U.S. patent application number 12/589059 was filed with the patent office on 2010-02-18 for advanced composite aerostructure article having a braided co-cured fly away hollow mandrel and method for fabrication.
Invention is credited to Gregory A. Allen, Anthony E. Reed.
Application Number | 20100038030 12/589059 |
Document ID | / |
Family ID | 41350865 |
Filed Date | 2010-02-18 |
United States Patent
Application |
20100038030 |
Kind Code |
A1 |
Allen; Gregory A. ; et
al. |
February 18, 2010 |
Advanced composite aerostructure article having a braided co-cured
fly away hollow mandrel and method for fabrication
Abstract
An article of fiber-reinforced composite material formed by
co-curing a lay up under a cycle of heat and pressure. The lay up
comprises a first uncured composite layer having at least one
uncured resin-impregnated laminate layer. At least one hollow
mandrel is provided comprised of a stiffened braided fabric and
having an upper, a lower and side surfaces. The mandrel is secured
on the upper surface of the first composite layer with the lower
surface thereof in engagement with the upper surface of the first
composite layer. A second uncured composite layer is positioned
over the upper and side surfaces of the hollow mandrel and at least
a portion of the upper surface the first uncured composite layer.
The second uncured composite layer has at least one uncured
resin-impregnated layer. After the cycle of heat and pressure in an
autoclave, the hollow mandrel is retained in the co-cured
article.
Inventors: |
Allen; Gregory A.;
(Prescott, AZ) ; Reed; Anthony E.; (Riverside,
CA) |
Correspondence
Address: |
GOODWIN PROCTER LLP;ATTN: PATENT ADMINISTRATOR
620 Eighth Avenue
NEW YORK
NY
10018
US
|
Family ID: |
41350865 |
Appl. No.: |
12/589059 |
Filed: |
October 16, 2009 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
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10440061 |
May 15, 2003 |
7625618 |
|
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12589059 |
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Current U.S.
Class: |
156/307.1 |
Current CPC
Class: |
B29D 22/00 20130101;
B32B 3/08 20130101; B29C 70/44 20130101; B29L 2031/3076 20130101;
B29D 99/001 20130101; B32B 5/28 20130101; B29C 70/543 20130101;
Y10T 428/1393 20150115; B32B 5/26 20130101; B29K 2105/246 20130101;
B32B 27/04 20130101; B29D 99/0007 20130101 |
Class at
Publication: |
156/307.1 |
International
Class: |
B32B 37/02 20060101
B32B037/02 |
Claims
1-16. (canceled)
17. A method of fabricating an article from fiber-reinforced
composite material comprising: assembling on a composite bond
surface a first uncured composite layer in the shape of at least
one surface of the article, the first composite layer having an
upper surface and at least one uncured resin-impregnated laminate
layer; providing at least one hollow mandrel, the at least one
hollow mandrel comprised of a braided fabric having upper, lower
and side surfaces; securing on the upper surface of the first
composite layer the hollow mandrel with the lower surface thereof
in engagement with the upper surface of the first composite layer;
assembling a second uncured composite layer on the bond surface
over the upper and side surfaces of the at least one hollow mandrel
and at least a portion of the upper surface the first uncured
composite layer, the second uncured composite layer having at least
one uncured resin-impregnated layer; and subjecting the lay up
assembled on the composite bond surface to a cycle of predetermined
pressure and temperature to cure the resin-impregnated layers so as
to form a unitary co-cured one-piece article, wherein the hollow
mandrel is retained in the co-cured article.
18. The method of claim 17, wherein the hollow mandrel is
impregnated with a resin.
19. The method of claim 17, wherein the hollow mandrel is pre-cured
to a predetermined shape.
20. The method of claim 17, wherein the cross-sectional shape of
the hollow mandrel varies along a longitudinal extent thereof.
21. The method of claim 17, wherein the step of providing at least
one hollow mandrel, comprises providing a pair of hollow mandrels,
and wherein the step of securing on the upper surface of the first
composite layer the hollow mandrel comprises securing the pair of
mandrels on the first composite layer in a laterally spaced-apart
configuration.
