DETAILED ACTION
Notice of Pre-AIA or AIA Status
The present application, filed on or after March 16, 2013, is being examined under the first inventor to file provisions of the AIA .
This Action is responsive to the Application filed 15 November 2024. Claims 1-20 are pending. Claims 1 and 17 have been written in independent form.
Priority
Receipt is acknowledged of certified copies of papers required by 37 CFR 1.55.
Claim Rejections - 35 USC § 103
The following is a quotation of 35 USC § 103 which forms the basis for all obviousness rejections set forth in this Office Action:
A patent for a claimed invention may not be obtained, notwithstanding that the claimed invention is not identically disclosed as set forth in section 102, if the differences between the claimed invention and the prior art are such that the claimed invention as a whole would have been obvious before the effective filing date of the claimed invention to a person having ordinary skill in the art to which the claimed invention pertains. Patentability shall not be negated by the manner in which the invention was made.
This application currently names joint inventors. In considering patentability of the claims the examiner presumes that the subject matter of the various claims was commonly owned as of the effective filing date of the claimed invention(s) absent any evidence to the contrary. Applicant is advised of the obligation under 37 CFR 1.56 to point out the inventor and effective filing dates of each claim that was not commonly owned as of the effective filing date of the later invention in order for the examiner to consider the applicability of 35 USC § 102(b)(2)(C) for any potential 35 USC § 102(a)(2) prior art against the later invention.
Claim(s) 1-3, 6, 9-17, and 19 is/are rejected under 35 USC § 103 as being unpatentable over Griess et al., U.S. Patent Application Publication 2020/0047867 A1 (hereinafter called Griess), and further in view of Dunleavy, U.S. Patent Application Publication 2010/0096067 A1 (hereinafter called Dunleavy).
Regarding claim 1, Griess teaches an elongate aircraft structural component (See e.g., TITLE; element 102 in all respective figures; ¶ [0008]) for an aircraft (See e.g., FIG. 4; ¶ [0001]), the elongate aircraft structural component comprising:
a web (See e.g., element 120 in all corresponding figures; ¶ [0009]) and a flange (See e.g., element 122 in all corresponding figures; ¶ [0009]),
wherein the flange extends from the web (See e.g., ¶ [0009]) and comprises layers of composite material (See e.g., ¶ [0124]), a length in a direction of a length of the elongate aircraft structural component (See e.g., FIGS. 6, 10, 14, 16 elements 122, 102), a width perpendicular to the length of the flange (See e.g., FIGS. 6, 10, 14, 16 element 122 and detailed in ¶ [0141], FIGS. 6-8 elements 190, 192, 194), and a thickness smaller than, and perpendicular to, both the width and the length of the flange (See e.g., FIGS. 6, 10, 14, 16 the thickness of element 122),
wherein the width of the flange varies along the length of the flange (See e.g., ¶s [0067], [0107]-[0108], & [0210]), whereby the flange comprises a wide flange region and a narrow flange region, the narrow flange region having a smaller width than the wide flange region (See e.g., FIGS. 5-8 elements 190, 192, 194; ¶s [0140]-[0142]), and
wherein the wide flange region comprises interconnecting the layers of composite material (See e.g., ¶ [0124]) along a path in a direction with at least a component parallel to the width of the flange (See e.g., FIGS. 5-8 element 108, which teaches a component parallel to the width of the flange, which is along a path in a direction with the wide flange region as set forth in the claim).
But Griess does not teach the wide flange region comprises stitching interconnecting the layers of composite material, the stitching extending along a path in a direction with at least a component parallel to the width of the flange.
However, Dunleavy teaches the wide flange region comprises stitching (See e.g., FIG. 9 element 66, where the stitches 66 connecting the flanges of the stiffeners 62 and 68 through the cross-shaped reinforcing strip 72 at the intersection teaches the wide flange region) interconnecting the layers of composite material (See e.g., ¶ [0003]), the stitching extending along a path in a direction with at least a component parallel to the width of the flange (See e.g., FIG. 9, where element 72 is a component parallel to the width of the flange 62 along a path direction; ¶s [0045] & [0049]).
