Plastic and nonmetallic article shaping or treating: processes – Combined
Reexamination Certificate
2000-12-28
2003-10-28
Staicovici, Stefan (Department: 1732)
Plastic and nonmetallic article shaping or treating: processes
Combined
C264S241000, C264S255000, C264S257000, C264S258000, C264S264000
Reexamination Certificate
active
06638466
ABSTRACT:
BACKGROUND OF THE INVENTION
1. Field of the Invention
This invention relates generally to composite material structures and, more specifically, to methods for manufacturing composite material structures. In one particular embodiment, this invention relates to single cure composite material structures formed with one or more separable and replaceable portions using resin transfer molding methods.
2. Background
Composite materials, such as carbon fiber present in an organic matrix, have been used to produce corrosion resistant and light weight structures. These structures typically weigh about 25% less than structures made of lightweight metals, such as aluminum, while at the same time offering similar strength to these metals. Composite structures have offered performance improvements in terms of lower weight, increased stiffness, and long fatigue life, and have been used to fabricate a wide variety of structures including, for example, aircraft structures (such as fuselage shell components, wing sections, tail sections, etc.). These composite structures have typically been manufactured by time consuming methods, such as hand placement and autoclave cure. For example, in one process for manufacturing hollow landing gear doors, resin impregnated prepreg materials have been hand laid-up on mandrels with a non-stick dividing material placed between sections of the prepreg material so as to isolate one wall of the landing gear door structure from the remaining walls of the structure during lay-up and cure, thus forming a removable wall section through which the mandrels could be later removed, and then the wall section replaced to form the finished door. However, traditional hand lay-up autoclave composite construction is labor intensive, uses high cost resin-impregnated raw materials that require special storage, and therefore often results in costly finished structures.
Many composite structures, such as aircraft wings and tails, are manufactured in multiple and separate parts so as to provide needed access to the interior of the structure for insertion of parts prior to final assembly. For example, traditional composite structure methods of producing a complex aircraft part such as a wing typically involve separate fabrication and curing of skeleton and skin components, and involve assembly after all internal parts are installed. Not only is this traditional method labor intensive, but the skins typically do not exactly match the skeleton, and customizing measures are usually needed during assembly to correct the gaps or interferences between the skins and the skeleton, requiring a substantial amount of time to achieve a uniform and consistent fit.
Single cure closed tool Resin Transfer Molding (“RTM”) is a method of forming composite structures in which one piece composite parts are cured under high temperature and high pressure in a closed tool using a single cure cycle. In a conventional RTM procedure, dry carbon fiber fabrics are applied to a mandrel, loaded into a tool, then injected with resin and cured to form a completed structure, thus eliminating the need for assembly of the traditional individually fabricated parts such spars and skins. For example, using one single cure RTM method in the manufacture of wing flaps, dry carbon fiber braided material is fitted over or wrapped around inner mandrels and placed inside a mold. The mold is closed around the mandrels and placed in a heated press. A vacuum is then drawn on the interior of the mold and a predetermined volume of heated resin injected into the mold to saturate the dry braided material. A relatively high pressure is applied to the resin, and the tool is heated to cure temperature. After the cure cycle is complete, the mold is removed from the press, and the mandrels and composite part are removed from the mold to cool, prior to removal of the mandrel from one of the open ends of the composite part.
Single cure closed tool RTM methods produce structures which have a relatively low labor content, and because dry fiber material is used no special storage of resin impregnated fiber material prepreg is required. Because a one piece composite structure is produced in a single step, considerable labor cost is saved by eliminating assembly steps. This is in part because of the ability to mold completed assemblies and also because parts may be cured in the molding tool, instead of requiring that the parts in the molding tool be move to an autoclave for curing. Use of a closed tool RTM method enables a closed structural section (e.g., box-like structures) to be fabricated in one piece, eliminating the assembly steps necessary when the spars and skins are fabricated separately. The tools for RTM curing typically allow for the tool itself to be heated and allow for pressure to be applied hydraulically to the resin whereas traditional methods of composite fabrication rely on the autoclave to apply heat and pressure.
Single cure closed tool RTM processes have been used to produce simple aircraft composite structures (e.g., wing flaps, drain mast) more quickly and efficiently than other methods. RTM processes have been tried using combinations of liquid injected resin and prepreg materials, and a combination of injected liquid resin and prepreg materials has been employed to produce a landing gear door. However, one factor that has hampered the use of closed tool RTM processes to produce larger and more complex hollow aircraft structures (e.g., wings and tails) is lack of access to the inside of the-completed composite structure. As previously described, access is typically required for such structures in order to install supporting structures such as ribs, and equipment such as fuel pumps and gages. For example, ribs may be desirable in an aircraft wing fuel tank to prevent the fuel from “sloshing” out towards the wing tip when the airplane rolls into a one wing down attitude. However, the current technology single cure cycle RTM part has no accessibility except from each end.
SUMMARY OF THE INVENTION
Disclosed herein are single cure RTM processes that may be employed to produce high quality composite structures having relatively lightweight and relatively low fabrication costs, and which offer the advantages of a single in-tool cure cycle while at the same time allowing internal equipment and/or internal structure to be added to a post cure structure. In one embodiment are methods that advantageously allow the use of single cure closed tool RTM process to fabricate relatively large and complex closed section structures such as an aircraft wing or tail fin (e.g., box like structures with a front spar, a rear spar, and with top and bottom skins) in one piece, while permitting full access to virtually any pre-determined part of the internal structure, thus eliminating extra assembly steps necessary when the spars and skins for such components are fabricated separately.
Advantageously, using the disclosed RTM process a relatively low cost one piece structure (e.g., wing or tail fin section) may be fabricated in a single cure cycle, but in a manner that allows access to the inside of the structure after cure (e.g., to add internal ribs to control fuel slosh/pressure and to add equipment such as fuel pumps and gauges). Surprisingly, the disclosed RTM process may be used to produce a composite structure having a separable part (e.g., separable top skin, separable bottom skin, or any other portion of the structure thereof) that is substantially separable from the rest of the structure (e.g., the skeleton) after cure. In one embodiment, the separable part (or any portion thereof) may be easily reconnected to the structure by virtue of the fact that it is molded as a matching part of the structure during the cure cycle.
When the disclosed single cure RTM processes are used to manufacture an aircraft wing or tail structure having a separable skin portion, the separable skin and the skeleton are cured in the same tool and with the same cure cycle. Therefore, the disclosed processes offer a major benefit over traditional processes that re
O'Keefe, Egan & Peterman, LLC
Raytheon Aircraft Company
Staicovici Stefan
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