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 .
Response to Amendment
Applicant filed a response and amended claim 1 on 06/27/2025. The claim objection is withdrawn in view of amendment to claim 1.
Response to Arguments
Applicant's arguments filed have been fully considered but they are not persuasive.
Applicant argues:
“The purpose of the composite material structural component disclosed in Hasegawa is
that unnecessary members of the product are not included to eliminate needless weight increase
of the aircraft. See paragraphs [0007] and [0009] of Hasegawa. Accordingly, in the composite
material structure of Hasegawa, the metal mesh 3 is attached to an outer surface of the outer plate
1 (base material 10) to function as a conductor for lightning strike protection and as a heat source
for forming the base material 10. This reduces the number of parts and eliminate needless
weight increase of the aircraft. See paragraphs [0035]-[0036] of Hasegawa.
The composite structure of Hardman comprises a heating layer 30 sandwiched between
insulation layers 28 and structural layers 25, 26. See FIG. 1 of Hardman.
If Hasegawa was modified with a laminate structure comprising structural layers,
insulating layers, and heating layers and using an integrally laminating technique as taught by
Hardman, the number of materials used would be greater than the what is originally disclosed by
Hasegawa ( e.g., the metal mesh 3 and the outer plate 1 ). Increasing the number of materials, in
turn, increases the weight of the aircraft. To this end, Hasegawa explicitly states that "an object
of the present invention is to eliminate needless weight increase of an aircraft." See paragraph
[0009] of Hasegawa. Therefore, the proposed modification set forth in the Office Action to
Hasegawa by Hardman renders the composite material structural component of Hasegawa
unsatisfactory for its intended purpose. See MPEP § 2143 (V) ("If a proposed modification
would render the prior art invention being modified unsatisfactory for its intended purpose, there
may be no suggestion or motivation to make the proposed modification." In re Gordon, 733 F.2d
900, 221 USPQ 1125 (Fed. Cir. 1984).).”
Examiner respectfully disagrees.
Hasegawa teaches the use of a conductive member as a heat source in manufacturing the structural component is inherently an unnecessary member in the product because the conductive member would be used only for the purpose as a heater during the manufacturing and would be left in the structure after completion of the structural component, providing no other purposes [0007], [0013]. Hasegawa teaches the use of a conductive member as a heater during manufacturing and additional purposes, such as a conductor for lightning strike protective, heat source for anti-icing and deicing, and/or the conductor for electromagnetic interference shielding, provides improved efficiency [0012]-[0013]. As taught by Hasegawa, the conductive member having at least two functions correspondingly reduces the weight of the aircraft body [0022]-[0023].
Hardman teaches a heated composite structure comprising structural layers, insulative layers, and a heat layer that is suitable for heating the plurality of players of the laminate [0017] and providing ice protection [0004].
In fact, Hardman teaches the primary ice protection system in use today are bleed air systems that re-directs a portion of the compressed air from the engine
before the air enters the combustion chamber. Hardman teaches the compressed
air is pressurized hot air that is used to heat up
surfaces of the aircraft to prevent the build-up of ice, and although bleed air systems are effective, they reduce the efficiency of the engines and increase the weight of the aircraft [0004].
The assembly of structural layers, insulative layers, and heating layers of Hasegawa in view of Hardman as modified are not needless weight added to an aircraft because, as taught by Hardman and desired by Hasegawa, the assembly provides multiple purposes, including heating for ice protection during use of the aircraft, thereby avoiding the use of an additional protection system that would reduce efficiency and increase weight of the aircraft.
Claim Rejections - 35 USC § 103
In the event the determination of the status of the application as subject to AIA 35 U.S.C. 102 and 103 (or as subject to pre-AIA 35 U.S.C. 102 and 103) is incorrect, any correction of the statutory basis (i.e., changing from AIA to pre-AIA ) for the rejection will not be considered a new ground of rejection if the prior art relied upon, and the rationale supporting the rejection, would be the same under either status.
