Note: Descriptions are shown in the official language in which they were submitted.
CA 02242050 1998-07-02
WO 97/25198 PCT/US97/00075
COMPOSITE HONEYCOMB
SAND YYICH STR UCT URE
Technical Field
The present invention relates to composite honeycomb sandwich
structure, and particularly to resin impregnated fabric sheets forming outer
skins adhered on opposed surfaces of a honeycomb core with an intermediate
barrier to eliminate resin flow from the skins to the core.
Background Art
Aerospace honeycomb core sandwich panels (having composite laminate
skins cocured with adhesives to the core through autoclave processing) fmd
widespread use today because of the high stiffness-to-weight (i.e., "specific
stiffness) and strength-to-weight (i.e., specific strength) ratios the panels
afford.
Typical honeycomb core sandwich panels are described in U.S. Patents
5,284,702; 4,622,091; and 4,353,947. Alteneder et al., Processing and
Characterization Studies of Honeycomb Composite Structures, 38th Int'l
SAMPE Symposiuni, May 10-13, 1993 (PCL Internal No. 200-01/93-AWA)
discusses common problems with these panels, including core collapse (i.e.,
core crush), skin laminate porosity, and poor tool surface fmish.
U.S. Patent 5,445,861 describes composite sandwich structure for sound
absorption (acoustic insulation) and other applications. The sandwich
structures have seven layers as follows:
(1) an outer skin;
(2) a small celled honeycomb or foam core;
(3) a frontside inner septum;
(4) a large celled middle honeycomb core;
(5) a backside, inner septum;
(6) a backside, small celled honeycomb or foam core; and
CA 02242050 1998-07-02
WO 97/25198 PCT/US97/00075
2
(7) an inner skin.
Tuned cavity absorbers in the middle honeycomb core absorb sound.
Performance of this structure suffers from resin flow to the cells of the
honeycomb cores during fabrication for the reasons already discussed and
because such flow alters the resonance of the structure.
Sunmarv of the Invention
With a high flow resin system, large amounts of resin can flow into the
core during the autoclave processing cycle. Such flow robs resin from the
laminate, introduces a weight penalty in the panel to achieve the desired
performance, and forces over design of the laminate plies to account for the
flow losses. The resin loss from the lamina.te plies also reduces the
thickness of
the cured plies which compromises the mechanical performance. To achieve
the desired performance and the corresponding laminate thickness, additional
plies are necessary with resulting cost and weight penalties. Because the
weight penalty is severe in terms of the impact on vehicle perfortnance and
cost
in modem aircraft and because the flow is a relatively unpredictable and
uncontrolled process, aerospace design and manufacture dictates that flow into
the core be eliminated or significantly reduced. In addition to the weight
penalty from resin flow to the core, we discovered that microcracking that
originated in the migrated resin could propagate to the bond line and degrade
mechanical performance. Such microcracking potential poses a catastrophic
threat to the integrity of the panel and dictates that flow be eliminated or,
at
least, controlled.
Flow from the laminates to the core occurs because of viscosity
reduction of the resin (i.e., thinning) at the elevated processing
temperatures.
Therefore, prior art attempts to solve the flow problem have generally focused
on retaining the ambient temperature viscosity of the resin at the curing
temperatures. For example, one might alter the processing cycle to initiate
CA 02242050 1998-07-02
WO 97/25198 3 PCT/US97/00075
curing of the resin during a slow heat-up, low pressure step to induce resin
chain growth before high temperature, high pressure completion. In this staged
cure cycle, one would try to retain the resin's viscosity by building
molecular
weight at low temperatures. Higher molecular weight resins have inherently
higher viscosity so they remain thicker and are resistant to damaging flow to
the core. Unfortunately, with a staged cure cycle, too much flow still occurs,
and the potential problems of microcracking still abound. Also, facesheet
porosity might increase beyond acceptable limits. Furthermore, a modified
cure cycle increases autoclave processing time. Increased processing time
translates to a significant fabrication cost increase with risk of rejection
of high
value parts at the mercy of uncontrolled and inadequately understood factors.
