Aeronautics and astronautics – Aircraft structure – Airship skin construction
Reexamination Certificate
2000-11-06
2002-09-17
Carone, Michael J. (Department: 3641)
Aeronautics and astronautics
Aircraft structure
Airship skin construction
C244S126000, C244S133000, C244S129600
Reexamination Certificate
active
06450450
ABSTRACT:
FIELD OF THE INVENTION
Pertains to a lightweight inexpensive fabrication that resists catastrophic failure from rapid thermal loading, enabling operation for a moderately extended period after the glass transition temperature, T
g
, has been reached. Specifically disclosed is a stitched composite panel that may comprise the skin of supersonic or hypersonic airframes.
BACKGROUND
Conventionally, surface skin components of airborne vehicles are fabricated of fiber reinforced plastic (FRP) compositions. These include glass fibers such as E-glass or S-glass, boron, carbon, and aramid fibers, the latter commercially familiar as KEVLAR®. The plastic matrix components of these composites include polyester and epoxy resins, polyimides, polyamides, polybutadiene resins and vinyl ester polymers. These plastic matrix resins are classified either as thermoplastics which can be molded repeatedly upon heating to a temperature at which they soften or thermosets which can be cured by cross-linking at anywhere from ambient to 400°-600° F., through chemical cross-linking, or, for a limited class, cured by light such as ultra-violet (UV).
Where structure is subjected to high rate thermal loading, thermosetting resins are employed in the matrix material. For less severe conditions, thermoplastic resins are employed although thermoset resins are used in most lay-ups.
In forming simple shapes such as flat sheets or rectangular polyhedra, lay-up and bagging techniques are straightforward. For example, Kirk-Othmer,
Encyclopedia of Chemical Technology
, 3
rd
Edition, 1981, Supplement Volume at pages 268-270, discloses a process for preparing a flat composite product by placing a carbon-fiber-epoxy pre-(resin)impregnated (prepreg) lay-up on a flat tool surface. The lay-up is covered with breather plies and a nylon bag is placed over the breather plies and sealed at its edges to the tool surface. Air is evacuated from the lay-up by pulling a vacuum on the bag. The assembly is placed in an autoclave, heated, and pressurized to cure.
A similar approach is followed in molding three-dimensional composite structures. For example, U.S. Pat. No. 3,962,394, issued to Hall, discloses a tubular mandrel coated with a resin fiber layer and surrounded by a compression sleeve formed of a thin film of nylon or rubber that is perforated with holes and split lengthwise. A layer of absorbent material is placed around the split compression sleeve and this assembly is surrounded by a bladder that is sealed at both ends to the tubular mandrel. The bladder is evacuated in order to cause the compression sleeve to compact the layers and expel trapped air and excess resin from the fiber-resin material through the holes in the compression sleeve.
Typically, 5 to 10 plies of fiber mats are laminated using intervening applications of resins to arrive at the final laminated panel. Conventionally, fibers within each fiber ply are oriented in a plane paralleling the surface of the ply. The fiber ply may be of a continuous strand type as in the case of a filament wound on a molding tool. Alternatively, the fiber material may be formed of a plurality of generally parallel continuous fibers, e.g. in the form of a “unidirectional tape,” or it may comprise fibers aligned in a unidirectional manner.
In addition, the fiber mat may take the form of a braided structure in which the fibers extend predominantly along one direction but are woven together, normally to provide an angle between strands, the “braid angle,” of about 15°-45°. A fabric in which fibers are interconnected by cross-strands intersecting at 90° is also used. The fiber layers may be in the form of prepregs in which fibers are impregnated with an uncured resin that is later cross-linked, providing the matrix material.
Various procedures are available for forming FRP composites. Fiber layers, are “laid-up” on a forming tool and resin is sprayed on them. Layers are added until the desired thickness is achieved and the resulting “lay-up” is then cured. The lay-up is pressed lightly to uniformly distribute the resin onto the fibers. Further, curing can occur under pressure. Reference Kirk-Othmer, Encyclopedia of Chemical Technology, 3rd Edition, 1981, Vol. 13, pp. 968-978 and Supplement Volume, pp. 260-281.
In some composites, it is desirable that resin material be uniformly applied throughout the thickness of the fiber plies. In others, pronounced layering, i.e., lamination, is desired. For example, U.S. Pat. No. 4,269,884, Fiber Reinforced Multi-Ply Stampable Thermoplastic Sheet, issued to Dellavecchia et al, May 26, 1981 discloses a process of forming a stampable thermoplastic sheet with several discrete layers. In the '884 patent, outer layers are formed of a thermoplastic resin that may contain optionally up to 50% of a filler and up to 45% of nonsiliceous fibers having a length ranging from about 0.01″-0.75″. The fibers are oriented in the plane parallel to the surface of the sheet. Below the outer layer, a molten resinous sheet is provided that allows the internal fiber mats to be impregnated by the resin. The fiber mats are supported by an internal screen. The fiber and resin layers pass through rollers that apply pressure of between 1000-1500 psi thus bonding the several layers and impregnating the fibers with the molten resin. The '884 patent prevents migration from one layer to the next, specifically the process prevents migration of the long reinforcing fibers to the outer resinous layer and migration of the short fibers, if present in the outer resin layer, into the reinforcing layer.
U.S. Pat. No. 5,871,604, Forming Fiber Reinforced Composite Product, issued to Hohman, on Feb. 16, 1999 provides improved lay-up methods wherein layers of fibrous reinforcing material and unset resinous matrix material are first interleaved. The matrix material is formed of unset resin and short reinforcing fibers with an average particle length of 0.2-0.6 mm. The fibers provide a weight ratio of short fibers-to-resin within the range 0.4-1. The reinforcing fibers are longer than the short fibers in the matrix material, ranging from 3 cm to the actual length of the final product. The fiber and matrix layers are integrated in order to cause the resin and short fibers that comprise the strengthened matrix material to enter the interstitial space of the longer reinforcing fibers.
As alternate fiber and matrix layers are built up, a pressure gradient is established from the outside of the built up layers to the forming surface that causes the resin to flow across the interfacial boundaries of the resin and fiber layers. This flow across the resin matrix-fiber interface promotes orientation of the short fibers in a direction across the interfaces so that the short fibers enter the interstitial spaces between the reinforcing fibers. After buildup, the resin is solidified, producing the FRP composite.
At speeds of Mach 3
+
, the free flight temperature profile (in the atmosphere) drives the design of the “skin” of an airframe such as may be used for a missile. Several design alternatives are available to address the “moderate” peak temperatures expected, i.e., >800° F. for ~15 minutes. Conventional designs use an ablative layer, such as cork, that carries no structural load but increases weight, bulk, and reduces payload for the same external dimensions that would be used for a subsonic design. To achieve an ablative material that would also minimize the loss of payload and other performance parameters, specialized costly materials requiring high cost secondary manufacturing processes have been proposed as one solution. One alternative that also addresses cost is to “adjust” the strength of a composite matrix to enable it to operate at or above the T
g
of its major component(s) for short periods, i.e., minutes as opposed to hours, with a nominal operational window of 15 minutes.
The composites soften at their T
g
so that if they are being used as a load-bearing member, they need to be “stiffened” to permit use at or above their T
g
, even for short durations.
Chuang Chun-Hua Kathy
MacDonald Frank D.
Stickler Patrick B.
Warren Ryon C.
Baugher Earl H.
Carone Michael J.
Foster Laura R.
Semunegus Lulit
The United States of America as represented by the Secretary of
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