Note: Descriptions are shown in the official language in which they were submitted.
CA 02349122 2001-05-30
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 part:ially by welding.
BACKGROUND INF'ORMATION
In the constri:iction 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 Paterit Publication DE 196 39 667 and corre-
sponding U. S. Patent 5,841,098, or German Patent Publication DE
198 44 035. Particularly, the stringers and frame members are
welded onto large formal: 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. Part:icularly 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
- 2 -
CA 02349122 2001-05-30
able to withstand a longitudinally or circumferentially extending
crack that spans or ex-tends 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 t:he crack is broken. Thus, the remaining
structure of the fuselage shell must be able to withstand the
requisite loacis, 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 boriding result in a greater total weight,
and involve greater costs and efforts in the fabrication proce-
dures. 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 stiffenincl profile members have been laser welded onto the
skin sheets.
Particularly, with a riveted or adhesively bonded junction be-
tween 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
- 3 -
CA 02349122 2001-05-30
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 loca-
tion of a stringer or frame member, it does not directly damage
the associated stringer or frame member. Therefore, the respec-
tive 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 stiffenirig 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 stiff-
ening profile member will fail, typically in a sudden rupturing
manner, which leads to a failure of the fuselage shell structure.
However, the f:act 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 structur-es in which the stiffening profile members are
- 4 -
CA 02349122 2001-05-30
welded onto the skin sheets. Namely, with such a welded junc-
tion, 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 effec-
tive 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 occur-
rence of such a crack. Namely, once such crack forms, it will
readily propaclate through both the skin and the stiffening pro-
file 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 predom-
inant 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,
- 5 -
CA 02349122 2001-05-30
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 exam-
ple, in the bottom or belly of the fuselage, while using riveted
or adhesively bonded stringers. on the sides and top of the fuse-
lage.
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 thick-
ness 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 ski_n sheet through the welded junc-
tion 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 resid-
ual strength of the fuselage shell structure after a crack has
formed in the skin. While such an integral protrusion or thick-
ening of the stiffening profile member aims to provide a certain
- 6 -
CA 02349122 2001-05-30
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 aircraft fuse-
lage shell component that has an increased residual strength
after a crack has formed. -in the skin thereof, under consideration
of a minimum structural weight of the shell component. It is a
further object of the invention to provide a shell component with
stiffening profile members welded onto the skin thereof that is
suitable for use in a.ll locations of the aircraft fuselage,
including the sides and the top of the fuselage shell. The
invention further aims to avoid or overcome the disadvantages of
the prior art, and to achieve additional advantages, as apparent
from the present specification.
The above objects have been achieved according to the invention
in a structural shell component for an aircraft fuselage, com-
prising a skin sheet as well as plural stiffening profile members
such as stringers or frame members, whereby the stiffening pro-
file members are at least: partially joined to the skin sheet by
means of welding. Particularly according to the invention, non-
unitary or non-integral strengthening elements are arranged on
and secured to the stiffening profile members, before the stiff-
ening profile members are welded onto the skin sheet. Each of
these strengthening elemerits is a separate, non-integral compo-
CA 02349122 2001-05-30
nent relative to the stiffening profile member onto which it is
secured, and preferably consists of a different material than the
stiffening profile member. In this inanner, the strengthening and
crack stoppincl effect of the strengthening elements can be opti-
mized or maximized, without unacceptably increasing the weight
or the costs of the finished structure. In other words, the
strengthening members can consist of material that is stronger
and lighter, but more costly, than that of the stiffening profile
members, for example.
Further, preferably, the strengthening elements are secured to
the stiffening profile members by a non-integral connection
method, or joining method, such as riveting or adhesive bonding.
Such a non-integral connection provides the crack stopping inter-
ruption that is necessary for preventing a crack from propagating
into the strengthening el.ements or thereby also further into the
stiffening profile members. Namely, if a crack propagates
through the welded joint from the skin sheet into the stiffening
profile member, it will not further propagate through the riveted
or adhesive joint into the strengthening element or elements.
Thereby the strengtheninq element or elements will maintain its
strength intact and hold together the stiffening profile member
at the location of the crack, which will inhibit the further
propagation of the crack in the stiffening profile member.
