U.S. patent application number 15/153216 was filed with the patent office on 2016-11-17 for method of forming composite structures.
The applicant listed for this patent is AIRBUS OPERATIONS LIMITED. Invention is credited to Steven EVANS, Jonathan PRICE.
Application Number | 20160332392 15/153216 |
Document ID | / |
Family ID | 53505857 |
Filed Date | 2016-11-17 |
United States Patent
Application |
20160332392 |
Kind Code |
A1 |
PRICE; Jonathan ; et
al. |
November 17, 2016 |
METHOD OF FORMING COMPOSITE STRUCTURES
Abstract
A method of forming composite components is described in which a
composite lay-up is created using varying fibre types across the
lay-up. This can adapt the lay-up to forming processes in regions
of the lay-up to be formed and uses fibre types which give the
greatest strength benefits in areas which do not need to be formed.
Regions not requiring forming may have binders in them activated
prior to a forming step or steps, after which the formed regions
may be impregnated with a matrix and the component cured.
Inventors: |
PRICE; Jonathan; (Bristol,
GB) ; EVANS; Steven; (Bristol, GB) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
AIRBUS OPERATIONS LIMITED |
Bristol |
|
GB |
|
|
Family ID: |
53505857 |
Appl. No.: |
15/153216 |
Filed: |
May 12, 2016 |
Current U.S.
Class: |
1/1 |
Current CPC
Class: |
B29C 70/30 20130101;
B32B 2307/50 20130101; B64C 3/00 20130101; B32B 2605/18 20130101;
B32B 2597/00 20130101; B29C 53/382 20130101; B29C 70/08 20130101;
B32B 37/10 20130101; C08J 2363/00 20130101; B32B 3/00 20130101;
B64C 2001/0072 20130101; C08J 5/042 20130101; B29C 70/34 20130101;
B64C 5/02 20130101; B32B 2405/00 20130101; B32B 27/06 20130101;
B64C 3/18 20130101; B29C 70/382 20130101; B29L 2031/3085 20130101;
B32B 27/00 20130101; B32B 2305/00 20130101; C08J 2300/22 20130101;
B32B 27/12 20130101; B32B 37/06 20130101; C08J 2300/24 20130101;
C08J 5/24 20130101; Y02T 50/40 20130101; B64C 1/06 20130101; B32B
3/04 20130101; Y02T 50/43 20130101; B29L 2031/3076 20130101; B32B
2305/72 20130101 |
International
Class: |
B29C 70/34 20060101
B29C070/34; B29C 70/38 20060101 B29C070/38; B29C 70/08 20060101
B29C070/08 |
Foreign Application Data
Date |
Code |
Application Number |
May 15, 2015 |
GB |
1508375.1 |
Claims
1. A method of forming a composite component, comprising the steps
of: a) providing a lay-up comprising a first region comprising dry
fibres and a second region comprising pre-preg fibres; b) forming
the lay-up to create a non-planar portion at the first, dry-fibre,
region; c) applying a matrix to at least the first, dry-fibre,
region; and d) curing the formed component to solidify the matrix
of the component.
2. A method according to claim 1, wherein a majority of the fibres
present in the thickness of the lay-up in the first region, prior
to application of the matrix, are dry fibres.
3. A method according to claim 1, wherein all of the fibres present
in the thickness of the lay-up in the first region, prior to
application of the matrix, are dry fibres.
4. A method according to claim 1, wherein a majority of the fibres
present in the thickness of the lay-up in the second region are
pre-preg fibres.
5. A method according to claim 1, wherein all of the fibres present
in the thickness of the lay-up in the second region are pre-preg
fibres.
6. A method according to claim 1, wherein a majority of the fibres
present in the thickness of the lay-up in the first region, prior
to application of the matrix, are dry fibres, and a majority of the
fibres present in the thickness of the lay-up in the second region
are pre-preg fibres.
7. A method according to claim 1, wherein prior to application of
the matrix, the lay-up comprises at least one region comprising dry
fibres, disposed between at least two regions comprising pre-preg
fibres.
8. A method according to claim 1, wherein the lay-up, prior to
application of the matrix, comprises a plurality of first regions
comprising dry fibres, disposed on different sides of a second
region comprising pre-preg fibres.
9. A method according to claim 1, wherein forming the component
comprises forming a channel having a closed bottom and an open top,
wherein the first, dry fibre region or regions, is/are disposed at
and/or adjacent at least one of the top edges of the channel.
10. A method according to claim 1, wherein the forming step
comprises forming a bend having an angle of more than 30 degrees,
preferably more than 60 degrees, more preferably more than 90
degrees.
11. A method according to claim 1, wherein the dry fibres are
provided only in, near, or adjacent to regions to be bent, formed
or folded in the forming process.
12. A method according to claim 1, wherein applying a matrix to the
dry fibre region includes a matrix injection or infusion
process.
13. A method according to claim 1, further comprising the step of
locally activating a first portion of the lay-up to activate a
binder or a thermoplastic layer in the first portion of the lay-up,
while leaving a second portion of the lay-up in an un-activated
state, prior to applying a matrix to the dry-fibre region.
14. A method according to claim 1, wherein the composite component
is a laminate composite component, and providing the lay-up
comprises providing a laminated lay-up with layers of dry-fibres
and layers of pre-preg fibres.
15. A method according to claim 14 wherein the laminated lay-up is
provided in step a), and then formed in step b) after step a).
16. A method according to claim 14, wherein the layers of
dry-fibres slide over one another as the lay-up is formed in step
b).
17. A method according to claim 1, wherein the dry-fibres slide
over one another as the lay-up is formed in step b).
18. A composite component formed according to the method of claim
1.