The following is a quotation of 35 U.S.C. 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.
The factual inquiries for establishing a background for determining obviousness under 35 U.S.C. 103 are summarized as follows:
1. Determining the scope and contents of the prior art.
2. Ascertaining the differences between the prior art and the claims at issue.
3. Resolving the level of ordinary skill in the pertinent art.
4. Considering objective evidence present in the application indicating obviousness or nonobviousness.
Claims 1, 4, 5, 8-10, 12, and 13 is/are rejected under 35 U.S.C. 103 as being unpatentable over Hasegawa (PG-PUB 2009/0107620) in view of Hardman (PG-PUB 2019/0322375).
Regarding claim 1, Hasegawa teaches a method comprising:
providing a composite preform that includes an electric heater (Figure 2, item 3; [0013]-[0014]; [0032]-[0033]), fiber reinforcement (Figure 2, item 10 and [0032]-[0033]), and thermoset material ([0008], [0019]-[0020]);
consolidating the composite preform to provide a composite aircraft component, the consolidating comprising heating the thermoset material using the electric heater to cure the thermoset material [0045],
wherein the electric heater is capable of functioning as a part of a thermal anti-icing system for melting and/or preventing ice accumulation an exterior surface of the composite aircraft component [0040] and [0043].
Hasegawa does not explicitly teach the electric heater and the fiber-reinforcement embedded within the cured thermoset material.
Hardman teaches a heated composite structure (Figure 4 and [0056]-[0057]) for use in ice protection [0054] comprising:
structural layers (125 and 126);
insulative layers (128); and
a heat layer (130).
Hardman teaches the layers are consolidated to form an integral laminate by heating the assembled plies under pressure [0052]-[0055]. Hardman teaches the structure is formed to wrap over a leading edge of a structure, such as a wing or nacelle [0055] and [0057].
Both Hasegawa and Hardman are drawn to the same field of endeavor pertaining to manufacturing heated composite structure for ice protection, such as a leading edge of a wing. It would have been obvious to one of ordinary skill in the art at the time of the effective filing date of the invention to modify the process of Hasegawa with the laminate configuration of Hardman, a known suitable configuration for forming a heated composite aircraft component. Accordingly, it would have been obvious to one of ordinary skill in the art at the time of the effective filing date of the invention to modify Hasegawa with a laminate comprising structural layers, insulating layers, and heating layers and using an integrally laminating technique as taught by Hardman to consolidate the layers to form the shaped composite of Hardman.
While Hardman teaches using thermoplastic composite material, one of ordinary skill in the art would have recognized the process of Hasegawa in view of Hardman as modified can be performed with the thermosetting composite material of Hasegawa because Hasegawa teaches embodiments of either thermoplastic and thermosetting composites and embodiments using thermoplastic composites can be used with thermosetting composite material [0045].
Accordingly, the process of Hasegawa in view of Hardman would teach consolidating the composite preform such that the electric heater and the fiber-reinforcement are embedded in the thermoset material.
Regarding claim 4, Hasegawa in view of Hardman teaches the process as applied to claim 1, wherein the providing of the composite preform comprises laying up the electric heater (Hardman, Figure 4, item 150) between a first layer and a second layer (Hardman, Figure 4, items 125 and 126; [0057]); the first layer includes a first portion of the fiber-reinforcement and a first portion of the thermoset material (Hardman, [0057] and Hasegawa, [0032]-[0033], [0045]); and the second layer includes a second portion of the fiber-reinforcement and a second portion of the thermoset material (Hardman, [0057] and Hasegawa, [0032]-[0033], [0045]).
Regarding claim 5, Hasegawa in view of Hardman teaches the process as applied to claim 1, wherein the heating layer comprises a heater including a plurality of electric heating elements embedded within the cured thermoset material (Hardman, [0057]).
