Office Action 2: Final Rejection (2024-12-06)
Rejection types: 102 - Lack of Novelty, 103 - Obviousness

CLAIMS:
AMENDMENTS TO THE CLAIMS

Please amend claims 1, 3, 4, 11, and 16 such that the status of the claims is as follows:

1\. (Currently Amended) A CMC component for a gas turbine engine comprising:an airfoil comprising:
a leading edge;
a trailing edge;
a suction sidewall; and
a 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.

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

3\. (Currently Amended) The component of claim 1, wherein the locally-thickened region is 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 is extends from 100% to at least 60% and not greater than 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 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\. (Currently Amended) 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 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 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.

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 .
Remarks
This communication is in response to the reply filed on 09/10/2024.
Claims 1-20 are currently pending and have been examined.
The reply renders the 35 U.S.C.§112(d) rejection moot and it is withdrawn.
The remarks regarding the availability of – US 11,905,847 (filed 21 Oct. 2022) are noted and the rejection of claims 10-15 is withdrawn and these claims are allowed.
In view of the claim amendments, the rejection of claims 1-9 and 16-20 is moot and withdrawn.  However, a new grounds of rejection is provided, supra, as required by the claim amendments.
Claim Rejections - 35 USC § 102
The following is a quotation of the appropriate paragraphs of 35 U.S.C. 102 that form the basis for the rejections under this section made in this Office action:
A person shall be entitled to a patent unless –
(a)(1) the claimed invention was patented, described in a printed publication, or in public use, on sale or otherwise available to the public before the effective filing date of the claimed invention.
(a)(2) the claimed invention was described in a patent issued under section 151, or in an application for patent published or deemed published under section 122(b), in which the patent or application, as the case may be, names another inventor and was effectively filed before the effective filing date of the claimed invention.
References cited in this section:
D1 - US 2015/0152734 (cited in response to amendment)
Claims 1 and 5-6 are rejected under 35 U.S.C. 102(a)(1) and/or 102(a)(2) as being anticipated by D1 (US 2015/0152734).
As to claim 1, D1 discloses a CMC component (fig. 1)  for a gas turbine engine  (fig. 1) comprising:
an airfoil (vane 30 in fig. 1 is shown in cross section in fig. 2; an annotated fig. 2 is provided for convenience) comprising:
a leading edge (54 in fig. 2);
a trailing edge (56);
a suction sidewall (52); and
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);
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
wherein the first thickness (T’) is greater than the second thickness (T; see annotated figure).
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As to claim 5, D1 discloses the component of claim 1 and further comprising:
a first platform (one of inner and outer bands 32, 34 in fig. 1); and
a second platform (the other of inner and outer bands 32, 34 in fig. 1),
wherein the airfoil (30) is connected to and extends between the first platform and the second platform (fig. 1).
As to claim 6, D1 discloses the component of claim 5, and further discloses wherein the component is a vane ([0020] the type of component, e.g., rotating blade, stator vane; also may be a rotating type [0034]).
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.
References cited in this section:
D1 - US 2015/0152734 (cited in response to amendment)
D2 - US 2017/0268345 A1 (previously presented)
Claims 1-8 and 16-19 are rejected under 35 U.S.C. 103 as being unpatentable over D2 (US 2017/0268345 A1) in view of D1 (US 2015/0152734).
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As to claim 1, D2 discloses a CMC component (figs. 2-3)  for a gas turbine engine  (fig. 1) comprising:
an airfoil (at least one airfoil 102) comprising:
a leading edge (114 in figs. 4-5);
a trailing edge (116);
a suction sidewall (120); and
a pressure sidewall (118) 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).
D2 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.
D1 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 D1 to modify D2 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 2, D2 as modified discloses the component of claim 1, and further discloses wherein the first thickness is up to 30% greater than the second thickness ([0034] 10% to 150% range disclosed).
As to claim 3, D2 as modified discloses the component of claim 1, and further discloses wherein the locally-thickened region (thicker portion 140 in figs. 3 and 5) is disposed between 0% and 50% span of the airfoil (fig. 3 shows 140 between 0% and 50% span).
As to claim 4, D2 as modified discloses the component of claim 1, and further discloses wherein the locally-thickened region is disposed between and 50% span and 100% span of the airfoil (fig. 5 shows another 140 between 50% and 100% span as the thick/thin is periodic from 106 to 104).
As to claim 5, D2 as modified discloses a first platform (one of 104 and 106 in fig. 3); and a second platform (the other of 104 and 106 in fig. 3), wherein the airfoil (103) is connected to and extends between the first platform and the second platform (fig. 3).