We eliminate resin (matrix) flow into the honeycomb core for sandwich
structure using high flow resin systems and results in reproducibility and
predictability in sandwich panel fabrication and confidence in the structural
performance of the resulting panel. We use a scrim-supported barrier film
between the fiber-reinforced resin composite laminates and the honeycomb
core. This sandwich structure is lighter for the same performance
characteristics than prior art panels because the resin remains in the
laminate
(skin) where it provides structural strength rather than flowing to the core
where it is worthless, introducing excess weight and potential panel failure.
We also generally use an unsupported film adhesive between the barrier film
and the laminates to bond the laminates to the barrier film. With these layers
(which might be combined into one product), they achieved improved
performance, retained the resin in the laminates and thereby reduced excess
resin that designers otherwise needed to design into the panels to account for
resin flow into the core, and reliably fabricated panels in which they had
structural confidence.
Core crush frequently occurred in the chamfer region of honeycomb
core when we cured a panel having a scrim-supported barrier film, particularly
CA 02242050 1998-07-02
WO 97/25198 4 PCT/US97/00075
when they tried to use lighter weight core materials. We can reduce core crush
in these panels by including a tiedown ply in contact with the core beneath
the
barrier film (and adhesive) because the tiedown ply reduced slippage of the
barrier film relative to the core during curing.
Controlling core slippage in the present invention allows us to use
lighter density honeycomb core to produce structures without costly scrap due
to core crush. We reduce manufacturing costs both by saving time, materials,
and rework/scrap and by improving the reliability of the manufacturing process
to produce aerospace-quality panels having the highest specific strength and
specific stiffness.
The added tiedown ply means that three or more tiedown plys will be
included in the fmal preform of the panel. In conventional practice, there
will
also be tiedown plys on the outer surfaces of the panel and possibly between
the laminate and the adhesive barrier film. Each tiedown ply extends
outwardly from the part beyond the net trim line of the fmished product.
Conventionally, the tiedown plies are secured individually and sequentially to
the layup mandrel with tape. Especially when using low density core it is
important to fix the relation of the plies to one another and to the mandrel.
Failure of the tape results in facesheet ply wrinkles or core crush. Core
crush
could still occasionally occur when the tiedown ply in contact with the core
puiled away from the tape securing it to the mandrel, slipping relative to the
other tiedown plies. The adhering strength of the tape alone was insufficient
to
overcome the forces acting on the core in a panel when we applied autoclave
pressure. We discovered how to adhere the tiedown plies to each other
reliably, easily, and inexpensively. Adhering the plies to each other
distributes
the forces acting on any individual ply among all the tiedown plies, reducing
the maximum force seen by the tape adhering the tiedown plies to the mandrel.
While described with respect to a composite honeycomb sandwich structure,
CA 02242050 1998-07-02
WO 97/25198 PCr/US97/00075
the adhering method is generally applicable to all applications involving
tiedown plies in composite construction.
Thus, in one aspect, the present invention relates to an improvement in
the manufacture of composite structure, especially composite honeycomb
sandwich structure, where tiedown plys are used to secure the part during
autoclave curing at elevated temperature and pressure. To lock the tiedown
plies together so that there is no movement of one ply relative to another, we
use a lower temperature curing adhesive to cure and to connect the several
plys
together during the early stages of autoclave curing prior to applying
pressure.
We apply the adhesive outside the net trim line for the part, so that it is
removed during fmishing of the part.
In another aspect, the present invention relates to adhering tiedown plys
to one another during the construction of composite structure, especially
during
the autoclave processing at elevated temperature and pressure of composite
honeycomb sandwich structure. The conventional practice of taping the
tiedown plys to the mandrel alone is unsatisfactory, because the taping must
be
sufficient to prevent slippage of any ply or of one ply relative to another.
