Office Action 3: Non-Final Rejection (2025-03-24)
Rejection types: 103 - Obviousness

CLAIMS:
AMENDMENTS TO THE CLAIMS

Please amend claims 1, 4, and 16 such that the status of the claims is as follows:

1\. (Currently Amended) A \[\[CMC\]\]ceramic matrix composite component for a gas turbine engine comprising:an airfoil comprising:
a leading edge;
a trailing edge;
a suction sidewall; and
a pressure sidewall joined to the suction sidewall at the leading edge and the trailing edge, the pressure sidewall comprising a locally- thickened region,
wherein the locally-thickened region has a first thickness, the first thickness being a maximum thickness of the pressure sidewall;
wherein the suction sidewall has a second thickness, the second thickness being a maximum thickness of the suction sidewall; \[\[and\]\] wherein the first thickness is greater than the second thickness; and wherein the airfoil comprises a plurality offiber plies.

2\. (Original) The component of claim 1, wherein the first thickness is up to 30% greater than the second thickness.

3\. (Previously Presented) The component of claim 1, wherein the locally-thickened region extends from 0% to at least 40% and not greater than 60% span of the airfoil.

4\. (Currently Amended) The component of claim 1, wherein the locally-thickened region extends from 100% to at least 60% and not eyond 40% span of the airfoil.

5\. (Original) The component of claim 1 and further comprising:a first platform; and
a second platform,

wherein the airfoil is connected to and extends between the first platform and the second platform.

6\. (Original) The component of claim 5, wherein the component is a vane.

7\. (Original) The component of claim 5, wherein the pressure sidewall comprises a first subset of plies.

8\. (Original) The component of claim 7, wherein the locally-thickened region comprises at least one additional ply.

9\. (Original) The component of claim 8, wherein the at least one additional ply is at least partially disposed within the first platform.

10\. (Original) A method of forming a \[\[CMC\]\] ceramic matrix composite component, the method comprising:identifying stress regions in the component exceeding a threshold value;
subsequently, forming a preform by:
laying up a first subset of plies over a mandrel;
laying up at least one reinforcing ply over the first subset of plies, the at least one reinforcing ply coextensive with the stress regions; and laying up a second subset of plies over the at least one reinforcing ply; and densifying the preform with a ceramic matrix.

11\. (Previously Presented) The method of claim 10, wherein the first subset of plies and the second subset of plies define an airfoil, the airfoil comprising:a leading edge;
a trailing edge;
a suction sidewall; and
a pressure sidewall.

12\. (Original) The method of claim 11, wherein the at least one reinforcing ply is disposed along the pressure sidewall between 0% span and 50% span.

13\. (Original) The method of claim 11, wherein the at least one reinforcing ply is disposed along the pressure sidewall between 50% span and 100% span.

14\. (Original) The method of claim 11 and further comprising: laying up at least one platform ply at an angle with respect to the at least one reinforcing ply, the at least one reinforcing ply being a projection of the at least one platform ply.

15\. (Original) The method of claim 10, wherein the step of densifying the preform with a ceramic matrix comprises one or a combination of chemical vapor infiltration, melt infiltration, and polymer infiltration and pyrolysis.

16\. (Currently Amended) A vane assembly for a gas turbine engine, the vane assembly comprising:an inner case;
an outer case; and
at least one \[\[CMC\]\] ceramic matrix composite (CMC) vane coupled between the inner case and the outer case, the at least one CMC vane comprising:an inner diameter platform;
an outer diameter platform; and an airfoil connected to and extending between the inner diameter platform and the outer diameter platform, the airfoil comprising:
a leading edge;
a trailing edge;
a suction sidewall; and
a pressuresidewalljoined to the suction sidewall at the leading edge andthe trailing edge, the pressure sidewall comprising a locally-thickened region extending from the inner diameter platform or the outer diameter platform toward a midspan

region pressure sidewall,wherein the locally-thickened region has a first thickness, the first thickness being a maximum thickness of the pressure sidewall;wherein the suction sidewall has a second thickness, the second thickness being a maximum thickness of the suction sidewall; \[\[and\]\] wherein the first thickness is greater than the second thickness; and wherein the airfoil comprises a plurality of fiber plies.

17\. (Original) The assembly of claim 16, wherein the first thickness is 30% greater than the second thickness.

18\. (Original) The assembly of claim 17, wherein the pressure sidewall comprises a first subset of plies.

19\. (Original) The assembly of claim 18, wherein the locally-thickened region comprises at least one additional ply.

20\. (Original) The assembly of claim 19, wherein the at least one additional ply is at least partially disposed within the inner diameter platform or the outer diameter platform.