As to claim 6, D2 further discloses wherein the component is a vane ([0019] the type of component, e.g., rotating blade, stator vane).
As to claim 7, D2 further discloses wherein the pressure sidewall comprises a first subset of plies ([0033] In certain embodiments, the airfoil wall 122 is generally formed from a plurality of plies 152, as shown in FIG. 6. However, in the hot zone 150 (i.e., where the thermal load is applied during use) includes additional plies 154 on the airfoil inner surface 103 such that the thicker section 150 corresponds to a contour on the airfoil inner surface 103. That is, additional plies 154 are placed on  the inner surface 103 to form the thicker section 140 in the hot area 150. However, the contour of the outer surface 103).
As to claim 8, D2 as modified discloses wherein the locally-thickened region comprises at least one additional ply ([0033] In certain embodiments, the airfoil wall 122 is generally formed from a plurality of plies 152, as shown in FIG. 6. However, in the hot zone 150 (i.e., where the thermal load is applied during use) includes additional plies 154 on the airfoil inner surface 103 such that the thicker section 150 corresponds to a contour on the airfoil inner surface 103. That is, additional plies 154 are placed on  the inner surface 103 to form the thicker section 140 in the hot area 150).
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As to claim 16, D2 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) 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).
D2 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.
D1 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 D1 to modify D2 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, D2 as modified discloses the assembly of claim 16, wherein the first thickness is 30% greater than the second thickness (see explanation for claim 2).
As to claim 18, D2 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, D2 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).
Claims 9 and 20 are rejected under 35 U.S.C. 103 as being unpatentable over D2 (US 2017/0268345 A1) in view of D1 (US 2015/0152734) and further in view of D3 (US 2023/0383657 A1 US 2023/0383657 A1, previously presented).
As to claim 9, D2 as modified discloses the component of claim 8 (see explanation for claim 8), but does not explicitly disclose the following which is taught by D3:
wherein the at least one additional ply (74 in fig. 5) is at least partially disposed within a first 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 D3 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.
As to claim 20, D2 as modified discloses the assembly of claim 19, but does not explicitly disclose the following which is taught by D3: 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 D3 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.
Allowable Subject Matter
Claims 10-15 are allowed.
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: REMARKS
This is in response to the Office Action of June 10, 2024, in which claim 12 was rejected under 35 U.S.C. § 112, claims 1-8 and 16-19 were rejected under 35 U.S.C. § 102(a)(1) and/or 102(a)(2) as allegedly being anticipated by D1 (US 2017/0268345 Al), claims 10-13 and 15 were rejected as allegedly being anticipated by D3 (US 11,905,847), and claims 9, 14, and 20 were rejected under U.S.C. § 103 as allegedly obvious over D1 and D3 in view of D2 (US 2023/0383657 Al). Claims 1, 3, 4, 11, and 16 are amended herein. Reconsideration and allowance in view of the amendments and arguments presented herein is respectfully requested.
Claim Rejections under 35 U.S.C.
Q
The Office Action rejected claim 12 under 35 U.S.C. § 112(d) as allegedly being of improper dependent form for failing to further limit the subject matter of the claim upon which it depends. The Examiner noted that claim 12 recites the last clause of claim 11. Claim 11 is amended herein to remove the last clause which is recited in claim 12. The Applicant respectfully requests that the rejection be withdrawn.
Claim Rejections under 35 U.S.C.
§
The Office Action rejected claims 1-8 and 16-19 under 35 U.S.C. § 102(a)(1) and/or 102(a)(2) as allegedly being anticipated by D1. Claims 1 and 16 are amended herein to clarify the requirement that the first thickness is the maximum thickness of the pressure sidewall and that the second thickness is the maximum thickness of the suction sidewall. Claims 3 and 4 are amended to require the locally-thickened region to extend from 0% to at least 40% but not greater than 60% span of the airfoil or from 100% to at least 60% but not greater than 40% span of the airfoil. Support for these amendments can be found, for example, in FIGs. 3 and 4 and    0023 and 0025 of the present application. Amended claim 1 recites the following, in relevant part (emphasis added):
an airfoil comprising:
a leading edge;
a trailing edge;
a suction sidewall; and
a pressure sidewall comprising a locally-thickened region, the 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.
Amended claim 16 recites the following, in relevant part (emphasis added):
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 pressure sidewall comprising a locally-thickened region extending from the inner diameter platform or extending from the outer diameter platform, 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.