Accordingly, because the prior art of Griess and Dunleavy teach that it was well-known, before the effective filing date of the claim invention, to use layers of composite materials to form elongate aircraft structural components, it would have been obvious to the skilled artisan in the art, having the prior art of Griess and Dunleavy before him, to modify the elongate aircraft structural component of Griess to incorporate the wide flange region comprises stitching interconnecting the layers of composite material, the stitching extending along a path in a direction with at least a component parallel to the width of the flange, as taught in the analogous art of Dunleavy. The skilled artisan in the art would have been motivated to make such a combination to achieve the predictable result of producing a stiffened panel with lines of stitching that greatly improve its ability to accommodate damage and to prevent the stiffeners from breaking up or becoming separated from the panel, e.g., in the event of an impact leading to the resin delaminating at the interface between the panel and the stiffening array, as suggested by Dunleavy (See e.g., ¶ [0047]).
Regarding claim 2, Griess, as modified by Dunleavy in the rejection of claim 1 hereinabove, further teaches wherein the layers of composite material (Griess See e.g., ¶ [0124]; Dunleavy See e.g., ¶ [0003]) comprise layers of fiber composite material (Griess See e.g., ¶ [0072]; Dunleavy See e.g., ¶ [0003]).
Regarding claim 3, Griess, as modified by Dunleavy in the rejection of claim 2 hereinabove, further teaches wherein the layers of fiber composite material (Griess See e.g., ¶ [0124]; Dunleavy See e.g., ¶ [0003]) comprise carbon fibers (Dunleavy See e.g., ¶ [0022]).
Regarding claim 6, Griess, as modified by Dunleavy in the rejection of claim 1 hereinabove, further teaches wherein the wide flange region (Griess See e.g., FIGS. 5-8 elements 122, 190, 192, 194; ¶s [0141], [0111], & [0067]) comprises a hole (Griess See e.g., FIG. 5 element 170; ¶ [0127]) for receiving a fastener (Griess See e.g., FIG. 5 element 172; ¶ [0127]) for fastening the flange to a skin section of the aircraft (Griess See e.g., ¶s [0003] & [0127]).
Regarding claim 9, Griess, as modified by Dunleavy in the rejection of claim 1 hereinabove, further teaches wherein the stitching (Dunleavy See e.g., FIG. 9 element 66) comprises a plurality of parallel stitchings interconnecting the layers of composite material (Dunleavy See e.g., FIG. 9; ¶s [0045] & [0049]).
Regarding claim 10, Griess, as modified by Dunleavy in the rejection of claim 1 hereinabove, further teaches wherein the stitching comprises a plurality of intersecting stitchings (Dunleavy See e.g., FIG. 9 elements 66).
Regarding claim 11, Griess, as modified by Dunleavy in the rejection of claim 1 hereinabove, further teaches wherein the direction of the path has equal components in a direction of the width of the flange and a direction of the thickness of the flange (Griess See e.g., FIGS. 5-8 elements 108 & 102, respectively).
Regarding claim 12, Griess, as modified by Dunleavy in the rejection of claim 1 hereinabove, further teaches wherein the direction of the path has a component parallel to the length of the flange (Griess See e.g., FIGS. 5-8 element 108).
Regarding claim 13, Griess, as modified by Dunleavy in the rejection of claim 1 hereinabove, further teaches wherein the elongate structural component is an aircraft stringer (Griess See e.g., element 102 in all respective figures; ¶s [0001], [0003], [0005], [0077], & [0079).
Regarding claim 14, Griess, as modified by Dunleavy in the rejection of claim 1 hereinabove, further teaches wherein the narrow flange region (Griess See e.g., ¶s [0140]-[0142]; Dunleavy See e.g., FIG. 9, the regions of elements 62 and 68 that are not located anywhere in the region where 62 and 68 intersect) also comprises stitching interconnecting the layers of composite material (Dunleavy See e.g., FIG. 9 element 66; ¶ [0003]).
Regarding claim 15, Griess, as modified by Dunleavy in the rejection of claim 1 hereinabove, further teaches an aircraft structural assembly, comprising:
the elongate aircraft structural component according to claim 1 (Griess please refer to the rejection of the elongate aircraft structural component in claim 1 hereinabove); and
a skin section affixed to the flange of the elongate aircraft structural component (Griess See e.g., ¶s [0003] & [0127]).
Regarding claim 16, Griess, as modified by Dunleavy in the rejection of claim 1 hereinabove, further teaches an aircraft (Griess See e.g., FIG. 4; ¶ [0001]) comprising: the aircraft structural assembly according to claim 15 (Griess please refer to the rejection of the aircraft structural assembly in claim 15 hereinabove).
