U.S. patent application number 11/790997 was filed with the patent office on 2007-09-06 for method of sealing core material.
This patent application is currently assigned to Toyota Motor Sales, U.S.A., Inc.. Invention is credited to Richard J. Gardiner, Michael K. Maxwell.
Application Number | 20070207323 11/790997 |
Document ID | / |
Family ID | 21689662 |
Filed Date | 2007-09-06 |
United States Patent
Application |
20070207323 |
Kind Code |
A1 |
Maxwell; Michael K. ; et
al. |
September 6, 2007 |
Method of sealing core material
Abstract
A molded composite structure and a method of manufacturing a
molded composite structure are disclosed. In one embodiment, an
aircraft wing panel and a method for manufacturing an aircraft wing
panel are disclosed. In another embodiment, the method of
manufacturing a molded composite structure uses a resin transfer
molding process.
Inventors: |
Maxwell; Michael K.; (Long
Beach, CA) ; Gardiner; Richard J.; (Murray,
UT) |
Correspondence
Address: |
FINNEGAN, HENDERSON, FARABOW, GARRETT & DUNNER;LLP
901 NEW YORK AVENUE, NW
WASHINGTON
DC
20001-4413
US
|
Assignee: |
Toyota Motor Sales, U.S.A.,
Inc.
|
Family ID: |
21689662 |
Appl. No.: |
11/790997 |
Filed: |
April 30, 2007 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
|
10000048 |
Oct 31, 2001 |
6557702 |
|
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11790997 |
Apr 30, 2007 |
|
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60254080 |
Dec 8, 2000 |
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Current U.S.
Class: |
428/414 ;
156/285; 428/411.1; 428/413; 428/421; 428/473.5 |
Current CPC
Class: |
Y10T 428/31515 20150401;
B29C 66/45 20130101; A63B 55/30 20151001; Y10T 428/31504 20150401;
B32B 37/146 20130101; B32B 7/12 20130101; A45C 5/06 20130101; B29C
70/865 20130101; Y10T 428/31511 20150401; A45C 5/14 20130101; Y10T
428/3154 20150401; A63B 55/408 20151001; A45C 13/36 20130101; A63B
55/404 20151001; B29C 70/48 20130101; B29C 2793/00 20130101; B29L
2031/3085 20130101; B32B 2605/18 20130101; Y10T 428/31721 20150401;
A45C 13/02 20130101; A45C 5/02 20130101 |
Class at
Publication: |
428/414 ;
156/285; 428/411.1; 428/413; 428/421; 428/473.5 |
International
Class: |
B32B 9/04 20060101
B32B009/04; B29C 65/00 20060101 B29C065/00 |
Claims
1-100. (canceled)
101. A method of sealing a core to be used in forming a composite
structure, the method comprising: applying a first thermoplastic
barrier film to a first layer of supporting material; applying an
adhesive on a surface of the core to form a first adhesive layer on
the core; applying the first thermoplastic barrier film and first
layer of supporting material on the core such that the first
adhesive layer is against the first stabilizing material and the
first layer of supporting material provides a vacuum path for
evacuation of the core; evacuating air from the core through the
vacuum path provided by the first layer of supporting material;
curing the first adhesive layer to seal the portion of the core
covered by the first thermoplastic barrier layer; machining the
core to a desired shape; applying a second thermoplastic barrier
film to a second layer of supporting material; applying an adhesive
on all unsealed surfaces of the core to form a second adhesive
layer on the unsealed surfaces of the core; applying the second
thermoplastic barrier film and second layer of supporting material
on the core such that the second adhesive layer is against the
second layer of supporting material and the second layer of
supporting material provides a vacuum path for evacuation of the
core; evacuating air from the core through the vacuum path provided
by the second layer of supporting material; and curing the second
adhesive layer to form a sealed core.
102. The method of claim 101, wherein the first and second layers
of adhesive are roll-coated adhesives.
103. The method of claim 102, wherein the first and second layers
of roll-coated adhesive include an epoxy resin.
104. The method of claim 101, wherein the first and second layers
of supporting material include fibrous material.
105. The method of claim 101, wherein the first and second layers
of supporting material include a ply of composite skin
material.
106. The method of claim 105, wherein the composite skin material
is impregnated with resin.
107. The method of claim 105, wherein the composite skin material
includes carbon fiber impregnated with an epoxy resin.
108. The method of claim 107, wherein the carbon fiber has been
wound so that substantially no gaps exist between adjacent bands of
fibers.
109. The method of claim 101, wherein the first and second
thermoplastic barrier films include a polyetherimide thermoplastic
barrier film.
110. The method of claim 101, wherein the first and second
thermoplastic barrier films include a polyvinyl fluoride film.
111. A method of stabilizing and sealing a core to be used in
forming a composite structure comprising: machining the core to a
desired shape; applying a thermoplastic barrier film to a layer of
supporting material; applying an adhesive on all surfaces of the
core to form an adhesive layer on the core; applying the
thermoplastic barrier film and first layer of supporting material
on the core such that the adhesive layer is against the first layer
of supporting material, the thermoplastic barrier film covers all
surfaces of the core, and the layer of supporting material provides
a vacuum path for evacuation of the core; evacuating air from the
core through the vacuum path provided by the first layer of
supporting material; and curing the adhesive layer to form a sealed
core.
112. The method of claim 111, wherein the layer of adhesive is a
roll-coated adhesive.
113. The method of claim 112, wherein the layer of roll-coated
adhesive includes an epoxy resin.
114. The method of claim 111, wherein the layer of supporting
material includes a fibrous material.
115. The method of claim 111, wherein the layer of supporting
material includes a ply of composite skin material.
116. The method of claim 115, wherein the composite skin material
is impregnated with resin.
117. The method of claim 115, wherein the composite skin material
includes carbon fiber impregnated with an epoxy resin.
118. The method of claim 117, wherein the carbon fiber has been
wound so that substantially no gaps exist between adjacent bands of
fibers.
119. The method of claim 111, wherein the thermoplastic barrier
film includes polyetherimide thermoplastic barrier film.
120. The method of claim 111, wherein thermoplastic barrier film
includes a polyvinyl fluoride film.
121. A sealed core for use in forming a composite structure, said
core comprising: a core substantially evacuated of air; a layer of
adhesive on all surfaces of the core; a layer of supporting
material on the adhesive layer; and a layer of thermoplastic
barrier film on the stabilizing material and covering all surfaces
of the core.
122. The sealed core of claim 121, wherein layer of adhesive is a
roll-coated adhesive.
123. The sealed core of claim 121, wherein the layer of roll-coated
adhesive includes an epoxy resin.
124. The sealed core of claim 121, wherein the layer of supporting
material includes fibrous material.
125. The sealed core of claim 121, wherein the layer of supporting
material includes a ply of composite skin material.
126. The sealed core of claim 125, wherein the composite skin
material is impregnated with resin.
127. The sealed core of claim 125, wherein the composite skin
material includes carbon fiber impregnated with an epoxy resin.
128. The sealed core of claim 127, wherein the carbon fiber has
been wound so that substantially no gaps exist between adjacent
bands of fibers.
129. The sealed core of claim 121, wherein the thermoplastic
barrier film includes polyetherimide thermoplastic barrier
film.
130. The sealed core of claim 121, wherein thermoplastic barrier
film includes a polyvinyl fluoride film.
Description
I. CROSS-REFERENCE TO RELATED APPLICATION
[0001] This application claims the benefit of U.S. Provisional
Application No. 60/254,080, filed Dec. 8, 2000, titled A SMOOTH
LAMINAR FLOW STRUCTURE AND METHOD OF FORMING SAME, the disclosure
of which is expressly incorporated herein by reference.
II. BACKGROUND OF THE INVENTION
[0002] A. Field of the Invention
[0003] The present invention relates to a molded composite
structure and a method of manufacturing a molded composite
structure.
[0004] B. Background of the Invention
[0005] In general, most airplanes comprise a number of components
such as a fuselage, an empennage, and wing structures. Wing
structures are particularly important in the construction of
airplanes because wing structures are the primary lift-producing
structures and perform some of the key functions for these
airplanes. For example, wing structures enable airplanes to take
off and land, to change speed, and to change direction, as well as
other functions. Furthermore, as one of the larger portions of the
aircraft, the overall aerodynamic properties of the aircraft
greatly depend on the wing structures. Finally, the cost of
manufacturing the wing structures has a large impact on the overall
manufacturing cost of these airplanes.
[0006] The ability of the wing structures to perform the functions
discussed above directly depends on the design and construction of
the wing structure. For example, the smoothness and weight of the
wing structures directly impacts the wing structures ability to
perform these functions.
[0007] In particular, the smoothness of the exterior of the wing
structures affects the ability of the aircraft to take off and
land, to change speed, and to change direction. If a wing structure
has an uneven or non-smooth surface, this can create unnecessary
drag, affecting the ability of the wing structures ability to
perform many of the key functions. The aircraft will not be able to
take off and land as easily, and it will be more difficult to alter
the speed and direction of the plane during flight.
[0008] The weight of the wing structures also impacts the ability
of the aircraft to take off and land, to change speed, and to
change direction. The heavier the wing structures are, the more
difficult it will be for the aircraft to take off and land.
Further, heavier wing structures also make it more difficult to
alter the speed and direction of the aircraft during flight.
[0009] The overall aerodynamic properties of an aircraft also
depend on the design and construction of the wing structures.
Ideally, airplanes are designed to create a smooth laminar flow of
air over the aircraft. The smoother the laminar flow of air, the
less energy is needed to fly the aircraft. This therefore reduces
the fuel costs for the plane. If the wing structures are not
designed to be aerodynamically sound, this smooth laminar flow will
be disrupted. For example, if the wing structures do not have a
smooth surface, added drag could result on portions of the wing
structures. This will therefore increase the amount of fuel needed
for flight.
[0010] The cost of manufacturing the wing structures also depends
on the design and construction of the aircraft. The cost of the
material used to manufacture the wing structures as well as the
labor costs of manufacturing the wing structures greatly impacts
the overall cost of the wing structures. Therefore, the design and
construction of wing structures are particularly important in the
manufacture of airplanes.
[0011] Currently, aircraft manufacturers use a number of different
methods to manufacture wing structures. One such process uses a
thin aluminum material to construct the structure. This process
involves the manufacture of sheets of aluminum, which are machined
and attached to one another to form the wing structures. Aluminum
provides an inexpensive source of wing structure material. However,
the manufacture of wing structures from aluminum is
labor-intensive. A large amount of time is spent in manufacturing
and assembling the aluminum sheets. In addition, while aluminum is
a light-weight metal, it is heavier than other non-metal materials
that could be used. Therefore, it causes the wing structure to be
unduly heavy. Finally, the mechanical attachments associated with
aluminum wing structures decrease the smoothness of the wing
structure.
