Note: Descriptions are shown in the official language in which they were submitted.
CA 02598765 2007-09-11
DESCRIPTION
METHOD AND APPARATUS FOR
MANUFACTURING COMPOSITE STRUCTURES
BACKGROUND OF THE INVENTION
1. Field of the Invention
This invention relates generally to composite material structures and, more
specifically,
to fiber placement methods and apparatus for manufacturing composite material
shells. In
particular, this invention relates to composite material shells, such as
aircraft fuselage
components, formed using fiber placement and other methods employing a
removable mandrel
and a bladder with or without an integral caul sheet.
2. Description of the Related Art
The fuselage of an airplane or other similar vehicle is generally a thin shell
of
revolution. In the case of an airplane, one of the significant loading
conditions for a fuselage is
circumferential tension resulting from internal pressure. Structurally it is
most efficient to carry
this loading with a structure having a continuous diameter or hoop without any
axial joints.
From a manufacturing sense each joint in a structure tends to add cost. Also,
from a
manufacturing sense each extra component or detail tends to add cost.
Composites have proven to be very useful materials, especially in the field of
aviation.
Weight is a very important and sensitive subject and any method to limit or
reduce it is
valuable. In addition, structures of composite materials are usually thinner,
allowing for
increased internal space or decreased area.
Composite materials, such as carbon fiber present in an organic matrix, have
been used
to produce corrosion resistant and light weight structures. These structures
typically weigh
about 25% less than structures made of lightweight metals, such as aluminum,
while at the same
time offering similar strength to these metals. As a result, composite
materials have been used
to fabricate a wide variety of structures including, most notably, aircraft
structures (such as
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fuselage shell components, wing sections, tail sections, etc.). However, these
composite
structures have typically been manufactured by time consuming application
methods, such as
hand placement.
In addition, when used to manufacture aircraft structures, such as fuselage
components,
composite structures have typically been manufactured in many separate parts,
such as fuselage
halves split down the longitudinal axis, that must be bonded or fastened
together, typically
using a flush joint. For example, in one method, a forward fuselage section
with four or more
separate composite structure components has been manufactured. Machining and
assembling of
flush joints into a single uniform component typically requires a substantial
amount of time to
achieve a uniform and consistent fl-u-s-IFjoint. In addition to extra time,
each flush joint adds
additional material and weight to the assembled aircraft component. Therefore,
the greater the
number of separate parts required to construct a single component, such as
fuselage component
sections, the larger the amount of time and the greater the amount of weight
added to the
assembled component.
In an effort to reduce composite part assembly time and to produce lighter
weight
composite parts, fiber placement (or tow placement) methods have been
developed. Such fiber
placement methods may include computer control integrated with a fiber
placement machine.
Operation of such a machine to place tow filaments on a mandrel to form
composite structures
is known in the art. Fiber placement methods involve the automated placement
(typically by
winding) of filaments (such as fibrous ribbons or tows which are pre-
impregnated with a
thermal set resin material such as epoxy) onto a mandrel to produce a
component, such as a
tube-shaped part. These fibers are typically placed at varying angles and in
segments of varying
width. A fiber tow is essentially a ribbon of carbon fiber, typically between
about 1/4" and
about 118" wide. Using a conventional fiber placement machine, multiple tows
are transported
to a movable payoff head and applied to a mandrel surface using a roller.
Typically, a payoff
head includes an automatic cutting system for cutting and restarting
individual tows. In
addition, typical fiber placement machines include heating devices to vary the
temperature and,
therefore, the properties of the tows as they are applied. Means for
controlling pressure applied
to the tows and mandrel during fiber placement are also typically employed.
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Although fiber placement processes may be used to produce composite structures
of
varying dimension and size more quickly and efficiently than other methods,
current fiber
placement techniques suffer from complications relating to mandrel
construction and removal
of the mandrel after fiber placement has occurred. In particular, segmented
mandrels have been
provided having segments that are joined during fiber placement and
disassembled after curing.
Segmented mandrel designs suffer from numerous problems, including expansion
of the
mandrel material during heat curing, leakage between mandrel segment joints,
and time and
effort involved in the assembly and disassembly of mandrel components.
In the construction of composite structures, and aircraft composite structures
in
particular, interior dimensions of a structure are of particular concern.
Although fiber
placement techniques have been used to produce aircraft fuselage shell
components, these shell
components have required cylindrical support frames and elongated longeron
support members
that serve to support the outer fuselage structure. These frames and longerons
are typically
wider or deeper than a composite fuselage wall of sandwich construction, and
therefore serve to
reduce the interior diameter of an aircraft fuselage. In other cases, fiber
placed composite
structures have been manufactured in non-continuous separate parts, such as
separate axial
fuselage half or quarter panels, that are assembled to form a single cross
sectional shape. These
structures suffer from the cost and weight problems described above for other
multi-piece
composite structure components.
Consequently, a need exists for simplified methods and apparatus for forming
relatively
large, single piece composite parts, such as aircraft -fuselage components. In
particular a need
exists for simplified methods of mandrel installation and removal. A need also
exists for a
method of manufacturing composite shell components, such as aircraft fuselage
parts, which do
not require internal frames or bracing and which have an increased internal
diameter.
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SUMMARY OF THE INVENTION
In one aspect, the present invention provides a tool useful in forming a
composite body. The tool has a mandrel body with an outer surface and one or
more
fluid openings are defined in the mandrel body outer surface. The tool also
includes a
mandrel body fluid system capable of supplying pressurized fluid and a vacuum
to
said one or more openings in the mandrel body outer surface. The tool also
includes a
bladder having an outer surface and an inner surface. The bladder is placeable
upon
the outer surface of the mandrel body. A caul sheet is also provided and has
an inner
surface and an outer surface, the inner surface of the caul sheet being
configured to
overlay the outer surface of the bladder and the outer surface of the caul
sheet being
configured to provide a surface for supporting an uncured composite body
thereon.
In another aspect, the invention provides a method useful in forming a
composite body, comprising:
providing a tool useful in forming a composite body, said tool comprising:
a mandrel body having an outer surface, said mandrel body having an
elongated shape and a longitudinal axis and having one or more fluid openings
defined in said mandrel body outer surface;
a mandrel body fluid system capable of supplying pressurized fluid and a
vacuum to said one or more openings in said mandrel body outer surface; and
a bladder fitted around the outer surface of the mandrel body,
wherein said mandrel body fluid system is capable of supplying pressurized
fluid or a vacuum to said one or more openings in said mandrel body outer
surface
such that said fluid flows through said one or more openings in a single
direction, said
single direction being either outward from said one or more openings, or
inward into
said one or more openings;
placing a plurality of fibers around said bladder to form an uncured body;
positioning the tool and the uncured body within a mold;
forming a seal between the bladder and the mold; and
drawing a vacuum between the bladder and the mold.
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A still further aspect of the invention provides a method useful in forming a
composite body, comprising the steps of:
providing a tool for use in forming a composite body, said tool comprising:
a mandrel body having an outer surface;
one or more fluid openings defined in said mandrel body outer surface;
a mandrel body fluid system capable of supplying pressurized fluid and a
vacuum to said one or more openings in said mandrel body outer surface; and
a bladder having an outer surface and an inner surface, the bladder having a
shape and dimensions complementary to the outer surface of the mandrel body
and the inner surface of the bladder being fitted around the outer surface of
the
mandrel body,
wherein said mandrel body fluid system is capable of supplying pressurized
fluid or a vacuum to said one or more openings in said mandrel body outer
surface
such that said fluid flows through said one or more openings in a single
direction, said
single direction being either outward from said one or more openings, or
inward into
said one or more openings;
placing a plurality of fibers around said tool to form an uncured body; and
removing the mandrel body from within the bladder prior to curing the
uncured body.
In a still further aspect, the invention provides a tool useful in forming a
composite body, including a mandrel having an outer surface, the outer surface
being
capable of receiving a bladder having inner and outer surfaces and a shape and
dimensions complementary to said outer surface of the mandrel body such that
the
bladder may be fitted around the outer surface of the mandrel body. One or
more
fluid openings are defined in the mandrel body outer surface. A mandrel body
fluid
system is provided and is capable of supplying pressurized fluid and a vacuum
to the
one or more openings in the mandrel body outer surface. The mandrel body fluid
system is capable of supplying pressurized fluid to the said one or more
openings so
as to reduce friction between the inner surface of the bladder and the outer
surface of
the mandrel body when removing the mandrel from the bladder. The mandrel body
fluid system is capable of supplying sufficient vacuum to secure the inner
surface of
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the bladder to the outer surface of the mandrel body while composite material
is
placed around the bladder and the mandrel body.
BRIEF DESCRIPTION OF THE DRAWINGS
FIG. 1 is a perspective view of a composite aircraft fuselage structure
according to one embodiment of the disclosed method and apparatus.
FIG. 2 is a cross-sectional view of a composite body wall having a sandwich
construction according to one embodiment of the disclosed method and
apparatus.
FIG. 3 is a perspective view of a mandrel body according to one embodiment
of the 20 disclosed method and apparatus.
FIG. 4 is a perspective view of an integral bladder/caul sheet according to
one
embodiment of the disclosed method and apparatus.
