U.S. patent application number 15/833317 was filed with the patent office on 2018-04-05 for system and method for fabricating a composite material assembly.
The applicant listed for this patent is LEARJET INC.. Invention is credited to Germain BELANGER, Alain LANDRY.
Application Number | 20180093752 15/833317 |
Document ID | / |
Family ID | 43038038 |
Filed Date | 2018-04-05 |
United States Patent
Application |
20180093752 |
Kind Code |
A1 |
LANDRY; Alain ; et
al. |
April 5, 2018 |
SYSTEM AND METHOD FOR FABRICATING A COMPOSITE MATERIAL ASSEMBLY
Abstract
A method for fabricating a composite material assembly includes:
a) providing an assembly system, b) laying down a first module on a
first mold, the first module comprising a first laminate covering a
first laminate support structure, c) laying down a second module on
a second mold, the second module comprising a second laminate
covering a second laminate support structure and extending over the
at least one removable insert, d) removing the at least one
removable insert from the second mold, and e) assembling the first
mold with the second mold while overlapping a section of the second
laminate extending over the at least one removable insert over the
first laminate.
Inventors: |
LANDRY; Alain;
(Beaconsfield, CA) ; BELANGER; Germain; (St.
Germain de Grantham, CA) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
LEARJET INC. |
Wichita |
KS |
US |
|
|
Family ID: |
43038038 |
Appl. No.: |
15/833317 |
Filed: |
December 6, 2017 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
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13577094 |
Sep 21, 2012 |
9873501 |
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PCT/IB2010/001724 |
Jul 13, 2010 |
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15833317 |
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61301754 |
Feb 5, 2010 |
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Current U.S.
Class: |
1/1 |
Current CPC
Class: |
B29C 65/5078 20130101;
B29C 33/26 20130101; B29C 70/30 20130101; B29C 65/505 20130101;
Y02T 50/43 20130101; B29C 31/08 20130101; B29C 70/088 20130101;
B29L 2031/3079 20130101; Y10T 428/13 20150115; B29C 66/73752
20130101; B29L 2031/3082 20130101; B29C 66/12822 20130101; B29C
66/721 20130101; B29C 66/12842 20130101; B64C 2001/0072 20130101;
B29C 66/543 20130101; B29C 70/54 20130101; B29C 31/085 20130101;
B29C 65/483 20130101; B64F 5/10 20170101; B29C 66/1162 20130101;
B64C 1/068 20130101; Y02T 50/40 20130101; B29C 66/72525 20130101;
B29C 65/5071 20130101; B29C 35/02 20130101; B29C 70/462
20130101 |
International
Class: |
B64C 1/06 20060101
B64C001/06; B29C 65/00 20060101 B29C065/00; B29C 31/08 20060101
B29C031/08; B29C 33/26 20060101 B29C033/26; B29C 65/50 20060101
B29C065/50; B29C 70/54 20060101 B29C070/54; B29C 70/46 20060101
B29C070/46; B29C 70/30 20060101 B29C070/30; B29C 70/08 20060101
B29C070/08 |
Claims
1.-11. (canceled)
12. A method for fabricating a composite material assembly
comprising the steps of: a) providing an assembly system
comprising: a first mold for receiving a first module made of
composite material, said first mold comprising: a first composite
material laminate support structure having first and second
opposite edges; and a first attachment interface for attachment of
the first mold to an adjacent mold; and a second mold for receiving
a second module made of composite material, said second mold
comprising: a second composite material laminate support structure
having first and second opposite edges; a second attachment
interface for attachment of the second mold to the first mold; and
at least one removable insert extending beyond at least one of said
first and second edges of the second mold; b) laying down the first
module on the first mold, the first module comprising a first
laminate covering the first laminate support structure; c) laying
down the second module on the second mold, the second module
comprising a second laminate covering the second laminate support
structure and extending over the at least one removable insert; d)
removing the at least one removable insert from the second mold; e)
assembling the first mold with the second mold while overlapping a
section of the second laminate extending over the at least one
removable insert over the first laminate.
13. The method according to claim 12, wherein the first and second
molds are portions of a cylindrical structure.
14. The method according to claim 12, wherein the first laminate
has a first interface profile shaped to fit into a complementary
second interface profile of the section of the second laminate
extending over the at least one removable insert and overlapping
over the first laminate, for forming a joint between the first and
second modules.
