FINAL 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 .
This is the fourth office action on the merits. This office action is in response to the amendment filed on 08/03/2026. Applicant has amended claims 1, 3, 5, 11, 13, and 18-19 and added claims 21-24. Claims 6-8, 10, 14-16, and 20 remain withdrawn from further consideration. Claims 1-3, 5, 11, 13, 18-19, and 21-24 are pending and examined.
Priority
Applicant’s claim for the benefit of a prior-filed application under 35 U.S.C. 119(e) or under 35 U.S.C. 120, 121, 365(c), or 386(c) is acknowledged. Applicant has not complied with one or more conditions for receiving the benefit of an earlier filing date under 35 U.S.C. 120 as follows:
The later-filed application must be an application for a patent for an invention which is also disclosed in the prior application (the parent or original nonprovisional application or provisional application). The disclosure of the invention in the parent application and in the later-filed application must be sufficient to comply with the requirements of 35 U.S.C. 112(a) or the first paragraph of pre-AIA 35 U.S.C. 112, except for the best mode requirement. See Transco Products, Inc. v. Performance Contracting, Inc., 38 F.3d 551, 32 USPQ2d 1077 (Fed. Cir. 1994).
The disclosure of the prior-filed application, Application No. 17/871,270, fails to provide adequate support or enablement in the manner provided by 35 U.S.C. 112(a) or pre-AIA 35 U.S.C. 112, first paragraph for one or more claims of this application. Application 17/871,270 has no disclosure of “a thermal transfer circuit comprising a closed loop and a pump configured to circulate a thermal routing fluid to communicate thermal energy into the working fluid, wherein the thermal routing fluid is different than the working fluid; a thermal transfer heat exchanger where thermal energy from the thermal routing fluid is communicated to the working fluid; and a primary heat exchanger where thermal energy from the primary energy conversion device is communicated to the thermal routing fluid” (claim 1, lines 8-14 and similarly for claim 11, lines 10-16 and claim 18, lines 3-6). Accordingly, claims 1-3, 5, 11, 13, 18-19, and 21-24 are not entitled to the benefit of the prior application filing date.
Claim Objections
Claim 18 is objected to because of the following informalities:
Claim 18, line 10: “the flow of working fluid” is believed to be in error for --a flow of the working fluid-- (see claim 18, line 5)
Claim 18, line 10: “a flow of cryogenic fuel” is believed to be in error for --a flow of a cryogenic fuel--
Appropriate correction is required.
Claim Interpretation
With respect to functional claim limitations and their interpretation, please refer to the Office Action mailed July 25, 2025.
Claim Rejections - 35 USC § 103
In the event the determination of the status of the application as subject to AIA 35 U.S.C. 102 and 103 (or as subject to pre-AIA 35 U.S.C. 102 and 103) is incorrect, any correction of the statutory basis (i.e., changing from AIA to pre-AIA ) for the rejection will not be considered a new ground of rejection if the prior art relied upon, and the rationale supporting the rejection, would be the same under either status.
The following is a quotation of 35 U.S.C. 103 which forms the basis for all obviousness rejections set forth in this Office action:
A patent for a claimed invention may not be obtained, notwithstanding that the claimed invention is not identically disclosed as set forth in section 102, if the differences between the claimed invention and the prior art are such that the claimed invention as a whole would have been obvious before the effective filing date of the claimed invention to a person having ordinary skill in the art to which the claimed invention pertains. Patentability shall not be negated by the manner in which the invention was made.
The factual inquiries for establishing a background for determining obviousness under 35 U.S.C. 103 are summarized as follows:
1. Determining the scope and contents of the prior art.
2. Ascertaining the differences between the prior art and the claims at issue.
3. Resolving the level of ordinary skill in the pertinent art.
4. Considering objective evidence present in the application indicating obviousness or nonobviousness.
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.
Claims 1-3, 5, 11, 13, 18-19, and 21-24 are rejected under 35 U.S.C. 103 as being unpatentable over Sen (US 2019/0249599 A1), in view of Sibilli (US 2023/0212983 A1) and Santini (US 2020/0191021 A1).