FIG. 5 is a perspective view of a clam shell mold according to one
embodiment of the disclosed method and apparatus.
FIG. 6 is a perspective view of a forward composite fuselage section and clam
shell mold halves according to one embodiment of the disclosed method and
apparatus.
FIG. 7 is a perspective view of an aft composite fuselage section and clam
shell mold halves according to one embodiment of the disclosed method and
apparatus.
CA 02598765 2007-09-11
FIG. 8 is a perspective view of molds for manufacturing clam shell mold halves
according to one embodiment of the disclosed method and apparatus.
FIG. 9 is a sequential illustrative flowchart showing process steps according
to one
embodiment of the disclosed method and apparatus.
FIG. 10 is a perspective view of an integral bladder/caul sheet installed on a
mandrel
according to one embodiment of the disclosed method and apparatus.
FIG. 11 is a cross-sectional view of an inner skin layer showing underlapping
and
overlapping skin halves according to one embodiment of the disclosed method
and apparatus.
FIG. 12 is a perspective view of a mandrel with integral bladder/caul sheet
installed on
a mandrel transportation dolly according to one embodiment of the disclosed
method and
apparatus.
FIG. 13 is a perspective view of a mandrel, mandrel removal fixture, and clam
shell
molds according to one embodiment of the disclosed method and apparatus.
FIG. 14 is another perspective view of a mandrel, mandrel removal fixture, and
clam
shell molds according to one embodiment of the disclosed method and apparatus.
FIG. 15 is a perspective view of a reusable bag according to one embodiment of
the
disclosed method and apparatus.
FIG. 16 is a partial perspective view of the aft end of a reusable bag
according to one
embodiment of the disclosed method and apparatus.
FIG. 17 is a cross-sectional view of a membrane and sealing elements according
to one
embodiment of the disclosed method and apparatus.
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FIG. 18 is a side view of a mandrel and bladder/caul sheet having forward and
aft
sealing loops according to one embodiment of the disclosed method and
apparatus.
FIG. 19 is a cross-sectional view of a sealing element mounted to bladder/caul
sheet
surfaces according to one embodiment of the disclosed method and apparatus.
FIG. 20 is a cross-sectional view of a membrane and sealing half according to
one
embodiment of the disclosed method and apparatus.
FIG. 21 is a perspective view of a clam shell mold half and clam shell sealing
loop
according to one embodiment of the disclosed method and apparatus.
FIG. 22 is a cross-sectional view of a clam shell seal half according to one
embodiment
of the disclosed method and apparatus.
FIG. 23 is a partial cross-sectional view of a bladder/caul sheet sealed to
clam shell
mold surfaces with mating seal elements according to one embodiment of the
disclosed method
and apparatus.
FIG. 24 is a cross-sectional view of a flexible membrane and seal elements
juxtaposed
with mated clam shell mold halves having corresponding mating seal elements
according to one
embodiment of the disclosed method and apparatus.
FIG. 25 is a cross-sectional view of a flexible membrane and seal elements
disposed in
mated relationship with mating clam shell mold halves having corresponding
mating seal
elements according to one embodiment of the disclosed method and apparatus.
FIG. 26 is a perspective view of an integral bladder/caul sheet and
corresponding
locating plugs and frames according to one embodiment of the disclosed method
and apparatus.
FIG. 27 is a cross-sectional view of an integral bladder/caul sheet, composite
shell,
locating plugs, and frames according to one embodiment of the disclosed method
and apparatus.
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DESCRIPTION OF SPECIFIC EMBODIMENTS
In embodiments of the disclosed method and apparatus, composite structures
having
high quality, light weight, and low fabrication costs are possible. These
characteristics are
achieved by, among other things, providing a single piece full cylinder
composite structure
design having no axial joints, composite shell outer surfaces which are tooled
to the outer mold
line, and relatively large composite structure components that reduce the
number of separate
components required for assembly. Furthermore, methods are provided for
fabrication of
l0 composite structure shells on a tool having an outer surface representing
the inside surface of
the fabricated structure to achieve minimum labor content.
Various embodiments of the disclosed method and apparatus are directed to
structural
designs. manufacturing processes, and tooling systems capable of producing
composite
structures at minimum cost and weight. In a typical embodiment, this
disclosure is directed to
fabrication of aircraft fuselage components. However, it will be understood
with benefit of this
disclosure that other embodiments of the disclosed method and apparatus are
possible,
including fabrication of aircraft components other than fuselage parts, and
fabrication of
structures for other types of aircraft and nonaircraft vehicles, and other
types of structures. It
will also be understood that hollow or closed shaped structures having non-
circular cross
sections, solid or only partially hollow structures, and/or structures having
separate axial halves
may also be constructed using one or more of the disclosed features. The
disclosed method and
apparatus offer particular advantages to those structures that benefit from
the characteristics of
light weight, strength and/or increased interior clearance and volume.
Advantageously, using the disclosed method and apparatus, relatively large
single piece
full cylinder composite structures may be manufactured including, for example,
a single piece
aircraft fuselage component including a nose section and a constant section.
By "full cylinder"
or "continuous skin" it is meant that the cross sectional diameter of a
tubular or other hollow
shaped structure is constructed in a single continuous piece (or hoop) without
any axial joints.
By "constant section" it is meant the part of a structure where the diameter
is constant over the
entire length. In such an embodiment, two composite laminated fuselage
components,
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including a single-piece forward section, may be cured at an elevated
temperature and pressure
and then joined with a circumferential splice at the aft end of the constant
section. In further
embodiments, large full cylinder components may be manufactured in a laminate
composite
sandwich to form a one-piece cylinder sandwich structure. Because components
are of full
cylinder construction, problems and costs associated with half section mating
are eliminated. In
addition, material waste during trimming and fitting is greatly reduced over
conventional
methods.
FIG. I shows the basic configuration of one embodiment of an aircraft fuselage
which
may be constructed using the methods and apparatus of the present disclosure.
In this case, the
overall fuselage is about 70 inches in diameter and is fabricated in two
sections. The forward
section 10 is about 24 feet long while the aft section 12 is about 12 feet
long. The forward
section includes both a nose section 10a of about 10.5 feet in length and a
constant section 10b
of about 13.5 feet in length. In this embodiment, forward section 10 and aft
section 12 are
fabricated separately and then joined or mated at joining area 11, typically
using a splice band.
Although FIG. 1 depicts a two-piece aircraft fuselage having a particular
shape, size,
and configuration, it will be understood with benefit of the present invention
that other shapes,
sizes and configurations of aircraft fuselages and other structures may be
constructed. It will
also be understood that an aircraft fuselage or other vehicular or non-
vehicular structure
constructed using the disclosed method and apparatus may be a single piece
structure or a
multiple piece structure having more than two sections. Specific examples of
other types of
aircraft components that may be fabricated include, .but are not limited to
flap panels. Specific
types of other vehicular structures that may be fabricated include, but are
not limited to space
vehicle structures, automobile structures, train structures, and boat
structures.
In the practice of the disclosed method, composite structures may be
fabricated using
automated fiber placement of filaments (or tows), or by other processes,
including but not
limited to fiber placement, filament winding or hand layup. Using fiber
placement methods,
composite materials may be laid up to form composite structures of single or
multiple
laminations. In the embodiment illustrated in FIG. 1, an aircraft fuselage
structure is typically
fabricated to be of a sandwich construction as shown in cross-sectional view
in FIG. 2.
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However, other types of composite bodies may be fabricated to be of such a
sandwich
construction. In the embodiment of FIG. 2, the composite shell has a cross-
section that
includes an inner skin 20, a honeycomb core 24, and an outer skin 22. In non-
reinforced areas,
inner skin 20 is typically of a thickness of from about 0.024 inches to about
0.034 inches, most
typically about 0.024 inches. Core 24 has a typical non-reinforced thickness
of from about 0.70
inches to about 0.75 inches, most typically about 0.71 inches. Outer skin 22
is a mirror image
of the inner skin. The total non-reinforced sandwich thickness for this
application is typically
constant, for example, at about 0.810 inches. It will be understood with
benefit of this
disclosure that the thickness of any or all three layers may be thicker or
thinner than those
dimensions given above. The thickness of any or all of these three layers may
also be varied for
purposes of local reinforcement or rigidity, and if desired in such a manner
that the overall wall
or sandwich thickness remains constant.
The above described embodiment allows, for example, a frameless aircraft
fuselage
body to be produced that is designed for a pressure differential of from about
8 psi to about
16 psi, most typically about 8.4 psi. In this example, the fuselage body
typically has a
maximum exterior diameter of about 70 inches and maximum interior diameter of
about 68:4
inches. This translates to a typical interior width of between about 4 inches
to about 5 inches
greater. and an interior volume of about 13% greater. than a comparable
aircraft fuselage body
having a similar outer diameter and utilizing frame and longeron construction.
In the
embodiment of FIGS. 1 and 2. two separate fuselage components are typically
fabricated and
assembled to produce an aircraft fuselage, for example, having an maximum
interior height of
about 5.8 ft. a maximum interior width of about 5.8 ft, and an interior length
of about 13.5 ft.
Although one aircraft fuselage embodiment having particular construction
details and
dimensions is shown and described above, it will be understood with benefit of
the present
disclosure that other fiber placed and/or composite structure embodiments
having varying
shapes, dimensions, and construction designs are also possible.