According to preferred embodiments of the invention, the
strengthening elements may be in the form of doubling or rein-
forcing members that are secured to the webs of the respective
- 8 -
CA 02349122 2001-05-30
stiffening profile members, or alternatively the strengthening
elements may comprise tension bands or cables that are secured
to the stiffening profile members so as to extend along the
respective lerigth thereof.
The inventive shell structure achieves the advantage that the
residual or remaining strength of the shell structure after a
crack has fornied in the skin thereof, is sufficient so that the
welded shell component can also be used in the side and top areas
of an aircraft. fuselage. Thus, it becomes possible to use such
welded fuselaqe shell components for the entire fuselage of an
aircraft, so that the use of riveted and adhesively bonded joints
between the stringers or frame members and the fuselage skin can
be avoided, in all areas of the aircraft rather than only the
lower belly of the fuselage. In this manner, the overall produc-
tion effort, costs and structural weight can be significantly
reduced in comparison to the use of conventional adhesively
bonded or riveted shell components.
Moreover, the inventive provision of a two-part, non-integral
structure of separate strengthening elements secured onto the
stiffening profile members achieves additional advantages over
the prior art provision of an integral thickening or protrusion
adjacent to the base of each stiffening profile member. Namely,
the use of separate or discrete strengthening elements allows the
strengthening elements to be made of a different material than
that of the stiffening profile members, which allows a greater
strength and a greater strength-to-weight ratio to be achieved,
- 9 -
CA 02349122 2001-05-30
without excessively increasing the costs. Also, the strengthen-
ing elements may be arranged with an oriented characteristic, for
example, in the manner of a tension band or cable that extends
along the length of the respective stiffening profile member, so
as to exert its strongest retaining forces in a direction that
is most effective for holding together the respective stiffening
profile member across a crack, in the event a crack should propa-
gate into the stiffening profile member. The non-integral joint
between the respective strengthening element and the stiffening
profile member provides better crack stopping isolation to pre-
vent the further propagation of a crack through or beyond such
a joint.
The inventive structure has thus solved or overcome all of the
prior art disadvantages of welded shell components, including
those that arise when using stiffening profile members having
thickened portions along the roots or bases thereof. If a pri-
mary crack develops in the fuselage skin, this crack might propa-
gate through t:he welded joint into the stiffening profile mem-
bers, but there the crack propagation will be delayed or entirely
stopped by the strengthening elements arranged according to the
invention on the stiffening profile members. This in turn has
the effect of stopping or hindering the propagation of the crack
further in the fuselage skin. The structure of interconnected
frame members and stringers remains substantially intact and
maintains its strength, so that the residual or remaining
strength after the initiation of a crack in the fuselage shell
is increased.
- 10 -
CA 02349122 2001-05-30
BRIEF DESCRIPTION OF THE DRAWINGS
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 cross-section through a structural
component according to a first embodiment of the in-
vention, in the area of a stringer welded onto a skin
sheet;
Fig. 2 is a schematic cross-section similar to that of Fig.
1, but showing a second embodiment;
Fig. 3 is a schematic cross-section of a structural component
according to a first embodiment of the invention, in
the area of a frame member welded onto a skin sheet;
Fig. 4 is a schematic cross-section similar to that of Fig.
3, but showing a structural component in the area of
a frame member according to a second embodiment;
Fig. 5 is a schematic cross-section generally similar to that
of Fig. 1, but showing the area of a stringer accord-
ing to a third embodiment;
Fig. 6 is an enlarged detail of the detail area VI indicated
in Fig. 5;
- 11 -
CA 02349122 2001-05-30
Fig. 7 is a schematic cross-section generally similar to that
of F'ig. 3, but showing the inventive structural compo-
nent: in the area of a f_rame member according to a
third embodiment;
Fig. 8 is a schematic cross-section generally similar to that
of F'ig. 4, but showing the inventive structure in the
area of a frame member according to the third embodi-
ment:;
Fig. 9 is a schematic cross-section of the starting condition
of a. stiffeninq profile member to be used according to
the invention in a fourth embodiment; and
Fig. 10 is a schematic cross-section showing the finished
structural component according to the invention in the
area. of the sl~iffening profile member according to
Fig. 9.
