Office Action 1: Non-Final Rejection (2024-06-10)
Rejection types: 102 - Lack of Novelty, 103 - Obviousness, 112 - Written Description/Enablement

CLAIMS:
CLAIMS:

1\. 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, the locally- thickened region being a maximum thickness region of the pressure sidewall,
wherein the locally-thickened region has a first thickness;
wherein a maximum thickness region of the suction sidewall has a second thickness; and
wherein the first thickness is greater than the second thickness.
2\. The component of claim 1, wherein the first thickness is up to 30% greater than the second thickness.
3\. The component of claim 1, wherein the locally-thickened region is disposed between 0% and 50% span of the airfoil.
4\. The component of claim 1, wherein the locally-thickened region is disposed between and 50% span and 100% span of the airfoil.
5\. 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\. The component of claim 5, wherein the component is a vane.

7\. The component of claim 5, wherein the pressure sidewall comprises a first subset of plies.
8\. The component of claim 7, wherein the locally-thickened region comprises at least one additional ply.
9\. The component of claim 8, wherein the at least one additional ply is at least partially disposed within the first platform.
10\. 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\. 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,
wherein the at least one reinforcing ply is disposed along the pressure sidewall between 0% span and 50% span.
12\. 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\. 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\. 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\. 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\. 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, the locally-thickened region being a maximum thickness region of the pressure sidewall,
wherein the locally-thickened region has a first thickness;
wherein a maximum thickness region of the suction sidewall has a second thickness; and

wherein the first thickness is greater than the second thickness.
17\. The assembly of claim 16, wherein the first thickness is 30% greater than the second thickness.
18\. The assembly of claim 17, wherein the pressure sidewall comprises a first subset of plies.
19\. The assembly of claim 18, wherein the locally-thickened region comprises at least one additional ply.
20\. 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 Application No. 18/236,263 filed on 08/21/2023.
Claims 1-20 are currently pending and have been examined.
Drawings
The drawings filed 08/21/2023 are acceptable.
Claim Interpretation Under 35 USC § 112
No claim elements in this application are presumed to invoke 35 U.S.C. 112(f).
Claim Rejections - 35 USC § 112
The following is a quotation of 35 U.S.C. 112(d):
(d) REFERENCE IN DEPENDENT FORMS.—Subject to subsection (e), a claim in dependent form shall contain a reference to a claim previously set forth and then specify a further limitation of the subject matter claimed. A claim in dependent form shall be construed to incorporate by reference all the limitations of the claim to which it refers.
The following is a quotation of pre-AIA  35 U.S.C. 112, fourth paragraph:
Subject to the following paragraph [i.e., the fifth paragraph of pre-AIA  35 U.S.C. 112], a claim in dependent form shall contain a reference to a claim previously set forth and then specify a further limitation of the subject matter claimed. A claim in dependent form shall be construed to incorporate by reference all the limitations of the claim to which it refers.
Claim 12 is rejected under 35 U.S.C. 112(d) or pre-AIA  35 U.S.C. 112, 4th paragraph, as being of improper dependent form for failing to further limit the subject matter of the claim upon which it depends, or for failing to include all the limitations of the claim upon which it depends.
Claim 12 merely recites the last clause of claim 11 and thus fails to further limit the subject matter of the claim upon which it depends.
Applicant may cancel the claim(s), amend the claim(s) to place the claim(s) in proper dependent form, rewrite the claim(s) in independent form, or present a sufficient showing that the dependent claim(s) complies with the statutory requirements.
NOTE: any prior art rejection provided below is made as best understood in view of the 35 U.S.C. 112(b) issues above.
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 2017/0268345 A1
D3 – US 11,905,847 (filed 21 Oct. 2022)
Claims 1-8 and 16-19  are rejected under 35 U.S.C. 102(a)(1) and/or 102(a)(2) as being anticipated by D1.
As to claim 1, D1 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 in fig. 5), the locally-thickened region (140) being a maximum thickness region of the pressure sidewall (fig. 5), wherein the locally-thickened region has a first thickness (fig. 5 shows thickness);
wherein a maximum thickness region (thinner portion 142 in fig. 4) of the suction sidewall (120) has a second thickness (142); and
wherein the first thickness (thicker portion 140) is greater than the second thickness (thinner portion 142; it is noted that 140 appears greater than 142).