Regarding claim 1, Sen teaches (Figs. 1-3) an aircraft propulsion system (as shown in Fig. 3) comprising:
a primary energy conversion device (16 – which comprises 22, 24, 26, 28, and 29) that uses a fuel (from 90 – Fig. 2) and air (from 24) to generate power (¶ [0046], ll. 1-3) and thermal energy (¶ [0045], ll. 15-17);
a bottoming cycle (100) where a working fluid (e.g., carbon dioxide – see ¶ [0073], ll. 9-10) is circulated within a closed circuit (105 – Fig. 3; circuit starts at 106 > 124 > 114 > 108 > 124 > 122, and finally back to 106) comprising a bottoming compressor section (106 – Fig. 3) and a bottoming turbine section (108 – Fig. 3), wherein the working fluid is compressed in the bottoming compressor section (106) and expanded through the bottoming turbine section (108) to generate shaft power (for electric machine 130 via shaft 112 – Fig. 3);
a thermal transfer circuit (116 – Fig. 3) comprising a closed loop (Fig. 3 shows circuit 116 as a closed loop) and a thermal routing fluid (¶ [0067], ll. 23-24: “heat exchange fluid”) to communicate thermal energy into the working fluid, wherein the thermal routing fluid is different than the working fluid (¶ [0067], ll. 28-30: “there is no mixing of the flows of heat exchange fluid and working fluid for the embodiment depicted”);
a thermal transfer heat exchanger (114 – Fig. 3) where thermal energy from the thermal routing fluid is communicated to the working fluid; and
a primary heat exchanger (120 – Fig. 3) where thermal energy from the primary energy conversion device (16) is communicated to the thermal routing fluid (¶ [0067], ll. 10-14), wherein the primary heat exchanger (120) is mounted to a structure of the primary energy conversion device (16) – (as shown in Fig. 3, primary heat exchanger 120 is mounted to HP turbine 28, which is a structure of 16) and the thermal transfer heat exchanger (114) is located within a nacelle (50 – Figs. 1-2) of an aircraft (¶ [0049], l. 10: “an aircraft…”) – (Fig. 3 shows thermal transfer heat exchanger 114 being part of bottoming cycle 100, which is shown in Figs. 1-2 as being mounted to outer casing 18 of turbomachine 16. Since outer casing 18 is located within nacelle 50, thermal transfer heat exchanger 114 is also located within nacelle 50) and the thermal transfer circuit (116) extends from the primary heat exchanger (120) to the thermal transfer heat exchanger (114) located within the nacelle (50) – (Fig. 3 shows thermal transfer circuit 116 extending from heat exchanger 120 to heat exchanger 114, which is located within nacelle 50 as discussed above).
However, Sen does not teach that the fuel is a cryogenic fuel.
Sibilli teaches (Fig. 2) a similar aircraft propulsion system (turbine engine 1 is a turbofan) comprising a primary energy conversion device (1) that uses a cryogenic fuel (from tank 20 – see ¶ [0049], ll. 1-2).
It would have been obvious to a person having ordinary skill in the art before the effective filing date of the claimed invention to modify Sen by using a cryogenic fuel instead of a traditional jet fuel, in order to enable better thermodynamic yields and enable CO2 emissions to be reduced to zero, as taught by Sibilli (¶ [0004], ll. 1-2 and ¶ [0005]), therefore providing:
the primary energy conversion device (Sen, 16) that uses a cryogenic fuel (Sen, fuel from 90 modified by Sibilli to be cryogenic) and air (Sen, from 24) to generate power and thermal energy.
However, Sen, in view of Sibilli, does not teach the thermal transfer circuit comprising a pump configured to circulate the thermal routing fluid.
Santini teaches (Fig. 4) a system comprising a similar primary energy conversion device (162) and a thermal transfer circuit (168, which transfer thermal energy between heat exchangers 107 and 166) comprising a pump (see annotated Fig. 4 on next page) configured to circulate the thermal routing fluid (as evidenced by the curved arrow within 168).
