Method for reducing core crush

Details Method for reducing core crush Method for reducing core crush

1998-10-01

2004-03-02

Jones, Deborah (Department: 1775)

Adhesive bonding and miscellaneous chemical manufacture

Surface bonding means and/or assembly means therefor

Chamber enclosing work during bonding and/or assembly

C156S285000, C156S286000, C156S087000, C156S090000

Reexamination Certificate

active

06698484

ABSTRACT:

TECHNICAL FIELD
The present invention relates to composite honeycomb sandwich structure having improved resistance to core crush. In a preferred embodiment, we adhere resin impregnated fabric sheets forming outer, opposed, skin surfaces to a chamfered honeycomb core, optionally, with an intermediate barrier film to eliminate resin flow from the skins to the core. We interleave a peripheral tiedown ply between the core and skins along the chamfer to reduce core crush.
BACKGROUND ART
Aerospace honeycomb core sandwich panels (having composite laminate skins cocured with adhesives to the core through autoclave processing) find 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. Pat. Nos. 5,284,702; 4,622,091; and 4,353,947, which we incorporate by reference. Alteneder et al.,
Processing and Characterization Studies of Honeycomb Composite Structures
, 38th Int'l SAMPE Symposium, 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 finish. We incorporate this article by reference.
As Hartz et al. described in U.S. Pat. No. 5,604,010 entitled “Composite Honeycomb Sandwich Structure,” with a high flow resin system, large amounts of resin can flow into the core during the autoclave processing cycle. Such flow robs resin from the laminate, introduces a weight penalty in the panel to achieve the desired performance, and forces over design of the laminate plies to account for the flow losses. The resin loss from the laminate 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 performance and cost in modern aircraft and because the flow is a relatively unpredictable and uncontrolled process, aerospace design and manufacture dictates that flow into the core be eliminated or significantly reduced. In addition to the weight penalty from resin flow to the core, 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 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 incorporate the Hartz et al. Patent by reference.
U.S. Pat. No. 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
(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. We incorporate this patent by reference.
The Hartz et al. process of eliminates 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. Hartz et al. 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. Hartz et al. 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.
We discovered that core crush frequently occurred in the chamfer region of honeycomb core when we cured a panel having a scrim-supported barrier film, particularly when we tried to use lighter weight core materials. We subsequently discovered that we could 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.
SUMMARY OF THE INVENTION
Our invention relates to a method for reducing core crush in composite honeycomb sandwich structure, especially panels of the general type Hartz et al. described. We incorporate one or more tiedown plys in contact with the core in its chamfer regions around the periphery to eliminate slippage of the skin over the core during autoclave curing, and, thereby, to eliminate core crush that results from the movement.
Our invention also relates to the resulting composite honeycomb sandwich structure. There, we usually minimize the weight by trimming the internal area of the tiedown ply(s) so that it frames and slightly overlaps the chamfer of the underlying core. By controlling core slippage, we are able to use the 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 tiedown ply also provides a path for egress of volatiles from the core and to equalize the pressure which permits us to maintain the correct pressures within the core to further reduce core crush.


REFERENCES:
patent: 4937125 (1990-06-01), Sanmartin et al.
patent: 5685940 (1997-11-01), Hopkins et al.

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