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and the fee set forth in 37 CFR 1.17(e) has been timely paid, the finality of the previous Office action has been withdrawn pursuant to 37 CFR 1.114.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 01/23/2026 has been entered.
Response to Arguments
Applicant’s arguments, see Pages 6-7 of the response, filed 01/23/2026, with respect to the rejection(s) made under 35 U.S.C. §103, have been fully considered and are persuasive. The previously cited prior art fails to disclose the added limitation to Claims 1, 11, and 19 of the plurality of cooling air passages having a mean diameter between 10-30 mils. Therefore, the rejection has been withdrawn. However, upon further consideration, a new ground(s) of rejection is made in view of Ohtsubo (US Patent No. 5,501,892).
Information Disclosure Statement
The information disclosure statement (IDS) was filed on 01/23/2026. The submission is in compliance with the provisions of 37 CFR 1.97. Accordingly, the information disclosure statement is being considered by the examiner.
Claim Objections
Claim 19 is objected to because of the following informalities: “leasing” in Claim 19, Line 6 should read “leading”. Appropriate correction is required.
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, if the differences between the claimed invention and the prior art are such that the claimed invention as a whole would have been obvious before the effective filing date of the claimed invention to a person having ordinary skill in the art to which the claimed invention pertains. Patentability shall not be negated by the manner in which the invention was made.
This application currently names joint inventors. In considering patentability of the claims the examiner presumes that the subject matter of the various claims was commonly owned as of the effective filing date of the claimed invention(s) absent any evidence to the contrary. Applicant is advised of the obligation under 37 CFR 1.56 to point out the inventor and effective filing dates of each claim that was not commonly owned as of the effective filing date of the later invention in order for the examiner to consider the applicability of 35 U.S.C. 102(b)(2)(C) for any potential 35 U.S.C. 102(a)(2) prior art against the later invention.
Claim(s) 1 and 11 is/are rejected under 35 U.S.C. 103 as being unpatentable over Ohtsubo (US Patent No. 5,501,892) in view of Tibbott (US Patent No: 9,206,697).
Regarding Claim 1: Ohtsubo discloses an airfoil component (Figure 8, No. 10) for a gas turbine engine (Column 1, Lines 20-22), comprising a cavity (2); a trailing edge portion of the airfoil component subject to elevated temperatures during operation of the gas turbine engine (Column 6, Lines 26-29); and at least one cooling air passage (3) extending from the cavity and passing through the trailing edge portion of the airfoil component (Figure 8), wherein the trailing edge portion of the airfoil component is subject to airflow and/or geometry constraints that prevent use of film cooling (Figure 8), wherein the at least one cooling air passage has a mean diameter D between 10 mil and 30 mil (Claim 3) and an intentionally roughened surface resulting from manufacture that augments convection-based cooling of the trailing edge portion of the airfoil component (Column 3, Lines 13-16; Column 4, Lines 36-37), and wherein the intentionally roughened surface has a surface roughness that results in a surface roughness average Ra to cooling passage mean diameter ratio (Ra/D) in a range of approximately 0.02-0.1 (Column 9, Lines 17-18 – surface roughness of 6 μm = 0.006 mm and mean diameter of 0.3 mm – 0.006/0.3 = 0.02). Ohtsubo, however, fails to disclose the cavity being configured to receive cooling air bled from a compressor of the gas turbine engine.
Tibbott teaches an airfoil component (Figure 3) for a gas turbine engine, the component comprising a cooling cavity (38) configured to receive cooling air bled from a compressor of the engine (Column 2, Lines 1-2).
Before the effective filing date of the claimed invention, it would have been obvious to one of ordinary skill in the art to make the cavity of Ohtsubo configured to receive cooling air bled from a compressor of the gas turbine engine, as taught by Tibbott, for the purpose of bringing in air to the airfoil component that is relatively cool compared to the gas temperature (Column 2, Lines 1-4).