We
discovered that we could adhere the plies cffectively to one another to reduce
maximum forces on the tape by applying a low temperature curing film
adhesive between the tiedown plys just outside the net trim line for the part.
In
the autoclave, this film adhesive melts and cures at a lower temperature than
the resin in the laminates so that it bonds the tiedown plies together prior
to
increasing the autoclave pressure at the higher temperature where the laminate
resin flows and cures. The film adhesive eliminates movement of the tiedown
plys relative to one another. In our preferred embodiment for a bismaleimide
(BMI) sandwich panel, we prefer to use an adhesive that cures at about 250 F
(121 C) for a BMI that cures around 375 F (191 C), and post-cures around
440
F.
CA 02242050 2006-05-17
5A
In a further aspect, the present invention relates to a composite honeycomb
sandwich
structure having improved resistance to core crush. The structure includes a
honeycomb core, having core cells and a peripheral chamfer. At least one
composite
laminate having plys of fibre-reinforced bismaleimide matrix resin is adhered
to the core.
A bondable grade polyimide barrier film is located between the laminate and
the core to
eliminate resin flow from the laminate into the core cells. An adhesive
between the
barrier film and the core bond the laminate and core. A supporting scrim
between the
adhesive and the core prevents sagging of the barrier film into the core
cells. A tiedown
ply of bismaleimide resin-impregnated woven fabric in contact with the chamfer
of the
core beneath the adhesive and scrim, eliminates slippage of the barrier film
relative to
the core and, in so doing, reduces core crush. Optionally, the tiedown ply may
have a
0/90 degree fibre orientation.
CA 02242050 1998-07-02
WO 97/25198 PCT/US97/00075
6
Brief Descriv_tion of the Drawings.
Fig. 1 illustrates a typical composite honeycomb sandwich structure.
Fig. 2 is a schematic, partial sectional view of the skin-core interface in
sandwich structure having a scrim-supported barrier film to prevent resin flow
from the skin to the core.
Fig. 3 is a schematic, partial sectional view of prior art honeycomb
sandwich structure, suffering resin flow to the core, using a supported film
adhesive without a barrier film.
Fig. 4 is another schematic, partial sectional view showing sandwich
structure with resin depletion in the skin, but where the resin is prevented
from
reaching the core with a bulging, unsupported barrier film.
Fig. 5 is a schematic, sectional elevation showing core crush of a
honeycomb sandwich panel caused by core and barrier film slippage.
Fig. 6 is another schematic, sectional elevation showing the use of a
tiedown ply to reduce core crush.
Fig. 7 is a graph of a typical autoclave curing cycle for making
composite honeycomb sandwich panels, showing that our tiedown adhesive
cures prior to the application of high pressure in the cycle.
Detailed Description of a Preferred Embodiment
As a frame of reference for this description, we will initially describe
typical composite honeycomb sandwich structure. Then, we will turn to our
invention of a method of reliably adhering the tiedown plies together.
A composite honeycomb sandwich panel minimizes, eliminates, or
significantly reduces resin flow from the laminates to the core, thereby
permitting a simpler processing cycle that is more robust for the manufacture
of
aerospace structure. Such a sandwich panel 100 (Fig. 1) generally has outer
facesheets or skins 102 adhered to a central honeycomb core 106. The finished
skins 102 comprise laminates of layers of fiber-reinforced organic matrix
resin
CA 02242050 1998-07-02
WO 97/25198 7 PCT/US97/00075
in a cured and consolidated composite form. The core 106 can be paper,
synthetic paper, metal, composite, or the like, as appropriate for the
application.
In panels of the present invention, we obtain higher specific strengths and
higher specific stiffnesses because we reduce core crush during autoclave
curing by incorporating at least one tiedown ply between the core 106 and skin
102 to reduce damaging slippage between the core and skin that otherwise
often occurs.