OFFICE ACTION:
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 .
Continued Examination Under 37 CFR 1.114
A request for continued examination under 37 CFR 1.114, including the fee set forth in 37 CFR 1.17(e), was filed in this application after final rejection.  Since this application is eligible for continued examination under 37 CFR 1.114, and the fee set forth in 37 CFR 1.17(e) has been timely paid, the finality of the previous Office action has been withdrawn pursuant to 37 CFR 1.114.  Applicant's submission filed on 02/05/2025 has been entered.
Remarks
This communication is in response to the reply filed on 09/10/2024.
Claims 1-20 are currently pending and have been examined.
In view of the claim amendments, the rejection of claims 1-9 and 16-20 is moot and withdrawn.  The indication of allowance as to claims 10-15 is withdrawn.
However, a new grounds of rejection is provided, supra, as required by the claim amendments.
Claim Rejections - 35 USC § 103
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 of this title, 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.
Claims 1-9 are rejected under 35 U.S.C. 103 as being unpatentable over US 2022/0268165 hereinafter D1, in view of US 11,162,432 B2, hereinafter D2.
As to claim 1
D1 discloses a ceramic matrix composite ([0044] The airfoil fairing 62 is fabricated of a fiber-reinforced composite 70; [0045] The fiber-reinforced composite 70 is a ceramic matrix composite, an organic matrix composite (OMC), or a metal matrix composite (MMC)) component for a gas turbine engine (airfoil fairing 62) comprising:
an airfoil (airfoil section 64 in fig. 3) comprising:
a leading edge (see annotated figure 3);
a trailing edge (see annotated figure 3);
a suction sidewall (see annotated figure 3); and
a pressure sidewall (see annotated figure 3) joined to the suction sidewall at the leading edge and the trailing edge (fig. 3);
wherein the suction sidewall (see annotated figure 3) has a second thickness (see annotated figure 3 referring to 8 layers), the second thickness being a maximum thickness of the suction sidewall (the circled area appears a maximum); and
wherein the airfoil comprises a plurality of fiber plies ([0049] the fiber plies 72 of the fiber-reinforced composite 70).
D1 does not explicitly disclose:
the pressure sidewall (see annotated figure 3) comprising a locally-thickened region, wherein the locally-thickened region has a first thickness, the first thickness being a maximum thickness of the pressure sidewall; and
wherein the first thickness is greater than the second thickness.
It is noted that D1 does disclose ([0049]) the tubes here are shown with three plies (for cavities 74a/74b) and one ply (cavity 74c), additional or fewer plies may be provided for strength and/or wall thickness considerations.
D2 teaches a nozzle segment having vanes where the thicknesses of one or more portions of the vane 52 (e.g., pressure side wall 66, suction side wall 68, leading edge wall 82, trailing edge wall 84, etc.) may be preferentially varied in accordance with expected (or estimated) operational forces (col. 5, lns. 36+) as doing so provides improved nozzle service life (the operational lifetime of the nozzle 32 may be increased, for example, by preferentially varying the thicknesses of various portions of the vanes 52, e.g. thicker in regions subject to higher forces, thinner in regions subject to lower forces, col. 5, ln. 4+).
It would be obvious to one of ordinary skill in the art to provide the pressure sidewall with a locally-thickened region, wherein the locally-thickened region has a first thickness, the first thickness being a maximum thickness of the pressure sidewall and wherein the first thickness is greater than the second thickness as taught by D2 and (suggested by D1) in order to provide improved service life (the operational lifetime of the nozzle 32 may be increased, for example, by preferentially varying the thicknesses of various portions of the vanes 52, e.g. thicker in regions subject to higher forces, thinner in regions subject to lower forces, col. 5, ln. 4+).
In addition, because D1 discloses selecting a number of plies for wall thickness consideration ([0049])  the recitation of the pressure sidewall with a locally-thickened region, wherein the locally-thickened region has a first thickness, the first thickness being a maximum thickness of the pressure sidewall and wherein the first thickness is greater than the second thickness would be obvious to one of ordinary skill as a matter of design choice.
As to claims 2-4
D1 as modified does not explicitly disclose:
[claim 2] wherein the first thickness is up to 30% greater than the second thickness;
[claim 3] wherein the locally-thickened region extends from 0% to at least 40% and not greater than 60% span of the airfoil; or
[claim 4] wherein the locally-thickened region extends from 100% to at least 60% and not beyond 40% span of the airfoil.
At the time of the invention, it would have been obvious to a person of ordinary skill in the art to choose the claimed dimensions as matter of mere design choice. It is noted that D1 does disclose ([0049]) the tubes here are shown with three plies (for cavities 74a/74b) and one ply (cavity 74c), additional or fewer plies may be provided for strength and/or wall thickness considerations.   Applicant has not disclosed that [claim 2] wherein the first thickness is up to 30% greater than the second thickness; [claim 3] wherein the locally-thickened region extends from 0% to at least 40% and not greater than 60% span of the airfoil; or [claim 4] wherein the locally-thickened region extends from 100% to at least 60% and not beyond 40% span of the airfoil  provides an advantage, is used for a particular purpose or solves a stated problem.  	At the top of MPEP 2144.04, it explains that various modifications, including changes in size and proportion, are “common practices which the court has held normally require only ordinary skill in the art and hence are considered routine expedients.” Specifically, §§IV. A cites, Gardner v. TEC Syst., Inc., 725 F.2d 1338, 220 USPQ 777 (Fed. Cir. 1984), cert. denied, 469 U.S. 830, 225 USPQ 232 (1984), wherein the Federal Circuit held that, where the only difference between the prior art and the claims was a recitation of relative dimensions of the claimed device and a device having the claimed relative dimensions would not perform differently than the prior art device, the claimed device was not patentably distinct from the prior art device.