As will be described below, frame details may be used to close out the
sandwich at
openings for doors and windows, and local reinforcements may be added within
the sandwich
as necessary to provide adequate strength for distributing loads such as occur
at points where
the wing joins the composite shell. Although FIG. 2 illustrates a three-laver
sandwich
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laminated composite cross-section having two skin layers and one core layer,
it will be
understood with the benefit of the disclosed method that a laminate composite
cross-section
may include only one layer, or may be a sandwich having four or more layers.
In addition it
will be understood that any combination of skin and core layers may be
employed including for
example, more than one core and greater or less than two skin layers.
In the practice of the disclosed method, any suitable fiber or tow material
for forming
composite structures using tow or fiber placement technology may be employed.
Such
materials include, but are not limited to, graphite, glass, "KEVLART"'", or
combinations thereof,
with a matrix material, such as epoxy, polyester, bismaleimide, polyimide
phenolic, or mixtures
thereof. In the embodiment illustrated in FIGS. 1 and 2, skins 20 and 22 are
typically a
combination of carbon fiber/epoxy prepregs applied in tape, fabric, and
continuous towpreg
forms. Prepreg is a term commonly used to refer to pre-impregnated. Specific
examples of
such carbon fiber/epoxy prepregs include "5276-1" and "195/P3" available from
CYTEC.
Although epoxy resins, such as "5276-1", are typically employed, other resins
may be used
including, but not limited to "8552-1", "E7K8", "E7T1-2". Most typically, a
high strength
carbon fiber with a toughened epoxy resin is employed. In addition, skins 20
and 22 may have
local reinforcements that are typically constructed of one or more plies of
fabric, tape or
towpreg applied in a size and orientation dictated by the loading condition.
Metallic
reinforcements, such as thin titanium sheet may also be laminated into the
structure, if desired.
As previously described, skins 20 and 22 are typically fabricated using
automated fiber
placement techniques, while honeycomb core 24 may be fabricated in a number of
ways,
including in metallic form using, for example, aluminum, stainless steel or
titanium foils, or in
composite form using, for example, glass fabric, graphite fabric, or "KEVLAR"
material in the
form of "NOMEXT"'". Typically, honeycomb composite cores employ phenolic resin
matrix
material. Typically, honeycomb core 24 is made from "NOMEX" available from,
for example,
Hexcel, Plascore. and Ciba-Geigy.
Although FIG. 2 illustrates an embodiment of the sandwich cross-section having
a
honeycomb core and two carbon fiber/epoxy skins, it will be understood with
benefit of the
present disclosure that a core may be of another material or construction
including, but not
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limited to any relatively low density material suitable for transferring shear
loading between
skins 20 and 22. Such other materials include, but are not limited to, rigid
foam and balsa
wood.
A. Fabrication of Composite Structures
In the disclosed method, composite structures are typically manufactured using
automated fiber placement methods, but may also be fabricated using various
other processes,
including but not limited to fiber placement, filament winding or hand layup.
When using
automated fiber placement methods, a band, typically consisting of up to 24
individual strands
of a unidirectional carbon/epoxy prepreg may be applied or wound under tension
onto the
surface of a tool, such as a mandrel. Individual strands are typically 6K and
have a width of
from about I /S inch to about 1 /4 inch, most typically about 1/8 inch. These.
tow strands
typically contain about 136 gm/m` of carbon fiber and have a. resin content of
about 38% by
weight. This tow configuration yields a thickness of about 0.005 inch.
However, other types of
tows including, but not limited to, 1K-12K towstrands may also be used.
Advantageously, use
of fiber placement is significantly faster and more accurate than traditional
hand layout of
composite material. In the practice of the disclosed method, fiber placement
is typically
performed using a fiber placement machine, such as a "SEVEN AXIS VIPERTM"
available from
Cincinnati Milacron. Among the other machines suitable for such fiber
placement include
machines produced by Ingersol. Advantageously, use of such fiber placement
machines with
the methods of the present disclosure allows fabrication of parts having
varying and complex
structural characteristics, such as aircraft fuselage parts and other aircraft
components in large
single pieces. Computer controls used with these machines are capable'of
stopping and re-
initiating application of fiber strands and of leaving necessary openings in a
part, so that the
openings do not have to be later formed in the part in separate steps. In
addition, design
changes may be easily implemented by changing the computer instructions.
Mandrel
In the practice of the disclosed method, tubular and other hollow composite
structures
are formed by placement of material on the outside surfaces of a form,
commonly referred to as
a mandrel. Material may be placed using automated fiber placement methods or
other
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processes, including but not limited to fiber placement, filament winding or
hand layup. FIG. 3
shows one embodiment of a mandrel body 40 suitable for manufacturing a forward
aircraft
fuselage section. As shown in FIG. 3, mandrel 40 typically includes a mandrel
shell 40a, an aft
mandrel support structure 40b, and a forward mandrel shaft 40c. Mandrel shell
40a includes
outline for windshield location feature 40d, smooth surface feature 40j, and
smooth surface
feature 40i. These features are for sealing the bladder/caul sheet 42 to the
mandrel shell surface.
Also shown in FIG. 3 are mandrel fluid openings 40e which are connected to a
mandrel
body fluid system (not shown). Mandrel body fluid system typically includes at
least one fluid
source and/or vacuum source, and one or more fluid supply lines coupled to
mandrel fluid
openings 40e and located within mandrel shell 40a. Although a plurality of
fluid openings 40e
are illustrated, it will be understood with benefit of the present disclosure
that as few as one
fluid opening may be employed, especially in the fabrication of smaller
structures. The mandrel
body fluid system may also include one or more fluid supply manifolds and
control valves
located within mandrel support structure 40b for controlling flow of fluid
and/or vacuum to the
fluid openings 40e. The mandrel body fluid system and mandrel fluid openings
40e may be
collectively referred to as,a mandrel body or mandrel fluid system. As
described below, the
mandrel body fluid system is used for facilitating composite structure removal
from mandrel 40
following composite material placement.
Manifold shell 40a may be constructed of any material suitably rigid for
forming a tool
or base for fiber placement of composites including, but not limited to
aluminum, steel, or
composite materials. Typically, manifold shell 40a is constructed of aluminum.
Aft mandrel
support structure 40b and forward mandrel shaft 40c are constructed of any
materials suitable
for supporting mandrel shell 40a, such as from a fuselage mandrel sling 40f as
shown in FIG. 3.
Aft manifold support structure 40b and forward mandrel shaft 40c are also
typically configured
to mate with other machine components used in the manufacture of a composite
structure, as
described below.
Although FIG. 3 shows a mandrel body of appropriate shape for forming an
aircraft
fuselage composite structure, it will be understood with benefit of this
disclosure that mandrel
bodies having shapes suitable for forming other aircraft fuselage designs, as
well as other
aircraft and non-aircraft components may also be used. It will also be
understood that other
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mandrel construction designs may be employed. For example, a collapsible
mandrel may be
employed with any aspect of the disclosed method and apparatus in the
fabrication of an aircraft
fuselage component (that is, a mandrel for forming both forward and aft
sections of the fuselage
in one piece).
In the embodiment of FIG. 3, fuselage mandrel body 40 is typically about 34
feet long.
Forward mandrel shaft 40c and aft mandrel support feature 40b are configured
to fit into
respective tail stock and head stock mounts of a fiber placement machine. In
this way mandrel
body 40 is placed in a fiber placement machine in a manner similar to mounting
in a lathe.
During fabrication of the composite structure, mandrel body 40 is rotated back
and forth as the
payoff head of a fiber placement machine applies a band of prepreg strands at
the selected
orientation and location.
Mandrel Body Fluid System
As described above, one embodiment of mandrel body 40 includes a fluid
floatation
system built into mandrel shell 40a to provide a means of pumping fluid
between the mandrel
and bladder during the bladder/caul sheet removal process. In a typical
embodiment, there are a
series of fluid holes or openings 40e drilled through mandrel shell 40a in a
pattern distributed
over the entire surface of mandrel 40a as illustrated in FIG. 3. Typically,
fluid holes 40e are
circular, having a diameter of from about 1/16 inch to about 1/4 inches and
are connected to
1/8 inch to 1 inch fluid supply lines disposed within mandrel shell 40a. In a
most typical
embodiment, approximately 250 circular holes 40e of about 1/8 inch diameter
are connected
with 1 /8 inch fluid supply lines that are manifolded. together in a way such
that holes 40e may
be selectively coupled to a vacuum source or pressurized fluid source, such as
pressurized air or
another suitable pressurized gas such as nitrogen, or a suitable liquid such
as soapy water
(which helps reduce friction). Fluid supply lines, fluid supply manifolds, and
fluid control
valves are constructed of suitable materials known in the art, typically those
capable of
withstanding pressures of from about 50 psi to about 500 psi. It will be
understood with benefit
of this disclosure that the mandrel body fluid system may comprise a greater
or lesser number
of openings than that described above. In addition, it will also be understood
that openings 40e
may be of shape and size other than that described above, including, for
example, openings that
are oval, rectangular, slot-shaped, etc. As described below, a vacuum source
may be used
CA 02598765 2007-09-11
ID
during fiber placement to secure a bladder tightly against the mandrel surface
so that it cannot
move relative to the mandrel, while pressurized fluid is typically used to
facilitate removal of a
composite part and bladder from mandrel body 40 after fiber placement.