Regarding claim 17, Griess teaches a method (See e.g., FIG. 19 block 2010) of fabricating an elongate aircraft structural component (See e.g., TITLE; element 102 in all respective figures; ¶ [0008]) for an aircraft (See e.g., FIG. 4; ¶ [0001]), the elongate aircraft structural component comprising:
providing a web (See e.g., element 120 in all corresponding figures; ¶ [0009]);
providing a flange (See e.g., element 122 in all corresponding figures; ¶ [0009]),
extending from the web (See e.g., ¶ [0009]) and comprising layers of composite material (See e.g., ¶ [0124]), a length in a direction of a length of the elongate aircraft structural component (See e.g., FIGS. 6, 10, 14, 16 elements 122, 102), a width perpendicular to the length of the flange (See e.g., FIGS. 6, 10, 14, 16 element 122 and detailed in ¶ [0141], FIGS. 6-8 elements 190, 192, 194), and a thickness smaller than, and perpendicular to, both the width and the length of the flange (See e.g., FIGS. 6, 10, 14, 16 the thickness of element 122),
wherein the width of the flange varies along the length of the flange (See e.g., ¶s [0067], [0107]-[0108], & [0210]), whereby the flange comprises a wide flange region and a narrow flange region, the narrow flange region having a smaller width than the wide flange region (See e.g., FIGS. 5-8 elements 190, 192, 194; ¶s [0140]-[0142]), and
interconnecting the layers of composite material (See e.g., ¶ [0124]) in the wide flange region (See e.g., ¶ [0124]) along a path in a direction with at least a component parallel to the width of the flange (See e.g., FIGS. 5-8 element 108, which teaches a component parallel to the width of the flange, which is along a path in a direction with the wide flange region as set forth in the claim).
But Griess does not teach providing stitching interconnecting the layers of composite material in the wide flange region, the stitching extending along a path in a direction with at least a component parallel to the width of the flange.
However, Dunleavy teaches the wide flange region comprises stitching (See e.g., FIG. 9 element 66, where the stitches 66 connecting the flanges of the stiffeners 62 and 68 through the cross-shaped reinforcing strip 72 at the intersection teaches the wide flange region) interconnecting the layers of composite material (See e.g., ¶ [0003]), the stitching extending along a path in a direction with at least a component parallel to the width of the flange (See e.g., FIG. 9, where element 72 is a component parallel to the width of the flange 62 along a path direction; ¶s [0045] & [0049]).
Accordingly, because Griess and Dunleavy teach that it was well-known, before the effective filing date of the claim invention, to use layers of composite materials to form elongate aircraft structural components, it would have been obvious to the skilled artisan in the art, having the prior art of Griess and Dunleavy before him, to modify the method of Griess to incorporate providing stitching interconnecting the layers of composite material in the wide flange region, the stitching extending along a path in a direction with at least a component parallel to the width of the flange, as taught in the analogous art of Dunleavy. The skilled artisan in the art would have been motivated to make such a combination to achieve the predictable result of producing a stiffened panel with lines of stitching that greatly improve its ability to accommodate damage and to prevent the stiffeners from breaking up or becoming separated from the panel, e.g. in the event of an impact leading to the resin delaminating at the interface between the panel and the stiffening array, as suggested by Dunleavy (See e.g., ¶ [0047]).
Regarding claim 19, Griess, as modified by Dunleavy in the rejection of claim 17 hereinabove, further teaches wherein the method comprises providing the wide flange region (Griess See e.g., FIGS. 5-8 elements 122, 190, 192, 194; ¶s [0141], [0111], & [0067]) with a hole (Griess See e.g., FIG. 5 element 170; ¶ [0127]) for receiving a fastener (Griess See e.g., FIG. 5 element 172; ¶ [0127]) for fastening the flange to a skin section of the aircraft (Griess See e.g., ¶s [0003] & [0127]).
Claim(s) 4-5 and 18 is/are rejected under 35 USC § 103 as being unpatentable over Griess, and further in view of Dunleavy, and further in view of LIU, CN-116462525-A, published 2023-07-21 (hereinafter called LIU).
Regarding claim 4, Griess, as modified by Dunleavy in the rejection of claim 2 hereinabove, further teaches wherein the layers of fiber composite material (Griess See e.g., ¶ [0124]; Dunleavy See e.g., ¶ [0003]) comprise unidirectional-fiber layers orientated along the length of the flange (Griess See e.g., FIGS. 6, 10, 14, 16 element 122; ¶ [0124]).
