Adhesive bonding and miscellaneous chemical manufacture – Methods – Surface bonding and/or assembly therefor
Reexamination Certificate
1999-02-19
2001-02-20
Ball, Michael W. (Department: 1733)
Adhesive bonding and miscellaneous chemical manufacture
Methods
Surface bonding and/or assembly therefor
C156S169000, C156S172000, C156S182000, C156S184000, C156S195000, C156S425000, C244S123800
Reexamination Certificate
active
06190484
ABSTRACT:
CROSS-REFERENCE TO RELATED APPLICATIONS
Not Applicable
STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT
Not Applicable
REFERENCE TO A MICROFICHE APPENDIX
Not Applicable
BACKGROUND OF THE INVENTION
1. Field of Invention
This invention describes a manufacturing process for a monolithic composite wing, more particularly, that reduces manufacturing cost, increases structural strength, enhances structural integrity and reduces structural weight. In another manner of speaking, this invention provides an automated fabrication method that increases production rate and reduces the life cycle cost of the aircraft.
2. Description of Prior Art
Composite materials are made of thin fibers, not necessarily limited to carbon, boron or kevlar, imbedded in an epoxy like resin. These composite materials possess, generally speaking, a high strength to weight ratio. That means, a small amount of composite material can support large forces when properly used. Hence, composite materials play a significant role in aerospace industries as well as in other commercial products.
Heretofore, the aircraft wing manufacturing processes used mechanical fasteners and manual lay-up of composite fabrics. This process is referred to as the built-up fabrication methodology, wherein upper and lower skins are attached to a framework of spars and ribs by means of mechanical fasteners. The disadvantages of the built-up fabrication process are the involvement of extensive manual labor and the need to assemble a large number of parts. Another serious drawback of the built-up process is that the load transmission efficiency is very low, because, the mechanical fasteners induce large stresses leading to delamination and breakdown of structural members.
The prior art abounds with examples of composite wing manufacturing. A number of patents disclosing typical inventions pertinent to the present invention will now be presented.
U.S. Pat. No. 3,962,506 to Edumond O. Dunahoo (Jun. 8, 1976) discloses a method for manufacturing multi-chambered, airfoil shaped, composite helicopter blades. The essence of this innovation lies in forming an airfoil shaped blade adjoining a number of flexible composite tubes that are then expanded by pressurization. Once all the cellular tubes are adjoined, a final wrapping of all the tubes is performed. Helicopter blades mostly experience radial tension and the tubes proposed in this concept adequately provide the needed strength. However, the same fabrication process cannot be applied to manufacture aircraft wing structures because of higher bending and torsion stiffness requirements. This fabrication process does not provide sufficient torsion stiffness for aircraft wings. Moreover, it is not possible to inflate aircraft wing cells to conform to the airfoil shape. This is the major drawback of this method.
U.S. Pat. No. 4,662,587 to Philip C. Whitener (May 5, 1987) and U.S. Pat. No. 4,565,595 to Philip C. Whitener (Jan. 1, 1986) disclose methods of manufacturing composite aircraft wing structures. The construction details of the invention are as follows:
Fabrication of complex multi-element spar-skin joints,
Assembly of non-integral mandrels that are required to form the cells conforming to local cross sections of the wing,
pre-fabrication of honeycomb spars,
pre-fabrication of top and bottom skins,
bonding of honeycomb spars and skins in a contiguously manner,
contiguously assembled framework is then wrapped around by pre-impregnated fibrous materials,
Pressurization of mandrels to apply pressure to the side walls and spars.
heating and curing the wing mold
Thus, the fabrication process involves a number of manual operations and hence precludes automation. A large number of fasteners are required to attach the skins to the honeycomb spars. This means, thicker skins and spars must be used to avoid excessive stress levels around fasteners. Other disadvantages are:
1. the structure is weak in torsion, because the skins and spars are joined by means of chemical glue that is weak in shear,
2. spar-skin joining process involves many elements and consequently many manual operations,
3. removable mandrels are required that increases manufacturing cost,
4. automation is applied only at the final wrapping and bonding process, while other operations involve intensive manual labor.
Hence, the art governed by the U.S. patent
4
,
565
,
595
yields a torsionally weaker structure and also increases manufacturing cost due to several manual operations in the wing assembly.
U.S. Pat. No. 5,216,799 to Paul Charnock et al., (Jun. 8, 1993) discloses a method of constructing a composite wing in which upper and lower skins are bonded to a preassembled framework of ribs and spars thus, obviating the need for metal fasteners and reducing assembly time. Essentially, this is the conventional approach used in the aerospace industry. No effort has been made to take advantage of composite fiber directional strength and minimum weight considerations. Hence, this invention falls short of being claimed as the economical one.
U.S. Pat. No. 5,735,486 to Matthias Piening, et al. (Apr. 7, 1998) discloses a method of constructing a wing comprising of upper and lower skins and spars wherein unidirectional stiffening members, called stringers, extending longitudinally, are mounted on the inside of the wing shells. The primary innovation lies in the construction of the stringers embedded into the wing shell. This fabrication process involves manual lay-up and assembly of numerous components.
U.S. Pat. No. 5,332,178 to Sam B. Williams (Jul. 26, 1994) discloses a method of fabricating a composite wing comprising of plurality of hallow triangular shaped spars arranged in juxtaposed parallel relation defining an airfoil with composite skins disposed about said assembled members. This is a novel concept except that it produces a non-optimal structure having a lower value of strength to weight ratio. Moreover, this construction results in lesser fuel volume and becomes difficult to manage the change in center of gravity as the fuel is dispensed in flight.
U.S. Pat. No. 5,496,002 to Rainer Schutze (Mar. 5, 1996) discloses a method of fabricating a wing comprising thin walled shells, pre-formed skins and circular rods as stringers. Thin walled shells of various diameters are arranged in such a manner as to fill the volume of the airfoil. In reality, this concept is a variation of another disclosure made in U.S. Pat. No. 5,332,178 to Sam B. Williams (Jul. 26, 1994). The disadvantages of this method are identical to those discussed above.
U.S. Pat. No. 4,538,780 to Richard D. Roe (Sep. 3, 1985) discloses a method for constructing ultra-light composite wing structures used in the design of missile fins and target drones. The fins are constructed from lightweight foam material and reinforced plastic skins. This approach does not apply to aircraft wing structures. It was with this knowledge of the foregoing state of technology that the present invention has been conceived and is now reduced to practice.
SUMMARY OF THE INVENTION
This invention describes a method of manufacturing monolithic composite wings without using mechanical fasteners. The fabrication process comprises the steps of:
1. Forming a spanwise elongated center wing box cell, (FIG.
3
), using tapered U shaped spars and plurality of riblets placed between spars at various span stations, including a pair of contoured skin-molds glued to the upper and lower surfaces of the spar caps. The skin-molds may be made up of contoured honeycomb sheets or a combination of woven composite fabrics including integrally formed stiffeners. Said framework of cell serves the dual role of a mandrel that provides airfoil shape as well as a load bearing structural member.
2. Wrapping, local heating, compacting and bonding said center box (cell) by means of resin impregnated composite tapes along judiciously chosen directions in plurality of layers along the wing span. The number of layers and their directions are predetermined to comply with the bending and torsion strength requirements.
3
Ball Michael W.
Rossi Jessica
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