Bladder and Caul Sheet
As described below, one embodiment of the disclosed method and apparatus
employs a
bladder with integral caul sheet section(s) that advantageously serves
multiple functions. These
functions include providing an inner seal for curing fiber placed parts,
providing a smooth outer
surface for forming smooth fiber placed part inner surfaces, and providing
locating features for
critical details, such as openings and other features formed in a fiber placed
part structure such
as an aircraft fuselage. Although typically employed in combination with caul
sheet surfaces to
form an integral bladder/caul sheet, it will be understood with benefit of the
present disclosure
that embodiments of the disclosed method and apparatus described herein may be
practiced
utilizing, among other things, a bladder without caul sheet sections, a
bladder in conjunction
with separate (non-integral) caul sheet sections, or a bladder in conjunction
with different sets
of interchangeable or detachable caul sheet sections. A bladder with or
without integral or non-
integral caul sheet sections may be employed with automated fiber placement
methods or with
various other processes, including but not limited to, fiber placement,
filament winding or hand
layup.
In one embodiment of the disclosed method, a flexible membrane or bladder is
employed between mandrel shell 40a and a fiber placed body formed by fiber
placement. This
flexible membrane serves several purposes, among which include forming an
intermediate
material to facilitate removal of an uncured fiber placed body (or composite
layup) part from
mandrel 40. The membrane also functions to seal the inside of a structural
composite from the
atmosphere within an autoclave and to transmit pressure (or compressive force)
uniformly to the
uncured composite laminate body during curing, as described below.
FIG. 4 illustrates one embodiment of a flexible membrane device used in the
disclosed
method and referred to as an integral bladder/caul sheet 42. As shown in FIG.
4. a flexible
membrane or bladder 43 is typically tubular and is designed to have a shape
and dimensions
complementary to the outer surface of a tubular shaped mandrel shell 40a so
that it may be
CA 02598765 2007-09-11
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indexed and fit directly on the surface of the mandrel shell 40a like the
finger of a glove.
Indexing is typically accomplished by providing matching holes in the
bladder/caul sheet 42
and mandrel 40. The bladder 43 is slipped over the mandrel 40 and the matching
holes aligned.
A tooling pin is then inserted through these holes and left there until the
bladder/caul sheet 42
is removed from the mandrel 40. There are typically two pins provided at the
aft end and two
pins provided at the forward end. Although mandrel shell 40a is depicted
having a cylindrical
shape, it will be understood with benefit of the present disclosure that a
bladder 43, caul sheet
section/s 44, and/or integral bladder/caul sheet 42 may be employed with
mandrels having
virtually any shape suitable for forming composite structures, including, for
example, square,
rectangular, oval, elliptical, irregular etc.
Typically, bladder 43 is constructed of butyl rubber, although it may be
constructed of
any suitably stretchable and resilient material including, but not limited to,
silicon rubber, nitrile
rubber, nylon film, or other elastomers. Also shown in FIG. 4 are caul sheet
sections 44. Caul
sheet sections 44 are typically relatively thin sheets of material that are
placed between the
bladder 43 and the uncured structural composite layup in order to smooth out
the interior
surface of a fiberplaced body that would otherwise conform to the surface of
bladder 43. Caul
sheet sections 44 may be any suitably smooth and rigid surface for creating an
uncured fiber
placed structure layup having a substantially smooth inner surface for
secondary bonding of
details. Typically, each caul sheet section 44 comprises a carbon/epoxy
laminate bonded to
bladder 43. However, caul sheet sections 44 may be constructed of any other
suitably smooth
and rigid or hard-surfaced materials including, but not limited to, fiberglass
and aluminum
sheet. It will be understood with benefit of this disclosure that any number
of caul sheet
sections may be employed, including one section or more than two sections. It
will also be
understood that a caul sheet section 44 may be coupled or bonded to a bladder
43 in any
suitable manner, including by pre-fabrication and attachment with adhesive, or
by laying up
uncured rubber and uncured prepreg fabric in a mold and then co-curing them
together. As
mentioned above, however, a bladder 43 may be employed with no caul sheet
sections, or with
non-integral or interchangeable caul sheet sections in the practice of the
disclosed method and
apparatus.
CA 02598765 2007-09-11
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Integral bladder/caul sheet 42 is configured to extend forward and aft of a
fabricated
composite structure so that it may be sealed directly to clamshell molds which
are used during
cure of the composite fiber placed body part as described below. As shown in
FIG. 4 in this
embodiment caul sheet sections 44 are typically carbon/epoxy laminates that
cover most of the
surface of rubber bladder 43 except for expansion spaces 46 (in this case,
located at the top and
bottom center lines of the mandrel body) and at end spaces 47 (located at the
forward and aft
ends of bladder 43). Expansion spaces 46 are zones where there is only rubber,
thereby
providing an expansion joint that allows a fiber place shell to expand out
against clam shell
molds during a curing process. In addition to expansion joints 46 located at
the center line areas
of the bladder 43, additional expansion spaces 46a may be positioned where
local contour
changes occur. In this embodiment, expansion spaces 46 and 46a are about 12
inches in width,
however, it will be understood with benefit of the present disclosure that
spaces having any
width suitable for allowing expansion to occur may be employed. Because the
entire inner
surface of rubber bladder 43 is a continuous layer of rubber that mates
tightly and securely with
the outer surface of mandrel 40, the bladder/caul sheet 42 provides rubber and
carbon/epoxy
layers laminated together to provide a continuous vacuum impervious shell.
In one embodiment, caul sheet sections 44 are located, laid up, and bonded to
the outer
mold line ("OML") surface of bladder 43 to provide a smooth bonding surface
for secondary
bonding of details and other assemblies to the inner mold line ("IML") of
fuselage shells. By
employing integral caul sheet sections 44, undesirable variations in a fiber
placed part that may
be formed in manufacturing processes employing only a bladder may be avoided.
These
variations may be caused in any soft area, such as where two pieces of core
unite so that the soft
and flexible surface of a conventional bladder tends to deform into the space
between the pieces
of core. thereby spreading them apart. In contrast to the relatively soft
surface of a bladder, the
relatively hard surface of caul sheet sections 44 does not tend to deform into
soft areas of a part,
instead tending to bridge across such soft spots.
Fabrication of bladder 43 and caul sheet sections 44 are typically performed
in caul
sheet molds as follows. The bladder outline surface is defined as a surface
concentric to the
OML but at a distance inside the OML equal to the composite part thickness
plus an allowance
for radial expansion. In the typical embodiment, the composite part thickness
is about 0.81
CA 02598765 2007-09-11
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inches and the expansion allowance is about 0.100 inches. In one embodiment,
bladder/caul
sheet 42 may be constructed using intermediate tools. First, a pair of solid
convex forms are
machined to the bladder/caul sheet, one representing the left side of the
fuselage, the other the
right side. On these forms, graphite epoxy shells are hand laid-up and cured.
Finally, the two
shells are mated together and the bladder/caul sheet laid-up and cured inside
them. Typically,
both the caul sheet and bladder materials are initially uncured and laid-up
simultaneously.
During the curing process, typically performed in an autoclave, the materials
bonded together to
form an integrated laminate (cure temperature of 350 F, cure pressure 100
psi). During
fabrication of caul sheet sections 44, scribe lines defining locations of
features such as
windshields, window and door locations may also be applied. Machined pads are
bonded to the
caul sheet surface in areas needed for attaching staged detail locators for
features such as cabin
window frames, cabin door frames, windshield frames, access ports, landing
gear doors, etc.
Positive locators (removable) may be provided, for example, for the following
staged or
precured details: windshield frame, window frames, cabin door frames,
emergency exit door
frames, etc. The outer surface of caul sheet sections 44 typically have
reference scribe lines
defining ply orientation of first ply down which is hand applied. As described
further below,
these machined pads are designed to fit into recesses in locator plugs that
may be secured to the
pads,and caul sheet section with any suitable securing device such as screws.
The locator plugs
are used to position the precured frames and they fill all the space between
the bladder/caul
sheet and the clam shell molds (two-piece) that is left open by the frame. A
removable spacer
or plug is provided to occupy the space where a splice ring will be installed
when the forward
and aft fuselage sections are joined.
Fiber Placement Devices
In the practice of the disclosed method, devices or machines suitable for
fiber placement
are typically employed to reduce labor and material waste. Any machine
suitable for fiber
placement may be used. Typically, a fiber placement machine employing a
multiple axis
numerical control ("NC") system is employed. Such a machine also typically has
a head stock
and tail stock for receiving forward mandrel shaft 40c and aft mandrel support
structure 40b of
mandrel 40. For example, a "CINCINNATI MILACRON VIPER FIBER PLACEMENT
MACHINETM using NC data generated from "CATIATM" model software is employed.