DETAILED DESCRIPTION OF PREFERRED EXAMPLE EMBODIMENTS AND OF THE
BEST MODE OF T'HE INVENTION
Figs. 1 and 2 each respectively show a portion of a structural
shell component and particularly a fuselage shell component 1 in
the area at which a stiffening profile member 2 is welded onto
a fuselage skin sheet 4. Particularly in Figs. 1 and 2, the
stiffening profile member 2 is a stringer 3 that runs in the
aircraft longitudinal direction of the aircraft fuselage struc-
- 12 -
CA 02349122 2001-05-30
ture. The overall structural shell component 1 to be used as a
fuselage shell of an aircraft includes a plurality of such
stringers 3 respectively extending in the aircraft longitudinal
direction and spaced apart from one another in the aircraft
circumferential direction. Generally, the stiffening profile
members 2 could be stringers, frame members, ribs, spars, etc.
The connectiori between each stringer 3 and the skin sheet 4 is
achieved by means of welding, such as laser beam welding, or
according to any other conventionally known technique, for exam-
ple as disclosed in the German Patent Publications 196 39 667 or
198 44 035 as mentioned above. Such a welded joint 4A between
the stringers 3 and the skin sheet 4 allows a simpler and more
economical fabrication, and additionally achieves a weight reduc-
tion in comparison to the prior typical methods using rivets or
adhesive for joining stringers onto the skin sheets.
In order to increase the residual or remaining strength of such
a welded structural shell. component 1 in the event a crack forms
in the skin sheet 4, each stringer 3 is strengthened by at least
one strengthening element 5 according to the invention. Particu-
larly, the stringers 3 are provided with the strengthening ele-
ments 5 before the stringers are welded onto the skin sheet or
sheets 4. Each stringer 3 in this example includes a stringer
main web 3A, a flange web 3B, and a rim web 3C. The strengthen-
ing elements 5 in the present embodiment are preferably length-
wise extending doubling or reinforcing members 6. At least one
of these doubling members 6 is secured onto at least one side of
- 13 -
CA 02349122 2001-05-30
the stringer niain web 3A in order to "double" or reinforce this
web. While Fi.g. 2 shows an embodiment with only one lengthwise
extending doubling member 6 secured onto one side of the stringer
main web 3A, Eig. 1 shows an embodiment in which two doubling
s members 6 are respectively secured onto the opposite sides of the
stringer main web 3A. Depending on the remaining strength of the
structural shell component that is to be achieved, or depending
on the arisincl loads, one or more lengthwise doubling members 6
can be used as necessary, in either the arrangement shown in Fig.
2 or the arrangement shown in Fig. 1.
The doubling members 6 preferably consist of a high strength
aluminum alloy or of fiber reinforced metal laminates including
plural layers of metal and of reinforcement fibers. In the
illustrated erribodiment of Figs. 1 and 2, the doubling members 6
are configured. as a simple rectangular sectional member such as
an I - profile member. The respective doubling members 6 are
secured onto the sides of the stringer main webs 3A by means of
adhesive bonding 15 or riveting 16, which is carried out before
the reinforced. stringers 3 are then welded onto the skin sheet
4. In this context, these stringers 3 may consist of any mate-
rial that is conventionally used for such stringers in aircraft
construction, and the adhesive bonding or riveting of the dou-
bling members 6 onto the stringers 3 can be carried out by any
conventional riveting or adhesive bonding techniques using any
conventionally known materials for such joining methods in the
field of aircraft construction.
- 14 -
CA 02349122 2001-05-30
The present invention is very simple to incorporate into existing
methods of aircraft construction, because the doubling members
6 are simply pre-secured onto the stringers 3 and then the rein-
forced stringers 3 may be handled in the usual manner for being
welded onto the skin sheet 4. This is simply achievable, because
in comparison, it would be very difficult and costly to secure
strengthening members onto the structural shell component 1 after
the stringers and frames had been welded onto the skin, espe-
cially in the context of a large format component.