19. An uncured composite preform comprising at least one first
region comprising dry fibres and at least one second region
comprising pre-preg fibres, wherein a majority of the fibres
present in a thickness of the preform in the first region are dry
fibres and a majority of the fibres present in a thickness of the
preform in the second region are pre-preg fibres, the lay-up being
formed, curved, bent or folded at the first, dry-fibre, region or
regions to provide at least one formed region adjacent the second,
pre-preg fibre, region or regions.
20. An aircraft comprising a composite component according to claim
18.
Description
FIELD OF THE INVENTION
[0001] The present invention relates to laminated composite
structures and components, particularly for the aerospace industry.
In particular, the invention relates to improved methods for
forming composite structures from laminated composite fibres, for
primary structures such as aircraft wing structures.
BACKGROUND OF THE INVENTION
[0002] Traditionally composite parts have been manufactured via
labour intensive hand lay-up process, by a skilled laminator. In
known methods, a base material for the lay-up, in the form of
either pre-preg or dry fibre composite material, is laminated into
a mould tool, which matches the geometry of the final component, so
that the base material is formed directly into the shape of the
final part. Using this approach enables complex geometries to be
achieved as the laminators' skill is used to tailor the material
into the contours of the component. However hand lay-up does not
enable high deposition rates of material.
[0003] In all market sectors there is a desire to reduce the
overall manufacture process time throughout all steps in the
production of a cured composite part. Particularly for large scale
or thick components having many plies, and particularly within the
Aerospace & Automotive sectors, this has resulted in the
development of automated deposition processes for all material
formats, such as: Automated Fibre Placement (AFP), Automated tape
lay-up (ATL) and Dry Fibre AFP (DAFP). However these complex
deposition systems have limitations, primarily with respect to the
geometrical shapes which they are able to create, due to the large
physical size of the end effector that delivers the material onto
the tool. For components where the geometry is "simple" and
generally flat, there is less of a problem with access for the end
effector. An example is in the formation of a composite wing skin
part in the aerospace sector in particular.
[0004] For more integrated structures, and for components with more
complex shapes, the size of the end effector can prevent it from
depositing inside cavities or recesses in the shape of the lay-up,
since the end effector may not fit or be able to reach between two
opposing walls of the feature or features, for example. This
necessitates further processing of the un-cured laminate structure
(also known as a preform) to generate the final shape. Typically
this additional processing is reliant on a method of forming, e.g.
in a press or a mould, the laminated preform into the desired shape
prior to curing. All forming process require the use of heat,
pressures and additional mould tooling or consumables, which adds
to the overall process time and cost.
[0005] However there are certain drawbacks associated with these
forming processes. One drawback is that there is a limit to the
geometrical shapes and thicknesses of a laminate that can be
"formed" without inducing unacceptable features in to the
structure, such as fibre deviations and wrinkles. This is a
significant problem for highly loaded structures, where the
laminate thickness can be over 20 mm thick in, for example, a wing
structure of an aircraft. Further, carbon fibres are stiff and
therefore they are difficult to bend and they do not stretch, which
makes forming them around corners a difficult task.
[0006] Different composite material formats can be more challenging
to form than others due to their inherent properties. Pre-preg
consists of fibre (unidirectional or woven) with a film of uncured
matrix already incorporated (pre-impregnated) into the sheet
material. The resin is sticky, therefore once plies of carbon are
laminated together into a stack, they adhere to each other and do
not easily slip relative to one another. The result is that forming
the pre-preg can cause wrinkling and other undesirable features in
the preform.
[0007] Fibres in dry fibre laminate preforms generally do not
adhere to each other and so are able to "slip" over each other,
unless specific binders are placed within the fibre stack and
activated by heat and/or pressure. Dry fibre preforms can come in
the form of Unidirectional fibres (UD), woven fibres, non-crimp
fabrics (NCF--generally layers of unidirectional fibres that are
assembled and stitched together), chopped strand mat (CSM), or any
other known form of fibres for structural composite
applications.
[0008] The terms resin, matrix or impregnation matrix can include
any type of polymer resin or polymer resin mixture presenting low
viscosity and which can be solidified by being polymerized for
general use in forming composite materials. Low viscosity is needed
in certain processes such that any kind of infusion or injection
process is possible. The viscosity of a matrix or resin in pre-preg
materials can of course be higher. As is known to the skilled
reader, a matrix can be injected under pressure, or infused by
drawing the matrix in under vacuum in a number of known processes
not described in detail herein.
[0009] The term "fibre" is used in the following description to
refer to any type of structural fibre such as carbon fibres, glass
fibres, aramid fibres, polyethylene (polyolefines), basalt or
natural fibres, as are generally used in composite materials.
[0010] In the particular case of the aerospace sector, the
predominant material for composite primary structure components is
carbon fibre pre-preg. Pre-preg currently has the highest
structural performance of all material formats and so is beneficial
in these implementations, where structural performance and
light-weighting is key.
[0011] Many parts of aircraft structures would traditionally be
manufactured from a number of separate parts which are then joined
together via mechanical fasteners to create the overall structure.
The mechanical bolting of joints in composite structures is not
efficient.
[0012] In the aerospace sector, utilisation of fibre composite
materials is increasing for primary structure wing applications.
Historically, primary wing structure components made from composite
materials have followed the philosophy of "black metal design",
where composite materials are used in place of metals for
components designed following the same design rules, shapes and
philosophies as are used for metals. In using this approach, the
greatest benefits of using composite material are not necessarily
fully exploited.
SUMMARY OF THE INVENTION
[0013] A first aspect of the invention provides a method of forming
a composite component, comprising the steps of: [0014] a) providing
a lay-up comprising a first region comprising dry fibres and a
second region comprising pre-preg fibres; [0015] b) forming the
lay-up to create a non-planar portion at the first, dry-fibre,
region; [0016] c) applying a matrix to at least the first,
dry-fibre, region; and [0017] d) curing the formed component to
solidify the matrix of the component.