The applicant has not disclosed that having the device with the claimed dimensions solves any stated problem or is for any particular purpose above the fact that this localized increase and thickness improves the mechanical properties of vane 54 while minimizing added weight and thickness and maintaining desired flow characteristics ([0025]). This rationale is indistinguishable from that of D1 which discloses ([0049]) the tubes here are shown with three plies (for cavities 74a/74b) and one ply (cavity 74c), additional or fewer plies may be provided for strength and/or wall thickness considerations.   It would have been obvious to one of ordinary skill in the art before the effective filing date of the claimed invention to have modified the device of D1 to have the claimed dimensions, because this modification does not patentably distinguish the claimed device, nor cause the device to perform differently, and thus the modification of the dimensions would have been obvious.
As to claim 5
D1 discloses a first platform (one of 66, 68); and a second platform (the other of 66, 68), wherein the airfoil (airfoil section 64 in fig. 3) is connected to and extends between the first platform and the second platform (fig. 2A, 2B).
As to claim 6
D1 discloses wherein the component is a vane ([0042] It is to be understood
that although the examples herein are discussed in context of a vane from the  turbine section, the examples can be applied to vanes in other portions of the engine 20).
As to claim 7
D1 discloses wherein the pressure sidewall comprises a first subset of plies (see annotated figure 3).
As to claim 8
D1 discloses wherein the locally-thickened region comprises at least one additional ply (see annotated figure 3 referring to 80 having 8 layers on one side).
As to claim 9
D1 discloses wherein the at least one additional ply is at least partially disposed within the first platform (see annotated figure 4).
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Claims 10-15 are rejected under 35 U.S.C. 103 as being unpatentable over D2, in view of US 11,530,614, hereinafter D3.
As to claim 10
D2 discloses a method of forming a ceramic matrix composite component (FIGS. 8B, and illustrate selected portions of the SA, SC airfoil fairing 62 during a fabrication process), the method comprising:
identifying stress regions in the component exceeding a threshold value (operational forces on the vane 52 may be determined, for example, via computer  modeling, physical testing, or other suitable analysis techniques. This may include, for example, modeling and analyzing the operational forces using an engineering simulation tool, modifying one or more of the thicknesses, and rerunning the analysis, col. 5, ln. 36+);
D2 does not explicitly disclose the following which is taught by D3 (see figs. 8A to 8C):
forming a preform by: laying up a first subset of plies (each set of cavity fiber plies 80 has at least one cavity fiber ply 80. Thus, although the tubes here are shown with three plies (for cavities 74a/74b) and one ply (cavity 74c), additional or fewer plies may be provided for strength and/or wall thickness considerations) over a mandrel (mandrel 96, col. 8, ln. 64+); laying up at least one reinforcing ply over the first subset of plies, the at least one reinforcing ply coextensive with the stress regions (coextensive plies as in figs. 3-7; each set of cavity fiber plies 80 has at least one cavity fiber ply 80. Thus, although the tubes here are shown with three plies (for cavities 74a/74b) and one ply (cavity 74c), additional or fewer plies may be provided for strength and/or wall thickness considerations); and  laying up a second subset of plies over the at least one reinforcing ply (each set of cavity fiber plies 80 has at least one cavity fiber ply 80. Thus, although the tubes here are shown with three plies (for cavities 74a/74b) and one ply (cavity 74c), additional or fewer plies may be provided for strength and/or wall thickness considerations); and densifying the preform with a ceramic matrix (the fiber-reinforced composite 70 is a ceramic matrix composite, an organic matrix composite (OMC), or a metal matrix composite (MMC). For instance, the ceramic matrix composite (CMC) is formed of ceramic fiber tows that are disposed in a ceramic matrix, col. 5, ln. 52+).
It would be obvious to one of ordinary skill in the art to use the form of manufacturing taught by D3 to form a component using the design technique disclosed by D2 as it would result in a component with tailored stress response. An invention created through use of a known technique to improve similar methods in the same way is obvious.
As to claim 11
D2 as modified teaches  wherein the first subset of plies and the second subset of plies define an airfoil, the airfoil comprising:
a leading edge (see annotated figure 4);
a trailing edge (see annotated figure 4);
a suction sidewall (see annotated figure 4); and
a pressure sidewall (see annotated figure 4).
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As to claims 12-13
D2 does not explicitly disclose:
[claim 11] the at least one reinforcing ply is disposed along the pressure sidewall between 0% span and 50% span, or
[claim 13] wherein the at least one reinforcing ply is disposed along the pressure sidewall between 50% span and 100% span.
At the time of the invention, it would have been obvious to a person of ordinary skill in the art to choose the claimed dimensions as matter of mere design choice.  Applicant has not disclosed that [claim 11] the at least one reinforcing ply is disposed along the pressure sidewall between 0% span and 50% span, or  [claim 13] wherein the at least one reinforcing ply is disposed along the pressure sidewall between 50% span and 100% span  provides an advantage, is used for a particular purpose or solves a stated problem.  	At the top of MPEP 2144.04, it explains that various modifications, including changes in size and proportion, are “common practices which the court has held normally require only ordinary skill in the art and hence are considered routine expedients.” Specifically, §§IV. A cites, Gardner v. TEC Syst., Inc., 725 F.2d 1338, 220 USPQ 777 (Fed. Cir. 1984), cert. denied, 469 U.S. 830, 225 USPQ 232 (1984), wherein the Federal Circuit held that, where the only difference between the prior art and the claims was a recitation of relative dimensions of the claimed device and a device having the claimed relative dimensions would not perform differently than the prior art device, the claimed device was not patentably distinct from the prior art device.
The applicant has not disclosed that having the device with the claimed dimensions solves any stated problem or is for any particular purpose above the fact that this localized increase and thickness improves the mechanical properties of vane 54 while minimizing added weight and thickness and maintaining desired flow characteristics ([0025]). It would have been obvious to one of ordinary skill in the art before the effective filing date of the claimed invention to have modified the device of D2 to have the claimed dimensions, because this modification does not patentably distinguish the claimed device, nor cause the device to perform differently, and thus the modification of the dimensions would have been obvious.
As to claim 14
D2 as modified teaches  laying up at least one platform ply at an angle with respect to the at least one reinforcing ply (82 is at an angle to 88), the at least one reinforcing ply being a projection of the at least one platform ply (as in fig. 4).
