skin
panel lA or rather on the structural component 1 to extend there-
~o along in the lengthwise direction of the aircraft so as to
stiffen and strengthen the fuselage shell. In this context, the
stringers 103, 104, 105 and 106 have preferably been fabricated
according to the fabrication process described above in connec-
tion with Fig. 2, or in any conventionally known manner, and are
then joined onto the structural component 1 by a welding process.
Frames 107 and 108 are arranged on the fuselage shell 100 extend-
ing in a direction perpendicular to the lengthwise direction of
the aircraft. These frames 107 and 108, among other things,
serve to carry the load introduction of loads from the tail unit
zo or empennage and the like, which is connected to the frames by
appropriate fixtures. The frames 107 and 108 are secured to the
stringers 103, 104, 105 and 106 by respective frame-clip-rivet
connections that are generally known in the art of aircraft
manufacturing, or any other known manner. For example, it is
2s similarly possible to connect the frames and/or clips to the
corresponding joined components and/or to the overall fuselage
structure 100 by means of welding.
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Fig. 4 shows a further embodiment of a semi-finished part 16,
which can be further used to manufacture a structural component
1 according to the invention. Particularly, the semi-finished
part 16 is manufactured as an extruded panel 16, which includes
s integrally formed stiffening or strengthening elements 18, e.g.
ribs or stringers 18, that are integrally formed on a base ele-
ment or base sheet 17, to provide a single integral large format
profile sectional member in the form of the extruded panel 16.
Thus, the single integrally extruded panel 16 can replace the
~o separate or individual and differently embodied structural ele-
ments, namely the skin and the stringers, that are typically used
in aircraft construction. The base element 17 forming the skin
17 is embodied in a thin-walled manner to meet the requirements
of the loads arising in the aircraft fuselage structure, while
~s the lengthwise extending stiffening elements 1BA to 18H act as
stringers and integrally stiffen and strengthen the resulting
structural component 1. The extruded panel 16 preferably con-
sists of a defect or damage tolerant alloy, for example AlMgSiCu
or AlMgLi, and is configured with a typical form that pertains
zo to aircraft fuselage shells, with regard to the thickness of the
base element 17, i.e. the skin 17, as well as the dimensions and
form of the stringers 18A to 18H.
Through the use of such an integral extruded panel 16 for an
aircraft fuselage shell, the number of individual parts is re-
z5 duced, since the skin and stringers together are integrally
formed in a single fabrication process, namely an extrusion
process, to provide the above described integral semi-finished
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CA 02317366 2000-09-OS
part 16. Thus, the present invention can totally avoid the need
for joining operations between conventional separate skin panels
and stringers, and also can avoid the formation of corrosion
attack locations such as rivet holes and joint gaps, which must
s otherwise be treated and sealed in a rather complex and costly
manner.
A further substantial advantage of this inventive embodiment is
that thickened portions of the fuselage skin can be provided
extending in the lengthwise direction, i.e. parallel to the
~o stringers 18, in any thickness as required up to the maximum
thickness limit of the extrusion technology being utilized.
Namely, such thicker areas of the fuselage skin are directly and
integrally extruded with the extruded panel 16, by simply provid-
ing the appropriately configured and dimensioned extrusion die.
15 Moreover, stepped differing thicknesses of the skin 17 as well
as a load dependent cross-section adaptation of the stringers 18A
to 18H can be achieved within limited regions of an extruded
panel 16 by means of chemical or mechanical milling.
Fig. 5 shows the joining of a plurality of extruded panels 16,
zo 16' and 16", which have each individually been fabricated by
extrusion as discussed above in connection with Fig. 4. Prefera-
bly, laser beam welding or friction contact welding, i.e. so-
called friction stir welding, processes are used for joining
these extruded panels 16, 16' and 16" to form a skin panel 1 of
z5 a fuselage shell or at least a part of a skin panel 1 ' . Particu-
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CA 02317366 2000-09-OS
larly, the joints are realized in the manner of butt weld joints
or seams 19 and 19'.
Fig. 6 illustrates a further variant of a process for fabricating
a skin panel 1" or at least a part of the same. At least one
s extruded panel 16 can be used as a supplement to or in combina-
tion with one or more sheet metal parts, for example an illus-
trated skin sheet 23 which has been fabricated from the sheet
metal semi-finished part 2. These components are joined together
along a butt weld joint 19" to form the skin panel 1". In this
~o context, any of the above described processes can be utilized.
According to the known differential construction technique,
individual stringers 20A, 20B, 20C and 20D are mounted on the
skin sheet 23, preferably by welding. In this context, the
stringers may be joined onto the skin sheet even before the
components 16 and 23 are joined to each other.
By using at least one extruded panel 16 as a supplement to or in
combination with any one or more of the sheet metal semi-finished
parts described above in connection with Figs . 1 to 3, the possi-
bilities for variations of the semi-finished parts being uti-
zo lized, and thereby the possibility of further improving the
adaptation of the finished structural component to the require-
ments at hand, are further increased. This is achieved in accor-
dance with the present inventive method of welding together a
plurality of smaller sheets having various thicknesses, along
z5 butt weld joints, so as to form the resulting structural compo-
nent.
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Although the invention has been described with reference to
specific example embodiments, it will be appreciated that it is
intended to cover all modifications and equivalents within the
scope of the appended claims. It should also be understood that
s the present disclosure includes all possible combinations of any
individual features recited in any of the appended claims.
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