[0012] Another current process uses wet lay-up composites to
manufacture wing structures. Composite materials are light and
inexpensive, and unlike aluminum, can produce a smooth structure.
However, like aluminum, constructing wing structures using wet
lay-up composite materials is labor-intensive and expensive. The
process of forming composite materials into the shape of a wing
structure with this method requires complicated machining and
tooling. Further, while composite materials are lighter than other
materials, the wet lay-up process requires the use of a large
amount of composite material. This increases the weight of the wing
structures and thereby affects the performance of the wing
structures.
[0013] Another current process to manufacture wing structures uses
hand laid-out prepreg. Like composite materials, hand laid-out
prepreg also results in a smooth structure. However, unlike
aluminum and composite materials, the construction of wing
structures from hand laid-out prepreg is not labor-intensive.
However, prepreg is very expensive. Therefore, this method greatly
increases the manufacturing costs.
[0014] Therefore, it is desirable to provide a molded composite
structure that utilizes inexpensive material, is constructed using
a nonlabor-intensive process, and provides a smooth laminar flow
surface.
[0015] Methods and structures in accordance with the invention
provide for a molded composite structure that is inexpensive, not
labor intensive to produce, and has a smooth laminar flow
surface.
III. SUMMARY OF THE INVENTION
[0016] A method consistent with the present invention provides a
method of manufacturing a molded composite structure, comprising:
preparing a material stack, wherein the material stack comprises a
core section having first and second opposing sides; preparing a
resin; preparing a mold; placing the material stack in the mold;
sealing the core section; infusing the mold and material stack with
the resin to form the structure; curing the structure; and removing
the structure from the mold.
[0017] Additional aspects of the invention are disclosed and
defined by the appended claims. It is to be understood that both
the foregoing general description and the following detailed
description are exemplary and explanatory only and are not
restrictive of the invention as claimed.
IV. BRIEF DESCRIPTION OF THE DRAWINGS
[0018] The accompanying drawings, which are incorporated in and
constitute a part of this specification, illustrate several
embodiments of the invention and together with the description,
serve to explain the principles of the invention.
[0019] In the drawings:
[0020] FIG. 1 is a block diagram illustrating the components for
manufacturing a molded composite structure consistent with an
embodiment of the invention;
[0021] FIG. 2A is a block diagram illustrating manufacturing a
molded composite structure consistent with an embodiment of the
invention, as shown in FIG. 1;
[0022] FIG. 2B is a block diagram illustrating laminate for
manufacturing a molded composite structure consistent with an
embodiment of the invention, as shown in FIG. 2A;
[0023] FIG. 2C is a block diagram illustrating core for
manufacturing a molded composite structure consistent with an
embodiment of the invention, as shown in FIG. 2A;
[0024] FIG. 3A is a block diagram illustrating a material stack for
manufacturing a molded composite structure consistent with an
embodiment of the invention, as shown in FIGS. 2A-2C;
[0025] FIG. 3B illustrates a material stack comprising laminate
materials for manufacturing a molded composite structure consistent
with an embodiment of the invention, as shown in FIG. 3A;
[0026] FIG. 3C illustrates a material stack comprising laminate and
core materials for manufacturing a molded composite structure
consistent with an embodiment of the invention, as shown in FIG.
3B;
[0027] FIG. 3D depicts a sealed core for manufacturing a molded
composite structure consistent with an embodiment of the invention,
as shown in FIG. 3C;
[0028] FIG. 3E depicts a material stack comprising a sealed core
for manufacturing a molded composite structure consistent with an
embodiment of the invention, as shown in FIG. 3D;
[0029] FIG. 4A is a block diagram illustrating structures that can
be manufactured from a mold consistent with an embodiment of the
invention, as shown in FIG. 1;
[0030] FIG. 4B is an illustrative section view of a mold 400 for
manufacturing a molded composite structure consistent with an
embodiment of the invention, as shown in FIG. 1;
[0031] FIG. 4C is a perspective view of mold elements for a wing
panel consistent with an embodiment of the invention, as shown in
FIG. 4B;
[0032] FIG. 4D is an illustrative section view of a mold for a wing
panel consistent with an embodiment of the invention, as shown in
FIG. 4C;
[0033] FIG. 4E is a plan view of a wing panel in a mold consistent
with an embodiment of the invention, as shown in FIG. 4D;
[0034] FIG. 5 is a block diagram illustrating components of resin
transfer molding for manufacturing a molded composite structure
consistent with an embodiment of the invention, as shown in FIG.
1;
[0035] FIG. 6A is a block diagram illustrating components for a
material process for manufacturing a molded composite structure in
accordance with one embodiment of the invention, as shown in FIG.
5;
[0036] FIG. 6B is a block diagram illustrating the components for a
material process for manufacturing a molded composite structure in
accordance with another embodiment of the invention, as shown in
FIG. 5;
[0037] FIG. 7 is a block diagram illustrating components for a
resin process for manufacturing a molded composite structure in
accordance with one embodiment of the invention, as shown in FIG.
5;
[0038] FIG. 8A is a block diagram illustrating a mold process for
manufacturing a molded composite structure in accordance with one
embodiment of the invention, as shown in FIG. 5;
[0039] FIG. 8B is a block diagram illustrating a mold process for
manufacturing a molded composite structure in accordance with
another embodiment of the invention, as shown in FIG. 5;
[0040] FIG. 8C is a block diagram illustrating a mold process for
manufacturing a molded composite structure in accordance with still
another embodiment of the invention, as shown in FIG. 5;
[0041] FIG. 8D is a block diagram illustrating a mold process for
manufacturing a molded composite structure in accordance with yet
another embodiment of the invention, as shown in FIG. 5;
[0042] FIG. 8E is a block diagram illustrating a mold process for
manufacturing a molded composite structure in accordance with yet
another embodiment of the invention, as shown in FIG. 5;
[0043] FIG. 9A is an illustrative section view of a mold for a wing
panel prepared and inspected consistent with an embodiment of the
invention, as shown in FIGS. 8A-8D;
[0044] FIG. 9B is an illustrative section view of a mold for a wing
panel with release agent consistent with an embodiment of the
invention, as shown in FIGS. 8A-8D;
[0045] FIG. 9C is an illustrative section view of a mold for a wing
panel with a material stack consistent with an embodiment of the
invention, as shown in FIGS. 8A-8D;
[0046] FIG. 9D is a cut-away view of a portion of a skin in a mold
for a wing panel consistent with an embodiment of the invention, as
shown in FIG. 9C;
[0047] FIG. 9E is an illustrative section view of a closed mold for
a wing panel loaded with a material stack consistent with an
embodiment of the invention, as shown in FIGS. 8A-8D;
[0048] FIG. 10A is a flow diagram illustrating an infusion process
in accordance with one embodiment of the invention, as shown in
FIG. 5;
[0049] FIG. 10B is a block diagram illustrating an infusion process
in accordance with another embodiment of the invention, as shown in
FIG. 5;
[0050] FIG. 10C is a block diagram illustrating an infusion process
in accordance with still another embodiment of the invention, as
shown in FIG. 5;
[0051] FIG. 10D is a block diagram illustrating an infusion process
in accordance with yet another embodiment of the invention, as
shown in FIG. 5;
[0052] FIG. 11A is an illustrative section view of a mold for a
wing panel in a confirmations stage consistent with an embodiment
of the invention, as shown in FIG. 10A;
[0053] FIG. 11B is an illustrative section view of a mold for a
wing panel in a vacuum stage consistent with an embodiment of the
invention, as shown in FIG. 11A;
[0054] FIG. 11C is an illustrative section view of a mold for a
wing panel in a heat stage consistent with an embodiment of the
invention, as shown in FIG. 11B;
[0055] FIG. 11D is an illustrative section view of a mold for a
wing panel in a cool down stage consistent with an embodiment of
the invention, as shown in FIG. 11C;
[0056] FIG. 11E is an illustrative section view of a mold for a
wing panel in an infusion stage consistent with an embodiment of
the invention, as shown in FIG. 11D;
[0057] FIG. 11F is an illustrative section view of a mold for a
wing panel in a hydrostatic equilibrium stage consistent with an
embodiment of the invention, as shown in FIG. 11E;
[0058] FIG. 11G is an illustrative section view of a mold for a
wing panel in a cure stage consistent with an embodiment of the
invention, as shown in FIG. 11F;
[0059] FIG. 11H is an illustrative section view of a mold for a
wing panel in a cool down stage after curing consistent with an
embodiment of the invention, as shown in FIG. 11G;
[0060] FIG. 11I is an illustrative section view of a mold for a
wing panel in a demold stage consistent with an embodiment of the
invention, as shown in FIG. 11H; and
[0061] FIG. 12 is a perspective view of a wing panel manufactured
consistent with an embodiment of the invention.
V. DESCRIPTION OF THE EMBODIMENTS
[0062] A. Introduction
[0063] Methods and structures in accordance with the present
invention will now be described with respect to an embodiment of a
molded composite structure, an aircraft wing panel. The invention
as claimed, however, is broader than wing panels and extends to
other molded composite structures, such as, for example, a full
wing structure, inserts, controls, empennages, fuselages, and
stabilizers. In addition, the invention as claimed, is broader than
aircraft structures and extends to automotive, forklift,
watercraft, and building structures.
[0064] B. Methods and Structures
[0065] FIG. 1 is a block diagram illustrating the components for
manufacturing a molded composite structure consistent with an
embodiment of the invention. As shown in FIG. 1, the integration of
a material (block 110), a mold (block 120), and a Resin Transfer
Molding ("RTM") process results in a molded composite structure 140
(for example, a wing panel).
[0066] Block 110 includes the selection and preparation of
materials to be used in manufacturing the molded composite
structure. Block 120 includes the preparation of a mold to form the
desired shape of the molded composite structure. RTM process 130
includes the placing of material 110 in mold 120, infusing material
110 with resin (not shown, but described in detail below), and the
curing of material 110 and the resin. Molded composite structure
140 represents the result of RTM process 130 using material 110 and
mold 120. For example, molded composite structure 140 may be a wing
panel. Molded composite structure 140 may also be another
structure. This implementation is merely exemplary, and other
implementations may also be used.
[0067] FIG. 2A is a block diagram illustrating manufacturing a
molded composite structure consistent with an embodiment of the
invention, as shown in FIG. 1. As shown in FIG. 2A, material 110
comprises at least one of the following exemplary materials:
laminate 210 and core 220. Material 110 may also comprise laminate
210, core 220, or some combination of laminate 210 and core 220.