As to claim 15
D2 as modified teaches wherein the step of densifying the preform with a ceramic matrix comprises one or a combination of chemical vapor infiltration, melt infiltration, and polymer infiltration and pyrolysis (D3 col. 9, ln. 22+ after lay-up, the fiber plies are infiltrated with the matrix material, such as by chemical vapor infiltration or chemical vapor deposition, followed by removal of the mandrels 96).
Claims 16-19 are rejected under 35 U.S.C. 103 as being unpatentable over US 2017/0268345 hereinafter D4, in view of US 2015/0152734, hereinafter D5.
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As to claim 16, D4 discloses a vane assembly (fig. 2) for a gas turbine engine (fig. 1), the vane assembly comprising:
an inner case (the engine of fig. 1 has an inner case, shown but not numbered);
an outer case (the engine of fig. 1 has an outer case, shown but not numbered); and
at least one CMC vane (at least one airfoil 102) coupled between the inner case and the outer case (fig. 2), the at least one CMC vane comprising:
an inner diameter platform (106 in fig. 3);
an outer diameter platform (104 in fig. 3); and
an airfoil (at least one airfoil 102) connected to and extending between the inner diameter platform and the outer diameter platform (fig. 2), the airfoil comprising:
a leading edge (114 in figs. 4-5);
a trailing edge (116);
a suction sidewall (120); and
a pressure sidewall (118) joined to the suction sidewall at the leading edge and the trailing edge (fig. 5), the pressure sidewall comprising a locally-thickened region (thicker portion 140 of 118 in fig. 5 is of a selected ‘region’),
wherein the locally-thickened region has a first thickness (fig. 5 shows thickness), the first thickness being a maximum thickness of the pressure sidewall (in selected region of sidewall 118 in fig. 5 the thickness is a maximum);
wherein the suction sidewall (120) has a second thickness (140), the second thickness being a maximum thickness of the suction sidewall (in fig. 5 the suction sidewall 120 thickness 140 appears a maximum) and
wherein the airfoil comprises a plurality of fiber plies (layers of plies 152 shown in fig. 6; In one embodiment, the CMC material used may be a continuous fiber reinforced CMC material. For example, suitable continuous fiber reinforced CMC materials include CMC materials reinforced with continuous carbon fibers, oxide fibers, silicon carbide monofilament fibers, and other CMC materials including continuous fiber lay-ups and/or woven fiber preforms).
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D4 does not explicitly disclose wherein the first thickness is greater than the second thickness.
An invention is obvious where some teaching, suggestion, or  motivation in the prior art would have led one of ordinary skill to modify the prior art reference or to combine prior art reference teachings to arrive at the claimed invention.
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D5 teaches an airfoil (vane 30 in fig. 1 is shown in cross section in fig. 2; an annotated fig. 2 is provided, supra, for convenience) which includes areas of different thicknesses on each of the suction and pressure sidewalls (a pressure sidewall (50) comprising a locally-thickened region (generally Z2 in fig. 2), the locally-thickened region (Z2), wherein the locally-thickened region (Z2) has a first thickness (fig. 2 shows thickness T and a thickness T’; T’ is the “first thickness”), the first thickness (T’) being a maximum thickness of the pressure sidewall ( T’ is the “first thickness” and is a maximum thickness) and wherein  the suction sidewall (52) has a second thickness (T), the second thickness being a maximum thickness of the suction sidewall (nominal suction sidewall thickness T in fig. 2 is a maximum thickness of the sidewall 52) and
the first thickness (T’) is greater than the second thickness (T; see annotated figure)).
It would be obvious to one of ordinary skill in the art to use the teachings of D5 to modify D4 to provide the wherein the first thickness is greater than the second thickness for the purpose of providing improved, localized thermal management properties ([0005]).
As to claim 17, D4 as modified discloses the assembly of claim 16, wherein the first thickness is 30% greater than the second thickness ([0034] 10% to 150% range disclosed).
As to claim 18, D4 as modified discloses the assembly of claim 17, and further discloses wherein the pressure sidewall comprises a first subset of plies (fig. 6 shows ply construction of the airfoil).
As to claim 19, D4 as modified discloses the assembly of claim 18, and further discloses wherein the locally-thickened region (thicker portion 140 in fig. 6) comprises at least one additional ply (154).
Claim 20 are rejected under 35 U.S.C. 103 as being unpatentable over D4 in view of D5 and further in view of D6 (US 2023/0383657 A1, previously presented).
As to claim 20, D4 as modified discloses the assembly of claim 19, but does not explicitly disclose the following which is taught by D6: wherein at least one additional ply (74 in fig. 5) is at least partially disposed within the inner diameter platform or the outer diameter platform (74 extends to platform 68 in fig. 5).
It would be obvious to one of ordinary skill in the art to use the teachings of D6 for the purpose of configuring fiber plies of the airfoil to facilitate mitigation of stress concentrations and thermal gradients ([0039]-[0044]).
An invention is obvious when created through some teaching, suggestion, or motivation in the prior art that would have led one of ordinary skill to modify the prior art reference or to combine prior art reference teachings to arrive at the claimed invention.
Conclusion
Applicant's amendment necessitated the new ground(s) of rejection presented in this Office action.  Accordingly, THIS ACTION IS MADE FINAL.  See MPEP § 706.07(a).  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 extension fee 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 date of this final action.
The prior art made of record on the attached PTOL-892 and not relied upon is considered pertinent to applicant's disclosure as each further discloses a state of the art.
The examiner has pointed out particular references contained in the prior art of record in the body of this action for the convenience of the applicant. Although the specified citations are representative of the teachings in the art and are applied to the specific limitations within the individual claim, other passages and figures may apply as well. Applicant should consider the entire prior art as applicable as to the limitations of the claims. It is respectfully requested from the applicant, in preparing the response, to consider fully the entire reference as potentially teaching all or part of the claimed invention, as well as the context of the passage as taught by the prior art or pointed out by the examiner.
Inquiry