22. The method of claim 17, further comprising: prior to the step
of subjecting the lay up assembled on the composite bond surface to
a cycle of predetermined pressure and temperature, assembling on a
composite bond surface a third uncured composite layer in the shape
of at least a second surface of the article, the third composite
layer having at least one uncured resin-impregnated laminate layer;
providing a second hollow mandrel, the second hollow mandrel
comprised of a braided fabric having upper, lower and side
surfaces; and securing on the upper surface of the first composite
layer the hollow mandrel with the lower surface thereof in
engagement with the upper surface of the first composite layer; and
assembling a fourth uncured composite layer on the bond surface
over the upper and side surfaces of the second hollow mandrel and
at least a portion of the upper surface the third uncured composite
layer, the fourth uncured composite layer having at least one
uncured resin-impregnated layer.
Description
BACKGROUND OF THE INVENTION
[0001] 1. Field of the Invention
[0002] The present invention relates generally to the field of
advanced composite aerostructure articles and methods of
fabrication, more particularly, to an advanced composite
aerostructure article having an integral co-cured fly away hollow
mandrel and a method of fabrication.
[0003] 2. Background Information
[0004] There is a growing trend in the aerospace industry to expand
the use of advanced composite materials for a diverse array of
structural and dynamic aerostructural applications because of the
strength-to-weight advantage provided by composite materials. One
particular application for the use of such advanced composite
materials lies in the fabrication of advanced composite articles
such as panels for nacelles for aircraft jet engine propulsion
systems and for control surface components, such as spoilers. Such
structural articles generally comprise inner and outer composite
skins, which are formed from composite materials such as graphite
or an aromatic polyamide fiber of high tensile strength that are
embedded in a resinous matrix, e.g., epoxy, having a honeycomb core
material interposed therebetween. One or more stiffening members
may be affixed to the outer skin and covered with an inner skin for
efficiently transmitting and/or reacting axial and/or bending loads
to which the component is subjected. Aerostructure articles, such
as spoiler components, have more than one external surface, i.e., a
lower and upper surface, having curvatures and often taper, at an
aft location, to a sharp edge. Manufacture of these articles
present additional challenges to current techniques for fabricating
composite structures.
[0005] There are two techniques currently employed for bonding
through autoclave processing a composite stiffening member in
combination with a composite structural panel: (1) the co-cured
bonding method; and (2) the secondary bonding method. Both methods
are disadvantageous in requiring costly non-recurring tooling
and/or costly recurring manufacturing steps.
[0006] A typical composite sandwich panel intended for use as an
aerostructure article is normally fabricated using two autoclave
cured inner and outer composite skins that are formed by using a
curing cycle with heat, pressure, and a unique tool for each skin.
A sandwich panel is then made up using a composite bond jig, tool
or fixture with the pre-cured face skin laid up on the bond jig
tool followed by a ply of film adhesive, a honeycomb aluminum or
non-metallic core of a given thickness, another ply of film
adhesive and finally the previously pre-cured inner skin. The bond
jig that is used to fabricate the sandwich panel is generally the
same tool that was used to create the outer composite skin. A
plurality of closure plies of uncured composite material are layed
up and the assembled sandwich panel are cured during their final
assembly stage. This sandwich panel is then vacuum bagged to the
composite bond jig and again cured in an autoclave under high
pressure and heat.
[0007] Thus, at least three very expensive and man-hour consuming
cure cycles have gone into the fabrication of this exceptionally
strong but lightweight composite/honeycomb core sandwich panel. At
least two different and expensive tools are needed in this process.
Manufacturing flow time is very long, energy use is high and the
manufacturing floor space required is excessive.
[0008] The co-curing method envisions curing the composite inner
and outer skins that are laid up with a layer of adhesive film and
honeycomb core in one cure cycle in the autoclave. A co-cured panel
is desirable in that it is less expensive to fabricate since only
one bond jig tool is required, only one cure cycle is needed, it is
less labor intensive, it requires less floor space to accomplish,
and a much shorter manufacturing flow time is achieved. However,
co-curing an aerostructure panel has never achieved widespread
acceptance because of the large loss of panel strength and
integrity that is lost due to the lack of compaction of the
composite plies placed over and under the honeycomb core details.
The composite plies dimple into the center of each core cell with
nothing but the cell walls to compact the composite skins. The only
way to overcome this knockdown factor is to add extra plies which
creates both unwanted weight and excess cost. Thus, because of
these constraints co-cured aerostructure panels are not widely
manufactured in the aerospace industry.