[0018] The method of the invention therefore provides a method of
forming a composite material by using varying fibre types across
the lay-up in order to adapt it to forming processes in regions of
the lay-up to be formed and to use fibre types which give the
greatest strength benefits in areas which do not need to be
formed.
[0019] The lay-up is provided in step a), and then formed in step
b) after step a), for instance by bending or folding the
lay-up.
[0020] Providing the lay-may up may comprise depositing pre-preg
fibres in an automated fibre depositing process and/or depositing
dry fibres using an automated dry fibre depositing process. This
can be done to improve a speed of creating the lay-up with mixed
fibre types.
[0021] Typically the composite component is a laminate composite
component, and providing the lay-up comprises providing a laminated
lay-up with layers of dry-fibres and layers of pre-preg fibres.
Typically the non-planar portion of the lay-up has more than 20
layers of dry-fibres--for instance in the case of a wing-box it may
have 30-50 layers of dry-fibres.
[0022] The lay-up is laminated in step a), and then formed in step
b) after step a), for instance by bending or folding the lay-up. In
other words all of the layers of dry-fibres are simultaneously
formed (for instance by bending or folding the lay-up) after they
have been stacked on top of each other to form to the lay-up. The
non-planar portion is formed after the laminated lay-up has been
provided--rather than bending/folding each ply of the lay-up one by
one as the layers are stacked on top of each other during step a).
The use of dry-fibres in the first, dry-fibre region enables the
laminate lay-up to be formed more easily. Typically the dry-fibres
and/or the layers of dry-fibres slide over one another as the
lay-up is formed to create the non-planar portion at the first,
dry-fibre, region.
[0023] A majority of the fibres, or layers of dry-fibre present in
the thickness of the lay-up in the first region, prior to
application of the matrix, may be dry fibres. Having a majority of
dry fibres helps adapt the material to being formed while maintain
some strength properties of other fibre types such as pre-preg in a
part of the thickness of the material.
[0024] Similarly a majority of the fibre layers present in the
thickness of the lay-up in the first region, prior to application
of the matrix, may be layers of dry fibres.
[0025] All of the fibres present in the thickness of the lay-up in
the first region, prior to application of the matrix, may be dry
fibres. This can provide the greatest degree of flexibility and
formability in the dry fibre region(s).
[0026] A majority of the fibres present in the thickness of the
lay-up in the second region may be pre-preg fibres. This allows
other forms of fibre to be accommodated if necessary.
[0027] Similarly a majority of the fibre layers present in the
thickness of the lay-up in the second region may be layers of
pre-preg fibres.
[0028] All of the fibres present in the thickness of the lay-up in
the second region may be pre-preg fibres. This can provide greatest
strength in the pre-preg fibre regions.
[0029] Prior to application of the matrix, the lay-up may comprise
at least one region comprising dry fibres, disposed between at
least two regions comprising pre-preg fibres.
[0030] The lay-up, prior to application of the matrix, may comprise
a plurality of first regions comprising dry fibres, disposed on
different sides of a second region comprising pre-preg fibres.
[0031] The lay-up, prior to application of the matrix, may comprise
a plurality of dry fibre regions disposed between plural pre-preg
fibre regions, wherein the method further comprises forming the
preform at the dry fibre regions prior to application of the
matrix.
[0032] Forming the component may comprise forming a channel having
a closed bottom and an open top, wherein the first, dry fibre
region or regions, is/are disposed at and/or adjacent at least one
of the top edges of the channel.
[0033] The forming step may comprise forming a bend in the lay-up
having an angle of more than 30 degrees, preferably more than 60
degrees, more preferably more than 90 degrees.
[0034] The dry fibres may be provided only in, near, or adjacent to
regions to be bent, formed or folded in the forming process.
[0035] Applying a matrix to the dry fibre region may include a
matrix injection or infusion process. The matrix injection process
may be, for example, a Resin Transfer Moulding (RTM) process; or a
Same Qualified Resin Transfer Moulding (SQRTM) process. The
infusion process may be, for example, a Liquid Resin Infusion (LRI)
process: or a Resin Film Infusion (RFI) process in which the lay-up
comprises layers of film and the matrix is applied to the first,
dry-fibre, region by infusion from the layers of film.
[0036] The method may further comprise the step of locally
activating a first portion of the lay-up to activate a binder or a
thermoplastic layer in the first portion of the lay-up, preferably
while leaving a second portion of the lay-up in an un-activated
state, prior to applying a matrix to the dry-fibre region.
[0037] The first portion may comprise dry fibres.
[0038] The second portion may comprise dry fibres or pre-preg
fibres or a combination of one or more regions of dry fibres and
one or more regions of pre-preg fibres.
[0039] The step of locally activating the first portion may
comprise applying localised heat and/or pressure to the first
portion.
[0040] The method may further comprise forming the second portion
of the lay-up after the first region has been activated and prior
to applying the matrix to the dry-fibre region.
[0041] The method may further comprise activating the second
portion of the lay-up after the second portion has been formed, by
local activation of binders in the second portion.
[0042] A further aspect of the invention provides a composite
component formed according to any aspect of the methods described
herein.
[0043] A further aspect of the invention provides an uncured
composite preform comprising at least one first region comprising
dry fibres and at least one second region comprising pre-preg
fibres, wherein a majority of the fibres present in a thickness of
the preform in the first region are dry fibres and a majority of
the fibres present in a thickness of the preform in the second
region are pre-preg fibres, the lay-up being formed, curved, bent
or folded at the first, dry-fibre, region or regions to provide at
least one formed region adjacent the second, pre-preg fibre, region
or regions.
[0044] A further aspect of the invention provides an aircraft
comprising a composite component as described herein.