15. The method according to claim 12, wherein the at least one
removable insert comprises a laminate overhang support surface,
said laminate overhang support surface being oriented at an offset
angle of at least 10.degree. with respect to a tangential direction
of the second laminate of the second mold, at the at least one of
said first and second edges of the second mold where the at least
one removable insert is positioned, towards an inner side of the
second mold.
16. The method according to claim 15, wherein the offset angle is
between 10.degree. and 15.degree..
17. The method according to claim 12, wherein the assembly system
further comprises a flexible elastomeric seal at a joint interface
between the first and second molds.
18. The method according to claim 12, further comprising the step
of, prior to step b), applying a release agent to the first and
second molds prior to layup of the first and second modules
thereon.
19. The method according to claim 12, further comprising the step
of f) curing the assembled first and second modules.
20. An aircraft fuselage comprising a composite material assembly
fabricated according to claim 12.
21. The aircraft fuselage according to claim 20, wherein the
fuselage is a solid laminate.
22. The aircraft fuselage according to claim 20, wherein the
fuselage is a sandwich structure.
23. The aircraft fuselage according to claim 20, wherein the
fuselage is a combination of a solid laminate in some locations and
a sandwich structure in other locations.
24. The aircraft fuselage according to claim 20, comprising at
least one component selected from the group comprising floor
attachments, cockpit windshields, cabin windows and passenger door
surrounding structures.
Description
FIELD OF THE INVENTION
[0001] The present invention generally relates to composite
materials. The present invention more specifically relates to a
system and method for fabricating a composite material
assembly.
BACKGROUND OF THE INVENTION
[0002] Composite material assembly, and more particularly fuselage
manufacturing through the use of multi-piece sections, typically
requires pre-solidification and cure of each piece prior to
assemble them with splices between individual sections or
portions.
[0003] The limitations of this methodology are: [0004] a minimum
two-step cure is required; [0005] additional mechanical fasteners
are required at splicing joints on primary structure components;
[0006] the methodology requires handling equipment and assembly
jigs (for out-of-mold operations); [0007] long fuselage
manufacturing time; [0008] over thickness at joints resulting in
stress concentration; [0009] increases in weight of assembly; and
[0010] surface preparation is required prior to bonding.
[0011] Various solutions for assembly of multi-piece sections have
been proposed in the prior art.
[0012] U.S. Pat. No. 7,459,048 discloses a method of manufacturing
a unitary section of an aircraft fuselage including steps of
disposing a thin layup mandrel element onto the outer shell surface
of a cylindrical inner mandrel shell to form a mandrel with a layup
surface. The method further includes steps of laying-up fibers onto
the layup surface while the mandrel rotates to form a unitary
pre-cured section of an aircraft fuselage.
[0013] WO 98/32589 discloses composite structures having a
continuous skin formed using automated fiber placement methods. The
multiple layers of fibers are placed on a fiber placement tool
including a mandrel body surrounded by a bladder. Uncured composite
structures are created by placing fibers around the fiber placement
tool as discontinuous segments that are capable of moving or
sliding in relation to each other in order to be expandable from
within. The uncured structures are then expanded against the other
surface of the molds by creating a vacuum between the bladder and
the molds.
[0014] U.S. Pat. No. 7,325,771 discloses structures and methods for
joining composite fuselage sections using spliced joints attaching
a first stiffener on a first composite part as well as a second
stiffener on a second composite part through a fitting. A strap is
then used to splice the first and second composite parts
together.
[0015] US 2006/0251847 discloses a method of joining composite
elements in which the bonding is done through the thickness of
fiber composite laminates in order to reduce interlaminar stresses
using non-interlocking and interlocking bonds.
[0016] US 2009/0148647 discloses a method of fabricating composite
structures by joining a plurality of composite modules along their
edges using scarf joints instead of using advance fiber placement
machines that require high capital investment and operating
costs.
[0017] However, there is still a need for a system and method for
fabricating composite material assemblies that facilitate assembly
of parts when forming structures while minimizing assembly
equipment costs.
SUMMARY OF THE INVENTION
[0018] An object of the present invention is to propose a system
and method that satisfies at least one of the above-mentioned
needs.