[0009] There are other particular problems when a honeycomb core
element is used to provide a stiffening element for an aerospace
article such as a fan cowl. As Hartz et al described in U.S. Pat.
No. 5,604,010 concerning a "Composite Honeycomb Sandwich
Structure," with a high-flow resin system, large amounts of resin
can flow into the core during the autoclave processing cycle. Such
flow robs resin from the laminate, introduces a weight penalty in
the panel to achieve the desired performance, and forces
over-design of the laminate plies to account for the flow losses.
To achieve the designed performance and the corresponding laminate
thickness, additional plies are necessary with resulting cost and
weight penalties. Because the weight penalty is severe in terms of
the impact on vehicle performance and costly in modern aircraft and
because the flow is a relatively unpredictable and uncontrolled
process, aerospace design and manufacture dictates that flow into
the core be eliminated or significantly reduced. In addition to the
weight penalty from resin flow to the core, it was discovered that
microcracking that originated in the migrated resin could propagate
to the bond line and degrade mechanical performance. Such
microcracking potential has a catastrophic threat to the integrity
of the panel and dictates that flow be eliminated or, at least,
controlled.
[0010] Unfortunately, the use of honeycomb core as a stiffener for
elements in an aerostructure article such as a fan cowl, or in a
structural panel has other deleterious effects, two of the greatest
drawbacks to aluminum core being its inherent significant cost and
corrosion. To minimize galvanic corrosion of the core caused by
contact with the face skins, isolating sheets are interposed
between such aluminum core and the face skins. Also, the aluminum
core has an inherent cost and also must be machined to a desired
shape in an expensive process. The honeycomb core may also be
subject to crush during manufacture and thereby limits the
pressures that may be used in autoclave processing. Also, the
honeycomb core if damaged in use has a spring back effect which
makes the detection of such damage more difficult.
[0011] Another method for reinforcing mandrels for stiffener
elements, such as hat sections, for aerospace advanced composite
structural panels involves the use of a composite stiffening member
formed over a polyimide foam mandrel which is fabricated by
machining a core mandrel to a desired shape. Obviously, the
machining of the core mandrel is expensive and time consuming and
further introduces the problem of properly bonding the core mandrel
to inner and outer skins.
[0012] A novel and useful technique for fabricating an advanced
composite aerostructure article is described in U.S. Pat. No.
6,458,309, and co-pending U.S. application Ser. No. 10/142,490,
filed May 5, 2002, both of which are incorporated by reference in
their entirety herein.
[0013] A need has arisen for a practical method of readily
producing stiffened, fiber-reinforced composite structures for
aerospace applications which are cost and labor efficient and which
save time in the fabrication process.
SUMMARY OF THE INVENTION
[0014] Accordingly, it is an object of the present invention to
provide a method for fabricating aerostructure advanced composite
articles that eliminates honeycomb core in stiffening elements,
provides a lighter weight assembly and is easier to repair.
[0015] Another object of the present invention is to reduce the lay
up cost of known advanced composite co-cure assemblies and to
increase assembly strength over such existing co-cure assembly by
being able to utilize advanced pressures in autoclave
processing.
[0016] Yet another object of the present invention is to improve
the quality of assembly of such co-cured advance composite
assemblies and thereby increase customer satisfaction.
[0017] A further object of the present invention is to provide a
process that provides an aerostructure component having a plurality
of external curved surfaces and that tapers to a narrow
dimension.
[0018] These and other objects of the invention, which will become
apparent with reference to the disclosure herein, are accomplished
by a new and improved method for fabricating an advanced composite
aerostructure article from fiber reinforced composite material and
incorporating a hollow stiffened graphite fabric mandrel that
becomes an integral part of such article.
[0019] In accordance with a preferred embodiment of the present
invention, an article of fiber-reinforced composite material formed
by co-curing a lay up under a cycle of heat and pressure is
provided, wherein the lay up comprises a first uncured composite
layer having at least one uncured resin-impregnated laminate layer.
At least one hollow mandrel is provided comprised of a pre-cured
stiffened braided or woven fabric and having an upper, a lower and
side surfaces and being secured on the upper surface of the first
composite layer with the lower surface thereof in engagement with
the upper surface of the first composite layer
[0020] A second uncured composite layer is positioned over the
upper and side surfaces of the hollow mandrel and at least a
portion of the upper surface of the first uncured composite layer.