PNG
media_image1.png
712
692
media_image1.png
Greyscale
As to claim 2, D1 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, D1 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, D1 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, D1 discloses the component of claim 1 and further comprising:
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, D1 discloses the component of claim 5, and further discloses wherein the component is a vane ([0019] the type of component, e.g., rotating blade, stator vane).
As to claim 7, D1 discloses the component of claim 5, and 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, D1 discloses the component of claim 7, and further 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).
PNG
media_image2.png
713
865
media_image2.png
Greyscale
As to claim 16, D1 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 in fig. 5) extending from the inner diameter platform or the outer diameter platform toward a midspan region pressure sidewall (fig. 3), the locally-thickened region being a maximum thickness region of the pressure sidewall (figs. 4 and 5), wherein the locally-thickened region has a first thickness (fig. 5 shows thickness);
wherein a maximum thickness region (thinner portion 142 in fig. 4) of the suction sidewall (120) has a second thickness (142); and
wherein the first thickness (thicker portion 140) is greater than the second thickness (thinner portion 142; it is noted that 140 appears greater than 142).
As to claim 17, D1 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, D1 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, D1 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 10-13 and 15 are rejected under 35 U.S.C. 102(a)(1) and/or 102(a)(2) as being anticipated by D3.
PNG
media_image3.png
475
709
media_image3.png
Greyscale
As to claim 10, D3 discloses a method of forming a CMC component, the method comprising:
identifying stress regions in the component exceeding a threshold value (col. 5, ln. 55 through col. 6, ln. 5, placing thickened regions to accommodate thermal stress implicitly requires determining stress regions);
subsequently, forming a preform by:
laying up a first subset of plies over a mandrel (a method of forming the airfoil 60. As shown in FIG. 8, fiber plies 64a are laid-up around a plurality of mandrels 72);
laying up at least one reinforcing ply over the first subset of plies, the at least one reinforcing ply coextensive with the stress regions (col. 7, lns. 14-29, the ceramic fiber plies 64a are used to locally build-up the thickness which provides improved cooling at areas of high thermal stress); and
laying up a second subset of plies over the at least one reinforcing ply (col. 7, lns. 14-29); and
densifying the preform with a ceramic matrix (col. 2, lns. 50-55, the forming includes i) fabricating a fiber preform of the airfoil around a plurality of mandrels, wherein the mandrel in the first sub-cavity is radially split, and ii) densifying the fiber preform with a ceramic matrix).
As to claim 11, D3 discloses the method of claim 10, and further discloses wherein the first subset of plies and the second subset of plies define an airfoil (fig. 4), the airfoil comprising:
a leading edge (62a);
a trailing edge (62b);
a suction sidewall (62d); and
a pressure sidewall (62c),
wherein the at least one reinforcing ply is disposed along the pressure sidewall between 0% span and 50% span (span S in fig. 4 shows 0-100% and fig. 5 shows additional plies at approximately 0-50 and 50-100%; see fig. 5).
As to claim 12, D3 discloses the method of claim 11, and further discloses wherein the at least one reinforcing ply is disposed along the pressure sidewall between 0% span and 50% span (span S in fig. 4 shows 0-100% and fig. 5 shows additional plies at approximately 0-50 and 50-100%; see fig. 5).
As to claim 13, D3 discloses the method of claim 11, and further discloses wherein the at least one reinforcing ply is disposed along the pressure sidewall between 50% span and 100% span (fig. 5).
As to claim 15, D3 discloses the method of claim 10, and further discloses wherein the step of densifying the preform with a ceramic matrix comprises one or a combination of chemical vapor infiltration (col. 2, lns. 50-55, the forming includes i) fabricating a fiber preform of the airfoil around a plurality of mandrels, wherein the mandrel in the first sub-cavity is radially split, and ii) densifying the fiber preform with a ceramic matrix), melt infiltration, and polymer infiltration and pyrolysis.
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 2017/0268345 A1
D2 – US 2023/0383657 A1 (filed 27 May 2022)
D3 – US 11,905,847 (filed 21 Oct. 2022)
Claims 9 and 20 are rejected under 35 U.S.C. 103 as being unpatentable over D1 (US 2017/0268345 A1), as applied to claims 1 and 16 above, in view of D2 (US 2023/0383657 A1).
As to claim 9, D1 discloses the  component of claim 8 (see explanation for claim 8), but does not explicitly disclose the following which is taught by D2:
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 D2 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, D1 discloses the assembly of claim 19, but does not explicitly disclose the following which is taught by D2: 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 D2 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.