[0019] According to the present invention, that object is
accomplished with a system for fabricating a composite material
assembly comprising: [0020] a first mold for receiving a first
module made of composite material, the first mold comprising:
[0021] a first composite material laminate support structure having
first and second opposite edges; and [0022] a first attachment
interface for attachment of the first mold to an adjacent mold; and
[0023] a second mold for receiving a second module made of
composite material, the second mold comprising: [0024] a second
composite material laminate support structure having first and
second opposite edges; [0025] a second attachment interface for
attachment of the second mold to the first mold; and [0026] at
least one removable insert extending beyond at least one of the
first and second edges of the second mold, wherein the first module
comprises a first laminate covering the first laminate support
structure, the second module comprises a second laminate covering
the second laminate support structure and extending over the at
least one removable insert, and wherein the at least one removable
insert is removed from the second mold prior to assembly of the
first mold to the second mold, and a section of the second laminate
extending over the at least one removable insert overlaps over the
first laminate after closing and assembly of the first mold onto
the second mold.
[0027] According to the present invention, there is also provided a
method for fabricating a composite material assembly comprising the
steps of: [0028] a) providing an assembly system comprising: [0029]
a first mold for receiving a first module made of composite
material, the first mold comprising: [0030] a first composite
material laminate support structure having first and second
opposite edges; and [0031] a first attachment interface for
attachment of the first mold to an adjacent mold; and [0032] a
second mold for receiving a second module made of composite
material, the second mold comprising: [0033] a second composite
material laminate support structure having first and second
opposite edges; [0034] a second attachment interface for attachment
of the second mold to the first mold; and [0035] at least one
removable insert extending beyond at least one of the first and
second edges of the second mold; [0036] b) laying down the first
module on the first mold, the first module comprising a first
laminate covering the first laminate support structure; [0037] c)
laying down the second module on the second mold, the second module
comprising a second laminate covering the second laminate support
structure and extending over the at least one removable insert;
[0038] d) removing the at least one removable insert from the
second mold; and [0039] e) assembling the first mold to the second
mold while overlapping a section of the second laminate extending
over the at least one removable insert over the first laminate.
[0040] The present invention provides means for manufacturing
one-piece composite components originating from more than one mold
while providing a structure that can be cured or solidified under
heat and vacuum in one step only, preferably with a composite
material in a pre-prep form which does not require autoclave
treatment.
[0041] A non-restrictive description of a preferred embodiment of
the invention will now be given with reference to the appended
drawings.
BRIEF DESCRIPTION OF THE DRAWINGS
[0042] FIGS. 1a to 1c are side cross-sectional views of the system
according to a preferred embodiment of the present invention,
showing an assembly sequence of a first monolithic laminate on a
first mold onto a second monolithic laminate of a second mold with
prior removal of a removable insert;
[0043] FIGS. 2a to 2c are side cross-sectional views of the system
according to another preferred embodiment of the present invention,
showing an assembly sequence of a first sandwich laminate on a
first mold onto a second sandwich laminate of a second mold with
prior removal of a removable insert, and a subsequent addition of a
layup splice;
[0044] FIGS. 3a to 3e are front views of a build sequence of a
tubular component using the system according to another preferred
embodiment of the present invention, using one removable insert per
mold;
[0045] FIGS. 4a to 4e are perspective views of the build sequence
of the tubular component shown in FIGS. 3a to 3e;
[0046] FIGS. 5a to 5c are front views of initial steps of a build
sequence of a tubular component using the system according to
another preferred embodiment of the present invention, with an
alternate distribution of removable inserts with respect to the
molds, with no insert on a first mold, one (1) insert on a second
mold and two (2) inserts on a third mold;
[0047] FIGS. 6a to 6c are perspective views of the build sequence
of the tubular component shown in FIGS. 5a to 5c;
[0048] FIG. 7 is a perspective view of a build sequence of a
fuselage component using the system according to another preferred
embodiment of the present invention and showing installation of
composite layup materials by personnel; and
[0049] FIGS. 8a and 8b are schematic views of a stepped-lap joint
interface and a scarf-joint interface respectively
PREFERRED EMBODIMENTS OF THE PRESENT INVENTION
[0050] An object of the present invention is to manufacture a
composite material assembly, such as, but not limited to, a tubular
profile structure from two or more longitudinal section components.
The whole assembly can be cured in one step in order to form a
one-piece tubular structure, such as, for example, a fuselage.
Hence, the components that will constitute the whole assembly are
joined before curing occurs and then the whole assembly is cured
through co-curing of these components, producing an end product
without any overly apparent seams.