The second uncured composite layer has at least one uncured
resin-impregnated layer. The hollow mandrel is retained in the
co-cured article.
[0021] According to an exemplary embodiment, the hollow mandrel of
the lay up is impregnated with a resin. The hollow mandrel of the
lay up may be pre-cured to a predetermined shape. At least one of
the upper and lower surface of the hollow mandrel may have a curved
shape. The hollow mandrel of the lay up may have a cross-sectional
area that varies along a longitudinal extent thereof.
[0022] The lay up may include multiple hollow mandrels secured to
the first composite layer in a laterally spaced-apart relation, and
the second composite layer covers the two laterally spaced-apart
hollow mandrels. At least one of the first and second uncured
composite layer may comprise a plurality of serially laid-up
uncured resin impregnated layers.
[0023] A third uncured composite layer may be provided having at
least one uncured resin-impregnated laminate layer. A second hollow
mandrel, comprised of a pre-cured stiffened braided or woven fabric
and having an upper, a lower and side surfaces, is secured on the
upper surface of the third composite layer with the lower surface
thereof in engagement with the upper surface of the third uncured
composite layer. A fourth uncured composite layer positioned over
the upper and side surfaces of the second hollow mandrel and at
least a portion of the upper surface the third uncured composite
layer, the fourth uncured composite layer having at least one
uncured resin-impregnated layer, wherein the second hollow mandrel
is retained in the co-cured article. The second composite layer is
positioned in engagement with the fourth composite layer. The first
and third composite layers may become the outer skin surfaces of
the aerostructure article.
[0024] The cross-sectional area of the hollow mandrel may be varied
as desired along the longitudinal extent thereof to provide a
preferred shape such as a tapered section, and to provide multiple
curved configurations.
[0025] In accordance with the invention, the objects of providing
an improved aerostructure article have been met. Additional
features of the invention will be described hereinafter which form
the subject of the claims of the invention. It should be
appreciated by those skilled in the art that the conception and the
disclosed specific embodiment may be readily utilized as a basis
for modifying or designing other structures and methods for
carrying out the same purposes of the present invention. It should
also be realized by those skilled in the art that such equivalent
constructions and methods do not depart from the spirit and scope
of the invention as set forth in the attached claims.
BRIEF DESCRIPTION OF THE DRAWINGS
[0026] For a fuller understanding of the nature and objects of the
invention, reference should be had to the following detailed
description taken in conjunction with the accompanying drawings in
which:
[0027] FIG. 1 is a perspective of a co-cured aerospace article that
has been fabricated by the method of a preferred embodiment of the
present invention.
[0028] FIG. 2 is a perspective view of a stage in the process of
fabricating the article of FIG. 1, in accordance with the present
invention.
[0029] FIG. 3 is a transverse sectional view of a portion of the
lay-up assembly in accordance with the present invention.
[0030] FIG. 4 is a side view of a mandrel comprising a braided
fabric useful for fabricating the article of FIG. 1 in accordance
with the present invention.
[0031] FIG. 5 is a perspective view of the mandrel of FIG. 4 in
accordance with the present invention.
[0032] FIG. 6 is a perspective view of a further stage in the
process of fabricating the article of FIG. 1, in accordance with
the present invention.
[0033] FIG. 7 is a perspective that illustrates the insertion of a
tube bag into a hollow mandrel in accordance with the present
invention.
[0034] FIG. 8 is a perspective that illustrates the sealing of a
tube bag to a vacuum bag and the sealing of the vacuum bag to a
bond jig in accordance with the present invention.
[0035] FIG. 9 is a perspective view of another stage in the process
of fabricating the article of FIG. 1, in accordance with the
present invention.
[0036] FIG. 10 is a perspective view of a another stage in the
process of fabricating the article of FIG. 1, in accordance with
the present invention.
[0037] FIG. 11 is a perspective view of a further stage in the
process of fabricating the article of FIG. 1, in accordance with
the present invention.
[0038] FIG. 12 is a perspective view of a still further stage in
the process of fabricating the article of FIG. 1, in accordance
with the present invention.