Besides
"CAT1A", any other software suitable for designing and storing numerical data
for NC
CA 02598765 2007-09-11
17 -
machining parts may be employed. In one embodiment of the disclosed method, a
fiber
placement machine capable of selective application of 24 rolls of 1/8 inch
wide and 5/1000 inch
thick slit tape or tow is employed. This allows. single pass application of a
tape with a
maximum width of 3 inches and a minimum width of 1/8 inch. Typically, a
machine having a
payoff head capable of cutting and restarting individual strands of tape is
employed. The payoff
head has a conforming roller (typically having I1 individual segments) for
application of tape
prepreg to a mandrel. Such a machine is also typically capable of controlling
the temperature of
the tape and roller pressure. In a typical embodiment the machine operates
along seven
different axes to allow the payoff head to follow the contour of a mandrel,
keeping it normal to
the mandrel surface during application of prepreg tape. Proper orientation of
the payoff head
(or tow placement head) in relation to the mandrel is controlled by a
computer. Such a machine
is also capable of controlling temperature and head roller pressure during
application of the tape
in order to control quality of the composite material (such as to eliminate
voids).
In the practice of the disclosed method, a fiber placement machine is
typically provided
with a head stock to receive the aft mandrel support structure 40b of a
mandrel 40 and the tail
stock to receive the forward mandrel shaft 40c of a mandrel 40. The head stock
and tail stock
are configured to allow the mandrel to be rotated back and forth and at
various speeds during
fiber placement. Rotation of the mandrel as well as other aspects of the fiber
placement system
are typically controllable by an NC system, such as one using "ACCRAPLACETMII
control
software (typically available from Cincinnati Milacron with purchase of one of
their machines),
in conjunction with design software, such as "CATIA".
Although a typical fiber placement machine and NC system have been described
above,
it will be understood with benefit of the present disclosure that any fiber
placement machine
suitable for application of prepreg tow to a mandrel may be employed. Further,
the structure
described herein could be fabricated entirely by hand layup techniques and/or
designed using
non-computerized manual design techniques known to those of skill in the art.
Clam Shell Molds
For molding and curing of a composite component, such as a fuselage, a clam
shell
mold 31 is typically employed as shown in FIG. 5 (in this case, for the
forward section of an
CA 02598765 2007-09-11
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aircraft
fuselage that is about 25 feet in length). Clam shell mold 31 typically
includes two
mold halves 30 and 32 in which an uncured composite body (such as a fuselage)
may be placed
to cure. After curing, the cured composite part may be removed from the mold,
as shown in
FIG. 6. FIG. 7 shows clam shell mold halves 36 and 38 for aft composite
fuselage section 34a.
In the illustrated embodiment, clam shell mold halves 30, 32, 36, and 38
provide a rigid
form upon which soft, pliable, uncured composite materials may be placed and
supported
during a curing process, which is typically carried out under controlled and
elevated temperate
and pressure. Inner surfaces 30a, 32a, 36a, and 38a of the clam shell molds
are developed to
provide a surface for molding the outer loft surface of a fuselage, and
therefore need to be as
accurate and smooth as possible. Advantageously, split clam shell molds as
those shown in
FIGS. 6 and 7 allow composite parts, such as fuselage shells, to be made
without an axial
splice. FIG. 7. illustrates how a clam shell mold may be split along the top
and bottom vertical
center lines in order to remove a part after cure.
Typically, clam shell mold halves 30, 32, 36, and 38 are made of carbon/epoxy
tooling
fabric. This material is desirable for this embodiment because it has
substantially the same
thermal coefficient of expansion as fuselage shells 34 and 34a which are
typically made of
carbon/epoxy. In manufacture, clam shell molds 30, 32, 36, and 38 are
typically laid up on
aluminum molds which have been NC machined to the contour of a fuselage outer
mold line,
although other mold materials are possible. One example of such aluminum molds
39a and 39b
is shown in FIG. 8. Using aluminum molds 39a and 39b, carbon/epoxy clam shell
molds 30,
32, 36, and 38 may be manufactured using graphite epoxy tooling prepreg which
cures at a
lower temperature (typically, about 120 F) than prepreg material for
manufacturing fuselage
composite sections. Clam shell molds 30, 32, 36 and 38 are then laid up on
aluminum molds
39a and 39b and cured in an autoclave. It will be understood that clam shell
molds may also be
made of any other material suitable for molding, such as glass fabric/epoxy,
that is compatible
with materials of the composite part being built.
In operation, respective halves of a clam shell mold are typically held
together with
clamps during molding and curing of a composite part section. These mold
halves are typically
configured to create a clearance of between about 0.010 inches and about 0.100
inches, most
CA 02598765 2007-09-11
L1 -
typically about 0.060 inches, between the interior surface of the mold halves
and exterior
surface of an uncured and unexpanded fiber placed part. Seals are typically
provided for sealing
the integral bladder/caul sheet combination 42 of FIG. 4 around a fuselage
composite section
34 to provide a sealed, leak-free system. In this way, a vacuum may be
maintained between
bladder/caul sheet 42 and the clam shell mold halves during molding and
curing, as described
below. For example, smooth sealing surfaces are provided at the forward and
aft ends of the
bladder/caul sheet 42. Similar surfaces are provided on the forward and aft
faces and sides of
the clam shell halves.
For the purpose of expanding the fiber placed shell the bladder/caul sheet and
clam
shells are sealed together by placing a_ sticky sealant tape on the smooth
seal surfaces and
covering the area between the surfaces with a flexible membrane that is
impermeable (usually a
commercially available nylon film). A vacuum pump is plumbed to a port{s)
mounted on the
clam shells and the air, or then gaseous or fluidized material evacuated from
the sealed volume
between the bladder/caul sheet and clam shell molds. When this is done, the
atmospheric
pressure outside the sealed volume is higher than inside and tends to push the
bladder/caul sheet
and clam shells together. Since the bladder rubber area is stretchable and the
fiber placed
fuselage shell is in an uncured state at this point, the external pressure
tends to expand the
bladder/caul sheet out towards the rigid clam shells. Depending of the
efficiency of the vacuum
pump the differential pressure may be as much as about 15 pounds per square
inch of area
(although less vacuum is acceptable). The net effect of this pressure is to
expand the caul sheet
away from the mandrel surface thus allowing the mandrel to be extracted. The
sealing
materials, sticky sealant tapes and nylon film are disposable and are usually
replaced after each
use.
Reusable or Permanent Bag
In another embodiment, when production rates are high enough it may be more
economical to accomplish the sealing process described in the previous
paragraph using a
permanent or reusable bag 100 as shown in FIG. 15. Reusable bag 100 is
typically placed over
bladder/caul sheet 42 after composite placement. For the reusable bag 100, the
sticky sealant
tape is replaced with bladder/caul sheet seal halves 108 and clam shell seal
halves 108a that are
bonded to bladder/caul sheet surfaces 42a and clam shell mold surfaces 33,
respectively, as
CA 02598765 2007-09-11
- L!
illustrated in FIGS. 19 and 22. As shown in FIGS. 15 and 16, mating bag seal
halves 106 and
106a for mating with respective seal halves 108 and 108a are bonded to
membrane material
104, which may be of any suitably flexible material such as rubber. Typically,
membrane
material 104 is a heavier flexible membrane (most typically fitted silicone
rubber sheet having a
thickness of about 0.125 inches thick). Although one particular embodiment of
reusable seal
halves 106, 106a, 108, and 108a are illustrated, these seal halves may be any
sealing elements
suitable for sealing bladder/caul sheet surfaces 42a to clam shell mold
surfaces 33. Typically,
reusable seals are provided for forming such a sealing function in a manner
similar, for
example, to the seals on a resealable sandwich bag.
In a most typical embodiment, bag seal elements 106 and 106a may be. any
suitable
reusable seal halves, such as silicon seals available from Bondline Products.
Mating seal halves
108 (such as from Bondline Products), may be provided (or bonded) at
appropriate mating
points on bladder/caul sheet surfaces 42a to form forward bladder/caul sheet
seal loop 109 and
aft bladder/caul sheet seal loop 110 as shown in FIG. 18. These seal loops
make a complete
loop around the forward and aft sections of the bladder/caul sheet surface
42a. In a similar
fashion, clam shell mold halve surfaces 33 may be provided (or bonded) with
mating seal
halves 108a (such as from Bondline Products) to form clam shell seal loops 112
as shown for
one clam shell mold half in FIG. 21. These seal loops make complete loops
around the outside
of the clam shell mold surfaces 33 of each mating clam shell mold half such
as, for example, by
bonding to bath tub flanges 33b as shown in FIGS.. 21, 22, 24 and 25. A cross
section of clam
shell sealing loops 112, showing clam shell seal halves 108a is illustrated in
FIGS. 22, 24 and
25. FIG. 25 shows a cross section of an assembled bag 100, with membrane 104
sealed to clam
shells 30 and 32 using seal halves 106a and 108a. In this way the bag 100
protects against
potential leaks in area 33a between the clam shell mold half mating surfaces
33.