Material 110 may also include other materials.
[0068] Laminate 210 includes any laminate material suitable for
forming a molded composite structure. Core 220 includes any
sandwich core materials. In one implementation, core 220 includes
sandwich core materials such as those used in spar structures and
those used as sandwich elements in a skin section between layers of
laminate. These implementations are merely exemplary, and other
implementations may also be used.
[0069] FIG. 2B is a block diagram illustrating laminate for
manufacturing a molded composite structure consistent with an
embodiment of the invention, as shown in FIG. 2A. As shown in FIG.
2B, several types of laminate 210 may be used in the manufacture of
a molded composite structure, such as a wing panel. In one
implementation, laminate 210 includes any fiber materials. For
example, laminate 210 may include carbon 230, fiberglass 240,
Kevlar 250, prepreg fiber 255, tackified fiber 257, or other types
of laminate 260, such as aramid fibers, or any combination of the
above mentioned laminates. These fibers may be used individually or
woven into a fabric or sheet. These implementations are merely
exemplary, and other implementations may also be used.
[0070] FIG. 2C is a block diagram illustrating core for
manufacturing a molded composite structure consistent with an
embodiment of the invention, as shown in FIG. 2A. As shown in FIG.
2C, core 220 includes foam 270, honeycomb 280, foam and honeycomb
290, or other 295. Foam 270 may be made from high temperature
thermo plastics that have been foamed. Honeycomb 280 may be made
from metal foils or plastic materials along with natural or
synthetic fibers formed into paper. Honeycomb 280 may also be made
from metallic materials, such as aluminum, stainless steel, or
titanium, or from non-metallic materials, such as aramid fibers or
paper. Honeycomb 280 resembles natural bee honeycomb. Foam and
honeycomb 290 includes any combination of foam 270 and honeycomb
280. Other 295 includes other types of core 220. These
implementations are merely exemplary, and other implementations may
also be used.
[0071] FIG. 3A is a block diagram illustrating a material stack for
manufacturing a molded composite structure consistent with an
embodiment of the invention, as shown in FIGS. 2A-2C. As shown in
FIG. 3A, a material stack 300 comprises layers of materials. In one
implementation, material stack 300 comprises one layer of material
302 and a second layer of material 304. However, material stack 300
may have any number of layers of material. In one implementation,
material 302 and material 304 are one of the materials described in
FIGS. 2A-2C. This implementation is merely exemplary, and other
implementations may also be used.
[0072] In one implementation, material 302 may be applied directly
on top of material 304 to form material stack 300 using any of a
number of well-known methods. In another implementation, an
adhesive layer (not shown) is applied between material 302 and
material 304. Material 302 and material 304 may be applied with a
specific orientation to increase the strength of material stack
300. These implementations are merely exemplary, and other
implementations may also used.
[0073] In one implementation, material 110 (described in FIG. 1)
includes material stacks, such as material stack 300. As described
in FIG. 1, material 110 is placed in mold 120, where it undergoes
RTM process 130 to form molded composite structure 140.
[0074] FIG. 3B illustrates a material stack comprising laminate
materials for manufacturing a molded composite structure consistent
with an embodiment of the invention, as shown in FIG. 3A. As shown
in FIG. 3B, a material stack 306 comprises layers of materials. In
one implementation, material stack 306 comprises one layer of
laminate 308 and a second layer of laminate 309. However, material
stack 306 may have any number of layers of material. In one
implementation, laminate 308 and laminate 309 are one of the
materials described in FIG. 2B. This implementation is merely
exemplary, and other implementations and materials may also be
used.
[0075] FIG. 3C illustrates a material stack comprising laminate and
core materials for manufacturing a molded composite structure
consistent with an embodiment of the invention, as shown in FIG.
3B. As shown in FIG. 3C, a material stack 310 comprises a layer of
core 314 surrounded by two layers of laminate 312 and 313. In this
implementation, the use of core 312 increases the strength of
material stack 310. In another implementation, the number of layers
of laminate 312 and 313 on one side of core 314 differs from the
number of layers as on the other side of core 314 (i.e., one side
has more or less layers than the other side). In one
implementation, core 314 and laminates 312 and 313 include those
materials described above in FIGS. 2B-2C. These implementations are
merely exemplary, and other implementations and materials may also
be used.
[0076] FIG. 3D depicts a sealed core for manufacturing a molded
composite structure consistent with an embodiment of the invention,
as shown in FIG. 3C. As shown in FIG. 3D, in one implementation,
core 328 may be sealed by thermoplastic barriers 322 and 323 to
create sealed core 320. In this implementation, adhesives 324 and
325 and support layers 326 and 327 are also included between the
thermoplastic barriers 322 and 323. This implementation is merely
exemplary, and other implementations and materials may also be
used.
[0077] In one implementation, core 328 is formed of one of the
materials described above in FIG. 2C. During RTM process 130
(described in FIG. 1), resin may intrude into core 328. In one
implementation, as shown in FIG. 3D, thermoplastic barriers 322 and
323 are used to seal core 328 to act as barriers and prevent
intrusion of resin into core 328. In one implementation,
thermoplastic barriers 320 and 321 are constructed of bondable
Teflon, Mylar, or Ultem. For example, Melenx 454, a type of
bondable Mylar, may be used. Further, thermoplastic barriers 322
and 323 may be formed of materials resistant to processing
pressures and temperatures so as to maintain seal of core 328, such
as, for example during RTM process 130 as described in FIG. 1.
These implementations are merely exemplary, and other
implementations and materials may also be used.
[0078] As shown in FIG. 3D, in one implementation, adhesives 324
and 325 are adhesives used to bond thermoplastic barriers 322 and
323 to core 328. In one implementation, adhesives 324 and 325 are
film adhesives comprising epoxy materials, such as epoxy #NB185
manufactured by Newport. Adhesives 324 and 325 may be applied
directly to core 328, directly to thermoplastic barriers 322 and
323, or on both core 328 and thermoplastic barriers 322 and 323.
These implementations are merely exemplary, and other
implementations may also be used.
[0079] As further shown in FIG. 3D, in one implementation, support
layers 326 and 327 are placed between core 328 and adhesives 324
and 325 to provide added strength to material stack 316. In one
implementation, support layers 326 and 327 are manufactured from
glass or scrim. In another implementation, support layers 326 and
327 are made of fiberglass, woven cloth, chopped matte, plastic
fibers, and/or organic fibers. In these implementations, adhesives
324 and 325 will bond core 328, support layers 326 and 327 and
thermoplastic barriers 322 and 323 together, respectively. These
implementations are merely exemplary, and other implementations and
materials may also be used.
[0080] In addition, during RTM process 130 (described in FIG. 1),
in one implementation, a vacuum may be drawn on material stack 320.
In this implementation, support layers 326 and 327 allow for a
vacuum path (not shown) for evacuation of core 328. This
implementation is merely exemplary, and other implementations may
also be used.
[0081] After application of these layers, core 328 is cured to seal
thermoplastic barriers 322 and 323 to core 328. In one
implementation, core 328 is cured at the same time that the molded
composite structure (i.e. wing panel) is cured. In another
implementation, core 328 may be cured prior to its use in the
manufacturing process. These implementations will be described in
more detail below. In addition, FIG. 3D depicts a core that has
been sealed on both sides. However, core 328 may be sealed on only
one side. These implementations are merely exemplary, and other
implementations may also be used.
[0082] FIG. 3E depicts a material stack comprising a sealed core
for manufacturing a molded composite structure consistent with an
embodiment of the invention, as shown in FIG. 3D. As shown in FIG.
3E, in one implementation, material stack 330 comprises core 339,
sealed by thermoplastic barriers 333 and 334 and including support
layers 337 and 338 and adhesives 335 and 336 to form sealed core
340. Laminates 318 and 319 surround sealed core 340. This
implementation is merely exemplary, and other implementations and
materials may also be used.
[0083] In one implementation, sealed core 340 is sealed as shown in
FIG. 3E (and described in FIG. 3D). Laminates 331 and 332 form the
outer layers of material stack 330. In this implementation, the
same number of layers of laminate 331 and 332 are used on either
side of sealed core 340 (e.g., one layer on each side). However, in
other implementations, the number of layers of laminate 331 and 332
on one side of sealed core 340 need not be the same as the number
of layers as on the other side of sealed core 340 (i.e., one side
may have more or less layers than the other side). These
implementations are merely exemplary, and other implementations may
also be used.
[0084] In another implementation, the laminate layers are applied
with a preferred fiber orientation on either side of the core,
resulting in added strength. This allows for the use of less layers
of laminate. In turn, this decreases the weight of the material
stack. This implementation is merely exemplary, and other
implementations may also be used.
[0085] FIG. 4A is a block diagram illustrating structures that can
be manufactured from a mold consistent with an embodiment of the
invention, as shown in FIG. 1. As shown in FIG. 4A, mold 120 may be
designed to construct a number of structures, including a panel
402, a wing 404, and other 406.
[0086] Panel 402 includes panels for wings and other structures.
Wing 404 includes a semi-span wing for an aircraft and a full-span
wing for an aircraft. A semi-span wings is a wing for one side of
the aircraft, for example, (i.e. a left or right wing). Therefore,
two semi-span wings could be constructed. A full-span wing is a
one-piece wing for both sides of the aircraft (i.e. a one piece
wing comprising both the left and right wing). Other 406 includes
any other structures, whether for an aircraft (such as fuselages,
ailerons, or flaps) or for other than aircraft (such as automotive,
forklift, watercraft, and building structures). In one
implementation, the shape of mold 120 determines both the external
and internal shape of a molded composite structure 140 such as
molded composite structure 140 in FIG. 1. This implementation is
merely exemplary, and other implementations may also be used.
[0087] FIG. 4B is an illustrative section view of a mold 400 for
manufacturing a molded composite structure consistent with an
embodiment of the invention, as shown in FIG. 1. As shown in FIG.
4B, in one implementation, mold 400 comprises a top outer shell 408
and a bottom outer shell 410. Further, mold 400 comprises O-ring
seals 412 and 414 and ports 416 and 418. Mold 400 may also include
other elements.
[0088] Top outer shell 408 and bottom outer shell 410 may determine
the external shape of the structure. For example, the interior
shape of top outer shell 408 and bottom outer shell 410 can be
designed to form the shape of any of the structures depicted in
FIG. 4A. In one implementation, top outer shell 408 and bottom
outer shell 410 are clamshell mold halves. In this implementation,
mold 400 may also contain internal mold elements (not shown, but
described in more detail in FIG. 4C). These internal mold elements
may form part of the interior shape of the structure. For example,
material 110 (not shown, but described below) may be placed around
internal mold elements (not shown, but described below) and within
top outer shell 408 and bottom outer shell 410 to form the
structure. These implementations are merely exemplary, and other
implementations may also be used.