[0039] FIG. 13 is a perspective view of another stage in the
process of fabricating the article of FIG. 1, in accordance with
the present invention.
[0040] FIG. 14 is a transverse sectional view of the lay-up
assembly in accordance with the present invention.
[0041] Similar numerals refer to similar parts in the drawing.
DETAILED DESCRIPTION OF THE INVENTION
[0042] Referring to the drawings in detail and in particular to
FIG. 1, the reference character 10 generally designates an
aerostructure article constructed in accordance with an exemplary
embodiment of the invention. The article 10 is an advanced
composite co-cured structure having a unitary skin 12 and 14 and
longitudinally extending spaced reinforcing hat sections 16, 18 and
20 for the illustrated exemplary part. The article 10 as shown
tapers from the forward portion 22 to the aft portion 24. The
illustrative example of the invention 10, shown in FIG. 1, is
representative of a spoiler unit for an aircraft wing, but the use
of the invention for the fabrication of aerostructure articles
extends to other shapes and applications.
[0043] Referring now to FIG. 2, the specific novel method of the
present invention will be described. A suitable lay-up mandrel or
composite bond jig (COBJ) 30 having a predetermined shape, such as
that used to provide the accurate shape of the skin 14 of the
article 10 illustrated in FIG. 1, is provided for receiving a
lay-up assembly for co-curing. A first uncured composite layer 32
is assembled on an upper surface of the COBJ 30 to provide the
shape of at least one surface of the composite aerostructure
article 10. The first uncured composite layer 32 having at least
one uncured resin-impregnated laminate layer that is generally a
graphite or aramid fabric laminate layer that is termed a
"pre-preg" in the advanced composite manufacturing industry. In the
exemplary embodiment, the first uncured composite layer 32 is
comprised of six uncured resin-impregnated laminate plies of
RMS-060 Type 2 material, which is a carbon fabric material in a
5-harness, balanced weave, 6000 end yarn (370 g/m.sup.2)
pre-impregnated with a 350.degree. F. cure epoxy resin. RMS-060
Type 2 materials are available from, for example, Hexel Corporation
and Cytec Corporation.
[0044] With reference to FIGS. 2 and 3, a pair of lightweight
hollow mandrels 40 and 42 are used to form hat sections 18 and 20
of the article 10 (illustrated in FIG. 1, above.) Mandrels 40 and
42 are layed-up on the first uncured composite layer 32 by a
suitable uncured tacking adhesive 44, shown in FIG. 3, such as BMS
5-154, which is a film adhesive consisting of a 350.degree. F. cure
epoxy resin coated to a final film weight of 0.03-0.08 PSF on a
woven fiberglass fabric. Such BMS 5-154 materials are available,
for example, from Cytec Corporation. The hollow mandrels 40 and 42
are, in effect, "tacked" to the first preform layer 32 by the
adhesive.
[0045] Mandrels 40 and 42 are manufactured from a stiffened,
braided material, such as carbon fiber with epoxy compatible
surface finish. The mandrel may be woven with 3K, 6K, or 12K carbon
fiber tow bundles in a variety of weave orientations or angles in
order to achieve various complex geometries and load carrying
capabilities. Custom tailored braided products are available from
several sources including: A&P Technology; Atlantic Research
Corporation; Fabric Development Inc.; Techniweave; and Textile
Products Inc. Prior to tacking to the first uncured composite layer
32, the braided material of mandrels 40 and 42 is impregnated with
a material, such as a resin material and then cured to a specific
shape. Suitable resins include two part room temperature epoxy
systems, 350.degree. F. cure epoxy resin compatible with RMS 060
pre-impregnated carbon fabric materials, and higher cure
temperature resin systems if required which are then cured to a
specific shape. The shape of mandrels 40 and 42 depends upon the
component being manufactured. As illustrated in FIGS. 4 and 5, the
hollow mandrel 40 is a pre-stiffened shape that longitudinally
extends along the upper surface of the layer 32. (Mandrel 42 is
substantially identical to mandrel 40 illustrated in FIGS. 4 and
5.) The cross-sectional shape of the hollow mandrel 40 may be
determined to provide optimal stiffening of the article 10. A
cross-sectional form for the hollow mandrel 40 in the exemplary
embodiment is rectangular (as illustrated in FIG. 5, above), but
other shapes such as square, trapezoidal or round could be provided
as desired. Another desirable feature of the hollow mandrel 40 is
that the cross-sectional shape may be varied along the longitudinal
extent of a hat section and to provide a curved upper surface 51
and/or lower surface 53. For example, if the article 10 were to be
a spoiler unit then the height of the hollow mandrel 40 is
decreased fore to aft to provide a tapered aft-portion of an
airfoil shape. The hollow mandrels 40 are pre-stiffened to the
extent required to provide a desired hat configuration during the
continuing fabrication of the article 10 as described
hereinafter.