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It would have been obvious to a person having ordinary skill in the art before the effective filing date of the claimed invention to modify Sen, in view of Sibilli, by including, in the thermal transfer circuit, a pump configured to circulate the thermal routing fluid, in order to provide a means for circulating the thermal routing fluid, as taught by Santini (evidenced by the curved arrow within circuit 168 in Fig. 4).
Regarding claim 2, Sen, in view of Sibilli and Santini, teaches the invention as claimed and as discussed above for claim 1, and the combination further teaches (Sen, Figs. 2-3) a fuel system (84 – Fig. 2) comprising a cryogenic fuel storage tank (¶ [0054], l. 9: “fuel source”) and a fuel flow path (92 – Fig. 2) for routing the cryogenic fuel (fuel from 90 modified by Sibilli to be cryogenic) to the primary energy conversion device (16), and a fuel/working fluid heat exchanger (122 – Fig. 3) where thermal energy from the working fluid is communicated to a flow of the cryogenic fuel (¶ [0069], ll. 22-25: “the engine sink heat exchanger 122 may additionally, or alternatively, be configured as a fuel heat exchanger, such as the exemplary heat exchanger 104B of FIG. 2”).
Regarding claim 3, Sen, in view of Sibilli and Santini, teaches the invention as claimed and as discussed above for claim 1, and the combination further teaches (Sen, Figs. 1-3) the primary energy conversion device (16) comprises a gas turbine engine (16) comprising a combustor (26) where the cryogenic fuel (from 90 – Fig. 2, modified to be cryogenic per Sibilli) is mixed with compressed air (from 24) and ignited to generate an exhaust gas flow (66 – Fig. 1), and the exhaust gas flow is expanded through a turbine section (28 and 30) to generate shaft power (for LP shaft 36) utilized to drive a propulsive fan (14 – Figs. 1 and 3), wherein the exhaust gas flow (66) is routed through the primary heat exchanger (120 – Fig. 3) for heating the thermal routing fluid (of circuit 116).
Regarding claim 5, Sen, in view of Sibilli and Santini, teaches the invention as claimed and as discussed above for claim 1, and Sen further teaches (Figs. 1-3) the thermal transfer heat exchanger (114 – Fig. 3) and a fuel/working fluid heat exchanger (122 – Fig. 3; see ¶ [0069], ll. 22-25: “the engine sink heat exchanger 122 may additionally, or alternatively, be configured as a fuel heat exchanger, such as the exemplary heat exchanger 104B of FIG. 2”) are located within a common aircraft structure (as shown in Fig. 3, both are part of bottoming cycle 100, which is shown in Figs. 1-2 as being located within outer casing 18 of turbomachine 16).
Regarding claim 11, Sen teaches (Figs. 1-3) an aircraft propulsion system (as shown in Fig. 3) comprising:
a core engine (16 – which comprises 22, 24, 26, 28, and 29) comprising a compressor (22 and 24), a combustor (26), and a turbine (28 and 29) where a fuel (from 90 – Fig. 2) is mixed with compressed air (from 24) from the compressor (22 and 24) in the combustor (26) and ignited to generate an exhaust gas flow (66 – Fig. 1) that is expanded through the turbine (28 and 29) to generate shaft power (for LP shaft 36);
a propulsive fan (14 – Figs. 1 and 3) coupled (via LP shaft 36) to be driven by the turbine (28 and 29);
a bottoming cycle (100) where a working fluid (e.g., carbon dioxide – see ¶ [0073], ll. 9-10) is circulated within a closed circuit (105 – Fig. 3; circuit starts at 106 > 124 > 114 > 108 > 124 > 122, and finally back to 106) comprising a bottoming compressor section (106 – Fig. 3) and a bottoming turbine section (108 – Fig. 3), wherein the working fluid is compressed in the bottoming compressor section (106) and expanded through the bottoming turbine section (108) to generate shaft power (for electric machine 130 via shaft 112 – Fig. 3);