[0089] Ports 416 and 418 are openings extending from the exterior
of mold 400 to the interior of mold 400. In one implementation,
ports 416 and 418 allow for the introduction of a material, such as
a resin, into mold 400. In another implementation, at least one of
ports 416 and 418 is attached to a vacuum (not shown) for creating
a vacuum inside mold 400. Ports 416 and 418 may also be used for
other functions. For example, ports 416 and 418 may also be capable
of being sealed. These implementations are merely exemplary, and
other implementations may also be used.
[0090] O-ring seals 412 and 414 allow mold 400 to be sealed upon
closure. By being precisely dimensioned, O-ring seals 412 and 414
can prevent significant leaks. In one implementation, O-ring seals
412 and 414 are rubber gaskets. However, other materials could be
used for O-ring seals 412 and 414. In addition, multiple O-rings, a
single O-ring, concentric O-rings, or other sealing methods may be
used.
[0091] FIG. 4C is a perspective view of mold elements for a wing
panel consistent with an embodiment of the invention, as shown in
FIG. 4B. As shown in FIG. 4C, in one implementation, mold elements
470 may be used to form a wing panel. Mold elements 470 comprise a
top outer shell 420 and a bottom outer shell 422, which form outer
mold line ("OML") tooling. Mold elements 470 also include a leading
edge mandrel 442, an internal bladder section 438, and a trailing
edge section 434, which form internal mold line ("IML") tooling.
Mold elements 470 further include a noseblock section 424, a
forward cabin area spar forming tooling 437, a middle insert
section 441, an aft cabin area spar forming tooling 439, and an end
plate 490.
[0092] As described above, OML tooling comprises top outer shell
420 and bottom outer shell 422. In one implementation, top outer
shell 420 and bottom outer shell 422 form the exterior shape of the
wing panel, as described in FIG. 4B.
[0093] As described above, the IML tooling comprises leading edge
mandrel 442, internal section 438, and trailing edge section 434.
In one implementation, leading edge mandrel 442, internal section
438, and trailing edge section 434 form the internal shape of the
wing panel as described in FIG. 4B.
[0094] Leading edge mandrel 442 forms the interior shape of the
leading edge of the wing panel. In one implementation, leading edge
mandrel 442 may be constructed of metallic materials such as
aluminum, nickel alloys, or Invar, or it may be constructed of
non-metallic materials. In this implementation, leading edge
mandrel 442 is solid, however, leading edge mandrel 442 may be
segmented (as in trailing edge section 434) or may be constructed
of bladders (as in internal section 438). In one implementation,
following cure of the wing panel (as described below), leading edge
mandrel 442 is removed from the structure. These implementations
are merely exemplary, and other implementations may also be
used.
[0095] Internal section 438 forms the internal section of the wing
panel. As shown in FIG. 4C, in one implementation, internal section
438 may also comprise an outboard bladder 464, a mid bladder 462,
and an inboard bladder 458. However, internal section 438 may
comprise any number of bladders or structures. In this
implementation, internal section 438 comprises bladders, however,
internal section 438 may be solid (as in leading edge mandrel 442)
or segmented (as in trailing edge section 434). In one
implementation, mid bladder 462 is used to form a fuel tank (not
shown). In another implementation, following cure of the wing
panel, bladders 458, 462, and 464 are removed from the structure.
In yet another implementation, any of bladders 458, 462, and 464
may be left in the structure following cure, and used as a fuel
tank. These implementations are merely exemplary, and other
implementations may also be used.
[0096] Outboard bladder 464 forms an outboard bay interior of the
wing panel. Mid bladder 462 forms a mid bay interior of the wing
panel. Inboard bladder 458 forms an inboard interior of the wing
panel. In one implementation, bladders 458, 462, and 464 are
elastomeric tooling. The use of elastomeric tooling allows for the
pressure within bladders 458, 462, and 464 to be altered during
processing. In one implementation, bladders 458, 462, and 464 are
constructed from silicone or polyethelene. These implementations
are merely exemplary, and other implementations and other materials
may also be used.
[0097] Trailing edge section 434 forms the interior shape of the
trailing edge of the wing panel. As shown in FIG. 4C, in one
implementation, trailing edge section 434 comprises insert sets
466, 468, 472, 474, 476, and 478. In this implementation, following
cure of the wing panel, insert sets 466, 468, 472, 474, 476, and
478 are removable from the structure. In this implementation,
insert sets 466, 468, 472, 474, 476, and 478 are multiple
interlocking hard tool elements, however, trailing edge section 434
may also comprise bladders (as in internal section 438). These
implementations are merely exemplary, and other implementations may
also be used.
[0098] In one implementation, hinge support ribs 460 (not shown)
are also included in the spaces between insert sets 466, 468, 472,
474, 476, and 478. Hinge support ribs 460 may provide support for
flaps and ailerons on the wing panel.
[0099] As described in FIG. 1, material 110 may be placed in mold
120 to form molded composite structure 140. Thus, with reference to
FIG. 4C, material 110 may be placed in and around mold elements 470
to form a wing panel. Nose block 424 is used to prevent material
110 from being pinched when top outer shell 420 and a bottom outer
shell 422 are closed. In one implementation, noseblock 424
comprises outboard insert 452 and inboard insert 454. In this
implementation, insert 452 and insert 454 are designed to mimic the
shape of leading edge mandrel 442. Insert 452 may be placed against
the long straight portion of leading edge mandrel 442. Insert 454
may be placed against the angled portion of leading edge mandrel
442. In one implementation, inserts 452 and 454 are constructed of
aluminum, nickel alloys, or Invar, or they may be constructed of
non-metallic materials. These implementations are merely exemplary,
and other implementations may also be used.
[0100] In another implementation a front spar 440 (not shown, but
shown in FIG. 4D) and a rear spar 436 (not shown, but shown in FIG.
4D) provide support for the wing panel and provide for the
connection of the wing panel to a fuselage of an aircraft. Front
spar 440 is located between leading edge mandrel 442 and internal
section 438. Rear spar 436 is located between internal section 438
and trailing edge section 434. Front spar 440 and rear spar 436 may
be box beam spars (as shown in FIG. 4D), I-beam spars, C-channel
spars, or any other type of spar. These implementations are merely
exemplary, and other implementations and other materials may also
be used.
[0101] In one implementation, spars 436 and 440 are constructed of
carbon-fiber. In another implementation, spars 436 and 440 include
core materials, such as foam core or honeycomb core. This core may
be sealed or unsealed. In another implementation, spars 436 and 440
are cured prior to being used in mold 470. However, spars 436 and
440 may be cured with the part (i.e. wing panel). Spars 436 and 440
may also include a bonding agent on the surface of spars 436 and
440. These implementations are merely exemplary, and other
implementations and other materials may also be used.
[0102] In one implementation, spars 436 and 440 (not shown, but
shown in FIG. 4D) extend beyond the end of the wing panel. This
allows spars 436 and 440 to be inserted in a fuselage to connect
the wing panel to the rest of an aircraft. In this implementation,
mold 422 extends beyond the length of the wing panel. As shown in
FIG. 4C, section 423 of top outer shell 420 and section 425 of
bottom outer shell 422 extend beyond the length of the wing
panel.
[0103] Forward cabin area spar forming tooling 437 and aft cabin
area spar forming tooling 439 are located in mold section 423.
Forward cabin area spar forming tooling 437 and aft cabin area spar
forming tooling 439 are used to support the portion of spars 436
and 440 (not shown, but shown in FIG. 4D) extending beyond the wing
panel during the curing process. Further, forward cabin area spar
forming tooling 437 and aft cabin area spar forming tooling 439
ensure that spars 436 and 440 do not shift during the cure process.
This implementation is merely exemplary, and other implementations
may also be used.
[0104] In one implementation shown in FIG. 4C, forward cabin area
spar forming tooling 437 comprises extraction block 497, wedge
block 498, lower insert 496, and upper insert 494. In addition,
forward cabin area spar forming tooling 437 may comprise other
elements. Upper insert 494 and lower insert 496 surround front spar
440 (not shown, but shown in FIG. 4D). Wedge block 498 is inserted
to press lower insert 496 against forward spar 440. Extraction
block 497 is used to further press wedge block 498 against lower
insert 496. These implementations are merely exemplary, and other
implementations may also be used.
[0105] Aft cabin area spar forming tooling 439 comprises extraction
block 482, wedge block 484, lower insert 488, and upper insert 486.
In addition, aft cabin area spar forming tooling 439 may comprise
other elements. Upper insert 486 and lower insert 488 surround rear
spar 436 (not shown, but shown in FIG. 4D). Wedge block 484 is
inserted to press upper insert 486 against rear spar 436.
Extraction block 482 is used to further press wedge block 484
against upper insert 486. These implementations are merely
exemplary, and other implementations may also be used.
[0106] Middle insert section 441 comprises mid bay top plate 493,
mid bay bottom plate 492, and bottom insert 470. In addition,
middle insert section 441 may comprise other elements. Middle
insert section 441 holds spars 436 and 440 (not shown, but shown in
FIG. 4D) in position during curing. Mid bay top plate 493 and mid
bay bottom plate 492 press lower insert 488 further against rear
spar 436 and upper insert 494 further against forward spar 440.
Bottom insert 470 further supports spars 436 and 440. These
implementations are merely exemplary, and other implementations may
also be used.
[0107] In one implementation, the components of forward cabin area
spar forming tooling 437, aft cabin area spar forming tooling 439,
and middle insert section 441 are constructed of aluminum, nickel
alloys, or Invar, or they may be constructed of non-metallic
materials. These implementations are merely exemplary, and other
implementations and other materials may also be used.
[0108] End plate 490 may be used to complete closure of mold
elements 470. In one implementation, end plate 490 seals mold
elements 470 such that a vacuum may be created inside of mold
elements 470. In one implementation, end plate 490 is constructed
of aluminum, nickel alloys, or Invar, or it may be constructed of
non-metallic materials. These implementations are merely exemplary,
and other implementations and other materials may also be used.
[0109] As described above, in one implementation, mold elements 470
may undergo curing along with the wing panel. In this regard, the
coefficient of expansion of mold elements 470 may be different from
each other or of the wing panel. Thus, during curing, mold elements
470 and the wing panel may expand more or less than one another.