[0046] As illustrated in FIGS. 3 and 6, a second uncured composite
layer 34 is then assembled on the COBJ 30 over the upper outer
surface of the hollow mandrels 40 and 42 and over at least a
portion of the first composite layer 32. In the exemplary
embodiment, second uncured composite layer 34 is a three-ply layer
of material such as RMS 060 Type 3. RMS 060, Type 3 material is a
carbon fabric in a 5 harness, balanced weave, 3000 end yarn (280
g/m.sup.2) configuration which is pre-impregnated with a
350.degree. F. cure epoxy resin. RMS 060 Type 3 material is used
over the mandrel because of its improved conformability for lay up
over the pre-cured mandrel and back down to the basic uncured skin.
It is available from Hexcel Corporation and Cytec Corporation.
[0047] Depending upon the composition and manufacturing process,
the cross-sectional shape of the hollow mandrels 40 and 42 may not
have sharp corners, but rather more rounded corners 48 as
illustrated in FIG. 3. While the strength of a hat section
fabricated with hollow mandrels having rounded corners is
sufficiently strong, the strength of such a hat section can be
improved by optionally providing, in effect, a fillet 50 along the
longitudinal length of the intersection of the bottom of the hollow
mandrel 40 and 42 and the first composite preform layer 32, as
illustrated in FIG. 3. This fillet 50 may be advantageously
provided by positioning a very small roll of graphite fiber fabric
material or a small roll of unidirectional material or
longitudinally extending strands of such material in the small
circumferentially extending gap provided by rounded corners 48 of
hollow mandrels 40 and 42 and the first uncured composite layer 32
as shown in FIG. 3. The graphite fiber material of the fillet 50
may either be uncured resin-impregnated graphite material or a
resin-free graphite material. In the exemplary embodiment, the
fillet 50 is fabricated from RMS-040 material. RMS 040 material is
unidirectional aligned carbon fiber tape (148/m.sup.2) which is
pre-impregnated with a 350.degree. F. cure epoxy resin. It is
available from Hexcel Corporation and Cytec Corporation. For
clarity, the assembly of first uncured composite layer 32, mandrels
40 and 42, second uncured composite layer 34, and fillets 50, may
be collectively referred to as first component 70 in FIG. 6.
[0048] The lay-up assembly (i.e., component 70) provided thus far
by the method of this invention, is then covered by suitable vacuum
bags 52 and 54 as shown in FIGS. 7 and 8. Tube bag 52 is positioned
within mandrels 40 (FIG. 7) and 42 (not shown). Vacuum bag 54 is
sealed to the tool periphery and to vacuum bag 52 using a suitable
sealant tape such that the inside of tube bag 52 and the outside of
bag 54 are exposed to the surrounding atmospheric pressure when
vacuum is applied between tube bag 52 and vacuum bag 54 (FIG. 8).
When vacuum is applied between bags 52 and 54 the surrounding
atmospheric pressure pushes the bag against all flat areas of lay
up 70, outside mandrels 40 and 42, and inside mandrels 40 and 42
such that the lay up is consolidated without collapsing the
stiffener shape provided by the mandrels. The consolidation under
vacuum is conducted for a specific period of time determined to
achieve the desired level of compaction. A typical compaction time
is 15 minutes or longer.
[0049] The process described above for FIGS. 2-8 is now repeated
for the second component 72 of the lay up assembly, illustrated in
FIGS. 9 and 10. As illustrated in FIG. 9, uncured composite layer
64 is placed on COBJ 62. Lightweight hollow mandrel 46 is layed-up
on the composite layer 64 by a suitable uncured tacking adhesive,
as illustrated in FIG. 3. An exemplary tacking adhesive is BMS
5-154, described above. Mandrel 46 is substantially identical to
mandrels 40 and 42, described above, and illustrated in FIGS. 4-5.