APPLICANT REBUTTAL: REMARKSThis Amendment is in response to the Office Action dated March 24, 2025, in which claims 1-20 are rejected under 35 U.S.C. §103 as allegedly being obvious over varying combinations of U.S. Pub. No. 2022/0268165 (D1), U.S. Pat. No. 11,162,432 (D2), U.S. Pat. No.11,530,614 (D3), U.S. Pub. No. 2017/068345 (D4), U.S. Pub. No. 2015/0152734 (D5), and U.S.Pub. No. 2023/03838657 (D6). The rejections are traversed in view of the amendments and arguments presented herein and allowance of the present application is respectfully requested.The Applicant notes a minor inconsistency: although the Office Action is identified as non-final, the conclusion of the Office Action states that the action is made final. For the purpose of this Amendment, the Applicant presumes that this first Office Action following the RCE of March 6th, 2025, is indeed intended to be non-Final.Examiner Interview Summary An applicant-initiated examiner interview was conducted telephonically on May 28, 2025, between Examiner Newton and the Applicant's representative Thea Reilkoff. Proposed amendments to the independent claims were discussed. Examiner Newton indicated the proposed amendments overcame the prior art of record. No agreement was reached with regard to patentability.Claim Rejections under 35 U.S.C. X103 Claims 1-9 Claims 1-9 are rejected as allegedly being obvious over D1 in view of D2.Independent claim 1 is presently amended to further define the structure of the component and ply layup. Claim 5, 6, 8, and 9 are cancelled to avoid redundancy. Amended independent claim 1 recites (emphasis added):A ceramic matrix composite vane for a gas turbine engine comprising:an airfoil comprising:a leading edge;a trailing edge;a suction sidewall; and a pressure sidewall joined to the suction sidewall at the leading edge and the trailing edge, the pressure sidewall comprising a locally-thickened region, a first platform; and a second platform, wherein the airfoil is connected to and extends between the first platform and the second platform;wherein the locally-thickened region extends from one of the first platform and the second platform to a midspan region of the airfoil and wherein the locally-thickened region has a first thickness, the first thickness being a maximum thickness of the pressure sidewall;wherein the suction sidewall has a second thickness, the second thickness being a maximum thickness of the suction sidewall;wherein the first thickness is greater than the second thickness; and wherein the airfoil comprises a plurality of fiber plies;wherein at least one of the plurality of plies is disposed only in the locally thickened region of the airfoil and one of the first platform and the second platform.Amended independent claim 1 requires at least one ply to be disposed only in the locally thickened region of the airfoil and the first or second platform, with the locally-thickened region provided on the pressure sidewall and extending from the first or second platforms to a midspan region of the airfoil. D1 discloses varying a number of plies used to form tubes defining cavities 74a-74c. The tubes extend a full span of the airfoil and form portions of each of the pressure sidewall and suction sidewall. D1 teaches varying the number of plies to vary the wall thickness of the tubes. There is no teaching or suggestion of providing a ply to form a locally thickened portion that extends only a portion of the span on the pressure sidewall. D2 discloses a vane formed by additive manufacturing. While Applicant acknowledges that D2 discloses adjusting a wall thickness based on expected operational stresses, D2 is uninformative with respect to a ply layup.Because the combination of D1 and D2 fails to teach or suggest all of the limitations of amended independent claim 1, it is respectfully submitted that amended independent claim 1 and all claims depending therefrom are in condition for allowance. Notice to that effect is respectfully requested.Claims 16-20 Independent claim 16 is rejected as allegedly being obvious over D4 in view of D5.Independent claim 16 is presently amended in a manner consistent with amended independent claim 1. Claims 19 and 20 are presently cancelled to avoid redundancy. It is respectfully submitted that the combination of D4 and D5 fails to teach or suggest all of the limitations of amended independent claim 16. Specifically, neither D4 nor D5 teaches or suggests that "the at least one of the plurality of plies is disposed only in the locally thickened region of the of the pressure sidewall and one of the inner diameter platform and the outer diameter platform." Furthermore, D4 does not teach or suggest "a locally-thickened region extending from the inner diameter platform or the outer diameter platform toward a midspan region" of the pressure sidewall having a maximum thickness greater than a maximum thickness of the suction sidewall and D5 fails to make up for this deficiency. D5 teaches an increased thickness T' in the region Z2 on a pressure sidewall but this thickness is the same as the maximum thickness T' on the suction sidewall shown in region Z1. As such, the maximum thickness on the pressure sidewall is not greater than a maximum thickness on the suction sidewall as required. D1 and D2 fail to make up for the deficiencies of D4 and D5 for the reasons discussed with respect to amended independent claim 1.Because no combination of the prior art of record teaches or suggests all of the limitations of independent claim 16, it is respectfully submitted that amended independent claim 16 and call claims depending therefrom are in condition for allowance. Notice to that effect is respectfully requested.Claims 10-15 Independent claim 10 is rejected as allegedly being obvious over D2 in view of D3.Independent claim 10 is directed to a method of manufacture. Independent claim 10 is presently amended to further define the structure of the component (vane) and further define the ply layup in a manner consistent with independent claims 1 and 16. Claim 11 is presently cancelled to avoid redundancy. Amended independent claim 10 recites, in relevant part:wherein the first subset of plies and the second subset of plies define the airfoil, the airfoil comprising a leading edge, a trailing edge, a suction sidewall, and a pressure sidewall;wherein the at least one reinforcing ply covers only a portion of the span of the airfoil and forms a portion of one of the first platform and the second platform; and wherein the second set of plies wraps around the airfoil.As previously discussed, D2 discloses a vane formed by an additive manufacturing method (3D printing, rapid prototyping, direct digital manufacturing, binder jetting, selective laser melting, and direct metal laser melting) and is, therefore, uninformative to a ply layup method to form a CMC component. D3 is the issued patent of D1. As discussed with respect amended independent claim 1, D1 (and thereby D3) does not teach or suggest a reinforcing ply that covers only a portion of the span of the airfoil and forms a portion of one of the platforms. None of the other prior art of record makes up for the deficiencies of D2 and D3 and as such, it is respectfully submitted that independent claim 10 and all claims depending therefrom are in condition for allowance. Notice to that effect is respectfully requested.New Claims Claims 21-23 are newly added to depend from claims 1, 10, and 16. Claims 21-23 recite "the at least one additional ply has curved edge disposed in a midspan region of the airfoil," as shown, for example, in FIGS. 3 and 5 of the present application. No new matter is added.Conclusion By addressing particular positions taken by the Patent Office in the above remarks, Applicant does not acquiesce to other positions that have not been explicitly addressed. In addition, Applicant's arguments for the patentability of a claim should not be understood as implying that no other reasons exist for the patentability of that claim. Nothing in this response should be construed as a concession of or an intent to concede any issue with regard to any claim, except as specifically stated in this response, and the amendment or cancellation of any claim does not necessarily signify concession of unpatentability of the claim prior to its amendment or cancellation. Applicant reserves the right to prosecute the rejected claims in further prosecution of this or related applications.In view of the amendments and arguments presented herein, pending claims 1-4, 7, 10, 12-18, and 21-23 are in condition for allowance. Notice to that effect is respectfully requested.If a telephone conference would be beneficial in resolving remaining issues, please contact the Applicant's attorney, Thea Reilkoff, at 612-337-7207.The Commissioner is hereby authorized to charge any additional fees which may be required under 37 C.F.R. 1.16 and 1.17, or credit any overpayment, to Deposit Account No.210279.Respectfully submitted, KINNEY & LANGE, P.A.Date:   June 20, 2025