Therefore, in one implementation, each of the elements of mold
elements 470 may be designed to prevent expansion or contraction of
the elements from damaging the wing panel or mold elements during
curing and subsequent cool down. This implementation is merely
exemplary, and other implementations may also be used.
[0110] FIG. 4D is an illustrative section view of a mold for a wing
panel consistent with an embodiment of the invention, as shown in
FIG. 4C. As shown in FIG. 4D, top outer shell 420 and bottom outer
shell 422 form the OML tooling as described in FIG. 4C. FIG. 4D
also shows the IML tooling described in FIG. 4C, including leading
edge mandrel 442, internal bladder section 438, and trailing edge
section 434. The IML tooling is located inside the OML tooling. In
addition, FIG. 4D shows front spar 440 positioned between leading
edge mandrel 442 and internal bladder section 438 and rear spar 436
positioned between internal bladder section 438 and trailing edge
section 434. In one implementation, material (not shown, but
described herein) is applied around the exterior of the IML tooling
and the spars to form the wing panel. This implementation is merely
exemplary, and other implementations may also be used.
[0111] FIG. 4D also shows noseblock section 424 as described in
FIG. 4C. As discussed above, noseblock section 424 is located next
to leading edge mandrel 442. In one implementation, mold 470 also
includes an integral tooling port 431. Port 431 runs from the
exterior of mold 470 to internal bladder section 438. In one
implementation, a pressure controlling device (not shown) is
connected to port 431 to alter the pressure within internal bladder
section 438. As described in FIG. 4C, internal section 438 may
comprise outboard bladder 464, mid bladder 462, and inboard bladder
458.
[0112] In one implementation, all three bladders may be connected
collectively to port 431. In another implementation, all three
bladders may be connected separately to port 431. Alternatively, in
still another implementation, inboard bladder 458 and mid bladder
462 are connected in series to one another. In this implementation,
only outboard bladder 464 and inboard bladder 458 would be
connected to port 431. In yet another implementation, port 431
would comprise multiple ports. In this implementation, each bladder
may have a corresponding port. Thus, any combination of bladder
connections and ports may be used to allow for control of the
pressure within the bladders. These implementations are merely
exemplary, and other implementations may also be used.
[0113] As shown in FIG. 4D, mold 470 also comprises ports 444 and
446. Ports 444 and 446 are similar to ports 416 and 418, as
described in FIG. 4B. In one implementation, ports 444 and 446 are
used to introduce material such as resin into mold 470. In another
implementation, ports 444 and 446 allow for the creation of a
vacuum inside mold 470. Mold 470 also contains O-ring seals 426 and
428, which are similar to O-ring seals 418 and 418 described in
FIG. 4B. In one implementation, O-ring seals 426 and 428 allow the
mold to be sealed. As described above, O-ring seals 426 and 428 may
constitute concentric O-rings or other sealing methods. This
implementation is merely exemplary, and other implementations may
also be used.
[0114] In another implementation, the geometry of the tooling is
designed to prevent fiber washout during resin infusion. As
described above, material stacks are applied around the IML tooling
and the spars. In one implementation, the material stacks have a
specific fiber orientation. As described above, this fiber
orientation provides for greater material strength. During the
infusion of resin into the mold, the force of the resin against the
fiber may cause the fibers to shift and thus alter the orientation.
This may decrease the strength of the material. However, the
tooling elements may be designed to prevent this fiber washout
through precise geometric controls. For example, by designing the
tool to precisely align with the internal mold elements and the
fibers, the shifting of the fibers from resin infusion is reduced.
This implementation is merely exemplary, and other implementations
may also be used.
[0115] FIG. 4E is a plan view of a wing panel in a mold consistent
with an embodiment of the invention, as shown in FIG. 4D. As shown
in FIG. 4E, mold 470 is used to form a wing panel 449. Wing panel
449 comprises skin 499, co-cured spars 440 and 436, a co-cured fuel
tank 450, and co-cured hinge support ribs 448. In one
implementation, skin 499 is a cured material stack as described in
FIGS. 3A-3E. In one implementation, co-cured spars 440 and 436 may
be loaded in the IML tooling as described in FIGS. 4C and 4D. In
one implementation, co-cured fuel tank 450 may be formed in wing
panel 449 as described in FIG. 4C. Piping (not shown) connects fuel
tank 450 to the engine (not shown) of an aircraft to provide fuel
to the engine. In one implementation, co-cured hinge supports 448
may be formed in the trailing edge of the wing, as described in
FIG. 4C. In one implementation, the OML tooling shown in FIG. 4D is
244 inches long and 70 inches wide at its widest point. In another
implementation, these elements are either co-cured, co-bonded,
and/or cured separately from one another. These implementations are
merely exemplary, and other implementations may also be used.
[0116] FIG. 5 is a block diagram illustrating components of resin
transfer molding for manufacturing a molded composite structure
consistent with an embodiment of the invention, as shown in FIG. 1.
As shown in FIG. 5, RTM process 130 comprises a material process
510, a resin process 520, a mold process 530, and an infusion
process 540. Material process 510 includes the preparation of
material stacks as described in FIGS. 2A-3D. Material process 510
is further described in FIGS. 6A-6B. Resin process 520 includes the
preparation of a resin to be infused into a material stack. Resin
process 520 is further described in FIG. 7. Mold process 530
includes the preparation of a mold as described in FIGS. 4A-4E.
Mold process 530 also includes the placement of a material stack in
the mold. Mold process 530 is further described in FIGS. 8A-9E.
Infusion process 540 includes the infusion of resin into the mold
and the curing of the resin and material to form a structure.
Infusion process 540 is further described in FIGS. 10A-11I. This
implementation is merely exemplary, and other implementations may
also be used.
[0117] FIG. 6A is a block diagram illustrating components for a
material process for manufacturing a molded composite structure in
accordance with one embodiment of the invention, as shown in FIG.
5. As shown in FIG. 6A, material process 510 includes
identification of laminate 610. In one implementation,
identification of laminate 610 includes the selection of any of the
laminates described in FIG. 2B. Next, identification of laminate
610 is followed by measure laminate 620. In one implementation,
measure laminate 620 includes the determination of the amount of
laminate to be used to make up each layer of laminate. This can be
determined based on the total desired weight of the laminate layer
or the number of plies of laminate to be used. Measure laminate 620
is followed by assemble laminate layers 630. In one implementation,
assemble laminate layers 630 includes the assembly of at least two
laminate layers.
[0118] Material process 510 also includes select core 640. Select
core 640 includes the selection of core material to be used. In one
implementation, core is any of those materials described in FIG.
2C. However, other materials may be used.
[0119] Select core 640 is followed by machine core 650. Machine
core 650 includes trimming, cutting, shaping, and preparing the
core material into a desired shape for placement in a material
stack.
[0120] Machine core 650 is followed by seal core 660. In one
implementation, seal core 660 includes the sealing of the core on
both sides. In another implementation, as described in FIG. 3D, a
support layer is placed on both sides of the core, an adhesive
layer is placed on both sides of the core, and a thermoplastic
barrier layer is placed on both sides of the core to seal the core.
As described above, the adhesive may be located on the
thermoplastic barriers, the core, or both. In another
implementation, the core would be cured to seal the thermoplastic
barrier layers around the core. This cure process may be performed
prior to formation of the material stack, following material stack
formation but prior to placement of the material stack in a mold,
or during cure of the wing panel. These implementations are merely
exemplary, and other implementations may also be used.
[0121] Seal core 660 is followed by form material stack 670. In one
implementation, form material stack 670 includes placing laminate
layers on either side of the sealed core to form a material stack.
In this implementation, material process 510 creates a material
stack as described in FIG. 3D. As further described in FIG. 3D, in
another implementation, either the same or a differing number of
laminate layers can be placed on either side of the core.
[0122] In another implementation, laminate layers are applied with
the same fiber orientation on either side of the core. Proper
alignment of the fibers can result in added strength. This allows
for the use of less layers of laminate. In turn, this decreases the
weight of the material stack. This implementation is merely
exemplary, and other implementations may also be used.
[0123] FIG. 6B is a block diagram illustrating the components for a
material process for manufacturing a molded composite structure in
accordance with another embodiment of the invention, as shown in
FIG. 5. As shown in FIG. 6B, similar to FIG. 6A, prepare material
510 includes identification of laminate 610, measure laminate 620,
and assemble laminate layers 630.
[0124] Material process 510 also includes prepare and seal first
side of core 680 occurs. Prepare and seal first side of core 680
includes the selection of the core to be used. In one
implementation, core can be any of the materials described in FIG.
2C. Prepare and seal first side of core 680 also includes the
sealing of one side of the core. In one implementation, a support
layer is placed on one side of the core and an adhesive layer is
placed on top of that support layer. In this implementation, a
thermoplastic barrier layer is then placed on top of the adhesive
layer. The thermoplastic barrier layer may then be cured to
complete the seal. This implementation is merely exemplary, and
other implementations may also be used.
[0125] Prepare and seal first side of core 680 is followed by
machine core 650. As in FIG. 6A, machine core 650 includes trimming
of the core into a desired shape for placement in the material
stack. However, in this implementation, the core is trimmed after
it has been sealed on one side. Sealing one side of the core
stabilizes the core. By stabilizing the core prior to machining
more detailed cutting and machining processes may be performed on
the core. For example, planning and certain router operations may
now be performed on the partially sealed core. This implementation
is merely exemplary, and other implementations may also be
used.
[0126] Machine core 650 is followed by prepare and seal second side
of core 690. In one implementation, prepare and seal second side of
core 690 includes sealing the remaining side of the core in the
same manner that the first side was sealed. This implementation is
merely exemplary, and other implementations may also be used.
[0127] Prepare and seal second side of core 690 is followed by form
material stack 670. As in FIG. 6A, in one implementation, form
material stack 670 includes the placement of the laminate layers on
either side of the sealed core to form the material stack. In this
implementation, there are even laminate layers. As described above,
material stack 670 may also include material stacks with uneven
laminate layers.
[0128] FIGS. 6A-6B have described material process 510, as shown in
FIG. 5. As shown in FIG. 5, material process 510 is followed by
resin process 520. Resin process 520 is described in FIG. 7.
[0129] FIG. 7 is a block diagram illustrating components for a
resin process for manufacturing a molded composite structure in
accordance with one embodiment of the invention, as shown in FIG.
5. As shown in FIG. 7, resin process 520 includes weigh 710, mix
720, heat 730, and de-gas 740. Resin process 520 then results in
mixed resin 750.