[0051] Referring to FIGS. a to 1c, according to a first preferred
embodiment of the present invention, a system 10 for fabricating a
composite material assembly is disclosed. The system 10 includes a
first mold 12 for receiving a first module 13 made of composite
material. The first mold 12 has a first composite material laminate
support structure 14 having first and second opposite edges 16, 18.
The first mold 12 also has a first attachment interface 20 for
attachment of the first mold 12 to an adjacent mold 22. The system
10 also comprises a second mold 22 for receiving a second module 23
made of composite material. The second mold 22 includes a second
composite material laminate support structure 24 having first and
second opposite edges 26, 28. The second mold 22 also has a second
attachment interface 30 for attachment of the second mold 22 to the
first mold 12.
[0052] The system 10 further comprises a removable insert 32
extending beyond the second edge 28 of the second mold 22. The
insert 32 is shaped such that it would contact the first mold 12 if
the first and second molds 12, 22 were attached together and would
prevent attachment therebetween if the insert 32 was present.
[0053] The first module 13 comprises a first laminate 34 covering
the first laminate support structure 14. The second module 23
comprises a second laminate 36 covering the second laminate support
structure 24 and extending over the removable insert 32. As better
shown in the transition between FIG. 1a and FIG. 1b, the removable
insert 32 is removed from the second mold 22 prior to assembly of
the first mold 12 to the second mold 22. As better shown in FIG.
1b, a section 38 of the second laminate 36 extending over the
removable insert 32 overlaps over the first laminate 34 after
closing and assembly of the first mold 12 onto the second mold 22.
At the initial closing of the molds 12, 22, the laminates 34, 36
are not cured or solidified, allowing the required tackiness,
softness and flexibility to ensure proper intermesh and layup at
the interface.
[0054] Preferably, the first laminate 34 has a first interface
profile 40 shaped to fit into a complementary second interface
profile 42 of the section of the second laminate 36 extending over
the removable insert 32. The second laminate 36 can therefore
overlap over the first laminate 34, and form a joint without an
overly apparent seam, between the first and second modules 13, 23.
Preferably, the interface profiles 40, 42 are chosen to form a
stepped-lap joint. During initial placement of the laminates 34, 36
on the molds 12, 22, the removable insert 32 provides the extension
surface that is required to form the stepped-lap joint.
[0055] Preferably, when the composite material assembly is a
tubular component, the final assembly results from two or more
joints. Given that the chosen type of joint for this application
must minimize any over thickness in order to obtain a uniform
structure thickness along the perimeter of circumference of the
assembly, it is preferable to use a type of joint that requires
superimposing two half-elements, preferably through a stepped-lap
interface as mentioned above. In other embodiments of the present
invention for fuselage applications, the stepped-lap interface, as
shown in FIG. 8A, at the sectional interface between modules could
be changed for a scarf-type lamination, as shown in FIG. 8B without
affecting the required constant thickness of the fuselage
structure.
[0056] As shown in FIG. 1b, manufacturing of a joint in accordance
with the present invention requires that one section 38 of the
laminate 36 overhangs temporarily and is therefore not supported
beyond the edge 38 of the mold 32. This overhanging configuration
is required for the period of time between removal of the insert 32
and closing of the molds 12, 22 for forming the assembly.
[0057] In order to allow closing of the molds 12, 22, each of the
half-elements of the complementary interface profiles 40, 42 to be
stacked must avoid contact with each other during the closing
movement of the molds 12, 22, as there can be a risk of localized
pre-adherence, before the two half-elements are positioned
correctly. Any incorrect positioning of the two sides of the
interface for the laminate could result in the formation of air
pockets and result in an abnormal discontinuity in the structural
laminate in the joint assembly zone.
[0058] In order to avoid this possibility of pre-adherence between
the two half-elements of the joint prior to the final closed
position of the molds 12, 22, the removable insert 32 preferably
has a geometrical form shaped to position the overhanging section
38 of the laminate 36, with the interface profile 42, above its
corresponding interface profile 40 on the other mold 12 without
incurring any contact or pre-adherence, after the insert 32 is
removed.
[0059] Preferably, the surface of the insert 32 on which the
overhanging section 38 of the laminated interface is resting has an
angular position of at least 10.degree. and preferably between
10.degree. and 15.degree. with respect to a tangential direction of
the second laminate 36 of the second mold 22, at the edge 28 of the
second mold 22 where the removable insert 32 is positioned, towards
an inner side of the second mold 22.