To prevent flow of resin from the composite laminate skin to the core,
we use an unsupported film adhesive 108 (Fig. 2), a barrier film 110, and a
scrim-supported film adhesive 112 between the skin 102 and the core 106 to
keep resin out of the cells 114 of the core 106.
Fig. 3 illustrates the core-filling problems that can result when a film
adhesive 112 is used alone without the barrier film 110 and film adhesive 108.
Cells 114 of the honeycomb fill with resin 118 which migrates from the
laminates and which thereby depletes the resin in the skin 102. Resin
depletion
impacts structural performance because it reduces ply thickness. Resin
depletion increases total weight since the cell resin 118 is simply waste. In
all
cases, uncontrolled resin flow and depletion makes the panel suspect,
especially to microcracking that can begin in the cell resin 118 during
thermal
cycling and migrate to the fiber-reinforced skin 102, especially at the bond
line
between the skin 102 and core 106.
Fig. 4 illustrates undesirable bulging that can occur if a barrier film 110
is used without a scrim-supported film adhesive 112 to try to eliminate cell
resin 118. Here, a waste resin bulge 120 protrudes downwardly into the cells
114 of the honeycomb core 106. While the resin is contained in the bulge 120,
the skin 102 is still depleted in resin. The flow of resin to bulge 120
imposes
structural performance and weight penalties comparable to the uncontrolled
condition illustrated in Fig. 3.
CA 02242050 1998-07-02
WO 97/25198 8 PCT/US97/00075
As shown in Fig. 2 with the film adhesive 108, barrier film 110, and
scrim-supported film adhesive 12, resin flow is checked without cell resin 118
or resin bulges 120. We discovered, however, that the barrier fiim produced a
slip plane between the laminate skins and the core which often resulted in
core
crush during the autoclave processing cycle. In 22 of 31 test panels, in fact,
we
experienced core crush in our initial trials. This rate of failure was
unacceptable from a cost and schedule perspective. Our tiedown plys in the
chamfer region reduce the frequency of or eliminate damaging core slippage
and the core crush attributable to such slippage.
For bismaleimide laminated skins made with RIGIDITE@ 5250-4-W-
IM7-GP-CSW, RIGIDITE 5250-4-W-IM7-GP-CSX, and RIGIDITE 5250-
4-W-IM7-GP-PW prepreg from Cytec Engineered Materials, Inc. (Cytec), the
film adhesive 108 preferably is 0.015 psf METLBOND 2550U adhesive, also
available from Cytec. The film adhesive provides additional resin to promote a
quality bond between the laminate and barrier film 110. The barrier film 110
preferably is a 0.001 inch thick, bondable grade, surface treated KAPTON
polyimide barrier film capable of withstanding the cure cycle to provide a
resin
impermeable membrane between the skin 102 and core 106. The scrim
preferably is fiberglass, "Style 104" fiber cloth and the film adhesive 112 is
0.06 psf METLBOND 2550G adhesive, available from Cytec. The scrim-
supported film adhesive prevents the barrier film from bulging into the core
cells, thereby retaining the resin in the laminate (i.e., skin layers) so that
the
cured ply thickness is maximized and thereby, we achieve maximum
performance at minimum weight for the panels.
The film adhesive 108, barrier film 110, and film adhesive 112 can be
purchased as a single item from Cytec as METLBOND 2550B-.082 36".
The plys of the skin 102 typically are prepregs of carbon fiber
impregnated with bismaleimide thermoset resin, although the present invention
applies to other resin systems. Tows might be used in place of the prepreg.
CA 02242050 1998-07-02
WO 97/25198 9 PCT/US97/00075
The film adhesive 108 should be tailored to achieve an adequate bond between
the skin 102 and barrier film 110. The honeycomb core generally is HRP
Fiberglass Reinforced Phenolic honeycomb available from Hexcel.