Regarding claim 8 and 9, Hasegawa in view of Hardman teaches the process as applied to claim 1,
the composite preform further includes a second electric heater (Hardman, [0073]-[0075]);
the second electric heater is embedded within the cured thermoset material (Hardman, Figure 4, 7-9 and Hasegawa, [0045]); and
the second electric heater is capable as a second part of the thermal anti-icing system for melting and/or preventing ice accumulation on the exterior surface of the composite aircraft component (Hardman, [0073] and [0078]).
Hardman teaches it may be desirable to vary the configuration of the resistive elements in the second heating zone to vary the heating characteristic of the second heating zone from the first heating zone [0073]. Hardman teaches, in the second heating zone, the resistive elements can have different dimensions such that when both first and second zone heaters are heated using the same voltage, heater with longer resistive elements would result in lower wattages supplied than the other [0073]. Therefore, due to the physical differences with the first and second electric heaters (Figures 7 and 8; [0073]-[0074], [0078]), the temperature to which the heaters heat up would be different.
Due to the differences in the size of resistive elements in the first and second heaters, the total resistance from the resistive elements in both heaters would vary during heating of the thermoset, resulting in different temperatures.
Regarding claim 8 and 10, Hasegawa in view of Hardman teaches the process as applied to claim 1, wherein the composite preform further includes a second electric heater (Hardman, [0073]-[0075]);
the second electric heater is embedded within the cured thermoset material (Hardman, Figure 4, 7-9 and Hasegawa, [0045]); and
the second electric heater is capable as a second part of the thermal anti-icing system for melting and/or preventing ice accumulation on the exterior surface of the composite aircraft component (Hardman, [0073] and [0078]).
Hardman teaches it may be desirable to use substantially identical configurations for two heating zones (Figure 7, item 270- second heating zone an item 290- third heating zone; [0072], [0074]), resulting in substantially the same level of heating. Therefore, the electric heater in the second and third zone would hat up to a common temperature during the heating of the thermoset material.
Regarding claim 12, Hasegawa in view of Hardman teaches the process as applied to claim 1, wherein the composite aircraft component is capable of use on an aircraft (Hasegawa, [0021]-0023], [0035]).
Hasegawa does not explicitly teach a nacelle inlet structure comprises the composite aircraft component; and the electric heater is located at a leading edge of the nacelle inlet structure.
Hasegawa in view of Hardman teaches forming a heated composite structure from a laminate comprising carbon fibers embedded with a thermoplastic matrix and heating layers [0034]-[0035]. Hardman teaches the heating layer comprising heating elements that operate as heating elements for ice protection in a variety of components, including nacelles and airfoils [0039], [0055]. Hardman teaches the structure can be used on forms a section of the nacelle that forms a leading edge around the intake of the jet engine [0056].
It would have been obvious to one of ordinary skill in the art at the time of the effective filing date of the invention to utilize the composite aircraft component of Hasegawa in an leading edge of a nacelle inlet structure, a known aircraft component in need of de-icing protection as taught by Hardman, to yield the predictable result of providing thermal anti-icing system for melting and/or preventing ice accumulation an exterior surface of the composite aircraft component as desired by Hasegawa.
Regarding claim 13, Hasegawa in view of Hardman teaches the process as applied to claim 1, wherein an aircraft wing comprises the composite aircraft component and the electric heater is located at a leading edge of the aircraft wing (Hasegawa, Figure 3 and [0037]).
Claim(s) 2 and 3 is/are rejected under 35 U.S.C. 103 as being unpatentable over Hasegawa (PG-PUB 2009/0107620) in view of Hardman (PG-PUB 2019/0322375), as applied to claim 1, in further view of Blackmore (PG-PUB 2003/0168164).
Regarding claim 2 and 3, Hasegawa in view of Hardman teaches the method as applied to claim 1, wherein consolidating comprises heating and applying pressure to the laminate (Hardman, [0053], [0059]).