The Office Action identified a suction sidewall 120 from D1 as analogous to the suction sidewall in claims 1 and 16. The Office Action identified a pressure sidewall 118 including a thicker region 140 from D1 as analogous to the pressure sidewall with a locally-thickened region in claims 1 and 16. The Office Action identified the thicker region 140 from D1 as analogous to the first thickness in claims 1 and 16. The Office Action identified a thinner portion 142 of the
suction sidewall 120 from D1 as analogous to the maximum thickness region and second thickness of the suction sidewall in claims 1 and 16.
D1 fails to teach or suggest a pressure sidewall with a locally-thickened region, the locally-thickened region having a first thickness that is the maximum thickness of the pressure sidewall and the suction sidewall having a second thickness that is the maximum thickness of the suction sidewall, wherein the first thickness is greater than the second thickness. Thinner portion 142 of suction sidewall 120 is not the second thickness, as it is not the maximum thickness of suction sidewall 120. The maximum thickness of suction sidewall 120 is thicker region 140, as shown in FIG. 5 below, making thicker region 140 the second thickness. D1 does not teach or suggest that the first thickness, thicker region 140 of pressure sidewall 118, is greater than the second thickness, thicker region 140 of suction sidewall 120.
FIG. S ofDl
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D1 further fails to teach or suggest the locally-thickened region extending from 0% to at least 40% but not greater than 60% span of the airfoil or extending from 100% to at least 60% but not greater than 40% span of the air foil as required in amended claims 3 and 4. As shown above in FIG. 3, thicker region 140 is not present at 0% or 100% span of the airfoil, and therefore cannot extend therefrom.
Because D1 fails to teach all the limitations of amended claim 1 and amended claim 16, it is respectfully submitted that amended claims 1 and 16 stand in condition for allowance. Notice to that effect is respectfully requested.
Claims 2-9 depend from claim 1 and claims 17-20 depend from claim 16. These claims are consequently allowable therewith at least by virtue of dependency. The rejections of claims 2-9 and 17-20 under 35 U.S.C. § 102 are moot in view of these amendments.
Dismissal of Prior Art
The Office Action rejected claims 10-13 and 15 under 35 U.S.C. § 102(a)(1) and/or § 102(a)(2) as allegedly being anticipated by D3. The Applicant respectfully traverses this rejection. The Applicant respectfully notes a rejection under § 102(a)(1) is improper as D3 was not published until 02/20/2024 and the present application was filed on 08/21/2023, before D3 was made public.
As provided in 35 U.S.C. § 102(b)(2)(C), a disclosure shall not be prior art under § 102(a)(2) if the subject matter disclosed and the claimed invention were assigned to the same person no later than the effective filing date of the claimed invention. This response serves as a clear and conspicuous statement by the applicant (or the applicant's representative of record) that the claimed invention of the application under examination and the subject matter disclosed in the U.S. patent document applied as prior art were owned by the same person or subject to an obligation of assignment to the same person not later than the effective filing date of the claimed invention sufficient to establish an exception under AIA 35 U.S.C. § 102(b)(2)(C). Both the present application and D3 were owned by RTX Corporation at the effective filing date. Therefore, it is respectfully submitted that D3 should be disqualified as prior art under 35 U.S.C. § 102(b)(2)(C). As a result, the rejections for claims 10-13 and 15 under 35 U.S.C. § 102 should accordingly be withdrawn.
Claim Re
j
ections under 35 U.S.C.
Q
The Office Action rejected claims 9 and 20 under 35 U.S.C. § 103 as allegedly being obvious over D1 in view of D2. The Office Action also rejected claim 14 under 35 U.S.C. § 103 as allegedly being obvious over D3 in view of D2. Dependent claims 9, 14, and 20 depend from independent claims 1, 10, and 16, respectively. In light of the amendments made to independent claims 1 and 16 and arguments presented, dependent claims 9, 14, and 20 are
allowable by at least virtue of dependency. As a result, the Applicant respectfully requests the rejections under 35 U.S.C. § 103 be withdrawn.