But Griess is silent regarding the layers of fiber composite material comprise unidirectional-fiber layers, having continuous fibers orientated along the length of the flange, and chopped strand layers, having discontinuous fibers orientated in a plurality of directions within a plane of each respective chopped strand layer.
However, LIU teaches that it was well-known for the layers of fiber composite material to comprise unidirectional-fiber layers, having continuous fibers, and chopped strand layers, having discontinuous fibers orientated in a plurality of directions within a plane of each respective chopped strand layer (See e.g., FIG. 6; page 5, the description for FIG. 6 located about midway on the page; page 8, the description for FIG. 6 located about midway on the page).
Accordingly, because the prior art of Griess, Dunleavy, and LIU teach that it was well-known, before the effective filing date of the claim invention, to use layers of composite materials on various aircraft structures, it would have been obvious to the skilled artisan in the art, having the prior art of Griess, Dunleavy, and LIU before him, to modify the elongate aircraft structural component of the combined invention of Griess and Dunleavy to incorporate the layers of fiber composite material comprise unidirectional-fiber layers, having continuous fibers, and chopped strand layers, having discontinuous fibers orientated in a plurality of directions within a plane of each respective chopped strand layer, as taught in the analogous art of LIU, and having these continuous fibers oriented along the length of the flange, as required by the instant claim. The skilled artisan in the art would have been motivated to make such a combination to achieve the predictable result of enhancing the bonding force between the layers and improving the toughness of the composite material, as suggested by LIU (See e.g., page 8, the description for FIG. 6 located about midway on the page).
Regarding claim 5, Griess, as modified by Dunleavy and LIU in the rejection of claim 4 hereinabove, further teaches wherein at least 50% of the layers of composite material are the unidirectional-fiber layers (LIU See e.g., FIG. 6) in the flange (Griess See e.g., element 122 in all corresponding figures; ¶ [0009]).
Regarding claim 18, Griess, as modified by Dunleavy in the rejection of claim 17 hereinabove, further teaches wherein the layers of fiber composite material (Griess See e.g., ¶ [0124]; Dunleavy See e.g., ¶ [0003]) comprise unidirectional-fiber layers orientated along the length of the flange (Griess See e.g., FIGS. 6, 10, 14, 16 element 122; ¶ [0124]).
But Griess is silent regarding the layers of fiber composite material comprise unidirectional-fiber layers, having continuous fibers orientated along the length of the flange, and chopped strand layers, having discontinuous fibers orientated in a plurality of directions within a plane of each respective chopped strand layer.
However, LIU teaches that it was well-known for the layers of fiber composite material to comprise unidirectional-fiber layers, having continuous fibers, and chopped strand layers, having discontinuous fibers orientated in a plurality of directions within a plane of each respective chopped strand layer (See e.g., FIG. 6; page 5, the description for FIG. 6 located about midway on the page; page 8, the description for FIG. 6 located about midway on the page).
Accordingly, because the prior art of Griess, Dunleavy, and LIU teach that it was well-known, before the effective filing date of the claim invention, to use layers of composite materials on various aircraft structures, it would have been obvious to the skilled artisan in the art, having the prior art of Griess, Dunleavy, and LIU before him, to modify the method of the combined invention of Griess and Dunleavy to incorporate the layers of fiber composite material comprise unidirectional-fiber layers, having continuous fibers, and chopped strand layers, having discontinuous fibers orientated in a plurality of directions within a plane of each respective chopped strand layer, as taught in the analogous art of LIU, and having these continuous fibers oriented along the length of the flange, as required by the instant claim. The skilled artisan in the art would have been motivated to make such a combination to achieve the predictable result of enhancing the bonding force between the layers and improving the toughness of the composite material, as suggested by LIU (See e.g., page 8, the description for FIG. 6 located about midway on the page).
Claim(s) 7-8 and 20 is/are rejected under 35 USC § 103 as being unpatentable over Griess, and further in view of Dunleavy, and further in view of Hallander et al., U.S. P 2017/0129207 A1 (hereinafter called Hallander).
Regarding claim 7, Griess, as modified by Dunleavy in the rejection of claim 6 hereinabove, further teaches the hole (Griess See e.g., FIG. 5 element 170; ¶ [0127]).
But neither Griess nor Dunleavy teaches the hole does not interrupt the stitching.