Hasegawa in view of Hardman does not explicitly teach consolidating further comprises applying pressure to the composite preform using tooling.
Hasegawa in view of Hardman does not explicitly teach consolidating further comprises applying pressure to the composite preform using a vacuum bag.
Blackmore teaches manufacturing composite structures from a stack of layers, including conductive layers in the form of a conductive fiber mat or tape [0058], [0080]; fibrous layers [0054]; and resin in the form of thermoplastic and/or thermosetting resin [0045], [0050]-[0051]. Blackmore teaches apply pressure to the stack using tooling (Figures 1-3) or a vacuum bag and open tooling (Figure 4 and 5) to shape and cure the laminate while heating is performed using the conductive layer(s) [0043], [0055], [0080]-[0081].
Hasegawa in view of Hardman teaches shaping and setting a preform using the heated layer for conducting heat but does not explicitly disclose the technique used to apply pressure and heating, prompting one of ordinary skill in the art to look elsewhere in the art. Both Hasegawa and Blackmore desire to manufacturing a composite by applying pressure and using an embedded layer as a heater. It would have been obvious to one of ordinary skill in the art to modify the process of Hasegawa with the known suitable techniques of shaping and curing as taught by Blackmore to yield the predictable result of applying pressure to form a composite while utilizing a conductive layer as an embedded heater.
Claim 5-7 is/are rejected under 35 U.S.C. 103 as being unpatentable over Hasegawa (PG-PUB 2009/0107620) in view of Hardman (PG-PUB 2019/0322375), as applied to claim 1, in further view of Linde (EP3575066, Machine Translation relied upon).
Regarding claim 5-7, Hasegawa in view of Hardman teaches the process as applied to claim 1.
Hardman teaches the heating layer comprises a heater including a plurality of electric heating elements embedded within the cured thermoset material (Hardman, [0057]), wherein the resistive elements can be carbon fibers (Hardman, [0066], [0069]-[0070]).
Hasegawa in view of Hardman does not teach:
the electric heater comprises a plurality of electric heating elements embedded within the cured thermoset material;
the electric heating elements are arranged in a grid; and
the electric heater comprises a carbon nanotube heater embedded within the cured thermoset material.
Linde teaches forming a component for an aircraft [0016] comprising assembling the component with at least one stiffening element or skin element and placing chains, strips, or cords of carbon nanotubes in or at a region for components [0021]-[0024]. Linde teaches during joining, the carbon nanotubes are used to provide local heating in the laminate [0035]-[0037]; [0089]-[0090]. Linde teaches the heating layer can be formed by preparing the carbon nanotubes in a grid [0024], [0073]-[0074], [0089]-[0090]
It would have been obvious to one of ordinary skill in the art at the time of the effective filing date of the invention to modify the process of Hasegawa in view of Hardman with a heating layer comprising a grid of carbon nanotubes, a known conductive material suitable for local heating within a laminate as taught by Linde, to yield the predictable result of providing embedded heating elements.
Claim(s) 11 is/are rejected under 35 U.S.C. 103 as being unpatentable over Hasegawa (PG-PUB 2009/0107620) in view of Hardman (PG-PUB 2019/0322375), as applied to claim 1, in further view of Hamilton (PG-PUB 2021/0162635).
Regarding claim 11, Hasegawa in view of Hardman teaches the process as applied to claim 1, wherein a plurality of sensors are used to monitor the conditions within the different zones and apply heat to the corresponding zones accordingly (Hardman, [0079]).
Hasegawa in view of Hardman does not teach monitoring the consolidation of the composite preform using a sensor; the composite preform further including the sensor; and the sensor embedded within the cured thermoset material.