[0130] The preparation of the resin in resin process 520 begins
with weigh 710. Weigh 710 includes the selection of the various
components to make up the resin. In one implementation, materials
are selected based on their ability to affect certain properties of
the resin, such as viscosity, strength, toughness, and gel cycle
time. In this implementation, following the selection of the
materials, a determination of the amount of each material to
include is made. This determination is made by weighing the
material. This implementation is merely exemplary, and other
implementations may also be used.
[0131] Weigh 710 is followed by mix 720. Mix 720 includes mixing of
the materials chosen in the weigh 710. Following mix 720 is heat
730. Heat 730 includes applying heat to the mixture to raise the
temperature of the mixture. Following heat 730 is de-gas 740.
De-gas 740 includes de-gassing of the resin after raising the
temperature of the resin to remove dissolved gasses or solvents
from mixed resin. In one implementation, de-gassing is achieved by
placing the material in a low-pressure environment. As the pressure
decreases, trapped gasses will boil to the surface of the material.
The pressure at which the materials are de-gassed should be at
least as low as the pressure used in the resin transfer molding
process. Otherwise, further de-gassing may occur during the resin
transfer molding process causing voids in the structure. These
implementations are merely exemplary, and other implementations may
also be used.
[0132] In another implementation, prior to mix 720, the materials
are separately heated and de-gassed. In yet another implementation,
no heat is applied at any point. These implementations are merely
exemplary, and other implementations may also be used.
[0133] In one implementation, weigh 710, mix 720, heat 730, and
de-gas 740 may be prepared specifically for the structure. For
example, resins V42, V43, SC32 from Applied Polymeric may be used.
In another implementation, instead of the use of a custom prepared
resin, an off-the-shelf resin may also be used. These
implementations are merely exemplary, and other implementations may
also be used.
[0134] FIG. 7 has described resin process 520, as shown in FIG. 5.
As shown in FIG. 5, following resin process 520 is mold process
530. Mold process 530 is described in FIGS. 8A-9E.
[0135] FIG. 8A is a block diagram illustrating a mold process for
manufacturing a molded composite structure in accordance with one
embodiment of the invention, as shown in FIG. 5. As shown in FIG.
8A, mold process 530 begins with preparation and inspection 810.
Preparation and inspection 810 includes the preparation of the mold
elements as described in FIGS. 4A-4E. It also includes the
inspection and fit checking of the mold elements to determine that
the mold will form the desired shape. Preparation and inspection
810 is further described in FIG. 9A.
[0136] Preparation and inspection 810 is followed by release agent
820. Release agent 820 includes application of a release agent to
the mold elements. This prevents the mold elements from adhering to
the formed structure, such as a wing panel, and from adhering to
other mold elements. Release agent 820 is further described in FIG.
9B.
[0137] Release agent 820 is followed by load material stack 830.
Load material stack 830 includes the placing of material in the
mold. In one implementation, material is created using material
process 510 as described in FIGS. 6A-6B. In one implementation,
material includes a material stack comprising a sealed core, a
material stack comprising laminate layers, a material stack
comprising a sealed core and laminate layers, a material stack
comprising an unsealed core and laminate layers, or a material
stack comprising a partially sealed core and laminate layers. As
described above, the number of laminate layers may be the same or
different on either side of the core. In addition, as described
above, the laminate layers may be applied so that the orientation
of the fibers provides for greater strength. These implementations
are merely exemplary, and other implementations may also be
used.
[0138] A material stack comprising only laminate layers may be used
in leading edge sections, integrating rib sections, and integrating
spar sections where core material may not be required. A material
stack comprising a sealed core and laminate layers may be used in
integrating rib sections and integrating spar skin sections where
core may be needed. A material stack comprising an unsealed core
sandwiched by laminate layers may also be used. The core will then
be sealed during the cure of the part. Load material stack 830 is
further described in FIGS. 9C-9D.
[0139] Load material stack 830 is followed by close mold 840. Close
mold 840 includes the closing of the mold around the material.
Close mold 840 is further described in FIG. 9E.
[0140] FIG. 8B is a block diagram illustrating a mold process for
manufacturing a molded composite structure in accordance with
another embodiment of the invention, as shown in FIG. 5. As shown
in FIG. 8B, this embodiment is identical to that described in FIG.
8A, except that load material stack 830 from FIG. 8A has been
replaced with load material stack with core and laminate 860 in
FIG. 8B. Load material stack with core and laminate 860 includes
the loading of a material stack comprising a core sandwiched by two
laminate layers. In one implementation, the core is either sealed
or unsealed. In another implementation, the laminate layers may be
identical in weight or ply count, but they are not required to be
so. Load material stack with core and laminate 860 is further
described in FIGS. 9C-9D. These implementations are merely
exemplary, and other implementations may also be used.
[0141] FIG. 8C is a block diagram illustrating a mold process for
manufacturing a molded composite structure in accordance with still
another embodiment of the invention, as shown in FIG. 5. As shown
in FIG. 8C, this embodiment is also identical to that described in
FIG. 8A, except load material stack 830 has been replaced with load
material stack with sealed core 870. Load material stack with
sealed core 870 includes placing of a sealed core material stack in
the mold. In one implementation, the sealed core also includes one
or more laminate layers as described in FIGS. 6A-6B. However, in
other implementations, the core need not be sandwiched by laminate
layers. Load material stack with sealed core 870 is further
described in FIGS. 9C-9D. These implementations are merely
exemplary, and other implementations may also be used.
[0142] FIG. 8D is a block diagram illustrating a mold process for
manufacturing a molded composite structure in accordance with yet
another embodiment of the invention, as shown in FIG. 5. As shown
in FIG. 8D, this embodiment is also identical to that described in
FIG. 8A except load material stack 830 has been replaced with load
material stack with laminate and sealed core 880. Load material
stack with laminate and sealed core 880 includes the placing a
material stack containing a sealed core sandwiched between laminate
layers in the mold. In one implementation, the material stack is
that described in FIG. 3D. However, other material stacks may be
used. Load material stack with laminate and sealed core 880 is
further described in FIGS. 9C-9D. These implementations are merely
exemplary, and other implementations may also be used.
[0143] FIG. 8E is a block diagram illustrating a mold process for
manufacturing a molded composite structure in accordance with yet
another embodiment of the invention, as shown in FIG. 5. As shown
in FIG. 8E, this embodiment is also identical to that described in
FIG. 8A except load material stack 830 has been replaced with load
material stack with laminate and unsealed core 890. Load material
stack with laminate and unsealed core 890 includes the placing a
material stack containing an unsealed core sandwiched between
laminate layers in the mold. In one implementation, the core is
sealed during the cure of the wing panel. Load material stack with
laminate and sealed core 880 is further described in FIGS. 9C-9D.
These implementations are merely exemplary, and other
implementations may also be used.
[0144] As shown in FIG. 8A, preparation and inspection 810 is the
first step in mold process 530. Preparation and inspection 810 is
described in FIG. 9A.
[0145] FIG. 9A is an illustrative section view of a mold for a wing
panel prepared and inspected consistent with an embodiment of the
invention, as shown in FIGS. 8A-8D. As shown in FIG. 9A, mold 900
includes a top clam shell half 901 and a bottom clamshell half 902
to form the OML tooling element as described in FIGS. 4A-4E. Mold
900 also includes a leading edge mandrel 910, an internal bladder
section 914, and a trailing edge section 918, which form the IML
tooling element as described in FIGS. 4A-4E. In addition, mold 900
includes a front spar 912 and a rear spar 916. As further described
in FIGS. 4A-4E, mold 900 also includes a noseblock section 904,
O-rings 906 and 920, an internal bladder port 915, and two ports
919 and 917. In one implementation of preparation and inspection
810, these elements are all inspected and checked to ensure that
the mold will form the desired shape and that it will close
properly. This implementation is merely exemplary, and other
implementations may also be used.
[0146] FIG. 9A describes preparation and inspection 810, as shown
in FIG. 8A. As shown in FIG. 8A, following preparation and
inspection 810 is release agent 820. Release agent 820 is described
in FIG. 9B.
[0147] FIG. 9B is an illustrative section view of a mold for a wing
panel with release agent consistent with an embodiment of the
invention, as shown in FIGS. 8A-8D. As shown in FIG. 9B, release
agent 921 is applied to the surfaces of the IML tooling element and
the interior of the OML tooling element. For example, in one
implementation, the interior of top clamshell half 901 and bottom
clamshell half 902 are treated with release agent 921 so that the
wing panel may be removed from mold 900 after curing. Also, in this
implementation, the exterior surfaces of leading edge mandrel 910,
internal bladder section 914, and trailing edge section 918 are
treated with release agent 921, so these elements may be removed
from the structure after curing. In addition, in this
implementation, release agent 921 is applied top the mating
surfaces of the OML tooling elements. This implementation is merely
exemplary, and other implementations may also be used.
[0148] Release agent 921 is a liquid or dry material that
facilitates removal of the part from the mold element surfaces
without damage to the part surface. In one implementation, release
agent 921 is a bond inhibiting agent. For example, Water Shield
from Zyvax may be used. This implementation is merely exemplary,
and other implementations may also be used.
[0149] FIG. 9B has described release agent 820, as shown in FIG.
8A. As shown in FIG. 8A, following release agent 820 is load
material stack 830. Load material stack 830 is described in FIGS.
9C-9D.
[0150] FIG. 9C is an illustrative section view of a mold for a wing
panel with a material stack consistent with an embodiment of the
invention, as shown in FIGS. 8A-8D. As shown in FIG. 9C, mold 900
comprises a skin 922 constructed of a material stack (as described
in FIGS. 6A-6B). The material stack may be loaded around the
outside of leading edge mandrel 910, front spar 912, internal
bladder section 914, rear spar 916, and trailing edge section 918
to form the skin 922. Mold 900 also includes internal port 915, top
clamshell half 901, bottom clamshell half 902, O-rings 920 and 906,
ports 917 and 919, and nose block 904. In one implementation, skin
922 comprises a material stack comprising laminate or core
sandwiched by laminate. This implementation is merely exemplary,
and other implementations may also be used.
[0151] As described above, in one implementation, the application
of skin 922 depends on the orientation of the fibers of the
laminate layers. In this implementation, the laminate layers are
placed in the mold such that the fibers are oriented to provide the
greatest strength. Mold 900 is designed to prevent alteration of
the orientation of the fibers during resin transfusion. This
implementation is merely exemplary, and other implementations may
also be used.