Figs. 3 and 4 respectively show portions of the structural shell
component 1 in. the area of a frame member 7 that extends in the
circumferential directiori of the aircraft fuselage. Thus, in the
present context, the stiffening profile member 2 is embodied as
a circumferential frame :member 7 which is welded onto the skin
sheet 4. According to the invention, strengthening elements 5
embodied as circumferentially extending doubling members 8 are
arranged on one side or on both sides of the frame member main
web 7A, while the frame member 7 further includes a flange web
7B and a rim web 7C. Particularly, Fig. 3 shows an arrangement
in which respective circumferentially extending doubling members
8 are glued or riveted onto both opposite sides of the frame
member main web 7A. On the other hand, Fig. 4 shows an arrange-
ment in which the frame member 7 includes an additional web 7D
that protrudes perpendicularly from the main web 7A, and the two
circumferentially extending doubling members 8 are secured onto
the two opposite sides of: this additional web 7D.
- 15 -
CA 02349122 2001-05-30
It is apparent: that the strengthening elements 5 in the form of
doubling members 8 can be arranged at any portion or location of
the frame member 7, depending on the profile shape of the frame
member 7 or generally the stiffening profile member 2, and de-
pending on the particular loading characteristic that will be
applied to the stiffening profile member. Similarly as mentioned
above in connection with the stringers 3, the gluing or riveting
of the circumferential doubling members 8 onto the frame member
7 is carried out before welding the frame member 7 onto the skin
sheet 4.
Figs. 5 to 8 show a structural shell component 1 with a further
embodiment of the reinforcement or strengthening of the stiffen-
ing profile members 2 in order to achieve an increase in the
residual strength of the welded shell structure. Particularly,
Figs. 5 and 6 show an embodiment of a stringer 30 that is
strengthened by a tension band or tension cable 31, which prefer-
ably consists of high strength steel or titanium alloys or fiber
reinforced composite materials. The tension band 31 is arranged
to extend alorig the length of the stringer web 32, i.e. in the
aircraft longitudinal direction. More particularly, the stringer
web 32 has a material thickening or thickened portion 33 protrud-
ing from one side thereof, with a through hole 34 extending in
the stringer lengthwise direction in this thickened portion 33,
for receiving the tension band 31 therein. Preferably, the
thickened portion 33 is provided in the lower half of the
stringer web 32, i.e. closer to the root or base along which the
stringer 30 is welded to the skin sheet 4. Alternatively, the
- 16 -
CA 02349122 2001-05-30
tension band could be ar:ranged in respective thickened portions
on both opposite sides of the respective stiffening profile
member.
The tension band 31 preferably has a polygon cross-sectional
shape, and is inserted irito the through hole 34 and then twisted
so that the tension band 31 becomes engaged with the wall of the
through hole _34. For this reason, square or rectangular cross-
sectional shapes of the tension band 31 are preferably used, so
that the angular edges of the tension band 31 can become engaged
in the through hole 34 as mentioned above. This ensures that a
relative sliding or shifting of the tension band 31 within the
through hole 34 and relative to the thickened portion 33 is
prevented, in the event of a crack forming and propagating into
the base area of the stringer 30. Thus, the tension band 31,
which remains intact and unaffected by the crack, holds together
the material thickened portion 33 on the opposite sides of the
crack and thereby hinders further propagation of the crack.
To enhance th_Ls effect of the securing or bonding between the
tension band 31 and the stringer 30, the complete stringer 30 is
heat treated, for example by a solution annealing process, in
order to improve the deformability of the material, and then
lateral clamping or pressing forces are applied to the thickened
portion 33 as indicated in Fig. 6, in order to deform the thick-
ened portion 3:3 and achie've a form-locked connection between the
tension band 31 and the thickened portion 33. In effect, the
- 17 -
CA 02349122 2001-05-30
thickened portion 33 is positively crimped onto the tension band 31.