Conclusion

NEXT 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

CITED REFERENCE SUMMARIES: Ref US20190383148A1: Claims:
- 15 . (canceled)
An airfoil for disposition and exposure to flow in a hot gas duct of a turbomachine, the airfoil comprising: a suction-side wall; a pressure-side wall with respect to exposure to the flow in the gas duct, the suction-side wall and pressure-side walls converging in an upstream end region toward a leading edge and converging in a downstream end region toward a trailing edge; and a cavity structure therein enclosed between a first inner wall surface of the suction-side wall and a second inner wall surface of the pressure-side wall when viewed in a cross-sectional plane axially parallel to a longitudinal axis of the turbomachine, wherein, when viewed in the cross-sectional plane, at least one of the first and second inner wall surfaces, normalized to a mean camber line of the airfoil as a reference, extends with a change in curvature into at least one of the upstream and downstream end regions in such a way that a clearance width of the cavity structure is increased at the at least one upstream and downstream end region.
a suction-side wall;
a pressure-side wall with respect to exposure to the flow in the gas duct, the suction-side wall and pressure-side walls converging in an upstream end region toward a leading edge and converging in a downstream end region toward a trailing edge; and
a cavity structure therein enclosed between a first inner wall surface of the suction-side wall and a second inner wall surface of the pressure-side wall when viewed in a cross-sectional plane axially parallel to a longitudinal axis of the turbomachine,
wherein, when viewed in the cross-sectional plane, at least one of the first and second inner wall surfaces, normalized to a mean camber line of the airfoil as a reference, extends with a change in curvature into at least one of the upstream and downstream end regions in such a way that a clearance width of the cavity structure is increased at the at least one upstream and downstream end region.
The airfoil as recited in claim 16 wherein the at least one first and second inner wall surfaces has a smooth shape at least in the region of the change in curvature when viewed in the cross-sectional plane.
The airfoil as recited in claim 16 wherein the suction-side or pressure-side wall forming the at least one first or second inner wall surface having the change in curvature has a wall thickness which is reduced by at least 5% and no more than 35% in the region of the change in curvature when viewed in the cross-sectional plane.
The airfoil as recited in claim 16 , wherein, when viewed in the cross-sectional plane, the change in curvature is located between the at least one upstream or downstream end region and a midpoint of the airfoil located centrally between the leading edge and the trailing edge as measured along the mean camber line.
The airfoil as recited in claim 16 , wherein, as considered with respect to a duct height of the gas duct, the at least one first or second inner wall surface extends with the change in curvature over an entire duct height.
The airfoil as recited in claim 16 , wherein, as considered with respect to a duct height of the gas duct, defined between 0% radially inwardly and 100% radially outwardly, the at least one first or second inner wall surface extends with the change in curvature at least over a region between 0% and 20%.
The airfoil as recited in claim 16 , wherein, as considered with respect to a duct height of the gas duct, defined between 0% radially inwardly and 100% radially outwardly, the at least one first inner wall surface extends with the change in curvature at least over a region between 80% and 100%.
The airfoil as recited in claim 16 , wherein the at least one first or second inner wall surface extends with a change in curvature both into the upstream end region and into the downstream end region.
The airfoil as recited in claim 16 , wherein both the first and second inner wall surfaces extend with a respective change in curvature into at least one of the upstream and downstream end regions.
The airfoil as recited in claim 16 , wherein the cavity structure forms a continuous cavity between the upstream and downstream end regions when viewed in the cross-sectional plane.
The airfoil as recited in claim 16 , wherein the cavity structure forms a continuous cavity between the first and second inner wall surfaces when viewed in the cross-sectional plane.
The airfoil as recited in claim 16 , wherein the suction-side and pressure-side walls are each in themselves uninterruptedly continuous so that side wall material between the respective first and second inner wall surfaces and a respective outer wall surface is free of discontinuities.
The airfoil as recited in claim 16 , wherein, when viewed in the cross-sectional plane, at least one of the first and second inner wall surfaces, normalized to a mean camber line of the airfoil as a reference, extends with the change in curvature into each of the two upstream and downstream end regions in such a way that a clearance width of the cavity structure is increased at each of the two end regions.
A turbine module having an airfoil as recited in claim 16 .
A method of operating an airfoil as recited in claim 16 in a turbomachine, the method comprising flowing a cooling fluid through the cavity structure during operation of the turbomachine.
The method as recited in claim 30 wherein the turbomachine is a jet engine.