BRIEF DESCRIPTION OF THE DRAWINGS
[0045] Embodiments of the invention will now be described with
reference to the accompanying drawings, in which:
[0046] FIG. 1 shows an aircraft which may incorporate components
formed according to the methods described herein;
[0047] FIG. 2 shows a schematic view of a prior art wing box
assembly;
[0048] FIG. 3 shows an assembly for a wing box according to the
methods of the present invention;
[0049] FIGS. 4A and 4B show forming steps performed on a component
in accordance with a method of the invention;
[0050] FIG. 5 is a schematic sectional view showing a pre-preg to
dry fibre transition at a corner region;
[0051] FIG. 6 shows a flat lay-up which can be folded to form the
corner region of FIG. 5;
[0052] FIG. 7 shows a male mould tool carrying a U-shaped
lay-up;
[0053] FIG. 8 shows the U-shaped lay-up transferred to a female
mould tool;
[0054] FIG. 9 shows the spar flanges formed using the female mould
tool as a mandrel;
[0055] FIG. 10 shows an alternative formed lay-up in which the spar
webs are pre-preg and only the flanges are dry fibre;
[0056] FIG. 11 shows a flat lay-up;
[0057] FIGS. 12 and 13 show a drape forming process applied to the
flat lay-up of FIG. 11;
[0058] FIG. 14 shows the U-shaped lay-up formed by the drape
forming process of FIG. 13 being prepared for further forming;
[0059] FIG. 15 shows the spar flanges being formed by rollers;
[0060] FIG. 16 shows an RFI lay-up arrangement;
[0061] FIG. 17 shows an RFI curing arrangement;
[0062] FIG. 18 shows an RTM curing arrangement; and
[0063] FIG. 19 shows an LRI curing arrangement.
DETAILED DESCRIPTION OF EMBODIMENT(S)
[0064] FIG. 1 shows an aircraft 1 into which components made
according to methods of the invention can be incorporated. The
invention relates to generally sheet-like, web-like, planar or
curved-planar components, which can form parts of the primary
structure of the aircraft, for example, forming parts of the
fuselage 11, parts of the vertical tail fin 12, or the horizontal
tail planes 13. Components of the invention may also be integrated
into a wing 14, elements of an engine housing 15, and even into
parts of the under carriage 16 and 17, or doors for the under
carriage. The invention relates to the stiffening of generally
sheet or web-like components such as wing skins, or generally
web-like features of bulkheads, spars or ribs of an aircraft, for
beams and bulkheads for any structure, as examples.
[0065] FIG. 2 shows a known assembly for forming a wing box of the
prior art. The wing box is made up of a lower cover 201, and an
upper cover 202, each of which is reinforced with a plurality of
stringers 211 and 212. A front spar 220 and rear spar 230 are
provided. A plurality of ribs 240 is also provided. The assembly
method generally comprises attaching a front spar and rear spar to
the lower cover using fixing means such as bolts or rivets. Rib
feet for attaching the ribs 240 to the lower cover 211 are attached
to the lower cover and ribs 240 can then be assembled on to the rib
feet and attached to the front 220 and rear 230 spars. Finally, the
upper cover 202 can be assembled on to the ribs 240 and front 220
and rear 230 spars to create the fully formed box. All of the
joints between parts generally include point fixing means such as
bolts or rivets and the stress concentrations created by such
fixing means can mean additional strength is required in the
components being fixed, which results in an increase in overall
weight.
[0066] With increasing use of composite materials in structural
components, such as those used in airframe structures and in wings
in particular, the most structurally efficient way to arrive at
light weight and lower cost composite structures is to increase the
integration of parts. Therefore, where possible, parts are
integrally formed as one piece rather than being formed separately
and joined together afterwards. Therefore it can be beneficial in
the development of composite structures to integrate increasing
levels of structural functionality via integration of the structure
into a lower number of component parts.
[0067] In the following described embodiment, wing spar and cover
components were combined into one part. This part can be fabricated
by depositing pre-preg material onto a male mandrel, to produce a
substantially U shaped preform. This preform can be removed from
the mandrel and the outward facing spar caps formed. Rib feet and
stringers are then added and the whole assembled preform is then
cured.
[0068] However, using standard known methods, it had not proved
possible to form the outward facing cap along the full length of
the wing with the large cross sectional thickness required for a
highly loaded structure.
[0069] The process of slipping and shearing of the fibre is
important to being able to "form" it into complex geometries (NB
all of the forming process will induce a degree of deviation into
each ply of the laminate). With pre-preg laminates that are adhered
together, sufficient slippage/shearing of the material is not
generally obtained.
[0070] Mixed fibre laminates as described herein overcome this
issue by locally changing the type of material incorporated within
the lay-up of the un-cured preform. In embodiments of the
invention, dry fibres are incorporated into the laminate in one or
more regions of a preform structure where it needs to be formed.
This can be achieved with dry fibre AFP or NCF material.
[0071] FIG. 3 shows an assembly for forming a U-box according to an
embodiment of the invention. The upper cover 302 comprising
stringers 312 is provided as illustrated, generally in a manner
similar to that provided for the assembly of FIG. 2. Similarly, a
plurality of ribs 340 is provided in a similar arrangement to the
known method illustrated in FIG. 2.
[0072] The methods described herein permit a substantially U-shaped
structure to be provided to form the lower part of the wing box
assembly. The substantially U-shaped lower wing box component
comprises a lower cover 301, which is integrally formed with front
spar 320. Stringers 311 are provided on the lower cover 301 in
substantially the same manner as provided for the prior art method
of FIG. 2. In the assemblies of FIG. 2 or 3, the stringers 211 or
311 can be bonded to, integrally formed with, or bolted to the
lower cover and/or the upper cover in a manner which is generally
known to the person skilled in the art and so is not described in
detail herein.