Claim 14 is rejected under 35 U.S.C. 103 as being unpatentable over D3 (US 11,905,847) in view of D2 (US 2023/0383657 A1).
As to claim 14, D3 discloses the method of claim 11 but does not explicitly disclose the following which is taught by D2: laying up at least one platform ply (74) at an angle with respect to the at least one reinforcing ply (72), the at least one reinforcing ply being a projection of the at least one platform ply (74 extends to platform 68).
It would be obvious to one of ordinary skill in the art to use the teachings of D2 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.
PNG
media_image4.png
1027
710
media_image4.png
Greyscale
Conclusion
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


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 .
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).
PNG
media_image1.png
606
959
media_image1.png
Greyscale
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).
PNG
media_image2.png
712
692
media_image2.png
Greyscale
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).
PNG
media_image3.png
713
865
media_image3.png
Greyscale
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

CITED REFERENCE SUMMARIES: 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 US20160003072A1: Claims:
An airfoil for a gas turbine engine comprising: a first layer forming a cavity having transitioning from a first thickness to a second thickness through a ply drop region; and a second layer secured to the first layer.
a first layer forming a cavity having transitioning from a first thickness to a second thickness through a ply drop region; and
a second layer secured to the first layer.
The airfoil according to claim 1 , comprising a space arranged between the first and second layers, and a filler is provided in the space.
The airfoil according to claim 2 , wherein the second layer terminates in ends forming a V-shape at a trailing edge of the airfoil, and the filler is provided between the first layer and second layer.
The airfoil according to claim 3 , wherein the second thickness is provided at a location between the first thickness and the filler.
The airfoil according to claim 2 , wherein the filler is provided near a leading edge of the airfoil.
The airfoil according to claim 1 , wherein each layer includes multiple plies.
The airfoil according to claim 6 , wherein the plies are constructed from a ceramic matrix composite bonded to one another by a resin.
The airfoil according to claim 7 , wherein the ceramic matrix composite is a silicon carbide material.
The airfoil according to claim 1 , wherein the airfoil is a vane.
The airfoil according to claim 9 , wherein the vane is a mid turbine frame vane.
The airfoil according to claim 9 , comprising a component passing through the cavity of the vane, the component adjacent to the first thickness.
The airfoil according to claim 9 , wherein a single cavity is provided in the airfoil.
A method of forming an airfoil comprising: wrapping a first layer about a mandrel and building a thickened area with the first layer relative to an adjacent area of the first layer; applying a filler over the thickened area; and wrapping second layer around the first layer and the filler.
wrapping a first layer about a mandrel and building a thickened area with the first layer relative to an adjacent area of the first layer;
applying a filler over the thickened area; and
wrapping second layer around the first layer and the filler.
The method according to claim 13 , comprising the step of placing the layers and the filler as an assembly in a mold.
The method according to claim 14 , comprising the step of removing the mandrel from the assembly to provide a cavity.
The method according to claim 13 , wherein the second layer is arranged in V-shape to provide the trailing edge.
The method according to claim 13 , wherein each layer includes multiple plies.
The method according to claim 17 , wherein the plies are constructed from a ceramic matrix composite bonded to one another by a resin.
The method according to claim 18 , wherein the ceramic matrix composite is a silicon carbide material.
Ref US11905847B1: [content unavailable]
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 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.
The airfoil as recited in claim 1 , wherein the at least one inner fiber ply includes an inner-most fiber ply.
The airfoil as recited in claim 1 , wherein the skin fiber ply circumscribes the internal airfoil cavity.
The airfoil as recited in claim 1 , wherein the at least one inner fiber ply has a first terminal end in the pressure side and a second terminal end in the suction side.
The airfoil as recited in claim 4 , wherein the first terminal end and the second terminal end are at an equivalent chord length position.
The airfoil as recited in claim 1 , wherein the at least one inner fiber ply includes multiple, non-consecutive fiber plies locally terminating at one of the leading end or the trailing end.
The airfoil as recited in claim 1 , wherein at a radial end of the airfoil section one or more of the fiber plies are turned out and form a platform.
The airfoil as recited in claim 1 , wherein one or more of the fiber plies have a braided or woven fiber configuration.
A gas turbine engine comprising: a compressor section; a combustor in fluid communication with the compressor section; and a turbine section in fluid communication with the combustor, the turbine section having airfoils disposed about a central axis of the gas turbine engine, each of the airfoils includes: 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 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.
a compressor section;
a combustor in fluid communication with the compressor section; and
a turbine section in fluid communication with the combustor, the turbine section having airfoils disposed about a central axis of the gas turbine engine, each of the airfoils includes: 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 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 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 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.