[0060] For fuselage applications, the required laminate
construction for the fuselage can be a monolithic configuration, as
shown in FIGS. 1a to 1c, or sandwich/core structure, as shown in
FIGS. 2a to 2c, or a combination of the two. Preferably, in the
case of a fuselage sandwich/core structure, as shown in FIGS. 2a to
2c, the initial laminated assembly produced with the removable
insert 32 and the molds 12, 22 is the same but this laminated
assembly may now be designated as an "outer skin" 74a. As shown in
FIG. 2b, the "outer skin" 74a receives a sandwich honeycomb core 72
followed by an "inner skin" 74b which could be of different
construction. The inner skin laminate 74b is terminated also at its
longitudinal edges 78, 79 by a stepped-lap geometry being in full
contact with the sandwich core 72 surface. Preferably, the
laminates are made of "out of autoclave" carbon-epoxy pre-preg and
the sandwich/core structure comprises a Nomex.TM. honeycomb core,
however, other materials may be used.
[0061] Preferably, the removable insert 32 is a structural element.
However, the removable insert may be an inflatable structure, or
any other retractable molding structure known to a person skilled
in the art.
[0062] Preferably, the attachment interfaces 20, 30 are hinge-type
interfaces. However, other types of attachment interfaces may be
used. Moreover, the attachment interfaces 20, 30 may comprise a cam
assembly in order to provide a sufficient amount of clearance for
the overhanging section 38 of the laminate 36 to avoid inadvertent
contact and pre-adherence with the other side of the interface.
[0063] Preferably, the system 10 further comprises a flexible
elastomeric seal at a joint interface between the first and second
molds 12, 22. The flexible elastomeric seal provides vacuum
integrity of the mold assembly needed for the curing procedure.
[0064] Preferably, a release agent is applied to the first and
second molds 12, 22 prior to laying down of the first and second
modules 13, 23 thereon. The release agent is preferably one of
three types: (i) liquid or paste, (ii) in the form of a plastic
film and (iii) of a permanent type such as a Teflon.TM. coating and
one skilled in the art can select the appropriate one for its
particular need. Additionally, other types of release agents may be
considered, The release agent is applied in each mold to allow
remolding of other modules after a curing step.
[0065] Preferably, the first and second molds 12, 22 are portions
of a cylindrical structure. The system can therefore be used to
form a curved assembly as shown in FIGS. 3a to 3e and FIGS. 4a to
4e. Preferably, the molds have a geometric shape adapted to form a
tubular-profiled structure and comprises at least two 180.degree.
sections or preferably three 120 sections.
[0066] Referring to FIG. 3a, when the molds are made of three
sections, a central mold 12 rests on the ground with the two other
molds 22, 52 placed adjacently.
[0067] Preferably, when the assembly molds 12, 22, 52 comprise
three sections to form a cylindrical structure, the removable
inserts can be positioned in different manners. In a preferred
embodiment of the present invention, one insert is associated with
each mold 12, 22, 52, as shown in FIGS. 3a to 3e, since such a
configuration allows for the manufacture of three (3) identical
molds/inserts. However, other configurations can be considered. For
example, no insert can be associated with the first mold 12, one
insert can be associated with the second mold 22, and two inserts
can be associated with the third mold 52, as shown in FIGS. 5a to
5c.
[0068] The closing sequence of the different molds 12, 22, 52 is
not influenced by the positioning and distribution of the inserts
among the different molds because a clearance zone has been
designed into the shape of the molds in order to position, within
this clearance zone, the overhanging section 38 of the laminate 36
to avoid contact between the two sides of the interface of the
assembled laminate interface during closing of the molds.