The supported film adhesive and barrier film layers in the sandwich
structure also function as corrosion barriers between the skin 102 and core
106
in the case where the core is metal, such as aluminum, and the skin includes a
galvanically dissimilar material, such as carbon fiber.
Additional information concerning preferred panels is presented in the
technical paper: Hartz et al., "Development of a Bismaleimade/Carbon
Honeycomb Sandwich Structure," SAMPE, March, 1996, which we incorporate
by reference. This paper describes both the Hartz et al. barrier film
improvement, the tiedown ply method, and the adhering method of the present
invention.
The Hartz-type panels provide mechanical and physical edgeband
properties equivalent to solid BMl/carbon laminate (cured at 0.59 MPa (85
psig)). Our tests confum that in our panels the edgeband cured-ply-thickness
is
equivalent to a solid laminate and that the edgeband 160 (Figs. 5 & 6) met the
requirements of the solid laminate nondestructive inspection specification.
The
edgeband and facesheet mechanical performance improved over results we
achieved with sandwich structure lacking the scrim-supported adhesive, barrier
film, adhesive combination. The flatwise tensile mechanical performance also
met design requirements.
Preconditioning the core to eliminate volatile evolution during curing by
heating the core to about 235 C(455 F), prior to laying up the sandwich
panel, especially for phenolic core, eliminates core-laminate disbonding
otherwise caused by outgassing from the core.
Core crush 200 (Fig. 5) occurs in the chamfer region 155 when the
barrier film 110 and core 106 slip relative to the facesheets 102 when
autoclave
pressure is applied and when the resin is melted. As shown in Fig. 5, the
CA 02242050 2005-08-02
WO 97/25198 10 PCTIUS97/00075
barrier films 110 and core 106 have moved toward the right to compress the
core in the chamfer region 155 to produce the core crush 200. The skin 102
has sagged in the edgeband region 160 where the core moved away.
Referring now to Fig. 6, the improved honeycomb sandwich panel
includes at least one tiedown ply 150 in contact with the core 106 along a
chamfer 155. Such a chamfer (i. e: an angled transition in the core, often at
the
edgeband 160) typically occurs around the periphery of the panel, but it might
also occur intermediate of the panel at join lines or hard points where
fasteners
or pass-throughs might be necessary in the assembled structure.
Typically we use a single ply 150 of carbon fiber or fiberglass fabric
with a conventional 0/90 fiber orientation in the fabrication of bismaleimide
panels having 5 or 81b/ft3 HRP core, like Hartz et al. describe. The tiedown
ply 150 functions to prohibit or to limit slippage of the skin relative to the
core
so as to reduce core crush otherwise attributable to the slippage. The tiedown
ply 150 anchors the core with the inherent roughness of the fabric when the
preform is heated during the autoclave processing cycle and the matrix resin
softens, melts, and, for high flow resins, essentially liquefies. With these
panels, we can save between 2.5-4 lb/ft3 of core because we can use lighter
density honeycomb core without suffering core crush. For a fighter, this
change can save as much as 25 lbs per vehicle.
As shown in Fig. 6, the tiedown ply 150 is a narrow, peripheral strip that
contacts the core 106 along at least a portion of the chamfer 155 for about 1
inch overlap with the core 106 and extends outward into the edgeband 160
beyond the trimline 165 of the part. The tiedown ply 150 might be on either
the flat side of the chamfer or the angled surface (which is how we show it in
Fig. 6). The key factor is that the tiedown ply 150 contact the core beneath
the
adhesive and barrier film 110 which is used to bond the laminate skin to the
core. The tiedown ply 150 is cutaway everywhere in the body of the part other
than a narrow peripheral area in the chamfer region, and forms a peripheral
CA 02242050 1998-07-02
WO 97/25198 PCT/US97/00075
frame around the edge of the panel. In this way, the tiedown ply 150 allows an
adhesive interface between the core 106 and the skins 102 in the panel region.