[0152] Further, in one implementation, the loading of the material
stack starts with the placement of the material on the interior of
the bottom clamshell half 902 to form the bottom portion of skin
922. After placement of the material stack, leading edge mandrel
910 would be placed in the mold, followed by front spar 912,
internal bladder section 914, rear spar 916, and trailing edge
section 918. The material stack would then be placed on top of the
IML tooling elements and the spars to form the top portion of skin
922. This implementation is merely exemplary, and other
implementations may also be used.
[0153] Additionally, in one implementation, skin 922 covers the
entirety of the IML tooling element and the spars with the
exception of portions of the trailing edge section 918. In this
implementation, skin 922 on the trailing edge section 918 is broken
to allow for the application of flaps and/or ailerons. However,
other implementations may include skin 922 that completely covers
the IML tooling element. These implementations are merely
exemplary, and other implementations may also be used.
[0154] As shown in FIG. 9C, skin 922 rests against noseblock
section 904. Noseblock section 904 prevents skin 922 from being
pinched by top clamshell half 901 upon closing of mold 900.
[0155] As also shown in FIG. 9C, a dotted box 951 is depicted along
the upper surface of skin 922. Dotted box 951 is described in FIG.
9D.
[0156] FIG. 9D is a cut-away view of a portion of a skin in a mold
for a wing panel consistent with an embodiment of the invention, as
shown in FIG. 9C. As shown in FIG. 9D, box 951 (from FIG. 9C)
comprises a cut-away of mold 900. In this cut-away, skin 922 is on
top of leading edge mandrel 910, front spar 912, internal bladder
section 914, rear spar 916, and trailing edge section 918. In one
implementation, skin 922 consists of material stacks containing a
core 924 and laminate 953. In this implementation, skin sections
954 and 959 above leading edge mandrel 910, front spar 912, and
rear spar 916 do not contain core 924. However, skin sections 958
and 960 above internal bladder section 914 and trailing edge
section 918 do contain core 924. This implementation is merely
exemplary, and other implementations may also be used.
[0157] Further, in another implementation, internal bladder section
914 forms a fuel tank 956. In this implementation, the material
stack containing core 924 is modified to provide greater strength
in the area around the fuel tank 956. This implementation is merely
exemplary, and other implementations may also be used.
[0158] Still further, in another implementation, spar caps 920 and
928 may be placed on spars 912 and 916. Spar caps 920 and 928 are
used to carry the structural load of the wing. Spar caps 920 and
928 may be co-cured or co-bonded with the wing panel. This
implementation is merely exemplary, and other implementations may
also be used.
[0159] FIGS. 9C-9D have described load material stack 830, as shown
in FIG. 8A. As shown in FIG. 8A, following load material stack 830
is close mold 840. Close mold 840 is described in FIG. 9E.
[0160] FIG. 9E is an illustrative section view of a closed mold for
a wing panel loaded with a material stack consistent with an
embodiment of the invention, as shown in FIGS. 8A-8D. As shown in
FIG. 9E, mold 900 includes top clamshell half 901 and bottom
clamshell half 902, which have been closed around noseblock section
904, skin 922, leading edge mandrel 910, front spar 912, internal
bladder section 914, rear spar 916, and training edge section 918.
Mold 900 also includes internal port 915, ports 917 and 919, and
O-rings 906 and 920. As described above, O-rings 906 and 920 may
comprise multiple O-rings or other sealing methods.
[0161] FIGS. 8A-9E have described mold process 530, as shown in
FIG. 5. As shown in FIG. 5, mold process 530 is followed by
infusion process 540. Infusion process 540 is described in FIGS.
10A-11I.
[0162] FIG. 10A is a flow diagram illustrating an infusion process
in accordance with one embodiment of the invention, as shown in
FIG. 5. As shown in FIG. 10A, infusion process 1000 begins with
confirming that the mold (such as mold 900) is properly loaded with
the correct material stack, confirming that the mold is closed,
confirming that mold is sealed, and leak checking the mold (stage
1002). This stage is further described in FIG. 11A. Next, vacuum is
applied to the interior of the mold (stage 1004). This stage is
further described in FIG. 11B. Next, heat is applied to the mold
(stage 1006). In one implementation, heat sufficient to cure a seal
core material in the material stack is used. This stage is further
described in FIG. 11C. Next, the mold is allowed to cool down
(stage 1008). Cool down includes lowering the temperature of the
mold in preparation for infusion. This stage is further described
in FIG. 11C. Next, resin is infused into the mold (stage 1010). In
this stage, resin is infused to fill any cavities in the material
stacks. This stage is further described in FIG. 11E. Next, a
hydrostatic equilibrium is achieved in the mold (stage 1012). A
hydrostatic equilibrium includes infusing the mold with resin until
the resin pressure going into the mold is equivalent to the resin
pressure coming out of the mold. This stage is further described in
FIG. 11F. Next, the mold is cured (stage 1014). Cure includes
application of heat to the mold under hydrostatic pressure. This
stage is further described in FIG. 11G. Next, the mold is allowed
again to cool (stage 1016). Cool down includes allowing the
temperature of the mold to decrease before removing the mold. In
one implementation, this stage is optional. This stage is also
further described in FIG. 11H. Finally, following cool down, the
structure (e.g. wing panel) and internal mold elements are removed
from the external mold elements and then the internal mold elements
are removed from the structure (stage 1018). This implementation is
merely exemplary, and other implementations may also be used. Some
of the other implementations are described in FIGS. 10B-10D.
[0163] FIG. 10B is a block diagram illustrating an infusion process
in accordance with another embodiment of the invention, as shown in
FIG. 5. As shown in FIG. 10B, infusion process 1050 is similar to
infusion process 1000 in FIG. 10A. However, in infusion process
1050, apply heat stage 1006 and cool stage down 1008 have been
replaced by apply heat stage 1020. In this implementation, apply
heat stage 1020 includes increasing the temperature of the mold to
a point sufficient to seal core material in the material stack but
not higher than the desired temperature for resin infusion. Thus,
this implementation does not require cool down 1008, as described
in FIG. 10A. This implementation is merely exemplary, and other
implementations may also be used.
[0164] FIG. 10C is a block diagram illustrating an infusion process
in accordance with still another embodiment of the invention, as
shown in FIG. 5. As shown in FIG. 10C, infusion process 1060 is
also similar to infusion process 1000 in FIG. 10A. However, in
infusion process 1060, apply heat stage 1006 and cool down stage
1008 are replaced by apply heat stage 1024 and apply heat stage
1026. In one implementation, apply heat stage 1024 includes
increasing the temperature of the mold to a point sufficient to
seal core material in the material stack but less than the
appropriate resin infusion temperature. Apply heat stage 1026
includes increasing the temperature of the mold to the proper
temperature for resin infusion. This implementation is merely
exemplary, and other implementations may also be used.
[0165] The embodiments in FIGS. 10A-10C demonstrate how the cure
temperatures of the adhesive used to seal core elements and the
appropriate temperature for resin infusion affect infusion
processes 1000, 1050, and 1060. Therefore, depending on the resin
and adhesive chosen, the appropriate temperature for resin infusion
may be lower than, higher than, or the same as the cure temperature
for the adhesive. Infusion process 1000 in FIG. 10A demonstrates
the situation where the appropriate temperature for resin infusion
is lower than the cure temperature for the adhesive. Therefore, as
shown in FIG. 10A, cool down stage 1008 is required before infusion
stage 1010. Infusion process 1050 in FIG. 10B demonstrates the
situation where the appropriate temperature for resin infusion and
the cure temperature for the adhesive is substantially the same.
Therefore, as shown in FIG. 10B, a stage is not needed after apply
heat stage 1020 because the mold is at the appropriate temperature
for infusion stage 1010. Infusion process 1060 in FIG. 10C
demonstrates the situation where the appropriate temperature for
resin infusion is higher than the cure temperature for the
adhesive. Therefore, as shown in FIG. 10C, a second apply heat
stage, i.e. apply heat stage 1026, is needed before infusion
1010.
[0166] FIG. 10D is a block diagram illustrating an infusion process
in accordance with yet another embodiment of the invention, as
shown in FIG. 5. As shown in FIG. 10D, infusion process 1070 is
also similar to infusion process 1000 in FIG. 10A. However, in
infusion process 1070, there is no apply heat stage 1006 and no
cool down stage 1008, as in infusion process 1000. Instead, in
infusion process 1070, infusion 1010 occurs after apply vacuum
1004. This can be done in this embodiment because the core
materials are cured prior to loading of the material stack in the
mold. For example, curing of the core material can be done during
preparation of the material stack prior to surrounding it with
laminate layers.
[0167] FIGS. 10B-10D illustrate a variety of implementations of
infusion process 1000. These implementations are merely exemplary,
and other implementations may also be used.
[0168] FIGS. 11A-11I now describe infusion process 1000 in more
detail. As shown in FIG. 10A, the first step in infusion process
540 is confirmations stage 1002. Confirmations stage 1002 is
described in FIG. 11A.
[0169] FIG. 11A is an illustrative section view of a mold for a
wing panel in a confirmations stage consistent with an embodiment
of the invention, as shown in FIG. 10A. As shown in FIG. 11A, mold
1100 shows a material stack 922 in a closed and sealed mold. Top
clam shell half 901 and bottom clamshell half 902 have been closed
around noseblock section 904, leading edge mandrel 910, front spar
912, internal bladder section 914, rear spar 916, and trailing edge
mandrel section 918. Mold 1100 also shows O-rings 906 and 920,
which help seal mold 1100 and ports 915, 917, and 919, which may
also help seal mold 1100.
[0170] In one implementation, confirmation is made that mold 1100
is properly loaded with the correct material stack, that the mold
1100 is closed, and that mold 1100 is sealed. Mold 110 may also be
leak checked. This implementation is merely exemplary, and other
implementations may also be used.
[0171] FIG. 11B is an illustrative section view of a mold for a
wing panel in a vacuum stage consistent with an embodiment of the
invention, as shown in FIG. 11A. As shown in FIG. 11B, a vacuum
source (not shown) is applied to mold 1110. In one implementation,
the vacuum source may be applied to one port, e.g. port 919, while
port 917 is closed. In another implementation, the vacuum may be
applied from port 917, while port 919 is closed. In still another
implementation, the vacuum may be applied to both ports 917 and
919. In one implementation, the vacuum level inside mold 900 is at
most 2 Torr. These steps may be performed using any vacuum source
capable of creating a sufficient vacuum. These implementations are
merely exemplary, and other implementations may also be used.
[0172] Following creation of a vacuum, mold 1110 is checked for
leaks. In one test protocol, leakage must be less than 5 inches of
Hg in a 5-minute period. However, other implementations and other
test protocols may be used. As shown in FIG. 11B, mold 900 includes
O-ring seals 906 and 920. In one implementation, O-ring seals 906
and 920 are dimensioned to minimize leakage after the mold is
closed. These implementations are merely exemplary, and other
implementations may also be used.