NEXT OFFICE ACTION: EXAMINER’S COMMENT
Response to Remarks
Applicant’s remarks, see pages 7-10, filed 6/20/2025, with respect to the claim rejection2 have been fully considered and are persuasive.  The claim rejection2 have been withdrawn.
The amendment cancelled claims 5,6,8,9,11,19 and 20 and added claims 21-23. Claims 1, 2, 3, 4, 7, 10, 12, 13, 14, 15, 16, 17, 18, 21, 22, 23 remain and are allowed.
Allowable Subject Matter
Claims 1, 2, 3, 4, 7, 10, 12, 13, 14, 15, 16, 17, 18, 21, 22, 23 are allowed. The following is an examiner’s statement of reasons for allowance: claims 1, 2, 3, 4, 7, 10, 12, 13, 14, 15, 16, 17, 18, 21, 22, 23 are allowed primarily because the prior art of record cannot anticipate Applicant’s claimed invention by a single reference nor render Applicant’s claimed invention obvious by the combination of more than one reference.
Any comments considered necessary by applicant must be submitted no later than the payment of the issue fee and, to avoid processing delays, should preferably accompany the issue fee.  Such submissions should be clearly labeled “Comments on Statement of Reasons for Allowance.”
Inquiry

CITED REFERENCE SUMMARIES: Ref US11162432B2: Claims:
A vane, comprising: an internal cavity configured to receive a flow of coaling fluid; a variable thickness wall adjacent the internal cavity; a variable thickness impingement plate separating the variable thickness wall from the internal cavity, the impingement plate including a plurality of apertures for directing the cooling fluid into an impingement cavity and against the variable thickness wall, wherein the impingement plate is configured to follow a contour of the variable thickness wall, and wherein the thickness of each section of the impingement plate is inversely proportional to a thickness of an adjacent section of the variable thickness wall; and a plurality of support beams extending between and physically connected to the impingement plate and the variable thickness wall, wherein the plurality of support beams are configured to maintain a constant separation distance between the impingement plate and the variable thickness wall and to provide structural support to the variable thickness wall.
an internal cavity configured to receive a flow of coaling fluid;
a variable thickness wall adjacent the internal cavity;
a variable thickness impingement plate separating the variable thickness wall from the internal cavity, the impingement plate including a plurality of apertures for directing the cooling fluid into an impingement cavity and against the variable thickness wall, wherein the impingement plate is configured to follow a contour of the variable thickness wall, and wherein the thickness of each section of the impingement plate is inversely proportional to a thickness of an adjacent section of the variable thickness wall; and
a plurality of support beams extending between and physically connected to the impingement plate and the variable thickness wall,
wherein the plurality of support beams are configured to maintain a constant separation distance between the impingement plate and the variable thickness wall and to provide structural support to the variable thickness wall.
The vane according to claim 1 , wherein the vane comprises a vane of a nozzle in a gas turbine system.
The vane according to claim 1 , wherein the variable thickness wall comprises one or more of a leading edge wall of the vane, a trailing edge wall of the vane, a pressure side wall of the vane, and a suction side wall of the vane.
The vane according to claim 1 , wherein the thickness of the variable thickness wall varies in accordance with expected operational forces on the vane.
The vane according to claim 4 , wherein the thickness of the variable thickness wall is 1.1 to 2.5 times a nominal wall thickness of the vane.
The vane according to claim 1 , further including a cooling channel formed in a thickened portion of the vane, wherein the cooling channel is fluidly coupled to the internal cavity.
The vane according to claim 1 , further including a channel for fluidly coupling the impingement cavity to at least one of an internal cavity of a diaphragm and a trailing edge cooling circuit of the vane.
The vane according to claim 7 , wherein the vane is formed integrally with the diaphragm.
The vane according to claim 1 , further including at least one channel for fluidly coupling the internal cavity to a wheel space of a gas turbine system.
A nozzle segment for a gas turbine system, comprising: an integrated nozzle and diaphragm, the nozzle including at least one vane, each vane including: an internal cavity configured to receive a flow of cooling fluid; a variable thickness wall adjacent the internal cavity; a variable thickness impingement plate separating the variable thickness wall from the internal cavity, the impingement plate including a plurality of apertures for directing the cooling fluid into an impingement cavity and against the variable thickness wall, wherein the impingement plate is configured to follow a contour of the variable thickness wall, and wherein the thickness of each section of the impingement plate is inversely proportional to a thickness of an adjacent section of the variable thickness wall; and a plurality of support beams extending between and physically connected to the impingement plate and the variable thickness wall, wherein the plurality of support beams are configured to maintain a constant separation distance between the impingement plate and the variable thickness wall and to provide structural support to the variable thickness wall, and wherein the thickness of the variable thickness wall varies in accordance with expected operational forces on the vane.
an integrated nozzle and diaphragm, the nozzle including at least one vane, each vane including: an internal cavity configured to receive a flow of cooling fluid; a variable thickness wall adjacent the internal cavity; a variable thickness impingement plate separating the variable thickness wall from the internal cavity, the impingement plate including a plurality of apertures for directing the cooling fluid into an impingement cavity and against the variable thickness wall, wherein the impingement plate is configured to follow a contour of the variable thickness wall, and wherein the thickness of each section of the impingement plate is inversely proportional to a thickness of an adjacent section of the variable thickness wall; and a plurality of support beams extending between and physically connected to the impingement plate and the variable thickness wall, wherein the plurality of support beams are configured to maintain a constant separation distance between the impingement plate and the variable thickness wall and to provide structural support to the variable thickness wall, and wherein the thickness of the variable thickness wall varies in accordance with expected operational forces on the vane.
an internal cavity configured to receive a flow of cooling fluid;