A cross section of bag sealing halves 106a and membrane 104 for forming a seal
with
sealing loopsõ 112 is illustrated in FIGS. 17, 24, and 25. A cross section of
bladder/caul sheet
sealing loops 109 and 110, showing bladder/caul sheet sealing halves 108 is
shown in FIG. 19. A
cross section of bag sealing halves 106 and membrane 104 for forming a seal
with sealing loops
109 and 110 is illustrated in FIG. 20. Mating seal elements 106 and 108 may be
sealed together
so as to create a seal between bladder/caul sheet 42 and clam shell
CA 02598765 2007-09-11
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23
mold surfaces 33, as shown in FIG. 23. Although one embodiment employs seal
halves 106
and 106a with male seal profiles, and seal halves '108 and 108a with female
seal halves, these
profile relationships may be reversed or mixed in any manner desirable. In
addition, multiple
seal loops and seal loops employing greater than one row of seal halves may
also be employed,
as may reusable bags having a single unitary seal section or multiple membrane
and seal
sections. Further, it will be understood with benefit of this disclosure that
a sealing bag similar
to reusable bag 100 may be constructed to be disposable.
In another embodiment, vacuum within sealed clam shell molds may be provided
with a
clam shell mold vacuum system including vacuum pump and plumbing.
B. Manufacture of Composite Sections
FIG. 9 graphically illustrates fabrication of a composite part according to
one
embodiment of the disclosed method and apparatus. Fabrication typically begins
with an
integral bladder/caul sheet placed on a mandrel body (902) around which fibers
are placed
(904). Next, clam shell molds are placed around the composite part (906 - 910)
and the mandrel
removed (912). The composite section part and clam shell molds are then
typically placed in an
autoclave for curing (914). After curing, the composite is removed from the
clam shell molds
(916) for further fabrication steps (such as, for example, those shown in
steps 918 - 920).
During manufacture, the steps of fiber placement through curing are typically
performed in a
clean room. Such a clean room is typically controlled for temperature,
humidity and
contaminants such as dirt, oils, etc. The following steps are illustrative of
one representative
embodiment of the disclosed method and apparatus. It will be understood with
benefit of the
present disclosure, that other embodiments are possible, including those
embodiments using
fewer, additional and/or differing materials and steps.
Fiber-Placement
In this embodiment, fabrication of a composite fuselage section on a mandrel
typically
consists of the basic steps of: (1) fiber-placement of an inner skin and
reinforcement plys; (2)
manual application of small reinforcement plies, pre-staged or C-staged close-
out frames for
windows and doors, and honeycomb core: and (3) fiber-placement of
reinforcement plys. Non-
crosslinked C-staged close-out frames are typically employed to facilitate
assembly with
CA 02598765 2007-09-11
24 -
uncured materials in the clam shell molds. However other staged frames, or no
frames at all are
also possible.
In this embodiment, a mandrel 40 with installed bladder/caul sheet 42 (as
shown in FIG.
10) is installed -into a fiber-placement machine such that forward mandrel
shaft 40c and aft
mandrel support structure 40b are received in tail stock and head stock,
respectively, of the
fiber-placement machine. Although an integral bladder/caul sheet is typically
employed, it will
be understood that fiber or composite material placement (by any suitable
method) may also be
accomplished using a bladder alone or a bladder in combination with non-
integral caul sheet
sections. Prior to fiber-placement, a vacuum is applied through the mandrel
fluid supply system
and openings 40e in mandrel 40 so that integral bladder/caul sheet 42 is held
tight to the
mandrel during fiber-placement. The surface of the bladder/caul sheet has been
coated with a
material, typically monocoat, that will not stick to the cured fuselage shell.
Next, a first ply of
inner skin is hand placed onto bladder/caul sheet 42. The first ply of the
inner skin is typically a
fabric prepreg of about .0085 inch/ply. with the fabric weave oriented at 45
to the axis of
rotation. The first ply fabric prepreg is typically vacuum bagged temporarily
to bladder/caul
sheet 42 so that it will adhere to it. Then carbon fiber/epoxy slit tape or
tow prepregs are placed
over the first ply to complete the inner skin 20 using the fiber-placement
machine. In this
embodiment, inner skin 20, core 24, and outer skin 22 are typically laid up in
the thicknesses
and dimensions as previously described. When slit tape is employed, it is
typically of about
0.125 inch from a parent tape manufactured to have about 136 g/m2 carbon
fiber. However,
other types and weights of slit tape or prepregs are possible. as are skin
constructions having
greater or lesser numbers of plies.
In this embodiment, a fiber-placement machine typically applies between l and
24
filaments of pliable uncured prepreg tape during placement. Each tape
typically has a thickness
about 511000 of an inch and a width of about 1 /8 inch. During fiber
placement, mandrel 40 and
the fiber placement roller head are moved in relation to each other so that
fiber filaments are
placed according to computer software (typically, "ACCRAPLACE" control
software and
"CATIA" design software). Typically, tow filaments or fibers are applied in
three basic
orientations relative to the longitudinal axis of an elongated composite part
(such as a fuselage,
so that 0 is oriented in a direction from forward to aft): in an axial
direction substantially
CA 02598765 2007-09-11
parallel to the longitudinal axis of the mandrel (most typically about 0 to
the longitudinal axis
of the mandrel), in a direction at an angle of between about 40 and about 50
(most typically
about 45 ) to the longitudinal axis of the mandrel, and in a direction roughly
perpendicular to
the longitudinal axis of the mandrel and coinciding with the outer diameter of
the mandrel (or at
an angle between about 85 and about 95 , most typically about 90 , to the
longitudinal axis of
the mandrel). This filament application orientation provides uniform strength
and structural
integrity to the finished composite part. In applying bands of fibers in the 0
and 45
directions, the bandwidth is typically narrowed by dropping one or more tows
as application
moves from larger diameter to smaller diameter sections of a part. Adjacent
bands are laid side
by side without overlaps to provide a constant thickness ply.
In a composite design including fibers or filaments that are oriented in
multiple
directions as described above, the filaments or fibers are typically placed so
that they form plies
of material that are symmetrical about the center of the thickness. This helps
ensure that
stresses are distributed uniformly so that the composite part won't tend to
warp during laminate
curing.
.During fiber-placement, precured laminate strips or stripping is typically
placed in the
expansion areas 46. and 46a which exist between separate caul sheet section 42
as shown in
FIG. 10. These precured laminate strips serve to act as a filler or bridge in
those areas where no
caul sheet material exists. Most typically these laminate strips are of a
thickness substantially
equivalent to the thickness of a ply of fabric. After fiber-placement, these
precured strips
become part of the composite section part.
In one embodiment of the disclosed method, the detailed design of the plies in
the inner
skin 20 is adjusted so that circumferentially directed fibers (those laid out
at about 90 ) are cut
and overlapped at opposite sides of a composite section ply to form a
continuous full cylinder
skin having two integral sections. Tow or tape plies placed in the direction
of the axis of
rotation are typically not overlapped since they offer little resistance to
circumferential growth,
however, any plies applied more than from about 10 to about 15 away from the
axis of
rotation should be overlapped in a similar manner. In the case of a composite
part. such as a
CA 02598765 2007-09-11
-~Z6-
fuselage, adjacent (or side by side) plies are typically laid out to form an
expansion area
oriented substantially parallel to the longitudinal axis of the part.
FIG. 11 is a cross-sectional representation of an inner skin ply 20x, showing
underlapping inner skin ply half 20a and overlapping inner skin ply half 20b.
Areas of overlap
(or circumferential expansion joints) are shown within circles and are
typically placed at
opposite sides of the composite inner skin shell ply 20f. These overlapping
areas provide a full
cylinder or continuous skin of fibers. These layers also allow for shell
expansion without
inducing substantial load or stretching into the fibers during curing, and
while at the same time
1.0 providing full structural strength after cure. For example, during cure
the prepreg material of
inner skin shell 20 is expanded outward toward the inner surface of clam shell
molds. This
occurs before the epoxy resin of the prepreg skin material has hardened or
cured. During
expansion, individual fibers of overlapping and underlapping sections of skin
ply halves 20a
and 20b are allowed to move or slide against or in relation to each other in
the direction
indicated in the circled areas 20c of FIG. 11, thus allowing expansion of
inner skin shell 20, for
example, against an inner surface of a clam shell mold. Similar overlapping
skin ply halves are
typically employed in the placement of an outer skin ply 22 for outer skin 22,
and for any
additional fiber ply skin layers present.
Advantageously, by providing discontinuous fiber segments juxtaposed in
relation to
each other so as to allow circumferential expansion, an uncured fiber placed
body having a full
cylinder continuous hoop skin may be expandable against the interior of a clam
shell mold
before and/or during curing. This expansion process tends to flatten out and
straighten
individual fibers, as well as to create a substantially uniform surface from
forced contact with
the mold interior surfaces. In addition, circumferential expansion of the
fiber placed body away
from a mandrel body serves to create clearance between the fiberplaced body
and associated
bladder, thereby facilitating removal of the body and bladder from the
mandrel. This expansion
occurs without substantial stretching or creation of residual stresses.
While FIG. II illustrates a inner skin ply having two overlapping expansion
joints 20c
and 20g at the top and bottom of skin ply 20x, it will be understood with
benefit of the present
disclosure that expansion joints 20c and 20g may be positioned at other
locations and utilized
CA 02598765 2007-09-11
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with other composite sandwich layers, symmetrically or unsymmetrically, around
the
circumference of skin ply 20x. It will also be understood that a varying
number of
circumferential expansion joints may be used including, for example, only one
overlapping
expansion joint, or three or more overlapping joints in a skin ply 20x.