[0073] The methods of the present invention allow the U-box
component 350 of FIG. 3 to be formed integrally from composite
materials. In a first example, the lay-up for the composite U-box
350 is formed directly in a female U-shaped mould. The lay-up is
formed in the mould using any form of known automated fibre
placement, automated tape lay-up and dry fibre automated fibre
placement. In a second example, the lay-up is formed directly onto
a male tool, as described in further detail below with reference to
FIG. 7, then transferred to a female U-shaped mould as shown in
FIGS. 8-10. In its initial laid-up form, the U-box component 350
does not yet comprise the spar caps shown as flange portions 361
and 362, which are eventually used to attach the upper cover 302 to
the front 320 and rear 330 spars. Using the method of the
invention, which will be described in greater detail in relation to
FIGS. 4A and 4B, the lower cover 301 and the spars 320 and 330, are
comprised of at least a majority of pre-preg fibres through the
thickness of the lay-up and preferably entirely of pre-preg fibres.
Pre-preg fibres have generally been found to have the greatest
structural performance and so are useful in structural components
such as a wing box illustrated in FIG. 3. In order to form the
flange portions 361 and 362 for the spar caps, it is necessary to
create a relatively sharp crease or fold in the lay-up prior to
curing. This is because the automated fibre placement technology
available is not necessarily well suited to laying up fibres in the
complex shapes shown in order to form the lower wing box part 350
illustrated in FIG. 3. As described above, pre-preg fibres do not
necessarily lend themselves well to being folded, especially in
thick profile sections, which are generally necessary for important
major structural parts such as the wing box assembly of FIG. 3.
Therefore, in the method of the invention, dry fibres are placed in
the region of the crease which forms the flanges 361 and 362 of the
lower wing box part 350. The dry fibres therefore provide a portion
of the lay-up in which the fibres are able to slide over one
another and therefore are better suited to being folded in the
manner required to form the flanges 361 and 362. Once the flange
sections 361 and 362 have been formed, then the overall component
350 can be cured. The forming operation utilised to fold the
flanges 361 and 362 will later be described in greater detail with
reference to FIGS. 8 and 9.
[0074] In an alternative method, one or more of the dry-fibre
regions of the lay-up may be locally activated to locally activate
a binder or a thermoplastic layer in one or more sub-regions of the
lay-up, preferably prior to forming the lay-up, in at least one
un-activated dry-fibre region or regions. Therefore, in the above
example, the dry fibres placed in the region of the crease which
forms the flanges 361 and 362 of the lower wing box part 350 may be
left un-activated, while dry-fibre regions which are adjacent the
areas to be formed, namely those which may appear in the lower
cover 301 and the spars 320 and 330, may be secured or stiffened,
or at least partially solidified by local activation. The
substantially planar portions of the flanges 361 and 362 may also
be locally activated, leaving only an at least predominantly
dry-fibre, region of the crease, in between the flanges 361 and 362
and the spars 320 and 330, un-activated and therefore still
formable. The crease may then be activated and subsequently
impregnated with matrix and cured after the forming step. The
forming operation may include the application of heat. If this heat
applied during the forming operation is sufficiently high to
activate the binder of the dry fibre region, then the forming
operation may also activate the binder in the dry fibre region to
leave a formed and activated region of dry fibres. This has the
advantage of adding stiffness to the pan during the forming
operation, which allows easier handling or manipulation of the
preform during subsequent steps. The methods described herein can
employ any binder suitable for use in a composite material to bind
the fibres prior to curing. Binders can come in different forms.
One example is marketed by Hexcel.TM. as Toughened NCF, Biaxial,
(45.degree./135.degree. & 135.degree./45.degree.), 2.times.268
gsm, IMA, V800E veil (6/6/0), which combines a binding and
toughening function. An alternative is a Toho-Tenax.TM. Toughened
UD-Woven Fabric--IMS65 fibre, 194 gsm, Toughened TA 1902-05 (5
gsm), EP05311 Binder (6 gsm), 1.27 m wide,
PBI_06-V8_05-IMS65-UD-0194-1270
[0075] Prior to curing, the dry fibres are impregnated with a
polymer matrix in order to complete the composite material which is
required in the initially dry fibre regions. Methods already exist
for impregnating and curing dry fibres in a mould. One such
available technology is known as same qualified resin transfer
moulding (SQRTM). SQRTM is a closed moulding method which combines
the processing of pre-preg materials and liquid moulding. SQRTM is
a development on standard resin transfer moulding (RTM) which is
also generally available and known to the skilled reader. RTM takes
a dry fibre preform and impregnates the dry fibre with matrix. Then
a chemical reaction is thermally activated by heat in the fibre mat
or lay-up and in the wall of the mould. RTM is generally a
closed-mould, vacuum-assisted process. In the RTM process, a fibre
preform or lay-up is placed into a mould cavity having the shape of
the desired part. The mould is then closed and a generally low
viscosity resin is pumped into the mould under pressure, displacing
the air within and around the fibres until the mould is filled.
After the fill cycle, the cure cycle starts, during which the mould
is heated and the resin polymerises to become a rigid plastic.
[0076] SQRTM follows a similar process, but substitutes the dry
fibres with a pre-preg lay-up. In SQRTM, pre-preg plies are
arranged within the mould, the mould is closed and then further
liquid resin is injected into the mould. Preferably, the injected
resin is the same resin that is found within the pre-preg and
provides the same mechanical properties. Differently to the RTM
process, in SQRTM, the injected resin is intended to fill cavities
around the part, but is not intended to impregnate the
pre-preg.
[0077] In the method of the present invention, a mixture of both
pre-preg fibres and dry fibres is employed at different regions in
the lay-up and so a combination of the effects generated in SQRTM
and traditional RTM is created, where in regions comprising dry
fibres, the matrix is impregnated in the dry fibres and in regions
where the lay-up comprises pre-preg fibres, the resin forms around
the pre-preg as in a standard SQRTM process. Alternative methods
for fibre impregnation and cure include Resin Film Infusion (RFI),
Liquid Resin Infusion (LRI) or any other suitable method known to
the skilled reader. The aforementioned SQRTM, RFI and LRI methods
will be described in greater detail below with reference to FIGS.