Regarding Claim 11: Ohtsubo discloses a method of making an airfoil component (Figure 8, No. 10) with augmented cooling for a gas turbine engine (Column 1, Lines 20-22), the method comprising forming the airfoil component with a cavity (2); manufacturing at least one cooling air passage (3) extending from the cavity and passing through a trailing edge portion of the airfoil component (Figure 8) subject to elevated temperatures during operation of the gas turbine engine (Column 6, Lines 26-29), wherein the trailing edge portion of the airfoil component is subject to airflow and/or geometry constraints that prevent use of film cooling (Figure 8), wherein the at least one cooling air passage has a mean diameter D between 10 mil and 30 mil (Claim 3) and an intentionally roughened surface resulting from manufacture that augments convection-based cooling of the trailing edge portion of the airfoil component (Column 3, Lines 13-16; Column 4, Lines 36-37), and wherein the intentionally roughened surface has a surface roughness that results in a surface roughness average Ra to cooling passage mean diameter ratio (Ra/D) in a range of approximately 0.02-0.1 (Column 9, Lines 17-18 – surface roughness of 6 μm = 0.006 mm and mean diameter of 0.3 mm – 0.006/0.3 = 0.02). Ohtsubo, however, fails to disclose the cavity being configured to receive cooling air bled from a compressor of the gas turbine engine.
Tibbott teaches an airfoil component (Figure 3) for a gas turbine engine, the component comprising a cooling cavity (38) configured to receive cooling air bled from a compressor of the engine (Column 2, Lines 1-2).
Before the effective filing date of the claimed invention, it would have been obvious to one of ordinary skill in the art to make the cavity of Ohtsubo configured to receive cooling air bled from a compressor of the gas turbine engine, as taught by Tibbott, for the purpose of bringing in air to the airfoil component that is relatively cool compared to the gas temperature (Column 2, Lines 1-4).
Claim(s) 4, 8-9, 14, and 18 is/are rejected under 35 U.S.C. 103 as being unpatentable over Ohtsubo and Tibbott as applied to claims 1 and 11 above, and further in view of Whittle (US Patent No: 10,711,621).
Regarding Claims 4 and 14: Ohtsubo, as modified by Tibbott, discloses the airfoil component of Claim 1 and the method of Claim 11; however, Ohtsubo fails to disclose the airfoil component being formed of a ceramic matrix composite (CMC).
Whittle teaches an airfoil component (Figures 1-3, No. 10) comprising a cooling air passage (62) with a diameter and a roughened surface (48, 60; Column 4, Lines 41-42 & 52), wherein the airfoil component is formed of a ceramic matrix composite (Column 5, Lines 3-4; Column 12, Lines 30-35).
Before the effective filing date of the claimed invention, it would have been obvious to one of ordinary skill in the art to make the airfoil component of Ohtsubo, as modified by Tibbott, out of a ceramic matrix composite, as taught by Whittle, for the purpose of providing the airfoil component with higher temperature capability, thus allowing for a reduction in cooling air flow used and consequently an increase in thermal efficiency and therefore reduced fuel burn (Column 12, Lines 30-35).
Regarding Claim 8: Ohtsubo, as modified by Tibbott and Whittle, discloses the component of Claim 4, wherein the at least one cooling air passage comprises a plurality of cooling air passages spaced apart from one another and extending from the cavity to the trailing edge of the airfoil (Ohtsubo: Figure 8).
Regarding Claim 9: Ohtsubo, as modified by Tibbott and Whittle, discloses the component of Claim 8, wherein the plurality of cooling air passages is substantially parallel and extends substantially orthogonal to the trailing edge of the airfoil (Ohtsubo: Figure 8).
Regarding Claim 18: Ohtsubo, as modified by Tibbott and Whittle, discloses the method of Claim 14, wherein manufacturing the at least one cooling air passage comprises manufacturing a plurality of cooling air passages spaced apart from one another and extending from the cavity to the trailing edge of the airfoil component (Ohtsubo: Figure 8).
Claim(s) 19 is/are rejected under 35 U.S.C. 103 as being unpatentable over Ohtsubo in view of Tibbott and Whittle.
Regarding Claim 19: Ohtsubo discloses an airfoil (Figure 8, No. 10) comprising a cavity (2); a leading edge and a trailing edge (Figure 8); and a plurality of cooling air passages (3) extending from the cavity and passing through the airfoil to the trailing edge of the airfoil (Figure 8), wherein the trailing edge of the airfoil is subject to airflow and/or geometry constraints that prevent use of film cooling (Figure 8), wherein the plurality of cooling air passages has a mean diameter D between 10 mil and 30 mil (Claim 3) and an intentionally roughened surface resulting from manufacture that augments convection-based cooling of the trailing edge portion of the airfoil component (Column 3, Lines 13-16; Column 4, Lines 36-37) and wherein the intentionally roughened surface has a surface roughness that results in a surface roughness average Ra to cooling passage mean diameter ratio (Ra/D) in a range of approximately 0.02-0.1 (Column 9, Lines 17-18 – surface roughness of 6 μm = 0.006 mm and mean diameter of 0.3 mm – 0.006/0.3 = 0.02). Ohtsubo, however, fails to disclose the airfoil being a CMC airfoil and the cavity being configured to receive cooling air bled from a compressor of the gas turbine engine.