Ref US3420502A: Claims:
IN AN AXIAL-FLOW TURBOMACHINE: AN AIRFOIL HAVING A SOIL OUTER WALL OF SUBSTANTIAL THICKNESS, SAID WALL THICKNESS VARYING FROM A MAXIMUM AT THE RADIALLY SPACED AIRFOIL ENDS TO A MIMIMUM ADJACENT THE LONGITUDINAL CENTER OF SAID AIRFOIL, SAID WALL HAVING LENDING AND TRAILING EDGE PORTIONS AND DEFINING A CENTRAL AIRFOIL CAVITY THEREWITHIN; A PLURALITY OF LONGITUDIALLY EXTENDING COOLING FLUID PASSAGES IN THE LENDING EDGE PORTION OF SAID WALL, THE OUTER ENDS OF SAID PASSAGES BEING OPEN TO THE OUTSIDE OF SAID AIRFOIL AT THE OUTER ENDS THEREOF TO RECEIVE A COOLING FLUID FROM THE OUTSIDE OF SAID AIRFOIL, THE INNER ENDS OF SAID PASSAGES OPENING INTO SAID CAVITY ADJACENT THE LONGITUDINAL MIDPOINT OF SAID AIRFOIL WITH AT LEAST ONE PAIR OF ENDS OF SAID PASSAGES EXTENDING FROM EACH OF THE ENDS OF SAID AIRFOIL BEING SPACED IN A DIRECTLY OPPOSED RELATIONSHIP, WHEREIN THE COOLING FLUID STREAMS FROM EACH AIRFOIL END COOL SAID LEADING EDGE WALL PORTION AND INTERACT TO CREATE A TURBULENT FLOW IN SAID CAVITY; AND A PLURALITY OF GENERALLY AXIALLY EXTENDING, RADIALLY SPACED TRAILING EDGE PASSAGES EXTENDING FROM SAID CAVITY THROUGH THE TRAILING EDGE WALL PORTION TO DISCHARGE THE COOLING FLUID STREAM IN SAID CAVITY THEREFROM
Ref US7487819B2: Claims:
A method comprising: injecting into a thin wall disposable core die a slurry having a viscosity of about 1to about 1,000 Pascal-seconds at room temperature when tested at a shear rate of up to 70 seconds â1 and a flow index of less than 0.6; the injecting being carried out at a hydrostatic pressure of about 0.07 to about 7 kilograms per centimeter square; wherein the thin wall disposable core die has an average wall thickness of about 1.5 to about 10 millimeters; restoring the thin wall disposable core die to its original position upon the removal of the hydrostatic pressure; curing the slurry to form a cured ceramic core; removing the thin wall disposable core die from the cured ceramic core; and firing the cured ceramic core to form a solidified ceramic core.
injecting into a thin wall disposable core die a slurry having a viscosity of about 1to about 1,000 Pascal-seconds at room temperature when tested at a shear rate of up to 70 seconds â1 and a flow index of less than 0.6; the injecting being carried out at a hydrostatic pressure of about 0.07 to about 7 kilograms per centimeter square; wherein the thin wall disposable core die has an average wall thickness of about 1.5 to about 10 millimeters;
restoring the thin wall disposable core die to its original position upon the removal of the hydrostatic pressure;
curing the slurry to form a cured ceramic core;
removing the thin wall disposable core die from the cured ceramic core; and
firing the cured ceramic core to form a solidified ceramic core.
The method of claim 1 , wherein the thin wall disposable core die has an average wall thickness of about 2 to about 8 millimeters.
The method of claim 1 , further comprising disposing the solidified ceramic core in a wax die; wherein the wax die comprises a metal.
The method of claim 3 , further comprising injecting wax between the solidified ceramic core and the wax die.
The method of claim 4 , further comprising cooling the injected wax to form a wax component with the solidified ceramic core enclosed therein.
wax component with the solidified ceramic core enclosed therein.
The method of claim 5 , further comprising immersing the wax component into a slurry; wherein the slurry comprises ceramic particles.
The method of claim 6 , further comprising subjecting the wax component to a firing process to create a ceramic outer shell.
The method of claim 7 , further comprising removing the wax from the wax component during the firing process.
The method of claim 7 , further comprising disposing molten metal into the ceramic outer shell to form a desired metal component.
The method of claim 9 , wherein the metal component is an airfoil.
Ref US20150152734: Claims:
1. 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.
2. 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.
3. 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.
4. 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.
5. 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.
6. 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.
7. 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.
8. 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.
9. 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.
10. 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.
11. 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.
12. 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.
13. 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.
14. 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 US20230383657A1: Claims:
An airfoil comprising: an airfoil section having an airfoil wall circumscribing an internal airfoil cavity and defining leading and trailing ends and pressure and suction sides, the airfoil wall being formed of a ceramic matrix composite (CMC) laminate comprised of fiber plies disposed in a ceramic matrix, the fiber plies including a skin fiber ply defining an exterior of the airfoil section and at least one inner fiber ply lining the skin fiber ply, the at least one inner fiber ply locally terminating at one of the leading end or the trailing end such that the airfoil wall has fewer of the fiber plies in the leading end or the trailing end than in each of the pressure side and the suction side.
an airfoil section having an airfoil wall circumscribing an int
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