In further detail for this embodiment, FIG. 11 is a cross-sectional view
looking down
longitudinal axis of one embodiment of a longitudinal composite part, such as
a fuselage
component. As shown in FIG. 11, underlapping fibers 20a are applied in the
circumferential
direction starting at 0 and wrapped 180 plus an overlap distance, and then
cut. Overlapping
fibers 20b for the second half of the ply are then started at 180 and wrapped
to 360 plus an
overlap distance, and then cut. Subsequent overlap areas 20c and 20g in
subsequent fiber plies
are rotated away from the 0 starting point far enough so the overlap areas do
not stack up on
one another (typically by about 5 , but may be any other angle suitable to
prevent expansion
area stacking) and form thickened areas in the shell.
In the practice of the disclosed method, an area of overlap having any
dimension
suitable for providing a circumferential expansion joint capable of allowing
expansion (as
described above) may be employed. However, an overlap of from about 1.0 inches
to about
1.75 inches, most typically of about 1.5 inches, is typically employed for
overlapping areas 20c
and 20g. That is, edges 20d of overlapping skin ply half 20b typically
overlaps edge 20e of
underlapping skin half 20a by from about 1.25 inches to about 1.75 inches,
most typically by
about 1.5 inches. Most typically, edges 20d and 20e form an area of overlap
that is symmetrical
about composite shell center axis 20f, as shown in FIG. 11. This overlap is
typically
accomplished by rotating the mandrel in one direction and placing or laying
out and then
cutting a fiber band to form underlapping skin half 20a between overlapping
areas 20c and 20g
as shown. The mandrel is then rotated in the opposite direction by the amount
of overlap
desired (typically about 1.5 inches). Then a new fiber band is initiated at
overlapping area 20g
and laid out to form overlapping skin half 20b by rotating the mandrel in the
first direction until
the circle is completed and the band overlaps underlapping skin half 20a at
overlapping area
20c.
CA 02598765 2007-09-11
28
When the shell is expanded the overlapped areas slide relative to one another
reducing
the length of overlap accordingly. When a prepreg is cured each ply overlap
area transfers load
from one side to the other making it act like a continuous path. Typically,
the cured overlap is
approximately one inch, although cured overlapping areas may be greater or
less than this.
Although FIG. - 11 illustrates a circumferential ply having two overlapping
areas, as many
overlap areas may be used in a circumferential ply as desired. When greater
than two
overlapping areas are employed, each additional overlapping area reduces the
amount of sliding
within the other overlapping areas during expansion.
In addition to the dimensions and overlapping configuration described above,
it will be
understood with benefit of this disclosure that the overlap of overlapping
areas 20c and 20g
may vary. In addition, the amount of overlap may vary from expansion joint to
expansion joint,
rather than being the same for each expansion joint. Thus, it may be possible
to have multiple
expansion joints in which one overlapping edge 20d overlaps underlapping edge
20e by greater
or less than the amount of overlap in another expansion joint. In addition, it
may also be
possible to form mixed overlapping shell segments in which one edge of a ply
segment overlaps
an adjacent ply segment and in which the other edge of the same ply segment
underlaps another
adjacent ply segment. It will also be understood that overlapping.
circumferential expansion
joints 20c may be employed in a single layer or skin, or may be simultaneously
employed in
individual layers or skins of a multiple layer or a multiple skin sandwich
composite section.
The expansion mechanism may be enhanced by heating tow filaments to reduce
resin
system viscosity and interfiber friction. This may be done, for example, by
heating the outer
surface of a heated mandrel (such as with a mandrel heating system) during the
expansion
process. Likewise, the interior of a clam shell mold may be heated during the
expansion
process (such as with a mold heating system). In some cases, both mold and
mandrel may be
heated. For an uncured fiber placed part, sufficient heat should be applied to
heat the tow
filament resin to a temperature of between about 100 F and about 150 F,
allowing a vacuum of
between about 10 psi and about 15 psi to expand the part against the inner
surface of a mold
shell.
CA 02598765 2007-09-11
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In areas of a composite structure requiring reinforcement, such as a wing
attachment
point of an aircraft fuselage, local reinforcements may be created during
fiber placement of
inner and outer skins, and/or may be hand placed between the fiber placement
of the skins.
Typically, for reinforcement areas of greater than or equal to about 4 inches
in length, an
integral local reinforcement is created in inner and/or outer skins' using the
fiber placement
machine. This is typically done by placing additional fibers in the
reinforcement areas to create
an area of the skin that is thicker than the remainder of the skin, typically
from about 10% to
about 200% thicker, and which tapers down in thickness in an outward
direction. Examples of
such larger reinforcement areas include areas for wing attachment, areas where
forward and aft
sections of a fuselage join, and, at places where various objects are mounted
For those
reinforcement areas of less than about 4 inches in. length, local
reinforcements are typically
hand placed using fabric pieces before and/or after placement of plugs, frames
and core.
Following fiber placement of inner skin 20 other composite shell layers may be
applied
including, for example, core materials and other prepreg carbon fiber
materials. Typically, a
honeycomb core 24 made from "NOMEX" or other suitable material is hand placed
on inner
skin 20, as shown in FIG. 2. Prior to hand placing honeycomb core 24, locating
plugs 200 and
frames 202 (typically C-staged frames) corresponding to openings in the
composite shell 10 are
typically hand placed as shown in FIGS. 26 and 27. The plugs 200 and frames
202 are
typically placed before core 24 to take advantage of the flexibility of the
core 24 in fitting it to
the frames. However plugs 200 and/or frames 202 may also be placed
simultaneously or after
placement of core 24. Such openings include doors, windows, avionics access
hatches, landing
gear door hatches, etc. Typically, frames 202 are hand placed and secured by
screws or bolts
(or other suitable securing devices) 206, onto features or positioning pads
204 present on the
bladder/caul sheet 42 underlying inner skin 20 as previously described, and as
shown in FIGS.
26 and 27.
As previously mentioned, conical positioning pads 204 are left exposed through
openings for plugs 200 and frames 202 left in inner skin 20, typically by
programmed NC
instructions to a fiber-placement machine. During fiber placement, inner skin
20 is placed so
that openings in the skin exist at the conical pads 204 bonded to caul sheet
section 44. This is
accomplished by cutting each tow at one edge of each pad 204, skipping over
the pad. and
CA 02598765 2007-09-11
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reinitiating it at the other edge of the pad. As previously described, these
conical pads are
configured to be received in corresponding openings defined in locating plugs
200, which are in
turn secured to caul sheet section 44 with securing devices 206. Locating
plugs are typically of
a thickness approximately equal to the wall thickness of the completed three
layer sandwich
composite shell, and may be constructed of the same material as the shell,
most typically of
carbon epoxy sandwich structure.
Once plugs 200 and frames 202 have been attached to pads 204 of a caul sheet
surface
44, honeycomb "NOMEX" core 24 may be hand placed over inner skin 20 in such a
way that
the preplaced frames 202 create openings in the core as shown in FIG. 27.
Once honeycomb core 24 is in place on inner skin 20, outer skin 22 may be
formed over
core 24 using a fiber-placement machine or other methods. Outer skin 22 is
typically formed
using fiber placement machine in a manner similar to inner skin 20, including
the use of
circumferential expansion joints and leaving openings corresponding to frames
202 and/or other
features. Following fiber-placement of outer carbon fiber skin 22, other
layers may be placed.
For example, a fabric layer containing metal filaments for lightning
protection is typically
applied. Plugs 200 are typically removed following curing of a composite part.
Typically, hand placement steps mentioned above are accomplished with the
mandrel 40
positioned in the machine or in a mandrel transportation dolly 40g as shown in
FIG. 12,
although any other suitable method may be employed.
Molding and Curing
In a typical embodiment of the disclosed method, once fiber-placement (by any
suitable
method) and hand layup of the composite shell layers is completed. Mandrel 40
and uncure
fiber placed composite shell 10 is then positioned or installed within clam
shells 30 and 32 as
shown in FIG. 14. This is typically accomplished by transferring the mandrel
40 and shell 10
to a mandrel extraction device, such as a mandrel extraction fixture 40h (as
shown in FIG. 13).
The lower clam shell half 32 is then rolled under mandrel 40 and the upper
clam shell half 30
placed on top of (or mated to) lower clam shell half 32. In the most typical
embodiment, a
vacuum bag (which may be disposable or a reusable permanent bag 100 as
previously
CA 02598765 2007-09-11
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described) is installed, sealing bladder/caul sheet 42 to the clam shells 30
and 32, and the sides
of the clam shells 30 and 32 as previously described. The vacuum holding
bladder/caul sheet
42 to mandrel 40 is then released. Next, a vacuum is drawn between
bladder/caul sheets 42 and
clam shell molds 30 and 32, typically using a vacuum pump connected to
openings or ports in
the clam shell molds. Vacuum pump and vacuum port plumbing may be accomplished
using
methods and devices known to those skilled in the art. Typically a vacuum of
about 15 psi is
applied. This vacuum, in conjunction with atmospheric pressure outside the
sealed bladder to
mold annulus, serves to draw integral bladder/caul sheet 42 out against the
inner skin of
uncured composite fiber placed shell 10, thereby expanding it out against the
interior surfaces
30a and 32a of clam shell molds 30 and 32. At the same time, the vacuum serves
to expand the
composite shell 10 and bladder/caul sheet 42 away from the mandrel surface. As
described
previously, heating may be used if desired to assist the expansion.