a thermal transfer circuit (116 – Fig. 3) comprising a closed loop (Fig. 3 shows circuit 116 as a closed loop) and a thermal routing fluid (¶ [0067], ll. 23-24: “heat exchange fluid”) to transfer thermal energy between the exhaust gas flow (66) and the working fluid, wherein the thermal routing fluid is different than the working fluid (¶ [0067], ll. 28-30: “there is no mixing of the flows of heat exchange fluid and working fluid for the embodiment depicted”);
a thermal transfer heat exchanger (114 – Fig. 3) where thermal energy from the thermal routing fluid is communicated to the working fluid;
a primary heat exchanger (120 – Fig. 3; note that this may be configured in a similar manner to 102A in Fig. 2 – see ¶ [0068], ll. 12-15) where thermal energy from the core engine (16) is communicated to the thermal routing fluid (¶ [0067], ll. 10-14), wherein the primary heat exchanger (120) is mounted to a turbine exhaust case (32 – Fig. 2) of the core engine (16) – (as shown in Fig. 2, primary heat exchanger 102A is mounted to the casing of exhaust section 32) and the thermal transfer heat exchanger (114) is located in an engine nacelle (50 – Figs. 1-2) of an aircraft (¶ [0049], l. 10: “an aircraft…”) – (Fig. 3 shows thermal transfer heat exchanger 114 being part of bottoming cycle 100, which is shown in Figs. 1-2 as being mounted to outer casing 18 of turbomachine 16. Since outer casing 18 is located in engine nacelle 50, thermal transfer heat exchanger 114 is also located in engine nacelle 50), and the thermal transfer circuit (116) extends from the turbine exhaust case (32 or where “120” is located in Fig. 3) to the thermal transfer heat exchanger (114) located in the engine nacelle (50) – (Fig. 3 shows thermal transfer circuit 116 extending from heat exchanger 120 to heat exchanger 114, which is located in engine nacelle 50 as discussed above);
a fuel system (84 – Fig. 2) comprising a fuel storage tank (¶ [0054], l. 9: “fuel source”) and a fuel flow path (92 – Fig. 2) for routing the fuel (from 90) to the core engine (16); and
a fuel/working fluid heat exchanger (122 – Fig. 3) where thermal energy from the working fluid is communicated to a flow of the fuel (¶ [0069], ll. 22-25: “the engine sink heat exchanger 122 may additionally, or alternatively, be configured as a fuel heat exchanger, such as the exemplary heat exchanger 104B of FIG. 2”).
However, Sen does not teach that the fuel is a cryogenic fuel and the fuel storage tank is a cryogenic fuel storage tank.
Sibilli teaches (Fig. 2) a turbine engine (1) comprising a combustor (16), wherein a tank (20) housing fuel is to be delivered (via lines 200 and 201) to the combustor (16), and
wherein the tank (20) is for storage of cryogenic fuel (¶ [0049], ll. 1-2).
It would have been obvious to a person having ordinary skill in the art before the effective filing date of the claimed invention to modify Sen by using a cryogenic fuel instead of a traditional jet fuel, in order to enable better thermodynamic yields and enable CO2 emissions to be reduced to zero, as taught by Sibilli (¶ [0004], ll. 1-2 and ¶ [0005]), therefore providing:
the core engine (Sen, 16)…where a cryogenic fuel (Sen, fuel from 90 modified by Sibilli to be cryogenic) is mixed with compressed air (Sen, from 24) from the compressor (Sen, 22 and 24) in the combustor (Sen, 26) and ignited to generate the exhaust gas flow (Sen, 66);
the fuel system (Sen, 84) comprising a cryogenic fuel storage tank (Sibilli, 20) and the fuel flow path (Sen, 92) for routing the cryogenic fuel to the core engine (Sen, 16); and
the fuel/working fluid heat exchanger (Sen, 122) where thermal energy from the working fluid is communicated to a flow of the cryogenic fuel.
However, Sen, in view of Sibilli, does not teach the thermal transfer circuit comprising a pump configured to circulate the thermal routing fluid.