[0069] Preferably, for assembly of cylindrical fuselage components,
among other applications, three molds 12, 22, 52 are provided. As
better shown in FIGS. 5a to 5c, the system 10 for fabricating a
composite material assembly comprises a third mold 52 for receiving
a third module made of composite material. The third mold 52
includes a third composite material laminate support structure 54
having first and second opposite edges 56, 58. The third mold also
has a pair of opposite third and fourth attachment interfaces 60,
62 for attachment of the third mold 52 to the first and second
molds 12, 22. The third mold 52 also has second and third removable
inserts 64, 65 extending beyond the first and second edges 56, 58
of the third composite material laminate support structure 54. The
first mold 12 comprises a fifth attachment interface 66 for
attachment of the first mold 12 to the third mold 52. The second
mold 22 comprises a sixth attachment interface 68 for attachment of
the second mold 22 to the third mold 52. A third laminate 70 covers
the third layup structure 54 and extends over the second and third
removable inserts 64, 65. The first, second and third removable
inserts 32, 64, 65 are removed from the second and third molds 22,
52 prior to assembly of the first, second and third molds 12, 22,
52. A section 71 of the third laminate extending over the second
removable insert 64 overlaps over the first laminate 34 after
closing and assembly of the third mold 52 onto the first mold 12.
Another section 73 of the third laminate extending over the third
removable insert 65 overlaps over the second laminate 36 after
closing and assembly of the third mold 52 onto the second mold 22.
As mentioned above, the distribution of the inserts 32, 64, 65
among the different molds as shown in FIGS. 5a to 5e may vary for a
selected assembly closure sequence and correspond, for example, to
the distribution of inserts 32, 64, 65 shown in FIGS. 3a to 3e. In
FIGS. 5a to 5e, the distribution of the inserts is such that first
mold 12 resting on the ground has no inserts and sections 71, 38
overlap over the first laminate which may be a more practical
sequence of assembly of the laminates in certain assembly
configurations.
[0070] According to the present invention, there is also provided a
method for fabricating a composite material assembly comprising the
steps of: [0071] a) providing an assembly system 10, as shown in
FIGS. 1a to 1c comprising: [0072] a first mold 12 for receiving a
first module 13 made of composite material, the first mold 12
comprising: [0073] a first composite material laminate support
structure 14 having first and second opposite edges 16, 18; and
[0074] a first attachment interface 20 for attachment of the first
mold 12 to an adjacent mold 22; and [0075] a second mold 22 for
receiving a second module 23 made of composite material, the second
mold 22 comprising: [0076] a second composite material laminate
support structure 24 having first and second opposite edges 26, 28;
[0077] a second attachment interface 30 for attachment of the
second mold 22 to the first mold 12; and [0078] at least one
removable insert 32 extending beyond the edge 28; [0079] b) laying
down the first module 13 on the first mold 12, the first module 13
comprising a first laminate 34 covering the first laminate support
structure 14; [0080] c) laying down the second module 23 on the
second mold 22, the second module 23 comprising a second laminate
36 covering the second laminate support structure 24 and extending
over the removable insert 32; [0081] d) removing the removable
insert 32 from the second mold 22; [0082] e) assembling the first
mold 12 with the second mold 22 while overlapping a section 38 of
the second laminate 36 extending over the removable insert 32 over
the first laminate 34.
[0083] Preferably, the method further comprises the step of f)
curing the assembled first and second modules 13, 23 in an oven.
When the method according to the present invention is used to
manufacture a fuselage assembly, considering the fact that the
entire composite structure of the fuselage has been realized in a
complete uncured state and that the composite structure is fully
assembled in a tubular profile, the entire fuselage assembly inside
the closed mold has to be solidified by putting it under vacuum and
heat inside a curing oven. Under only one "heat and pressure cycle"
the pre-preg laminate and adhesive will cure and solidify to
generate a one-piece tubular section of fuselage without an overly
apparent seam. It is however understood by one skilled in the art
that any appropriate curing process is possible pursuant to the
invention.
[0084] Preferably, the one-piece section of fuselage produced using
the system or method may integrate or comprise floor attachment
members, a cockpit windshield, cabin windows and passenger door
surrounding structures. All of these features may be all cured in
one step only. The system and method according to the present
invention can be used for any portion of a flying vehicle which
possesses a tubular profile with a need to be co-cured for reducing
any overly apparent seam, such as any cabin of an aircraft.
[0085] Referring to FIG. 7, the system and method according to the
present invention can be used for manufacturing of one-piece
fuselage sections and facilitate the layup of composite pre-prep
material on the molds 12, 22, 52 in an almost horizontal position,
thus reducing the counter effect of gravity when compared to a
tubular or cylindrical molds.
[0086] Although preferred embodiments of the present invention have
been described in detail herein and illustrated in the accompanying
drawings, it is to be understood that the invention is not limited
to these precise embodiments and that various changes and
modifications may be effected therein without departing from the
scope or spirit of the present invention.
* * * * *