[0173] In one implementation, a vacuum source may be exposed to
internal bladder section 914 using port 915 to create a low
pressure condition inside internal bladder section 914. In another
implementation, pressure may be applied to internal bladder section
914 via port 915. A pressure source (not shown) may control the
pressure inside internal bladder section 914 using port 915. These
implementations are merely exemplary, and other implementations may
also be used.
[0174] FIG. 11C is an illustrative section view of a mold for a
wing panel in a heat stage consistent with an embodiment of the
invention, as shown in FIG. 11B. As shown in FIG. 11C, mold 1120
may be heated and pressed. In one implementation, heat is applied
to the exterior of top clamshell half 901 and bottom clam shell
half 902. The heat can be applied using an oven, an autoclave, a
press, or any other method of applying heat to an object. In an
implementation using an autoclave, the autoclave presses top
clamshell half 901 and bottom clamshell half 902 together. In this
implementation, ports 917 and 919 are closed and port 915 is open
to allow the internal bladder section 914 to vent to the autoclave
atmosphere. Therefore, pressure exists on both sides of skin 922.
In an implementation using a press, the press presses top clamshell
half 901 and bottom clam shell half 902 together. In this
implementation, ports 917 and 919 would be closed and compressed
gas would be placed in internal bladder section 914 using port 915.
A press with heated plates may also be used. In one implementation,
this stage may be used to cure the adhesive in a material stack,
thereby sealing a core material. These implementations are merely
exemplary, and other implementations may also be used.
[0175] FIG. 1D is an illustrative section view of a mold for a wing
panel in a cool down stage consistent with an embodiment of the
invention, as shown in FIG. 11C. As shown in FIG. 11D, mold 1130
may be cooled, after the heating stage, as described in FIG. 11C.
In this implementation, heat is dissipating from top clamshell half
901 and bottom clam shell half 902. In this implementation, the
temperature of halves 901 and 902 are lowered to a temperature
appropriate for resin infusion. The appropriate temperature will
depend on the choice of resin. In one implementation, resin is
infused at 130 degrees Fahrenheit at 3 atmospheres. Cooling may be
accomplished in this implementation by actively cooling the mold or
by allowing ambient atmosphere to gradually cool down the mold.
These implementations are merely exemplary, and other
implementations may also be used.
[0176] FIG. 11E is an illustrative section view of a mold for a
wing panel in an infusion stage consistent with an embodiment of
the invention, as shown in FIG. 11D. As shown in FIG. 11E, mold
1140 may be infused with resin 1102. In one implementation, resin
1102 is infused through port 917 and evacuated through port 919. In
this implementation, resin is infused at 130 degrees Fahrenheit.
This implementation is merely exemplary, and other implementations
may also be used.
[0177] In one implementation, the resin infusion process begins
with port 917 being open and with a vacuum being applied to port
919. In one implementation, port 919 has a trap mechanism (not
shown) to allow a vacuum to be created in mold 1140 during infusion
of resin 1102. A pump (not shown) infuses resin 1102 into mold 1140
through port 919 at a specified pressure while a vacuum continues
to be applied to port 919. In one implementation, this resin
pressure is 45 psi. However, this pressure can range from 10-200
psi. As shown in FIG. 11E, in these implementations, pressure
gauges 1104 and 1106 may be located at ports 917 and 919. When
resin 1102 is initially pumped into mold 1140, the pressure at port
917 will be the pressure at which resin 1102 is being introduced.
However, the pressure at port 919 will reflect a low pressure
reading due to the vacuum source, as shown in FIG. 11E.
[0178] As more resin enters port 917, some resin may start to
evacuate port 919. In this implementation, at the point that resin
begins filling port 919, a vacuum is no longer applied at port 919.
As resin 1102 begins to exit port 919, the pressure measured at
port 919 will increase. This implementation is merely exemplary,
and other implementations may also be used.
[0179] In another implementation, pressure may be applied through
port 915 to internal bladder section 914 to balance the pressure in
internal bladder section 914 against the pressure of resin 1102
entering mold 1140. This halts the resin flow in the mold.
Pressurizing internal bladder section 914 removes excess resin,
consolidates laminate layers, and minimizes voids. These
implementations are merely exemplary, and other implementations may
also be used.
[0180] As described above, infusion of resin 1102 may cause a shift
in the fiber orientation of the material stack. In one
implementation, as described above, the geometry of the tool is
precisely controlled to reduce alteration of the fiber orientation.
In addition, in this implementation, the pressure inside mold 1140
may also be controlled to offset the effect of the pressure of
resin infusion. This will also reduce alteration of the fiber
orientation. This implementation is merely exemplary, and other
implementations may also be used.
[0181] FIG. 11F is an illustrative section view of a mold for a
wing panel in a hydrostatic equilibrium stage consistent with an
embodiment of the invention, as shown in FIG. 11E. As shown in FIG.
11F, mold 1150 may be placed into a hydrostatic equilibrium. In one
implementation, hydrostatic equilibrium is obtained when the resin
pressure entering port 917 equals the resin pressure coming out of
mold 1150 at port 919, as shown on pressure gauges 1104 and 1106.
This condition indicates that all the cavities within the material
stack have been filled with resin 1102. In this implementation,
mold 1150 will be held at hydrostatic equilibrium for a few minutes
to ensure that all cavities have been filled. The amount of time to
hold equilibrium, depends on the size and shape of the part. In one
implementation, equilibrium is held from 30 minutes to one hour.
This implementation is merely exemplary, and other implementations
may also be used.
[0182] FIG. 11G is an illustrative section view of a mold for a
wing panel in a cure stage consistent with an embodiment of the
invention, as shown in FIG. 11F. In one implementation, heat is
applied to cure resin 1102 in mold 1160. In this implementation,
the outer mold elements of mold 1160 are clamped together (not
shown) and placed in an oven and heated (not shown). In another
implementation, mold 1160 is placed in an autoclave where heat and
pressure is applied. In still another implementation, heated
platens are pressed against either side of mold 1160. Additionally,
other methods described in FIG. 11C may be used to cure resin 1102.
These implementations are merely exemplary, and other
implementations may also be used.
[0183] The temperature to which mold 1160 is heated depends on the
material stack and resin. In one implementation, mold 1160 is
heated to 270 degrees Fahrenheit to cure the structure and then to
300 degrees Fahrenheit to post-cure the structure. Post curing
allows for increased strength in the structure. This implementation
is merely exemplary, and other implementations may also be
used.
[0184] FIG. 11H is an illustrative section view of a mold for a
wing panel in a cool down stage after curing consistent with an
embodiment of the invention, as shown in FIG. 11G. As shown in FIG.
11H, mold 1170 may be cooled after the cure stage, as described in
FIG. 11G. In this implementation, heat is dissipating from mold
1170. Additionally, other methods described in FIG. 11D may be used
to cool down mold 1170. This implementation is merely exemplary,
and other implementations may also be used.
[0185] FIG. 11I is an illustrative section view of a mold for a
wing panel in a demold stage consistent with an embodiment of the
invention, as shown in FIG. 11H. As shown in FIG. 11I, mold 1180
may be disassembled, or demolded, to remove the molded composite
structure, such as a wing panel. In one implementation, mold 1180
may be demolded when it reaches 180 degrees Fahrenheit either
during or after cool down. In this implementation, the wing panel
is removed from top clamshell half 901 and bottom clam shell half
902. The wing panel comprises the co-cured skin 922 and the
co-cured spars 912 and 916. In addition, this wing panel comprises
leading edge mandrel 910, internal bladder section 914, and
trailing edge mandrel section 918 in its interior. In this
implementation, these are removed. Other implementations may be
used.
[0186] With regard to FIG. 11I, as described above, the elements of
mold 1180 may have a different coefficient of thermal expansion
than one another and/or the molded composite structure. Thus, upon
cool down, as described in FIG. 11H, mold 1170 (in FIG. 11H) or
mold 1180 (in FIG. 11I) could contract in such a way as to damage
the mold or the molded composite structure. In one implementation,
mold 1180 may accommodate thermal expansion. In another
implementation, tooling may be allowed to shrink relative to the
molded composite structure and allow the molded composite structure
to move in the tool, so that the molded composite structure is not
placed under strain or stress. For example, in this implementation,
the molded composite structure may expand relative to the tooling
during cool down. Similarly, in this implementation, the root end
of the molded composite structure may move relative to the main
body of the tooling to relieve stress during cool down. This
implementation is merely exemplary, and other implementations may
also be used.
[0187] FIGS. 10A-11I have described infusion process 540, as shown
in FIG. 5. As shown in FIG. 5, infusion process 540 results in the
creation of a structure, such as a wing panel. An example of a wing
panel is shown in FIG. 12.
[0188] FIG. 12 is a perspective view of a wing panel manufactured
consistent with an embodiment of the invention. As shown in FIG.
12, in one implementation, a wing panel 1200 comprises skin 1230,
co-cured spars 1210 and 1220, and co-cured ribs 1270 for support of
hinges for ailerons or flaps (not shown). In this embodiment, all
of the elements of wing panel 1200 are manufactured according to
the described processes, e.g. RTM process 130 as described in FIG.
5 (and otherwise described herein). Additionally, in this
implementation, other elements may be formed in wing panel 1200.
For example, a fuel tank 1260 could be formed in the structure.
Fuel tank ribs 1240 could also be included to be co-cured with the
rest of the structure to support fuel tank 1260. Other elements may
also be formed in wing panel 1200. This implementation is merely
exemplary, and other implementations may also be used.
[0189] In one implementation, skin 1230 will result in a smooth
laminar flow of air over wing panel 1200. In this implementation, a
smooth laminar flow of air includes a streamlined flow of a fluid
(i.e. air) over wing panel 1200 with little turbulence. This
implementation is merely exemplary, and other implementations may
also be used.
[0190] Other embodiments of the invention will be apparent to those
skilled in the art from consideration of the specification and
practice of the invention disclosed herein. It is intended that the
specification and examples be considered as exemplary only, with a
true scope and spirit of the invention being indicated by the
following claims.
VI. CONCLUSION
[0191] As described above, therefore, other embodiments of the
invention will be apparent to those skilled in the art from
consideration of the specification and practice of the invention
disclosed herein. It is intended that the specification and
examples be considered as exemplary only, with a true scope and
spirit of the invention being indicated by the following claims and
their equivalents. In this context, equivalents mean each and every
implementation for carrying out the functions recited in the
claims, even if not explicitly described therein.
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