Traditionally, in making a Hartz-type panel, we use four complete cover
sheet tiedown plies 175 in an effort to anchor the layers and the core, and we
show all these plies in Fig. 6. These traditional plies 175 were commonly used
in sandwich panel fabrication prior to introducing the Hartz-type barrier
film,
and we commonly use them all, although we believe we can now eliminate all
but the outer plies and the peripheral, core contacting tiedown ply 150. That
is,
we would use three total plies rather than five, as Fig. 6 shows.
The tiedown plies 150 and 175 extend through the edgeband 160 beyond
the net trim line 165 to anchoring points that we tape to the layup mandrel.
To
further prevent slippage of the tiedown plies, we have incorporated a low
curing (i.e. 121 C for BMI panels) film adhesive 180 between the tiedown plies
just outside the net trim line of the part. The film adhesive 180 eliminates
movement of one ply relative to the others when we apply pressure during the
autoclave curing cycle. Curing at a temperature of about 100 - 150 F below the
curing temperature of the laminate resin, the tiedown adhesive cures before we
need to increase the autoclave pressure and the cured adhesive bonds the
tiedown plys to one another. Using the adhering method eliminates relative
movement of the plys and eliminates facesheet wrinkles and core crush that
otherwise can occur.
The tiedown method saves material, reduces cost, and saves weight,
because it use the "picture frame" peripheral tiedown ply 150 (with the
traditional, internal sheets omitted). The normal tiedown procedure entails
plys
on the outer surfaces of the skins and internally between the skin and
underlying adhesive (Fig. 5). A traditional tiedown system will fail without
the
"picture frame" ply because the barrier film 110 pennits the core to slip. The
Corbett and Smith method will fail occasionally without the adhering method
of the present invention.
CA 02242050 1998-07-02
WO 97/25198 12 PCT/US97/00075
For lightweight core (i. e. 5-8 lb/ft3) with the bismaleimide prepreg and
adhesive system previously described, we hold the chamfer angle to 20 2
By "chamfer" we mean an angled, cut region (a ramp) of the honeycomb
core tapering from fiill thickness to no thickness with a steady slope. A
chamfer is used at the edge band of a composite honeycomb sandwich panel to
provide a smooth transition between the structural body of the panel that has
the embedded honeycomb and a connecting edge band lacking any honeycomb
core. The method of the present invention allows us to use much steeper
chamfer angles than traditional practices often require if one is to avoid
core
crash without one tiedown ply. While we prefer a 20 chamfer, we believe that
we could increase the angle to whatever angle suited the panel design
requirements.
By "autoclave processing" we mean the cycle of elevated temperature
and pressure applied to the panel to consolidate and cure resin in the
laminate
while bonding or otherwise adhering the cured laminate to the honeycomb core.
Our preferred cycle is illustrated in Fig. 7. Our adhesive for the tiedown
plies
cures at about 250 F (121 C) so it cures prior to the increase in autoclave
pressure that can introduce relative motion between layers in the panel.
If core crush occurs, the damage to the panel is generally so extensive
that repair is impossible so the part is scrapped. The cost of today's
advanced
composite resins and reinforcing fibers requires a process that virtually
eliminates core crush. Otherwise, the processing costs are prohibitive. With
panels being designed as close to the design edge as possible, core crush is a
significant issue. The method of the present invention reduces cores crush and
ply movement or wrinkling.
While we have described preferred embodiments, those skilled in the art
will readily recognize alterations, variations, and modifications, which might
be
made without departing from the inventive concept. Therefore, interpret the
claims liberally with the support of the full range of equivalents known to
those
CA 02242050 1998-07-02
WO 97/25198 13 PCT/US97/00075
of ordinary skill based upon this description. The examples are given to
illustrate the invention and are not intended to limit it. Accordingly, define
the
invention by the claims and limit the claims only as necessary in view of the
pertinent prior art.
