Stock material or miscellaneous articles – Structurally defined web or sheet – Honeycomb-like
Reexamination Certificate
1998-09-14
2001-01-30
Lorin, Francis J. (Department: 1775)
Stock material or miscellaneous articles
Structurally defined web or sheet
Honeycomb-like
C156S292000, C156S382000
Reexamination Certificate
active
06180206
ABSTRACT:
TECHNICAL FIELD
The present invention relates to aircraft fixed leading edges that contain composite honeycomb sandwich panels, and to an improved manufacturing process to reduce core crush in these panels. In a preferred embodiment, a prepreg ply having woven fabric as the reinforcement is positioned between the honeycomb core and resin impregnated fabric sheets (prepregs) that butt an erosion strip and that form the lower skin. The ply extends over the erosion strip and can be tied down to reduce core crush by fixing the relative position of the core and first ply in the laminated skins.
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 their high stiffness-to-weight (i.e., “specific stiffness) and strength-to-weight (i.e., specific strength) ratios. Typical honeycomb core sandwich panels are described in U.S. Pat. Nos. 5,604,010; 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 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. I incorporate these patents and article by reference.
U.S. Pat. No. 5,445,861 by Newton et al 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.
As Hartz et al. described in U.S. Pat. No. 5,604,010, large amounts of resin can flow into the core of sandwich structure during the autoclave processing cycle of a high flow resin system. 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 reduced the thickness of the cured plies which compromises the mechanical performance. To achieve the desired performance and the corresponding laminate thickness, additional plies were necessary with resulting cost and weight penalties. Because the weight penalty was severe in terms of the impact on vehicle performance and cost in modern aircraft and because the flow was a relatively unpredictable and uncontrolled process, aerospace design and manufacture dictated that flow into the core be eliminated or significantly reduced. In addition to the weight penalty from resin flow to the core in these panels, microcracking that originated in the migrated resin could propagate to the bond line and degrade mechanical performance. Such microcracking potential posed a catastrophic threat to the integrity of the panel and dictated that flow be eliminated or, at least, controlled.
Flow of resin from the laminates to the core occurred because of viscosity reduction of the resin (i.e., thinning) at the elevated processing temperatures. Therefore, prior art attempts to solve the flow problem generally focused on retaining the ambient temperature viscosity of the resin at the curing temperatures. For example, the processing cycle was altered 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, manufacturers tried 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 occurred, and microcracking still was a concern. Also, facesheet porosity increased 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.
The Hartz et al. process of U.S. Pat. No. 5,604,010 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 allowed designers to reduce the margin of safety, and reliably fabricated panels in which they had structural confidence.
Smith and Corbett of Boeing discovered that core crush frequently occurred in the chamfer region of honeycomb core when curing a Hartz-type panel having a scrim-supported barrier film, particularly when using lighter weight core materials. They reduced 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. The tiedown ply extended beyond the margin of the part to a place where it could be adhered to the layup mandrel. They described their invention in U.S. patent application Ser. No. 08/616,903. The Smith and Corbett method uses one or more tiedown plys in contact with the core (at least 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 such movement.
Using tiedown plies in Hartz-type panels allowed Boeing to minimize the weight of the panels. Weight is reduced by using lower density core and by trimming the internal area of the tiedown plies so that they frame the core and only slightly overlap the chamfer of the underlying core. By controlling core slippage with the tiedown plies, lighter density honeycomb core is used to produce structures without costly scrap due to core crush. Manufacturing costs are minimized 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 plies also provide a path for egress of volatiles from the core and to equalize the pressure between the core and autoclave.
SUMMARY OF THE INVENTION
Composite sandwich panels for fixed leading edges on modern transport aircraft are manufactured with reduced core crush by including a full-surface tiedown prepreg ply as the facing ply between the core and lower skin laminate that abuts the panel's erosion strip. With this ply arrangement, the core cannot slide over the lower skin laminate during autoclave curing. The facing ply extends over the erosion strip into the margin of the part and can be adhered to the mandrel. The facing ply (and overlying plies t
Hammar John C.
Lorin Francis J.
The Boeing Company
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