a plurality of support beams extending between and physically connected to the impingement plate and the variable thickness wall, wherein the plurality of support beams are configured to maintain a constant separation distance between the impingement plate and the variable thickness wall and to provide structural support to the variable thickness wall, and wherein the thickness of the variable thickness wall varies in accordance with expected operational forces on the vane.
The nozzle segment according to claim 10 , wherein the variable thickness wall comprises one or more of a leading edge wall of the vane, a trailing edge wall of the vane, a pressure side wall of the vane, and a suction side wall of the vane.
The nozzle segment according to claim 10 , wherein the thickness of the variable thickness wall is 1.1 to 2.5 times a nominal wall thickness of the vane.
The nozzle segment according to claim 10 , further including a cooling channel formed in a thickened portion of the vane, wherein the cooling channel is fluidly coupled to the internal cavity.
The nozzle segment according to claim 10 , further including a channel for fluidly coupling the impingement cavity to at least one of an internal cavity of the diaphragm and a trailing edge cooling circuit of the vane.
The nozzle segment according to claim 10 , further comprising at least one channel for fluidly coupling the internal cavity of the vane to a wheel space of the gas turbine system.
Ref US20150152734A1: Claims:
A turbine airfoil for a gas turbine engine, the turbine airfoil comprising: an outer peripheral wall comprising: an external surface, the outer peripheral wall enclosing an interior space; and a concave pressure sidewall and a convex suction sidewall joined together at a leading edge and at a trailing edge; wherein the outer peripheral wall has a varying wall thickness which incorporates a locally-thickened wall portion; and a film cooling hole comprising a shaped diffuser exit passing through the outer peripheral wall within the locally-thickened wall portion.
an outer peripheral wall comprising: an external surface, the outer peripheral wall enclosing an interior space; and a concave pressure sidewall and a convex suction sidewall joined together at a leading edge and at a trailing edge; wherein the outer peripheral wall has a varying wall thickness which incorporates a locally-thickened wall portion; and
an external surface, the outer peripheral wall enclosing an interior space; and
a concave pressure sidewall and a convex suction sidewall joined together at a leading edge and at a trailing edge;
wherein the outer peripheral wall has a varying wall thickness which incorporates a locally-thickened wall portion; and
a film cooling hole comprising a shaped diffuser exit passing through the outer peripheral wall within the locally-thickened wall portion.
The turbine airfoil of claim 1 , wherein the film cooling hole further comprises an upstream metering portion which communicates with the interior space of the airfoil and a divergent downstream portion which communicates with the external surface of the turbine airfoil.
The turbine airfoil of claim 1 , wherein the locally-thickened wall portion is defined by a discrete element protruding from an inner surface of the outer peripheral wall.
The turbine airfoil of claim 1 , wherein the outer peripheral wall comprises a tapered portion incorporating both a relatively smaller thickness and a relatively larger thickness, and the locally-thickened wall portion is defined by the relatively larger thickness.
The turbine airfoil of claim 1 , wherein the locally-thickened wall portion is defined by one of the sidewalls which is thicker than the other of the sidewalls.
The turbine airfoil of claim 1 , further comprising a rib extending between the pressure sidewall and the suction sidewall, wherein the rib and portions of the sidewalls adjacent to the rib cooperate to define two or more cavities within the interior space, and wherein one of the portions of the sidewalls defines the locally-thickened wall portion.
The turbine airfoil of claim 1 , wherein the airfoil is part of a turbine vane and extends between an arcuate outer band and an arcuate inner band.
The turbine airfoil of claim 1 , wherein the airfoil is part of a turbine blade comprising root and a tip, and the outer peripheral wall tapers in thickness from a maximum value at the root to a minimum value at the tip.
The turbine airfoil of claim 8 , wherein the outer peripheral wall comprises a first locally-thickened portion at the root and a second locally-thickened portion at the tip, wherein the first and second locally-thickened portions have equal thickness.
A turbine blade for a gas turbine engine, the turbine blade comprising: an airfoil comprising a root and a tip, the airfoil defined by an outer peripheral wall comprising: an external surface, the outer peripheral wall enclosing an interior space; and a concave pressure sidewall and a convex suction sidewall joined together at a leading edge and at a trailing edge, wherein the outer peripheral wall tapers in thickness from a maximum value at the root to a minimum value at the tip, wherein the outer peripheral wall further comprises a first locally-thickened portion at the root and a second locally-thickened portion at the tip, the first and second locally-thickened portions having equal thickness; and first and second film cooling holes each comprising a shaped diffuser exit, the first film cooling hole passing through the outer peripheral wall within the first locally-thickened portion and the second film cooling hole passing through the outer peripheral wall within the second locally-thickened portion.
an airfoil comprising a root and a tip, the airfoil defined by an outer peripheral wall comprising:
an external surface, the outer peripheral wall enclosing an interior space; and a concave pressure sidewall and a convex suction sidewall joined together at a leading edge and at a trailing edge, wherein the outer peripheral wall tapers in thickness from a maximum value at the root to a minimum value at the tip, wherein the outer peripheral wall further comprises a first locally-thickened portion at the root and a second locally-thickened portion at the tip, the first and second locally-thickened portions having equal thickness; and
a concave pressure sidewall and a convex suction sidewall joined together at a leading edge and at a trailing edge,
wherein the outer peripheral wall tapers in thickness from a maximum value at the root to a minimum value at the tip,
wherein the outer peripheral wall further comprises a first locally-thickened portion at the root and a second locally-thickened portion at the tip, the first and second locally-thickened portions having equal thickness; and
first and second film cooling holes each comprising a shaped diffuser exit, the first film cooling hole passing through the outer peripheral wall within the first locally-thickened portion and the second film cooling hole passing through the outer peripheral wall within the second locally-thickened portion.
