Cutting by use of rotating axially moving tool – Processes – Bit detachable
Reexamination Certificate
1999-09-10
2001-07-03
Bishop, Steven C. (Department: 3722)
Cutting by use of rotating axially moving tool
Processes
Bit detachable
C408S003000, C408S016000, C408S043000, C408S046000, C408S234000, C409S202000, C409S212000
Reexamination Certificate
active
06254317
ABSTRACT:
CROSS-REFERENCE TO RELATED APPLICATIONS
(Not Applicable)
STATEMENT RE: FEDERALLY SPONSORED RESEARCH/DEVELOPMENT
(Not Applicable)
BACKGROUND OF THE INVENTION
Military and commercial jet aircraft manufacturing techniques are well-known in the art. Crucial to such manufacture is the construction of the fuselage assembly, which is widely regarded as the most complex and costly aspect of aircraft construction, particularly in relation to the construction of large, commercial aircraft. Generally, such process includes making certain barrel and aft duct assemblies with additional substructure and installing subsystems thereonto. Ultimately, skin panels are fastened thereabout, typically by a multiplicity of rivets. With respect to the latter, it is widely recognized that such process is extremely costly and labor intensive due to the high quantity of fastener holes to be drilled in this particular section of the aircraft. In this regard, it is not uncommon for more than 1,500 fastener holes be drilled in order to attach the exterior skin components to only one side of the fuselage assembly of a single aircraft.
Additionally problematic with the construction of the fuselage assembly of such aircraft is the incorporation of interference-fit holes in a metal substructure forming the fuselage, and clearance-fit holes in the composite skin affixed thereto. In this regard, and unlike the majority of other fastener holes formed on aircraft, such bifurcated hole size system is deemed desirable insofar as weight savings can be substantially accomplished. Specifically, such two-step hole arrangement enables a lighter substructure to be utilized without compromising strength, due to the interference-fit holes, while preventing skin delamination, via the use of clearance-fit holes.
Notwithstanding the benefits of such design and resultant lightweight structure formed thereby, such assembly process has proven to be more difficult insofar as the skin components must be repeatedly removed from the structure so the clearance-fit holes formed thereon may be opened up by hand. Accurate skin location then becomes an issue as the skin is unloaded and reloaded on the assembly, making mechanization of such manufacturing process difficult to achieve.
Accordingly, there is a substantial need in the art for an assembly method, and in particular a systematic drilling method for use in the construction of fuselage assemblies that enables a multiplicity of fastener holes to be formed about the fuselage assembly in a simultaneous manner. There is an additional need in the art for such method that can simultaneously produce dissimilar-type holes, and in particular, interference-fit holes in the metal substructure of such fuselage assembly and clearance-fit holes in the composite skin to be attached thereto, continuously about such fuselage assembly in multiple planes. There is yet further a need in the art for such a method that can provide for such automated formation of fastener holes throughout a fuselage assembly that further includes quality assurance mechanisms to verify aircraft skin position and orientation relative the fuselage assembly when affixed thereto via such mechanized process.
BRIEF SUMMARY OF THE INVENTION
The present invention specifically addresses and alleviates the above-identified deficiencies in the art. In this regard, the present invention is directed to a systematic drilling method for use in forming a plurality of fastener holes upon an aircraft substructure, and in particular fuselage assemblies, along multiple axes. Such method preferably incorporates a triple gantry drilling system adapted to simultaneously affix a multiplicity of skin components to the metal substructure of a fuselage assembly during the construction thereof. According to a preferred embodiment, the gantry system utilized in such method includes upper, left and right side gantries wherein each gantry includes a robotic arm having an automatic tool changing system formed thereon. The latter preferably includes an attachment for holding a plurality of drill mechanisms exchangeable with a plurality of equipment and tools to selectively and controllably implement automated countersinking, sealing, and fastening, respectively. Preferably, such combination of such gantries will enable fastener holes to bore about five (5) dissimilar axes about both sides and upper surface of the fuselage. Specifically, each drilling system of each gantry will be capable of moving up and down along a first axis, side to side along a second axis, inwardly and outwardly relative the fuselage assembly along a third axis, rotating about a fourth axis, and rotating with a specified degree of pitch or tilt according to a fifth axis.
In a more highly preferred embodiment, there is further provided on the drill head assembly a vision system having a closed circuit television with a vision recognition system, the latter being integrated with a workstation main controller. An optical probe system is further incorporated into the drill head assembly to ensure quality control and verify aircraft skin position and orientation with respect to respective one of the gantries of such system. In this respect, such camera will assist mechanics to view the actual drilling process and verify substructure position as part of the method of the present invention.
It is therefore an object of the present invention to provide a systematic drilling method for use in the construction of fuselage assemblies that enables a multiplicity of fastener holes to be simultaneously formed thereabout.
Another object of the present invention is to provide a systematic drilling method that enables interference-fit holes to be formed within the metal substructure of a fuselage assembly while at the same time enabling clearance-fit holes to be formed in the composite skin affixed thereto.
Another object of the present invention is to provide a systematic drilling method that further provides for accurate skin placement location during the manufacture of the fuselage assembly.
Still further objects of the present invention are to provide systematic drilling methods that provide for rapid, mechanized manufacture of fuselage assemblies with greater precision, substantially less labor and sufficiently greater durability than prior art methods and manufacturing systems.
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Anderson Terry J.
Bishop Steven C.
Hoch, Jr. Karl J.
Northrop Grumman Corporation
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