Expansion serves to create a clearance of between about 0.4 inches and about
0.10
inches, most typically about 0.060 inches, between the outer surface of
mandrel 40 and the
inside surface of the bladder 43. This clearance is typically somewhat larger
than the initial
clearance between the interior surfaces 30a and 32a of the clam shell molds
and the exterior
surface of the fiber placed shell 10, because uncured fiber placed composite
shell 10 is not
completely compacted until the autoclave curing process is complete. For the
same reason, it is
also typically larger than the initial clearance between the outer surface of
mandrel 40 and the
inner surface of bladder 43. It will be understood with benefit of this
disclosure that the
clearance may not be uniform and may therefore not be present around the
entire outer surface
of mandrel 40.
Next, mandrel 40 is removed from clam shell molds 30 and 32, leaving uncured
shell 10
and integral caul sheet/bladder 42 within the clam shell molds. It has been
found that even with
a vacuum established between the clam shell molds and integral bladder/caul
sheet 42, mandrel
40 is typically not freely detachable from the interior of the bladder/caul
sheet 42. Typically,
there is at least some adhesion or friction between mandrel 40 and
bladder/caul sheet 42 during
mandrel removal. Forces resisting mandrel removal typically are proportional
to the surface
area of the mandrel and any irregularities in thickness or shape of the
tooling components that
may exist. Therefore, in order to facilitate removal of the mandrel, fluid
pressure may be
CA 02598765 2007-09-11
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applied to the mandrel body fluid system. This produces fluid flow through
openings 40e in the
surface of mandrel 40, thereby tending to provide a fluid bearing that tends
to lift or separate the
inner surface of bladder 43 from the outer surface of mandrel shell 40a,
allowing uncured
composite fiberplaced shell 10 and integral caul sheet/bladder 42 to be
floated off mandrel 40.
Typically, from about 10 psi to about 40 psi of air is applied to accomplish
mandrel removal.
During this process, air typically leaks out at the ends of the bladder/caul
sheet 42.
Removal of mandrel 40 from caul sheet/bladder 42 is typically facilitated
using mandrel
removal fixture 40h illustrated in FIG. 13. Mandrel removal fixture (or
platform) 40h is
designed to be capable of supporting mandrel 40 and fiber placed composite
shell 10 (typically
on their side) on one end in cantilever fashion within clam shells 30 and 32
using aft mandrel
support structure 40b. Mandrel removal fixture 40h is designed so that the
longitudinal axis of
the mandrel 40 and accompanying fiber placed composite shell 10 may be
adjusted at pivot
point 50 so that they are parallel with the floor 60 for removal. As shown in
FIG. 13, in one
embodiment clam shell mold half 32 fits on a cart 62 that is mounted on rail
64 that is aligned
with the axis of the mandrel 40, composite shell 10 and clam shell molds 30
and 32. Typically,
alignment of mandrel 40 and clam shells 30 and 32 on extraction cart 62 are
adjusted if
necessary to ensure that the assembly of bladder/caul sheet 42, composite
shell 10. and clam
shells 30 and 32 may be rolled away without binding. When the components are
properly
aligned, fluid pressure (typically compressed air) is applied to the mandrel
body fluid supply
system to provide floatation between the mandrel 40 and bladder/caul sheet 42.
At this point,
the assembly including the bladder/caul sheet 42, composite shell 10, clam
shells 30 and 32,
and cart 62 are rolled away leaving mandrel 40 on extraction fixture 40h.
Next, uncured fiber
placed composite shell 10 and clam shell molds 30 and 32 are transferred
(typically using an
autoclave cure cart) into an autoclave for curing, as described below.
Advantageously, a second
integral bladder/caul sheet 42 may be applied to mandrel 40 and mandrel 40 may
be returned to
the fiber placement machine for placement of the next composite shell.
Next, fiber placed composite shell 10 and clam shell molds 30 and 32 are
transferred,
typically with the same about 15 psi vacuum drawn between integral
bladder/caul sheet 42 and
clam shells 30 and 32 as described above, into a controlled environment for
curing and
consolidation of laminates. Typically a controlled curing environment is
provided by an
CA 02598765 2007-09-11
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autoclave filled with a gas, such as nitrogen or carbon dioxide, that won't
support combustion.
Once placed in the autoclave, the clam shell vacuum system is typically
attached to the
autoclave vacuum system and the autoclave door sealed shut. Curing of the
composite fuselage
shell 10 is then accomplished by increasing nitrogen pressure in the autoclave
to a prescribed
level (typically -from about 40 psi to about 100 psi, most typically about 45
psi with a
honeycomb sandwich structure), and by increasing the autoclave temperature to
between about
340 F and about 360 F, most typically about 350 F in a prescribed manner
{i.e., heating rate
and schedule) to complete chemical reaction of the particular resin system
(typically epoxy for
this embodiment). The bladder/caul sheet to clam shell vacuum may be released
after autoclave
pressure reaches about 15 psi or above. To obtain good part expansion and a
good conformal
outer surface on a fiber placed part, a total differential of between about 40
psi and about 50 psi,
most typically about 45 psi, is typically employed between the clam shell mold
vacuum (within
the bladder/clam shell space) and the autoclave pressure (outside the
bladder).
During the curing process, among other things, resin is hardened and any
trapped air is
eliminated from between fiber plies or laminates . Typically, temperature
within the autoclave
is increased to between about 340 F and about 360 F, most typically about 350
F for curing.
However, it will be understood the specific combination of time within the
autoclave,
temperature, and pressure may be specified or varied as necessary to develop
the desired quality
laminate within a chosen resin system.
In this aircraft fuselage embodiment, an autoclave molding process, with its
flexible
membrane, is typically desirable to other molding techniques (such as
compression molding).
This is because the parts are large and/or may have complex contours. In
addition, the use of a
flexible membrane allows the pressure in the autoclave to act uniformly on the
surface of a
composite part, and allows the most uniform temperature distribution
throughout the part during
cure. However, although an autoclave process with flexible membrane has been
described
above for molding and curing of a fiber placed composite shell, it will be
understood with
benefit of the present disclosure that other molding and curing processes may
be successfully
employed to mold and/or cure composite shells or other composite parts
produced using the
disclosed method. It will also be understood that the clam shell molds,
integral bladder/caul
sheet, mandrel, associated fluid systems, and other aspects of the disclosed
method and
CA 02598765 2007-09-11
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apparatus may be employed individually or together to mold and cure composite
parts formed
using other fabrication techniques.
Bladder/Caul Sheet Removal and Re-Installation
Once composite fuselage shell 10 has been cured within the autoclave, it is
removed
along with clam shell molds 30 and 32 from the autoclave. The vacuum is then
released from
the clam shell vacuum system, and the bladder/caul sheet 42 carefully removed
from inside of
the fuselage shell 10. In one embodiment, a bladder/caul sheet removal fixture
is typically used
to remove bladder/caul sheet 42 from fuselage shell 10. The bladder/caul sheet
removal fixture
includes vacuum or suction cups mounted on double acting air cylinders and
extending radially
in position such that they may be deployed within the interior of an integral
bladder/caul sheet
42. Once positioned within the bladder/caul sheet 42, the double acting air
cylinders are
energized so that vacuum cups are placed against the inner surface of bladder
42. Next, the
bladder/caul sheet removal fixture is removed along the longitudinal axis of
fuselage shell 10
such that bladder/caul sheet 42 is removed by the vacuum cups from fuselage
shell 10 and then
slid out. A bladder/caul sheet installation fixture is typically fitted over
the outside surface of
an integral bladder/caul sheet 42 and expanded so that it may be slid back
onto a mandrel for
reuse.
It will be understood with benefit of the present disclosure that the
disclosed method and
apparatus may be employed without the use of bladder/caul sheet removal and/or
installation
fixtures, and that a bladder/caul sheet may be installed on a mandrel or
removed from a
composite shell using any suitable method, including methods that are
destructive to the
bladder/caul sheet 42 (in which case it may not be reused). It will also be
understood that the
methods and apparatus described above may be employed with a bladder having no
integral
caul sheet.
Following removal of bladder/caul sheet 42 from composite shell 10, clam shell
mold
halves 30 and 32 are separated and composite shell 10 removed for further
finishing. Such
further finishing may include, for example, attachment to other composite
shell components,
such as for example another composite section using methods known to those of
skill in the art.
CA 02598765 2007-09-11
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While the invention may be adaptable to various modifications and alternative
forms,
specific embodiments have been shown by way of example and described herein.
However, it
should be understood that the invention is not intended to be limited to the
particular forms
disclosed. Rather, the invention is to cover all modifications, equivalents,
and alternatives
falling within the spirit and scope of the invention as defined by the
appended claims.
Moreover, the different aspects of the disclosed structures and methods may be
utilized in
various combinations and/or independently. Thus, the invention is not limited
to only those
combinations shown herein, but rather may include other combinations.