Santini teaches (Fig. 4) a system comprising a similar core engine (162) and a thermal transfer circuit (168, which transfer thermal energy between heat exchangers 107 and 166) comprising a pump (see annotated Fig. 4 below) configured to circulate the thermal routing fluid (as evidenced by the curved arrow within 168).
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It would have been obvious to a person having ordinary skill in the art before the effective filing date of the claimed invention to modify Sen, in view of Sibilli, by including, in the thermal transfer circuit, a pump configured to circulate the thermal routing fluid, in order to provide a means for circulating the thermal routing fluid, as taught by Santini (evidenced by the curved arrow within circuit 168 in Fig. 4).
Regarding claim 13, Sen, in view of Sibilli and Santini, teaches the invention as claimed and as discussed above for claim 11, and Sen further teaches (Figs. 1-3) the thermal transfer heat exchanger (114 – Fig. 3) and the fuel/working fluid heat exchanger (122 – Fig. 3) are located within a common aircraft structure (as shown in Fig. 3, both are part of bottoming cycle 100, which is shown in Figs. 1-2 as being located within outer casing 18 of turbomachine 16).
Regarding claim 18, Sen teaches (Figs. 1-3) a method of operating an aircraft propulsion system (as shown in Fig. 3) comprising:
communicating thermal energy from a heat source (combustion gases 66 – Fig. 1) into a thermal routing fluid (¶ [0067], ll. 23-24: “heat exchange fluid”) circulating within a closed loop thermal transfer circuit (116 – Fig. 3 shows circuit 116 as a closed loop);
transferring thermal energy from the thermal routing fluid into a working fluid (via heat exchanger 114 – Fig. 3) of a bottoming cycle (100) where the heated working fluid is circulated within a closed circuit (105 – Fig. 3; circuit starts at 106 > 124 > 114 > 108 > 124 > 122, and finally back to 106) comprising a bottoming compressor section (106 – Fig. 3) and a bottoming turbine section (108 – Fig. 3), wherein the working fluid is compressed in the bottoming compressor section (106) and expanded through the bottoming turbine section (108) to generate shaft power (for electric machine 130 via shaft 112 – Fig. 3); and
cooling the flow of working fluid with a flow of fuel (from 90 – Fig. 2) within a fuel/working fluid heat exchanger (122 – Fig. 3) where thermal energy from the working fluid is communicated to the flow of the fuel (¶ [0069], ll. 22-25: “the engine sink heat exchanger 122 may additionally, or alternatively, be configured as a fuel heat exchanger, such as the exemplary heat exchanger 104B of FIG. 2”).
However, Sen does not teach that the fuel is a cryogenic fuel.
Sibilli teaches (Fig. 2) a similar aircraft propulsion system (turbine engine 1 is a turbofan) comprising a primary energy conversion device (1) that uses a cryogenic fuel (from tank 20 – see ¶ [0049], ll. 1-2).
It would have been obvious to a person having ordinary skill in the art before the effective filing date of the claimed invention to modify Sen by using a cryogenic fuel instead of a traditional jet fuel, in order to enable better thermodynamic yields and enable CO2 emissions to be reduced to zero, as taught by Sibilli (¶ [0004], ll. 1-2 and ¶ [0005]), therefore providing:
cooling the working fluid flow (from bottoming cycle 100 of Sen, Fig. 3) with a cryogenic fuel (Sen, fuel from 90 modified by Sibilli to be cryogenic ) within the fuel/working fluid heat exchanger (Sen, 122) where thermal energy from the working fluid is communicated to the cryogenic fuel flow.
However, Sen, in view of Sibilli, does not teach communicating thermal energy from the heat source into the thermal routing fluid circulating within the closed loop thermal transfer circuit with a pump.
Santini teaches (Fig. 4) a system comprising: communicating thermal energy from a heat source (gas turbine engine 162) into a thermal routing fluid circulating (as evidenced by the curved arrow within 168) within a closed loop thermal transfer circuit (168, which transfer thermal energy between heat exchangers 107 and 166) with a pump (see annotated Fig. 4 on next page).