Hamilton teaches a process for curing a composite article, wherein a heat source may be embedded in or attached to the composite article [0067]. Hamilton teaches composite article may be electrically conductive and be directly resistively or inductively heated; in effect the composite article acting as a heat source [0067]-[0068]. Hamilton teaches a temperature for detecting temperature-related property proximal to the heat source [0059], [0075]-[0083], wherein the temperature may be attached to a surface of or embedded within a composite article [0063]. Hamilton teaches using the temperature sensor to monitor curing [0095].
Hamilton teaches it has been proposed to embed means for detecting temperature within composite articles to accurate monitor temperature deep within a composite structure, but this would add cost and complexity of manufacture and for many applications, it is not desirable for the thermocouple to be present in the final composite article [0012].
Hasegawa in view of Hardman teaches incorporating individual sensors in the composite aircraft component for the purpose of monitoring the conditions within the different zones during application of ice protection, etc. Because Hasegawa in view of Hardman requires the final composite article to have the sensors, it would have been obvious to one of ordinary skill in the art at the time of the effective filing date of the invention to improve the process of Hasegawa in view of Hardman with the technique of using the temperature sensors of Hasegawa in view of Hardman for an additional functionality of monitoring temperature during curing as taught by Hamilton, thereby providing improved and controlled curing without any additional complexity or cost to manufacturing.
Claim 14 is/are rejected under 35 U.S.C. 103 as being unpatentable over Hasegawa (PG-PUB 2009/0107620) in view of Hardman (PG-PUB 2019/0322375), as applied to claim 1, in further view of Epstein (US 8,444,093).
Regarding claim 14, Hasegawa in view of Hardman teaches the process as applied to claim 1, wherein the composite aircraft component is capable of use on an aircraft (Hasegawa, [0021]-0023], [0035]), including on a leading edge of a wing (Hasegawa, [0027]).
Hasegawa in view of Hardman does not explicitly teach a nacelle inlet structure comprises the composite aircraft component; and the electric heater is located at a leading edge of the nacelle inlet structure.
Epstein teaches an airplane leading edge de-icing apparatus (Figure 1)
Epstein teaches a second heat diffuser side 50 may be variously embodied, for example, the second heat diffuser side 50 may be formed such that when joined to the airplane 20 it covers virtually any desired amount of the first and second wings 24, 26, tailfin 30 and first and second horizontal stabilizers 32, 34 (Col 3, ln 14-45).
It would have been obvious to one of ordinary skill in the art at the time of the effective filing date of the invention to utilize the composite aircraft component of Hasegawa in a same manner in an leading edge for an aircraft stabilizer, a known aircraft component in need of de-icing protection as taught by Epstein, to yield the predictable result of providing thermal anti-icing system for melting and/or preventing ice accumulation an exterior surface of the composite aircraft component as desired by Hasegawa.
Conclusion
THIS ACTION IS MADE FINAL. Applicant is reminded of the extension of time policy as set forth in 37 CFR 1.136(a).
A shortened statutory period for reply to this final action is set to expire THREE MONTHS from the mailing date of this action. In the event a first reply is filed within TWO MONTHS of the mailing date of this final action and the advisory action is not mailed until after the end of the THREE-MONTH shortened statutory period, then the shortened statutory period will expire on the date the advisory action is mailed, and any nonprovisional extension fee (37 CFR 1.17(a)) pursuant to 37 CFR 1.136(a) will be calculated from the mailing date of the advisory action. In no event, however, will the statutory period for reply expire later than SIX MONTHS from the mailing date of this final action.
Any inquiry concerning this communication or earlier communications from the examiner should be directed to HANA C PAGE whose telephone number is (571)272-1578. The examiner can normally be reached M-F, 9:00-5:30.
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If attempts to reach the examiner by telephone are unsuccessful, the examiner’s supervisor, Phillip Tucker can be reached on 5712721095. The fax phone number for the organization where this application or proceeding is assigned is 571-273-8300.
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/HANA C PAGE/Examiner, Art Unit 1745
/MICHAEL A TOLIN/Primary Examiner, Art Unit 1745