Tibbott teaches an airfoil component (Figure 3) for a gas turbine engine, the component comprising a cooling cavity (38) configured to receive cooling air bled from a compressor of the engine (Column 2, Lines 1-2).
Before the effective filing date of the claimed invention, it would have been obvious to one of ordinary skill in the art to make the cavity of Ohtsubo configured to receive cooling air bled from a compressor of the gas turbine engine, as taught by Tibbott, for the purpose of bringing in air to the airfoil component that is relatively cool compared to the gas temperature (Column 2, Lines 1-4).
Whittle teaches an airfoil (Figures 1-3, No. 10) comprising a cooling air passage (62) with a diameter and a roughened surface (48, 60; Column 4, Lines 41-42 & 52), wherein the airfoil is a ceramic matrix composite (CMC) airfoil (Column 5, Lines 3-4; Column 12, Lines 30-35).
Before the effective filing date of the claimed invention, it would have been obvious to one of ordinary skill in the art to make the airfoil of Ohtsubo, as modified by Tibbott, a CMC airfoil, as taught by Whittle, for the purpose of providing the airfoil component with higher temperature capability, thus allowing for a reduction in cooling air flow used and consequently an increase in thermal efficiency and therefore reduced fuel burn (Column 12, Lines 30-35).
The limitation “is machined through the leading edge and/or the trailing edge to the cavity with predetermined tooling at a predetermined rotational speed and feed rate” is being treated as a product-by-process limitation; that is, the plurality of cooling air passages is made by machining. Product-by-process claims are limited ONLY to the structure implied by the cited steps, NOT to the manipulation of the recited steps. It has been held that if the product in the product-by-process claim is the same as or obvious from a product of the prior art, the claim is unpatentable, even though the prior product was made by a different process, and the burden shifts to the applicant to show an unobvious difference. See MPEP 2113.
Claim(s) 1-2 and 11-12 is/are rejected under 35 U.S.C. 103 as being unpatentable over Subramanian (US Publication No: 2019/0106990) in view of Ohtsubo.
Regarding Claim 1: Subramanian discloses an airfoil component (Figures 1-3, No. 30) for a gas turbine engine (Claim 15), comprising a cavity (24) configured to receive cooling air bled from a compressor of the gas turbine engine (Paragraph [0027], Lines 1-9); a trailing edge portion (20) of the airfoil component subject to elevated temperatures during operation of the gas turbine engine; at least one cooling air passage (22) extending from the cavity and passing through the trailing edge portion of the airfoil component (Figure 3; Paragraph [0027], Lines 1-4), wherein the trailing edge portion of the airfoil component is subject to airflow and/or geometry constraints that prevent use of film cooling (Figures 1-3), wherein the at least one cooling air passage has a mean diameter and an intentionally roughened surface (Figure 6, No. 37) resulting from manufacture that augments convection-based cooling of the trailing edge portion of the airfoil component (Paragraph [0035], Lines 3-8), and wherein the intentionally roughened surface has a surface roughness (Figure 6; Paragraph [0035], Lines 4-8). Subramanian, however, fails to disclose the mean diameter being between 10-30 mils and a surface roughness average to cooling passage mean diameter being in a range of approximately 0.02-0.1.
Ohtsubo discloses an airfoil component (Figure 8, No. 10) for a gas turbine engine (Column 1, Lines 20-22), comprising a cavity (2); a trailing edge portion of the airfoil component subject to elevated temperatures during operation of the gas turbine engine (Column 6, Lines 26-29); and at least one cooling air passage (3) extending from the cavity and passing through the trailing edge portion of the airfoil component (Figure 8), wherein the at least one cooling air passage has a mean diameter D between 10 mil and 30 mil (Claim 3) and an intentionally roughened surface resulting from manufacture that augments convection-based cooling of the trailing edge portion of the airfoil component (Column 3, Lines 13-16; Column 4, Lines 36-37) and wherein the intentionally roughened surface has a surface roughness that results in a surface roughness average Ra to cooling passage mean diameter ratio (Ra/D) in a range of approximately 0.02-0.1 (Column 9, Lines 17-18 – surface roughness of 6 μm = 0.006 mm and mean diameter of 0.3 mm – 0.006/0.3 = 0.02).