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It would have been obvious to a person having ordinary skill in the art before the effective filing date of the claimed invention to modify Sen, in view of Sibilli, by including a pump to provide communicating thermal energy from the heat source into the thermal routing fluid circulating within the closed loop thermal transfer circuit with the pump, in order to provide a means for circulating the thermal routing fluid, as taught by Santini (evidenced by the curved arrow within circuit 168 in Fig. 4).
Regarding claim 19, Sen, in view of Sibilli and Santini, teaches the invention as claimed and as discussed above for claim 18, and Sen further teaches (Figs. 1-3) wherein thermal energy (from combustion gases 66 – Fig. 1) is generated by a core engine (16) and transferred into the thermal routing fluid within a primary heat exchanger (120 – Fig. 3; note that this may be configured in a similar manner to 102A in Fig. 2 – see ¶ [0068], ll. 12-15) that is located at a turbine exhaust case (32 – Fig. 2) of the core engine (16) – (as shown in Fig. 2, primary heat exchanger 102A is mounted to the casing of exhaust section 32), and wherein the thermal transfer circuit (116) extends from the primary heat exchanger (120) to a thermal transfer heat exchanger (114 – Fig. 3) located in an engine nacelle (50) of an aircraft (¶ [0049], l. 10: “an aircraft…”) – (Fig. 3 shows thermal transfer heat exchanger 114 being part of bottoming cycle 100, which is shown in Figs. 1-2 as being mounted to outer casing 18 of turbomachine 16. Since outer casing 18 is located in engine nacelle 50, thermal transfer heat exchanger 114 is also located in engine nacelle 50. Fig. 3 shows thermal transfer circuit 116 extending from heat exchanger 120 to heat exchanger 114, which is located within engine nacelle 50 as discussed earlier).
Regarding claim 21, Sen, in view of Sibilli and Santini, teaches the invention as claimed and as discussed above for claim 1, and Sen further teaches (Figs. 1-3) wherein the bottoming compressor section (106 – Fig. 3) and the bottoming turbine section (108 – Fig. 3) are located in the nacelle (50 – Figs. 1-2) of the aircraft (Fig. 3 shows bottoming compressor section 106 and bottoming turbine section 108 being part of bottoming cycle 100, which is shown in Figs. 1-2 as being mounted to outer casing 18 of turbomachine 16. Since outer casing 18 is located in nacelle 50, bottoming compressor section 106 and bottoming turbine section 108 are also located in nacelle 50).
Regarding claim 22, Sen, in view of Sibilli and Santini, teaches the invention as claimed and as discussed above for claim 11, and Sen further teaches (Figs. 1-3) wherein the bottoming compressor section (106 – Fig. 3) and the bottoming turbine section (108 – Fig. 3) are located in the engine nacelle (50 – Figs. 1-2) of the aircraft (Fig. 3 shows bottoming compressor section 106 and bottoming turbine section 108 being part of bottoming cycle 100, which is shown in Figs. 1-2 as being mounted to outer casing 18 of turbomachine 16. Since outer casing 18 is located in engine nacelle 50, bottoming compressor section 106 and bottoming turbine section 108 are also located in engine nacelle 50).
Regarding claim 23, Sen, in view of Sibilli and Santini, teaches the invention as claimed and as discussed above for claim 1, and Sen further teaches (Figs. 1-3) wherein the bottoming compressor section (106 – Fig. 3) and the bottoming turbine section (108 – Fig. 3) are located in a structure (outer casing 18, or alternatively, nacelle 50 – Figs. 1-2) of the aircraft that is different from the structure (HP turbine 28) of the primary energy conversion device (16) to which the primary heat exchanger (120) is mounted (outer casing 18/nacelle 50 is different from HP turbine 28).
