Stock material or miscellaneous articles – Coated or structually defined flake – particle – cell – strand,... – Staple length fiber
Reexamination Certificate
2001-03-01
2004-06-01
Jones, Deborah (Department: 1775)
Stock material or miscellaneous articles
Coated or structually defined flake, particle, cell, strand,...
Staple length fiber
C428S035700, C428S036100, C428S036900, C428S188000, C052S793100
Reexamination Certificate
active
06743504
ABSTRACT:
BACKGROUND OF THE INVENTION
1. Field of the Invention
The present invention relates generally to the field of composite structures and, more particularly but not by way of limitation, to composite structures having skins separated and stiffened by hollow hats, each of which incorporates an integral co-cured fly away hollow mandrel used in laying up and curing the structure.
2. Prior Art
There is a growing trend in the aerospace industry to expand the use of advanced composite materials for a diverse array of structural and dynamic aerostructural applications because of the strength-to-weight advantage provided by composite materials. One particular application for the use of such advanced composite materials, such as graphite or an aromatic polyamide fiber of high tensile strength that are embedded in a resinous matrix, e.g., an epoxy, is for airfoil structures that are composed of skins separated and stiffened by a honeycomb core material. In the instance of an aerospace article such as a fan cowl, one or more stiffening members are affixed to the outer skin and covered with an inner skin for efficiently transmitting and/or reacting axial and/or bending loads to which the fan cowl is subjected.
There are two techniques currently employed for bonding through autoclave processing a composite stiffening member in combination with composite face layers: (1) the secondary bonding method, and (2) the co-cured bonding method. Both methods are disadvantageous in requiring costly non-reusable tooling and/or costly and tedious manufacturing steps.
A typical composite sandwich panel intended for use as an aerostructure part is normally fabricated using two autoclave-cured inner and outer composite skins that are formed by using a curing cycle with heat, pressure, and a unique tool for each skin. A sandwich panel is then made up using a composite bond jig, tool or fixture with the pre-cured face skin laid-up on the bond jig tool followed by a ply of film adhesive. A honeycomb aluminum or non-metallic core of a given thickness is placed on the face skin, another ply of film adhesive is applied, and finally the previously pre-cured inner skin is placed on the adhesive film. The bond jig that is used to fabricate the sandwich panel is usually the same tool that was used to create the outer composite skin. A plurality of closure plies of uncured composite material are laid up. Next, the assembled sandwich panel is cured during its final assembly stage. The entire sandwich panel is then vacuum bagged to the composite bond jig and again cured in an autoclave under high pressure and heat.
Thus, at least three very expensive and labor intensive fabrication and cure cycles have gone into the production of the exceptionally strong and lightweight composite honeycomb core sandwich panel. At least two different and expensive tools are needed in this process. Manufacturing flow time is very long, energy use is high, and the manufacturing floor space required is considerable.
The second method referred to above, the co-curing method, involves curing the composite inner and outer skins that are laid up with a layer of adhesive film and honeycomb core in one cure cycle in the autoclave. A co-cured panel is desirable in that it is less expensive to fabricate—only one bond jig tool is required, only one cure cycle is needed, the method is less labor-intensive, less floor space is required, and a much shorter manufacturing flow time is achieved. However, co-curing an aerostructure panel has never achieved wide-spread acceptance because of a large loss of panel strength and integrity, which is due to the lack of compaction of the composite plies placed over and under the honeycomb core. The composite plies dimple into the center of each core cell with nothing but the cell walls to compact the composite skins. The only way to overcome this “knockdown” characteristic is to add extra plies, which creates both unwanted weight and added cost. Thus, because of these constraints co-cured aerostructure panels are not widely manufactured in the aerospace industry.
There are other particular problems when a honeycomb core element is used to provide a stiffening element for an aerospace component. As Hartz et al. have described in U.S. Pat. No. 5,604,010 concerning a “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 skin plies to account for the resin losses to the honeycomb core. To achieve the designed 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 costly in modern aircraft and because the resin 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, it has been learned that micro-cracking that originated in the migrated resin can propagate to the bond line and degrade mechanical performance. Such micro-cracking potential has a catastrophic threat to the integrity of the panel and dictates that flow be eliminated or at least controlled.
Unfortunately, the use of a honeycomb core as a stiffener for elements in a aerostructure component, such as a structural panel, has other deleterious effects. Two of the greatest drawbacks to an aluminum core are its inherent significant cost and susceptibility to corrosion. To minimize galvanic corrosion of the core caused by contact with the face skins, isolating sheets are interposed between the aluminum core and the face skins. Also, the aluminum core is expensive and also must be machined to a desired shape in a costly process. The honeycomb core may also be subject to crush during manufacture, which imposes a limit on the pressures that may be used in autoclave processing. Thus, the processing of an aerospace advanced composite article is limited to an autoclave pressure of not greater than 45 psi, rather than a higher pressure that would increase the strength of the resultant advanced composite article. Also, the honeycomb core, if damaged in use, has a spring-back property, which makes the detection of such damage more difficult.
In providing reinforcing mandrels for stiffener elements, such as hat sections, for aerospace advanced composite structural panels, it is also known to provide a composite stiffening member in the form of a polyamide foam mandrel fabricated by machining a core mandrel to a desired shape. Obviously, the machining of the core mandrel is expensive and time consuming and further introduces the problem of properly bonding the core mandrel to the inner and outer skins.
Therefore, a great need has arisen for a practical method of readily producing stiffened, fiber-reinforced composite structures useful in the construction of integrally stiffened components for aerospace applications, which are cost and labor efficient and which save time in the fabrication process.
Accordingly, it is an object of the present invention to provide a method for fabricating aerostructure advanced composite articles that eliminates a honeycomb core as a spacing and stiffening element, provides a lighter weight assembly, and is easier to repair. Another object of the present invention is to reduce the lay-up cost of known advanced composite co-cure assemblies by at least 15% and to increase assembly strength over previously known co-cure assembly methods by being able to utilize high pressures in autoclave processing. Yet another object of the present invention is to improve the quality of co-cured advance composite assemblies and thereby increase customer satisfaction. A further object of the present invention is to provide a process that provides an assembly that can be manufactured in one ma
Allen Gregory A.
Belisario David J.
Boss Wendy
Goodwin & Procter LLP
Jones Deborah
Rohr Inc.
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