However, Hallander teaches that it was well-known that the hole does not interrupt the stitching (See e.g., FIG. 13b elements W & 15; ¶s [0055], [0091], [0166], [0185]-[0186], [0251] & [0264], where the disclosure of composite material with stitched elements (i.e., 15 and 17), and elements 17 hold and keep the distance between elements 15, and the separations between elements 15 creating a window (W) through which a bolt hole will be drilled, teach the hole does not interrupt the stitching).
Accordingly, because the prior art of Griess, Dunleavy, and Hallander teach that it was well-known, before the effective filing date of the claim invention, to use and secure layers of composite materials on various aeronautical composite structures using various techniques, it would have been obvious to the skilled artisan in the art, having the prior art of Griess, Dunleavy, and Hallander before him, to modify the elongate aircraft structural component of the combined invention of Griess, Dunleavy, and Hallander to incorporate the hole does not interrupt the stitching, as taught in the analogous art of Hallander. The skilled artisan in the art would have been motivated to make such a combination to achieve the predictable result of optimizing the bolt bearing strength, which will also improve the strength within the pre-preg plies and also improve the bolt bearing-, open hole compression-, interlaminar shear-intralaminar tension- and pure tension strength, as suggested by Hallander (See e.g., ¶ [0264]).
Regarding claim 8, Griess, as modified by Dunleavy and Hallander in the rejection of claim 6 hereinabove, further teaches wherein the stitching (Hallander See e.g., FIG. 13b elements 15) surrounds the hole (Hallander See e.g., FIG. 13b element W).
Regarding claim 20, Griess, as modified by Dunleavy in the rejection of claim 19 hereinabove, further teaches the providing the wide flange region (Griess See e.g., FIGS. 5-8 elements 122, 190, 192, 194; ¶s [0141], [0111], & [0067]) with a hole (Griess See e.g., FIG. 5 element 170; ¶ [0127]).
But neither Griess nor LIU teaches providing the hole such that the hole does not interrupt the stitching.
However, Hallander teaches that providing the hole such that the hole does not interrupt the stitching (See e.g., FIG. 13b elements W & 15; ¶s [0055], [0091], [0166], [0185]-[0186], [0251] & [0264], where the disclosure of composite material with stitched elements (i.e., 15 and 17), and elements 17 hold and keep the distance between elements 15, and the separations between elements 15 creating a window (W) through which a bolt hole will be drilled, teach providing the hole such that the hole does not interrupt the stitching) was well-known.
Accordingly, because the prior art of Griess, Dunleavy, and Hallander teach that it was well-known, before the effective filing date of the claim invention, to use and secure layers of composite materials on various aeronautical composite structures using various techniques, it would have been obvious to the skilled artisan in the art, having the prior art of Griess, Dunleavy, and Hallander before him, to modify the method of the combined invention of Griess, Dunleavy, and Hallander to incorporate providing the hole such that the hole does not interrupt the stitching, as taught in the analogous art of Hallander. The skilled artisan in the art would have been motivated to make such a combination to achieve the predictable result of optimizing the bolt bearing strength, which will also improve the strength within the pre-preg plies and also improve the bolt bearing-, open hole compression-, interlaminar shear-intralaminar tension- and pure tension strength, as suggested by Hallander (See e.g., ¶ [0264]).
Conclusion
It is noted that any citation to specific pages, columns, lines, or figures in the prior art references and any interpretation of the references should not be considered to be limiting in any way. A reference is relevant for all it contains and may be relied upon for all that it would have reasonably suggested to one having ordinary skill in the art. In re Heck, 699 F.2d 1331, 1332-33, 216 U.S.P.Q. 1038, 1039 (Fed. Cir. 1983) (quoting In re Lemelson, 397 F.2d 1006, 1009, 158 U.S.P.Q. 275, 277 (C.C.P.A. 1968)).
Any inquiry concerning this communication or earlier communications from the Examiner should be directed to TERRI L FILOSI whose telephone number is (571)270-1988. The Examiner can normally be reached Monday-Friday 7:00 AM -3:30 PM EST.
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If attempts to reach the Examiner by telephone are unsuccessful, the Examiner’s supervisor, Timothy D Collins can be reached at 571-272-6886. The fax phone number for the organization where this application or proceeding is assigned is 571-273-8300.
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/TERRI L FILOSI/
Examiner
Art Unit 3644
19 March 2026
/TIMOTHY D COLLINS/Supervisory Patent Examiner, Art Unit 3644