17-19.
[0078] FIGS. 4A and 4B illustrate in greater detail regions of the
lower wing box cover 350 of FIG. 3 which can benefit from the
features of the present invention.
[0079] FIG. 4A illustrates a portion of the lower wing box 350,
including a planar part 352 of the lower cover 301 where it meets
the spar 320. There are two ways in which the initial angle in the
lower wing box cover can be formed at the line 400. One manner in
which this profile can be formed, is if the pre-preg lay-up is
deposited into a mould having a complimentary shape, so that the
component is laid-up on the inside of the mould and the automated
deposition of the pre-preg forms the desired angle in the initial
structure. In this manner, a lay-up comprising pre-preg fibre
material having pre-formed angles in it is provided. However,
alternatively, the angle at line 400 may also be created by
depositing dry fibres at the region of the angle 400 and pre-preg
fibres either side of it, to form the generally planar parts 351
and 352 of the lay-up. A dry fibre region is therefore formed
between adjacent pre-preg fibre regions. It will be apparent to the
skilled reader, that the generally planar parts 351 and 352 may
still incorporate features, thickness changes and slight variations
in profile, but they are generally of a planar form, having a
significant extent in a direction of their respective planes, and a
thickness in a perpendicular direction to their respective planes
sufficiently small as to give them a substantially sheet or
web-like configuration. The methods of the invention therefore
provide substantially planar regions of a preform which are
primarily made from pre-preg fibre material, while in between the
pre-preg regions, dry fibre regions are provided to permit the
desired forming procedures.
[0080] A further alternative lay-up arrangement is shown in FIG. 5
in which the lower cover 301 (including the planar part 352)
comprises layers of pre-preg fibres shown in solid line, and the
rest of the lay-up comprises layers of dry-fibres shown in dashed
line (both the spar 320 and the flange 362 being formed entirely
from layers of dry-fibres).
[0081] The number of layers shown in FIG. 5 is schematic only. In
the case of FIG. 5 the non-planar portion of the lay-up is shown
with only nine dry-fibre layers, but more typically it will have
more than 20 layers of dry-fibres--for instance it may have 30-50
layers of dry-fibres. In the thickest part of the cover the number
of pre-preg fibre layers may be as high as 120.
[0082] FIG. 5 illustrates a transition 430 between the layers of
pre-preg fibres (referred to below as "pre-preg plies") and the
layers of dry-fibres (referred to below as "dry fibre plies") where
they butt up against each other. This transition can be provided in
a number of different ways to what is shown in FIG. 5. For example,
the transition can be achieved by gradually dropping off the
pre-preg plies along a ramp distance 440, and replacing the
pre-preg plies with dry fibre plies. The ply transition 430 is
staggered so as to ultimately achieve at least a majority of dry
fibres at the region of the fold 400. The transition represented
within FIGS. 5 and 6 may not be representative of the true ramp
distance of the actual transition within the region of the fold
400.
[0083] In an alternative method, alternating layers of dry fibre
and pre-preg can be interwoven and the proportion of pre-preg to
dry fibre can gradually transition from 100% pre-preg, through
around 50% pre-preg and 50% dry fibre, to up to around 100% dry
fibre, as seen through the thickness of the material, at the point
at which the fold 400 is eventually to be formed.
[0084] However, it will be appreciated that providing a majority of
dry fibres and a relatively low proportion of pre-preg fibres in
the region to be formed can also realise at least some of the
benefits of the invention. For example, pre-preg fibres may be
provided at the outer extent of the fold to be formed. By providing
at least a proportion of dry fibres in the thickness of the lay-up,
layers of fibre in the lay-up can slide over one another to enable
the forming of desired features in a preform. A preform can
therefore beneficially comprise both pre-preg and dry fibres
through a thickness of the lay-up.
[0085] Therefore, in the case of the component of FIG. 4A, the
initial fold 400 may be provided by folding a substantially planar
or sheet-like lay-up in a direction of arrow 410. This folding step
may be carried out after locally activating a binder in one or more
of the substantially planar regions of the lay-up.
[0086] Rather than being laid-up into a female mould or onto a male
mould to form a U-shape, the lower wing box 350 may be laid up as a
substantially flat sheet which is then folded to form both fold
lines 400, 401. FIG. 6 illustrates the lay-up as a substantially
flat sheet before it has been formed to generate the folds 400 or
401. The layers are stacked on top of each other to form the stack
of FIG. 6 which is subsequently folded. FIG. 4B shows the forming
of a second fold 401 in order to form, for example, the flange
portions 361 and 362 of the U-box component 350 of FIG. 3.
[0087] Again, as described in relation to FIG. 4A, in the region of
the fold line 401, dry fibres can be deposited so that at least a
proportion, or a majority, or preferably the entirety, of the
thickness of the lay-up in the region of the fold 401 comprises dry
fibres. The entire thickness of the lay-up may at this region 401
be comprised of dry fibres. Therefore, in order to form the flanged
portions 420, a fibre based preform can be folded at a region 401
in which the lay-up comprises primarily dry fibres. In this manner,
a lay-up can be provided comprising pre-preg fibres in regions
where the lay-up has been deposited substantially in its final form
for curing, so needing no forming prior to the curing procedure.
Further, the lay-up can comprise at least a majority of dry fibres
in regions where the lay-up is to be further formed or folded to
arrive at the eventual shape of the component before the curing
stage.
[0088] In the example described above, the planar part 351 shown in
FIG. 4A (which forms both the spar 320 and the flange 362) is
formed entirely of dry fibre with no pre-preg fibres. In an
alternative example, the spar 320 and flange 362 may be formed from
pre-preg (like the lower cover 301) and only the regions of the
fold lines 400, 401 are formed from dry fibre, with transitions
between the dry fibre and pre-preg as shown in FIGS. 5 and 6.