Figs. 7 and 8 generally relate to an embodiment using a tension
band similarly to Figs. 5 and 6, but in particular show a tension
band 71 arrancied in a frame member 70. Similarly as described
above in connection with Figs. 5 and 6, a tension band 71 con-
sists of high strength steel or titanium alloys or fiber rein-
forced materials, and is arranged in a through hole 74 extending
in the frame member lerigthwise direction (i.e. the aircraft
circumferential direction) in a material thickened portion 73
protruding from one side of the frame member web 72, and prefera-
bly on a lower half thereof closer to its base or root that is
welded to the skin sheet 4. The tension band 71 preferably has
a polygon cross-section, for example a quadrangular or square
cross-section. When the tension band 71 is inserted into the
through-hole '74 and then twisted, the tension band 71 becomes
engaged in the through hole 74, so as to prevent a relative
shifting or displacement of the material thickened portion 73
relative to the tension band 71, in the event that a crack propa-
gates into the base of the frame web 72 and the thickened portion
73. The entire frame member 70 is heat treated, for example
preferably by a solution annealing process, and then the thick-
ened portion 73 is pressed or crimped as described above in
connection with Fig. 6 i.n order to achieve a form-locking connec-
tion between the thickened portion 73 and the tension band 71.
Figs. 9 and 10 show a further advantageous embodiment of a rein-
forced stiffening profile member 2 for a structural shell compo-
- 18 -
CA 02349122 2001-05-30
nent 1 having an increased residual strength according to the
invention. Iri this embodiment, the stiffening profile member 2
is a special stringer 35 that includes a split or slitted web 36
including two web legs 36A and 36B at the base or root end of the
stinger that will be welded onto the skin sheet 4. A receiver
opening 37 for receiving a tension band 38 therein is provided
at the end of the split or opening of the split web portion 36
where the two web legs join each other. Fig. 9 shows the
stringer 35 in its starting configuration before the insertion
of the tension band 38 -.herein. The split web portion 36 is
configured so as to form a clamp including the receiver opening
37 into which the tension band 38 is inserted. To insert the
tension band 38, it must be pushed past a protruding edge of
material 39 along the mouth of the receiver opening 37, so that
the tension band 38 is forcefully clipped into place and held in
position in the receiver opening 37 by the protruding edge or lip
of material 39.
The tension band 38 preferably has a substantially round cross-
section and is provided with a roughened outer surface. While
other cross-sectional shapes of the tension band 38 are also
useable, a round cross-section is preferred because it can be
easily and economically fabricated, and is also readily available
in the form of' tension cables or the like. The provision of a
roughened surface is one possibility in order to hinder or pre-
vent the relative movement between the tension band 38 and the
stringer 35, to delay or prevent the propagation of a crack
further through the stringer 35 in the manner described above.
- 19 -
CA 02349122 2001-05-30
The tension band 38 may be made of the above mentioned high
strength steel or titanium alloys or fiber-reinforced composite
materials.
After the tension band 38 has been pressed into the receiver
opening 37, then the stringer 35 is subjected to a heat treat-
ment, such as a solution annealing treatment, in order to improve
the deformability of the:material. After the annealing step, the
split-open web portion 36 is pressed closed, thereby tightly
clamping and pressing the receiver opening 37 around the tension
band 38, so as to achieve a form-locking or crimped connection
of the tensior.L band 38 within the stringer 35. Thereafter, the
finished stringer 35 can. be welded onto the skin sheet 4 in a
substantially conventional manner. The resulting finished ar-
rangement of the stringer 35 with the tension band 38 clampingly
held therein, and being welded onto the skin sheet 4 is shown in
Fig. 10.
Although the invention has been described with reference to
specific example embodiments, it will be appreciated that it is
intended to caver all modifications and equivalents within the
scope of the appended claims. It should also be understood that
the present disclosure includes all possible combinations of any
individual features recited in any of the appended claims. While
it is not expressly stated in the above description, it should
be understood that the inventive reinforcement can be provided
for all of the stiffening profile members of a fuselage shell,
or only for each alternating second stiffening profile member,
- 20 -
CA 02349122 2001-05-30
for example, depending on the overall strength requirements of
the finished fuselage shell. In some applications, an adequate
overall residual strength might be achieved by providing the
inventive reir-forcement for only some of the stringers and frame
members, especially at critical locations.
- 21 -