Before the effective filing date of the claimed invention, it would have been obvious to one of ordinary skill in the art to make the mean diameter of Subramanian between 10-30 mils and the surface roughness average to mean diameter ratio in a range of approximately 0.02-0.1, as taught by Ohtsubo, for the purpose of providing the airfoil component with a high flexural strength (Column 8, Lines 64-67; Column 9, Lines 1-18).
Regarding Claim 2: Subramanian, as modified by Ohtsubo, discloses the component of Claim 1, wherein the component is metallic (Subramanian: Paragraph [0006], Lines 1-8).
Regarding Claim 11: Subramanian discloses a method of making an airfoil component (30) with augmented cooling for a gas turbine engine (Claim 15), comprising forming the airfoil component with a cavity (24) configured to receive cooling air bled from a compressor of the gas turbine (Paragraph [0027], Lines 1-9); and manufacturing at least one cooling air passage (22) extending from the cavity and passing through a trailing edge portion (20) of the airfoil component subject to elevated temperatures during operation of the gas turbine engine (Figure 3; Paragraph [0027], Lines 1-4), wherein the trailing edge portion of the airfoil component is subject to airflow and/or geometry constraints that prevent the use of film cooling (Figures 1-3), wherein the at least one cooling air passage has a mean diameter (Paragraph [0033], Lines 13-16) and an intentionally roughened surface (37) resulting from the manufacturing that augments convection-based cooling of the portion of the component (Paragraph [0035], Lines 1-8). Subramanian, however, fails to disclose the mean diameter being between 10-30 mils and a surface roughness average to cooling passage mean diameter being in a range of approximately 0.02-0.1.
Ohtsubo discloses an airfoil component (Figure 8, No. 10) for a gas turbine engine (Column 1, Lines 20-22), comprising a cavity (2); a trailing edge portion of the airfoil component subject to elevated temperatures during operation of the gas turbine engine (Column 6, Lines 26-29); and at least one cooling air passage (3) extending from the cavity and passing through the trailing edge portion of the airfoil component (Figure 8), wherein the at least one cooling air passage has a mean diameter D between 10 mil and 30 mil (Claim 3) and an intentionally roughened surface resulting from manufacture that augments convection-based cooling of the trailing edge portion of the airfoil component (Column 3, Lines 13-16; Column 4, Lines 36-37) and wherein the intentionally roughened surface has a surface roughness that results in a surface roughness average Ra to cooling passage mean diameter ratio (Ra/D) in a range of approximately 0.02-0.1 (Column 9, Lines 17-18 – surface roughness of 6 μm = 0.006 mm and mean diameter of 0.3 mm – 0.006/0.3 = 0.02).
Before the effective filing date of the claimed invention, it would have been obvious to one of ordinary skill in the art to make the mean diameter of Subramanian between 10-30 mils and the surface roughness average to mean diameter ratio in a range of approximately 0.02-0.1, as taught by Ohtsubo, for the purpose of providing the airfoil component with a high flexural strength (Column 8, Lines 64-67; Column 9, Lines 1-18).
Regarding Claim 12: Subramanian, as modified by Ohtsubo, discloses the method of Claim 11, wherein the airfoil component is formed of metal alloy (Subramanian: Paragraph [0006], Lines 1-8; Paragraph [0045], Lines 1-9).
Conclusion
Any inquiry concerning this communication or earlier communications from the examiner should be directed to MICHAEL L SEHN whose telephone number is (571)270-3564. The examiner can normally be reached M-F 8:30 AM-6 PM, every other Friday off.
If attempts to reach the examiner by telephone are unsuccessful, the examiner’s supervisor, Courtney Heinle can be reached at 571-270-3508. The fax phone number for the organization where this application or proceeding is assigned is 571-273-8300.
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/MICHAEL L SEHN/Primary Examiner, Art Unit 3745