Plastic and nonmetallic article shaping or treating: processes – Direct application of fluid pressure differential to... – Producing multilayer work or article
Reexamination Certificate
1997-01-29
2004-02-17
Staicovici, Stefan (Department: 1732)
Plastic and nonmetallic article shaping or treating: processes
Direct application of fluid pressure differential to...
Producing multilayer work or article
C264S257000, C264S258000, C264S512000, C156S173000, C156S175000
Reexamination Certificate
active
06692681
ABSTRACT:
BACKGROUND OF THE INVENTION
1. Field of the Invention
This invention relates generally to composite material structures and, more specifically, to fiber placement methods and apparatus for manufacturing composite material shells. In particular, this invention relates to composite material shells, such as aircraft fuselage components, formed using fiber placement and other methods employing a removable mandrel and a bladder with or without an integral caul sheet.
2. Description of the Related Art
The fuselage of an airplane or other similar vehicle is generally a thin shell of revolution. In the case of an airplane, one of the significant loading conditions for a fuselage is circumferential tension resulting from internal pressure. Structurally it is most efficient to carry this loading with a structure having a continuous diameter or hoop without any axial joints. From a manufacturing sense each joint in a structure tends to add cost. Also, from a manufacturing sense each extra component or detail tends to add cost.
Composites have proven to be very useful materials, especially in the field of aviation. Weight is a very important and sensitive subject and any method to limit or reduce it is valuable. In addition, structures of composite materials are usually thinner, allowing for increased internal space or decreased area.
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. As a result, composite materials have been used to fabricate a wide variety of structures including, most notably, aircraft structures (such as fuselage shell components, wing sections, tail sections, etc.). However, these composite structures have typically been manufactured by time consuming application methods, such as hand placement.
In addition, when used to manufacture aircraft structures, such as fuselage components, composite structures have typically been manufactured in many separate parts, such as fuselage halves split down the longitudinal axis, that must be bonded or fastened together, typically using a flush joint. For example, in one method, a forward fuselage section with four or more separate composite structure components has been manufactured. Machining and assembling of flush joints into a single uniform component typically requires a substantial amount of time to achieve a uniform and consistent flush joint. In addition to extra time, each flush joint adds additional material and weight to the assembled aircraft component. Therefore, the greater the number of separate parts required to construct a single component, such as fuselage component sections, the larger the amount of time and the greater the amount of weight added to the assembled component.
In an effort to reduce composite part assembly time and to produce lighter weight composite parts, fiber placement (or tow placement) methods have been developed. Such fiber placement methods may include computer control integrated with a fiber placement machine. Operation of such a machine to place tow filaments on a mandrel to form composite structures is known in the art. Fiber placement methods involve the automated placement (typically by winding) of filaments (such as fibrous ribbons or tows which are pre-impregnated with a thermal set resin material such as epoxy) onto a mandrel to produce a component, such as a tube-shaped part. These fibers are typically placed at varying angles and in segments of varying width. A fiber tow is essentially a ribbon of carbon fiber, typically between about ¼″ and about ⅛″ wide. Using a conventional fiber placement machine, multiple tows are transported to a movable payoff head and applied to a mandrel surface using a roller. Typically, a payoff head includes an automatic cutting system for cutting and restarting individual tows. In addition, typical fiber placement machines include heating devices to vary the temperature and, therefore, the properties of the tows as they are applied. Means for controlling pressure applied to the tows and mandrel during fiber placement are also typically employed.
Although fiber placement processes may be used to produce composite structures of varying dimension and size more quickly and efficiently than other methods, current fiber placement techniques suffer from complications relating to mandrel construction and removal of the mandrel after fiber placement has occurred. In particular, segmented mandrels have been provided having segments that are joined during fiber placement and disassembled after curing. Segmented mandrel designs suffer from numerous problems, including expansion of the mandrel material during heat curing, leakage between mandrel segment joints, and time and effort involved in the assembly and disassembly of mandrel components.
In the construction of composite structures, and aircraft composite structures in particular, interior dimensions of a structure are of particular concern. Although fiber placement techniques have been used to produce aircraft fuselage shell components, these shell components have required cylindrical support frames and elongated longeron support members that serve to support the outer fuselage structure. These frames and longerons are typically wider or deeper than a composite fuselage wall of sandwich construction, and therefore serve to reduce the interior diameter of an aircraft fuselage. In other cases, fiber placed composite structures have been manufactured in non-continuous separate parts, such as separate axial fuselage half or quarter panels, that are assembled to form a single cross sectional shape. These structures suffer from the cost and weight problems described above for other multi-piece composite structure components.
Consequently, a need exists for simplified methods and apparatus for forming relatively large, single piece composite parts, such as aircraft fuselage components. In particular a need exists for simplified methods of mandrel installation and removal. A need also exists for a method of manufacturing composite shell components, such as aircraft fuselage parts, which do not require internal frames or bracing and which have an increased internal diameter.
SUMMARY OF THE INVENTION
In one aspect, this invention is a tool for use in forming a composite body. The tool includes a mandrel body having an outer surface and a bladder having outer and inner surfaces. The bladder has a shape and dimensions complementary to the outer surface of the mandrel body so that the bladder may be fitted around the outer surface of the mandrel body. The tool also includes at least one caul sheet section having an inner surface coupled to the outer surface of the bladder so that the caul sheet section overlays at least a portion of the outer surface of the bladder.
In another aspect, this invention is a tool for use in forming a composite body. The tool includes a mandrel body having an outer surface and one or more fluid openings defined in the mandrel body outer surface. The tool also includes a mandrel body fluid system capable of supplying pressurized fluid or a vacuum to the one or more openings in the mandrel body outer surface.
In another aspect, this invention is a method of forming a composite body including the steps of providing a mandrel body having an outer surface, an elongated shape, and a longitudinal axis. The method includes placing a plurality of fibers on the outer surface of the mandrel body to form an uncured body. In this method, the fibers are placed around the mandrel body in a plurality of discontinuous segments juxtaposed in relation to each other. The discontinuous segments are capable of moving in relation to each other so that the uncured body is expandable from within.
In another aspect, this invention is a method of forming a composite body, including the steps of providing a tool and placing a
Lenzen, Jr. Glenn H.
Raytheon Aircraft Company
Schubert William C.
Staicovici Stefan
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