[0089] As described above, the whole component may be cured in a
single operation after the forming step or steps. Further, binders
in one or more dry-fibre portions of the component may be locally
activated before the forming step or steps, to secure or stiffen
the lay-up in regions where the forming is not to be carried out.
This can provide formable regions in between activated sections of
the component or lay-up. This can ease manipulation of the lay-up
in the forming steps, since the activated sections will remain in
their desired shape while the forming of the un-activated sections
takes place.
[0090] FIG. 7 illustrates the initial lay-up operation onto a male
tool 370. In FIG. 7 and subsequent drawings, pre-preg fibre regions
are indicated in solid line and dry-fibre regions are indicated in
dashed line.
[0091] As has been previously discussed, the lay-up can be
deposited onto the male tool 370 using any form of known automated
fibre placement, automated tape lay-up or dry fibre automated fibre
placement. The laid-up preform comprises at least a majority of
pre-preg fibres at the cover region 301 and at least a majority of
dry fibres at the spar regions 320 and 330.
[0092] FIG. 7 shows the step of providing the laminated lay-up, and
FIGS. 8 and 9 show the subsequent forming stage for the lower wing
box 350. The U-shaped lay-up is transferred into a complimentary
female mould 380. The female mould is designed to allow the spar
flange regions to protrude from the top of the mould. The female
mould is shaped to contain the cover and spar regions of the
preform. The spar flange regions 361 and 362 are unbounded by the
female mould to permit the folding operation.
[0093] The spar flange regions 361 and 362 protruding from the
female mould are folded to create the spar flange 361 and 362 of
the lower wing box 350. This may be performed by a rolling
operation. Alternatively, this may be performed by a pressing
operation. It will be apparent to the skilled reader, that other
suitable forming methods may be used.
[0094] FIG. 10 illustrates the final forming stage for a lower wing
box of an alternative embodiment, in which the spar regions 320 and
330 comprise at least a majority of pre-preg fibres (shown in solid
line), and only the spar flange regions 361 and 362 are dry-fibre
(shown in dashed line).
[0095] FIGS. 11-15 illustrate an alternative method by which the
lower wing box 350 can be formed. In this case the lower wing box
is deposited as a substantially planar or sheet-like lay-up. The
lower wing box shape is subsequently provided by performing a
series of folding operations.
[0096] Depositing a generally flat sheet of fibres as shown in FIG.
11, be they dry fibres or pre-preg fibres, is in general the
fastest way in which a lay-up can be formed by known automated
depositing technologies. It can therefore be advantageous to create
a product from a single flat sheet lay-up, which is then formed to
provide the eventual preform for the component. The component can
then be cured as a single part in its eventual folded form. The
embodiment of FIGS. 11 to 15 shows a method by which the lower wing
box 550, can be formed as a single part from a substantially flat
sheet lay-up or preform, according to an embodiment of the present
invention.
[0097] FIG. 11 shows a substantially flat sheet preform, which
comprises a number of sub-regions having differing properties. The
lay-up 500 comprises a first spar flange region 501, which is a
substantially planar region formed primarily of pre-preg fibres. A
substantially linear formable region 502 is located adjacent the
spar flange 501, and is formed primarily of dry fibres to enable
that region to be formed or folded in accordance with the methods
of the invention. Adjacent that first formable region 502 is a spar
region 503. Adjacent the spar region 503 is a second formable
region 504, which will again be primarily or wholly formed from dry
fibres throughout most or all of the thickness of the lay-up. A
cover region 505 is formed in between two formable regions 504 and
506. The cover region is substantially planar and primarily formed
from pre-preg fibres. As will be appreciated, the cover region may
include certain variations in thickness and ridges or channel like
formations to allow the attachment of ribs, for example, or to
accommodate other internal components to the wing, but is generally
of planar form and so is suited to being deposited using pre-preg
fibres without a need for significant further forming operations to
form the cover. A further spar region 507 is provided in between
folding regions 506 and 508. Like the first spar region 503, this
spar region is formed primarily from pre-preg fibres and can
incorporate variations in thickness and slight contours to allow
incorporation of internal features of the wing box inside the two
covers. A second spar flange region 509 is formed adjacent folding
region 508 and this region is again formed primarily or wholly from
pre-preg fibres and has a generally planar form.
[0098] As can be seen, a generally flat lay-up can therefore be
provided, which comprises a plurality of substantially planar
regions formed from pre-preg fibres, and one or more formable
regions 502, 504, 506, 508, which are formed either wholly or
primarily from dry fibres, to permit folding or forming of the
lay-up in the formable regions. As has been described above, at
least a portion of one or more of the substantially planar regions,
or any region where forming is not required, may be locally
activated, preferably with the application of heat and/or pressure,
to activate a binder in those regions prior to any or all of the
forming steps described in relation to FIGS. 12-15 being carried
out. After forming, any or all of the dry fibre regions can be
locally activated and/or provided with a matrix via known methods
as described herein, and subsequently cured.
[0099] FIG. 11 illustrates a first step in the alternative method,
in which the substantially flat lay-up is deposited onto a lay-up
surface. The lay-up surface comprises an upper surface of a male
tool 571 and two removable sidewall sections 573. The male tool 571
is mounted on a support 572 which is carried by a base 590. The
lay-up may be deposited using any form of known automated fibre
placement, automated tape lay-up or dry fibre automated fibre
placement.
[0100] Next, the sidewall sections 573 are removed as shown in FIG.
12 and the lay-up 500 and tool components 571 and 572 are vacuum
bagged to enable the lay-up to be formed via vacuum assisted drape
forming. The vacuum bagging arrangement comprises a bagging film
593, adhesive tape 594 which secures the edge of the bagging film
593 to sweeper blocks 595, and a vacuum port 591.