Fillets 50, as illustrated in FIG. 3, may be placed in corners 48
of mandrel 46. As illustrated in FIG. 10, uncured composite layer
66 is then assembled on the COBJ 62 over the upper outer surface of
the hollow mandrel 46 and over at least a portion of the first
composite layer 64. Component 72 in FIG. 10 comprises the third
uncured composite layer 64, mandrel 46, and the fourth uncured
composite layer 66. The vacuum procedure described above and
illustrated in FIGS. 7-8 is repeated to form the uncured composite
layer 66 over the mandrel 46 and uncured composite layer 64.
[0050] As illustrated in FIGS. 11, 12, and 13, the two components
70 and 72 of the lay up assembly are arranged for curing. In
particular, second component 72 may be moved as indicated by arrow
A (FIG. 12), and positioned such that the mandrel 46 is positioned
between laterally spaced-apart mandrels 40 and 42 (FIG. 13).
Moreover, second uncured composite layer 66 of component 72 is
placed in engagement with second uncured composite layer 34 of
component 70 as shown in FIG. 13.
[0051] With reference to FIG. 14, uncured composite layers 76 and
78 are layed-up over the side portions of mandrels 40 and 42, which
assist in joining layer 66 and 34. In the exemplary embodiment,
layers 76 and 78 comprise three-plies of RMS-060 Type 2. In similar
fashion, layers 80 and 82 may be added to the inside forward edge
of the mandrels, and if required may wrap in a 90 degree angle to
form a localized forward flange that is tied into the mandrel.
These type of layers are represented by 80 and 82 may be added to
any of the mandrels in order to locally increase stiffness or
thickness of the forward end of the mandrel for attachment of
hardware at a future assembly stage. In the exemplary embodiment,
layers 80 and 82 comprise four plies of RMS 060 Type 2 fabric.
[0052] The lay up assembly, arranged as shown in FIG. 14, is vacuum
bagged for autoclave cure. Vacuum tube bags, similar to vacuum tube
bags 52 in FIG. 7, described above, are placed within mandrels 40,
42, and 46, and within the open spaces between the opposing skin
stiffeners as depicted by 84 and 86 in FIG. 14. A non-bondable
separator film is placed over the lay up and a breather fabric is
placed over the non-bondable material. Finally, an outer bag is
placed over the entire lay up, and is sealed to the bond tool and
the individual tube bags, as previously described above. The lay up
assembly arranged on the COBJ 30 is then placed in a suitable
autoclave and subjected to a suitable curing cycle having an
elevated temperature and elevated pressure over a desired time
period to provide for curing of the layers 32, 34 and 64, 66 and
the hollow mandrels 40, 42, and 46 into a unitary co-cured one
piece aerostructure article. The increased pressure within the
autoclave will be balanced by the pressure within the hollow
mandrels 40, 42, and 46 since the expanded tube bags will conform
to the interior shape of the hollow mandrels 40, 42, and 46 (as
illustrated in FIG. 7-8).
[0053] The pressure used for curing the aerostructure article 10 in
the autoclave can be within a range of 65-75 psig for the curing
cycle for resin-impregnated graphite or aramid fabric material.
Other higher pressures could be used for the curing cycle, as may
be desired for other advanced composite materials. The use of an
increased pressure provides for greater quality of the article and
increased physical properties. The curing process will provide for
the resins of the first composite preform layers 32 and 64 and
second composite preform layers 34 and 66 to penetrate the graphite
fabric material of the hollow mandrels 40, 42, and 46 and the
tacking adhesive flows as well to provide an integral co-cured
article.
[0054] After the curing cycle has been completed in the autoclave,
the COBJ 30 is removed from the autoclave and the vacuum bags are
removed from the article 10 and the mandrels 40, 42, and 46. The
article 10 may then be removed from the COBJ 30 and suitably
trimmed for further use.
[0055] The present disclosure includes that contained in the
appended claims as well as that of the foregoing description.
Although this invention has been described in its preferred forms
with a certain degree of particularity, it is understood that the
present disclosure of the preferred form has been made only by way
of example and numerous changes in the details of construction and
combination and arrangement of parts and method steps may be
resorted to without departing from the spirit and scope of the
invention.
* * * * *