The gas turbine engine as recited in claim 9 , wherein there are at least four of the fiber plies, and the at least one inner fiber ply includes an inner-most fiber ply.
The gas turbine engine as recited in claim 10 , wherein the skin fiber ply circumscribes the internal airfoil cavity.
The gas turbine engine as recited in claim 11 , wherein the at least one inner fiber ply has a first terminal end in the pressure side and a second terminal end in the suction side.
The gas turbine engine as recited in claim 12 , wherein the first terminal end and the second terminal end are at an equivalent chord length position.
The gas turbine engine as recited in claim 14 , wherein at a radial end of the airfoil section one or more of the fiber plies are turned out and form a platform.
The gas turbine engine as recited in claim 9 , wherein one or more of the fiber plies have a braided or woven fiber configuration.
Ref US11905847: Claims:
1. An airfoil comprising:
an airfoil wall having a span between first and second radial ends and defining leading and trailing edges, pressure and suction sides each joining the leading and trailing edges, and an internal cavity,
the internal cavity being divided into at least first and second sub-cavities extending over the span,
the first sub-cavity defining a venturi tube, the venturi tube having a convergent-divergent geometry such that a cross-sectional area of the venturi tube decreases from the first radial end to a location along the span at which the cross-sectional area is at a minimum and the cross-sectional area increases from the location of the minimum to the second radial end, and
airfoil wall being formed of a ceramic matrix composite having ceramic fiber plies that locally build-up a thickness of at least one of the pressure or suction sides of the airfoil wall at the location of the minimum such that the venturi tube tapers in a tangential direction along the span from the first radial end to the location of the minimum and expands in the tangential direction along the span from the location of the minimum to the second radial end.
2. The airfoil as recited in claim 1, wherein the first radial end is 0% of the span, the second radial end is 100% of the span, and the location of the minimum is in a range of 40% to 60% of the span.
3. The airfoil as recited in claim 1, wherein the cross-sectional area of the venturi tube has a maximum, and the minimum differs from the maximum by 25% to 75%.
4. The airfoil as recited in claim 1, wherein the ceramic fiber plies locally build-up the thickness of both of the pressure and suction sides of the airfoil wall at the location of the minimum.
5. The airfoil as recited in claim 1, wherein the ceramic fiber plies locally build-up the thickness of only one of the pressure or suction sides of the airfoil wall at the location of the minimum.
6. A gas turbine engine comprising:
a compressor section;
a combustor in fluid communication with the compressor section; and
a turbine section in fluid communication with the combustor, the turbine section having airfoils disposed about a central axis of the gas turbine engine, each of the airfoils includes:
an airfoil wall having a span between first and second radial ends and defining leading and trailing edges, pressure and suction sides each joining the leading and trailing edges, and an internal cavity,
a rib joining the pressure and suction sides and dividing the internal cavity into at least first and second sub-cavities extending over the span, the rib being arced from the first radial end to the second radial end so as to define a curvature,
the first sub-cavity defining a venturi tube, the venturi tube having a convergent-divergent geometry such that a cross-sectional area of the venturi tube decreases from the first radial end to a location along the span at which the cross-sectional area is at a minimum and the cross-sectional area increases from the location of the minimum to the second radial end, and
due to the curvature of the rib s the venturi tube tapering in a chordal direction along the span from the first radial end to the location of the minimum and expanding in the chordal direction along the span from the location of the minimum to the second radial end.
7. The gas turbine engine as recited in claim 6, wherein the airfoil wall is formed of ceramic matrix composite.
8. The gas turbine engine as recited in claim 6, wherein the first radial end is 0% of the span, the second radial end is 100% of the span, and the location of the minimum is in a range of 40% to 60% of the span.
9. The gas turbine engine as recited in claim 6, wherein the cross-sectional area of the venturi tube has a maximum, and the minimum differs from the maximum by 25% to 75%.
10. The gas turbine engine as recited in claim 6, wherein the rib is convex in an aft chordal direction.
11. The gas turbine engine as recited in claim 6, wherein the cross-sectional area of the venturi tube is bound by the rib, the trailing edge, and the pressure and suction sides of the airfoil wall.
Ref APP-18_236263: Claims:
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.
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 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 secon
[Truncated to input limit]