[0101] FIG. 13 shows a first, drape forming step in the forming
process for the lay-up of FIG. 11. In this first step, the spar
regions 503 and 507 are drawn about the male tool 572 by a vacuum
applied to the vacuum port 591. This first forming operation forms
first and second spar regions of the lay-up, 503 and 507
respectively, which are disposed at an angle with respect to the
cover region 505, by deformation of the formable regions 504 and
506. Again, any dry-fibre regions of the areas requiring no forming
may be activated with heat and/or pressure before the forming step
takes place.
[0102] FIG. 14 shows an intermediate step between the drape forming
step and a final forming step. The preform is removed from the
bagging arrangement and the support 572 is removed so that the spar
flange portions of the lower wing box preform 550 protrude from the
male tool 571.
[0103] In a final forming step, as illustrated in FIG. 15, the spar
flange portions 501 and 509 are folded to obtain the lower wing box
550. In this case the spar flanges 501 and 509 are folded by
rollers 600. Alternatively, the spar flanges may be formed by
pressing or as part of a vacuum-assisted drape forming process
[0104] In an alternative method, the lower wing box preform may be
removed from the male tool 571 and placed into a complimentary
female mould (similar to the female mould 380 shown in FIGS. 8-10)
prior to the final forming step.
[0105] As will be appreciated by the skilled reader, internal
stringers and/or ribs may be incorporated, similarly to those
illustrated in respect of FIG. 3. The spar flange regions 501 and
509 can be fixed or bonded to the upper cover 302 to create an
enclosed box structure. It is then possible to infuse matrix into
the dry fibre regions and cure them to form the final cured part.
Prior to infusing the dry fibre regions, any so far un-activated
regions may be activated prior to application of the matrix, to
stiffen or secure them prior to the application of the matrix and
curing steps. Pre-preg fibre regions may be cured simultaneously
with the matrix-infused dry fibre regions.
[0106] After the lay-up has been formed, matrix is applied to at
least the first, dry-fibre, region; and then the formed component
is cured to solidify the matrix of the component. As mentioned
previously, the matrix may be applied by RTM, Resin Film Infusion
(RFI) or Liquid Resin Infusion (LRI) for example. FIGS. 16 to 19
are schematic drawings showing these methods in more detail.
[0107] FIG. 6 shows a lay-up suitable for use with a SQRTM, RTM or
LRI process, whereas FIG. 16 shows a lay-up suitable for use in an
RFI process. In the case of FIG. 6 the planar part 351 is formed
entirely of dry fibre. In the case of FIG. 16 the, the layers of
dry fibre are interleaved with layers of semi-solid resin film 450
(each film 450 illustrated in FIG. 16 by a line of crosses).
Alternatively, a single layer of resin film may be provided for
multiple layers of dry fibre.
[0108] FIG. 17 illustrates the RFI curing process in which a
preform 601 is laid up with semi-solid resin film 602 and then
vacuum bagged into a female mould 603. The vacuum bagging
arrangement comprises a bagging film 604, a vacuum line 605 and an
adhesive tape 606. The lay-up is then heated to infuse the
dry-fibres with the resin matrix material from the semi-solid resin
film. A positive pressure may be applied to the lay-up during the
curing process, by performing the operation within an autoclave.
Alternatively, the vacuum can be omitted when performing this
operation within an autoclave. The female mould 380 shown in FIGS.
8-10 may provide the female mould for an RFI process as shown in
FIG. 17.
[0109] FIG. 18 illustrates an alternative method in which the lower
wing box is impregnated and cured using Resin Transfer Moulding
(RTM). The RTM apparatus comprises a lower female mould 603 and an
upper tool 608, and a chamber 610 between the upper mould and lower
tool. The upper tool has channels 609 to transfer resin 607 from a
melt pot into the chamber 610. The lay-up is placed within the
chamber and applied with heat and pressure to wet the lay-up and
subsequently cure the composite. Alternatively, the cure can be
performed at an ambient temperature. The female mould 380 shown in
FIGS. 8-10 may provide the female mould for the curing and
impregnation process shown in FIG. 18.
[0110] In a third alternative method, the lay-up is infused and
cured using Liquid Resin Infusion (LRI), as shown in FIG. 19. The
LRI apparatus comprises a mould 603 and a vacuum bagging
arrangement; comprising a bagging film 604 and an adhesive tape
606. The LRI apparatus further comprises a vacuum pump 611, a resin
tap 612 and a resin pot 613. The preform is initially vacuum bagged
into the mould and, once the preform has been bagged, the resin tap
is opened. The vacuum pump draws resin 607 from the resin pot 613
into the preform so as to infuse the preform with liquid resin. The
preform is then cured via heating at ambient pressure within an
oven. The female mould 380 shown in FIGS. 8-10 may provide the
female mould for the curing and infusion process as shown in FIG.
19.
[0111] As will be appreciated, the illustrated method permits the
forming of a closed wing box structure with minimal separate parts.
This reduces the number of fixing operations required, as well as
reducing the number of fixing elements, bonding regions, or regions
in which bolts or rivets are required to hold parts together. This
all reduces the number of manufacturing and assembly operations
which are required, reduces overall weight, and increases general
structural integrity of the structure, since fixing means such as
bolts and rivets create stress concentrations which can reduce the
overall efficiency of the structure. As will be evident to the
skilled reader, although the illustrated embodiment relates to the
forming of a wing box, the forming method could be applied to any
component having two or more substantially planar regions which are
separated by a formed region, such as a fold or curved section
formed in a forming process. The method could also be applied to
any component having substantially planar regions separated by
curves, bends or folds. The methods can also be applied to any
generally tubular component having a plurality of generally planar
sides separated by generally longitudinal folds.
[0112] Although the invention has been described above with
reference to one or more preferred embodiments, it will be
appreciated that various changes or modifications may be made
without departing from the scope of the invention as defined in the
appended claims.
* * * * *