The turbine blade of claim 10 , wherein one of the first film cooling hole and the second film cooling hole comprises an upstream metering portion which communicates with the interior space of the turbine blade and a divergent downstream portion which communicates with the external surface of the turbine blade.
The turbine blade of claim 10 , wherein the outer peripheral wall further comprises a tapered portion incorporating both a relatively smaller thickness and a relatively larger thickness, and the locally-thickened wall portion is defined by the relatively larger thickness.
The turbine blade of claim 10 , wherein the locally-thickened wall portion is defined by one of the sidewalls which is thicker than the other of the sidewalls.
The turbine blade of claim 10 , further comprising a rib extending between the concave pressure sidewall and the convex suction sidewall, wherein the rib and portions of the sidewalls adjacent to the rib cooperate to define two or more cavities within the interior space, and wherein one of the portions of the sidewalls defines the locally-thickened wall portion.
Ref US20170268345A1: Claims:
An airfoil of a turbine engine having a radial direction extending away from an engine axis, comprising: an airfoil wall having an airfoil outer surface and an airfoil inner surface, the airfoil extending radially from a first end to a second end, wherein the airfoil defines a cooling channel interior to the inner surface, and wherein a thickness is defined between the airfoil outer surface and the airfoil inner surface, with the thickness varying in the radial direction from the first end to the second end along at least one radial cross-section of the airfoil.
an airfoil wall having an airfoil outer surface and an airfoil inner surface, the airfoil extending radially from a first end to a second end, wherein the airfoil defines a cooling channel interior to the inner surface, and wherein a thickness is defined between the airfoil outer surface and the airfoil inner surface, with the thickness varying in the radial direction from the first end to the second end along at least one radial cross-section of the airfoil.
The airfoil as in claim 1 , wherein the airfoil outer surface defines a leading edge portion, a trailing edge portion, a pressure side wall, and a suction side wall, and wherein the radial cross-section of the airfoil is on the leading edge portion.
The airfoil as in claim 1 , wherein the airfoil defines at least one area of higher heat flux on the airfoil outer surface, and wherein the airfoil wall is thicker in the at least one area than on radially surrounding areas of the airfoil wall.
The airfoil as in claim 3 , wherein the airfoil wall comprises a ceramic matrix composite formed from a plurality of plies, and wherein a thicker section is defined by having more plies than adjacent sections.
The airfoil as in claim 4 , wherein the thicker section corresponds to a thermal load on the airfoil during use of the turbine engine.
The airfoil as in claim 4 , wherein the thicker section comprises additional plies on the airfoil inner surface such that the thicker section corresponds to a contour on the airfoil inner surface.
The airfoil as in claim 4 , wherein the airfoil extends radially between an inner band and an outer band such that the airfoil forms a nozzle segment.
The airfoil as in claim 7 , wherein a plurality of thicker sections and thinner sections are defined in the radial cross-section from the inner band to the outer band.
The airfoil as in claim 1 , further comprising at least one rib extending from a first point on the airfoil inner surface to a second point on the airfoil inner surface.
The airfoil as in claim 1 , further comprising: a strut disposed within the radial cooling channel.
a strut disposed within the radial cooling channel.
The airfoil as in claim 10 , wherein the strut comprises a metal.
The airfoil as in claim 10 , wherein the strut has an outer surface that defines a plurality of apertures that provide for fluid communication to the inner radial cooling passage.
The airfoil as in claim 1 , wherein the airfoil comprises a plurality of film holes in fluid communication with the radial cooling channel.
The airfoil as in claim 13 , wherein the plurality of film holes provide for bore cooling of the airfoil of at least one of the pressure side wall, the suction side wall, a leading edge portion, or a trailing edge portion.
The airfoil as in claim 13 , wherein the plurality of film holes provide for film cooling of the trailing edge portion of the airfoil.
The airfoil as in claim 1 , wherein the airfoil is a rotor blade.
A turbine nozzle of a turbine engine, comprising: an outer band; an inner band; and an airfoil extending from the outer band to the inner band; wherein the airfoil comprises an airfoil wall having an airfoil outer surface and an airfoil inner surface, and wherein a thickness is defined between the airfoil outer surface and the airfoil inner surface, with the thickness varying along at least one radial cross-section of the airfoil in a radial direction defined from the outer band to the inner band.
an inner band; and
an airfoil extending from the outer band to the inner band; wherein the airfoil comprises an airfoil wall having an airfoil outer surface and an airfoil inner surface, and wherein a thickness is defined between the airfoil outer surface and the airfoil inner surface, with the thickness varying along at least one radial cross-section of the airfoil in a radial direction defined from the outer band to the inner band.
The turbine nozzle as in claim 17 , wherein a thicker section on the airfoil wall corresponds to a thermal load on the airfoil during use of the turbine engine.
The turbine nozzle as in claim 17 , wherein the airfoil defines at least one area of higher heat flux on the airfoil outer surface of a leading edge portion, and wherein the airfoil wall is thicker in the at least one area than on radially surrounding areas of the airfoil wall.
The turbine nozzle as in claim 19 , wherein the airfoil wall comprises a ceramic matrix composite formed from a plurality of plies, and wherein a thicker section is defined within the leading edge portion by having more plies than adjacent sections.
Ref US20220268165A1: [content unavailable]
Ref US20220268165: Claims:
1. A vane arc segment comprising:
an airfoil fairing having a fairing platform and a hollow airfoil section extending there from, the hollow airfoil section defining an airfoil profile and surrounding an internal cavity, the fairing platform defining a gaspath side and a non-gaspath side, the airfoil fairing being formed of a fiber-reinforced composite comprised of fiber plies, the fiber plies including at least one cavity fiber ply that is arranged as a tube that circumscribes the internal cavity, the at least one cavity fiber ply extending through the fairing platform and defining at least a portion of an upstand
[Truncated to input limit]