Regarding claim 24, Sen, in view of Sibilli and Santini, teaches the invention as claimed and as discussed above for claim 18, and Sen further teaches (Figs. 1-3) wherein the bottoming compressor section (106 – Fig. 3) and the bottoming turbine section (108 – Fig. 3) are located in an engine nacelle (50 – Figs. 1-2) of an aircraft (¶ [0049], l. 10: “an aircraft…”) – (Fig. 3 shows bottoming compressor section 106 and bottoming turbine section 108 being part of bottoming cycle 100, which is shown in Figs. 1-2 as being mounted to outer casing 18 of turbomachine 16. Since outer casing 18 is located in engine nacelle 50, bottoming compressor section 106 and bottoming turbine section 108 are also located in engine nacelle 50), and wherein the thermal transfer circuit (116 – Fig. 3) extends from a primary heat exchanger (120 – Fig. 3; note that this may be configured in a similar manner to 102A in Fig. 2 – see ¶ [0068], ll. 12-15) mounted to a turbine exhaust case (32 – Fig. 2) of a core engine (16) – (as shown in Fig. 2, primary heat exchanger 102A is mounted to the casing of exhaust section 32) to a thermal transfer heat exchanger (114 – Fig. 3) located in the engine nacelle (50) – (Fig. 3 shows thermal transfer circuit 116 extending from heat exchanger 120 to heat exchanger 114, which is located in engine nacelle 50 as discussed above).
Response to Arguments
Applicant’s arguments regarding the new limitations in claims 1, 11, and 21-24 have been considered but are moot in view of the new ground(s) of rejection, necessitated by Applicant's amendments. To the extent possible, Applicant's arguments have been addressed in the body of the rejections at the appropriate locations.
Regarding Applicant’s argument that “Sen does not teach a thermal transfer circuit that extends from a primary heat exchanger at the core or turbine exhaust case to a thermal transfer heat exchanger located in the nacelle” (REMARKS, last line of pg. 7 to line 2 of pg. 8), Examiner respectfully disagrees as outlined in the 35 U.S.C. 103 rejections above. Sen teaches in Fig. 3 a thermal transfer circuit 116 that extends from primary heat exchanger 120 (which is located at the core of turbomachine 16 and at the turbine exhaust case 32, as shown in Fig. 2 – note that 120 is equivalent to 102A as discussed in ¶ [0068], ll. 12-15) to a thermal transfer heat exchanger 114, which is located in nacelle 50 (Fig. 3 shows thermal transfer heat exchanger 114 being part of bottoming cycle 100, which is shown in Figs. 1-2 as being mounted to outer casing 18 of turbomachine 16. Since outer casing 18 is located in nacelle 50, thermal transfer heat exchanger 114 is also located in nacelle 50).
Applicant further argues “nor does it locate the bottoming compressor and turbine sections in the nacelle while the primary heat exchanger remains at the core” (lines 2-3 of pg. 8). Similar to thermal transfer heat exchanger 114, bottoming compressor and turbine sections 106 and 108 are also part of bottoming cycle 100, which is shown in Figs. 1-2 as being mounted to outer casing 18 of turbomachine 16. Since outer casing 18 is located in nacelle 50, bottoming compressor and turbine sections 106 and 108 are also located in nacelle 50.
Conclusion
THIS ACTION IS MADE FINAL. Applicant is reminded of the extension of time policy as set forth in 37 CFR 1.136(a).
A shortened statutory period for reply to this final action is set to expire THREE MONTHS from the mailing date of this action. In the event a first reply is filed within TWO MONTHS of the mailing date of this final action and the advisory action is not mailed until after the end of the THREE-MONTH shortened statutory period, then the shortened statutory period will expire on the date the advisory action is mailed, and any nonprovisional extension fee (37 CFR 1.17(a)) pursuant to 37 CFR 1.136(a) will be calculated from the mailing date of the advisory action. In no event, however, will the statutory period for reply expire later than SIX MONTHS from the mailing date of this final action.
Any inquiry concerning this communication or earlier communications from the examiner should be directed to HENRY NG whose telephone number is (571)272-2318. The examiner can normally be reached M-F 9:30 AM - 6:30 PM.
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/HENRY NG/ /GERALD L SUNG/ Primary Examiner, Art Unit 3741 Examiner, Art Unit 3741
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