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and the fee set forth in 37 CFR 1.17(e) has been timely paid, the finality of the previous Office action has been withdrawn pursuant to 37 CFR 1.114.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 12/04/2025 has been entered.
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 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.
Claims 1, 6-8, 10, 11, 16-17, and 20 are rejected under 35 U.S.C. 103 as being unpatentable over Mazeaud (5163285) in view of Knapp (US-Pub 2015/0089955), and Huber (5782076).
Regarding claim 1, Mazeaud discloses an aircraft propulsion system (1, fig 1), comprising: a compressor section (6,7, and 9, fig 1) configured to produce compressed air; a combustor section (4, fig 1); a turbine section (5, fig 1) having a vane stage (17, fig 1) and a rotor stage (18, fig 1), the vane stage includes a high pressure turbine (HPT) inlet guide vane stage (17 is an inlet guide vane) and the rotor stage includes an HPT first rotor stage (18 is the first rotor), wherein the HPT inlet guide vane stage is disposed aft of the combustor section and the HPT first rotor stage is disposed aft of the HPT inlet guide vane stage; a fuel source configured to contain a non-hydrocarbon fuel (col 3, lines 5-11); a heat exchanger (21, fig 1) having an air passage having an air inlet (connection to 20a, fig 1) and an air outlet (connection to 20b, fig 2), the air passage configured to permit a passage of compressed bleed air therethrough, and a fuel passage ( col 3, lines 5-10, there would be a passage to let fuel pass through the heat exchanger, as it has to get through the heat exchanger to the fuel injector somehow) configured to permit a passage of the non-hydrocarbon fuel therethrough, wherein the air passage and the fuel passage are fluidically separate; a compressor bleed air passage (20a, fig 1) configured to receive said compressed air bleed off a core flow path (10, fig 1) aft of the compressor and upstream of the combustor section (the compressor rotor ends at 7, and the air is taken out of 10 which is defined as a duct which contains stationary vanes 11 to align the flow, it is known to have a diffusion duct after the compressor to align the flow after exiting the compressor and before entering the combustor, so the bleed air passage is after the compressor and upstream of the combustor), the compressor bleed air passage in fluid communication with the air inlet of the heat exchanger air passage; and a first cooling air passage (20b, fig 1) configured to receive a first portion of said compressed air exiting the air outlet of the heat exchanger air passage and configured to direct said first portion of compressed bleed air exiting the air outlet of the heat exchanger air passage to the HPT inlet guide vane stage and to the HPT first rotor stage (col 3, lines 25-40) from a radially inward position.
Mazeaud does not disclose wherein the bleed passage is aft of the compressor section, a first portion of said compressed bleed air is directed to the heat exchanger air passage, and a second portion of said compressed bleed air is directed to a control valve and away from the turbine section, wherein the vane stage further includes an HPT second vane stage, wherein the HPT second vane stage is disposed aft of the HPT first rotor stage; and the cooling air passage is configured to direct said compressed bleed air exiting the air outlet of the heat exchanger air passage to the HPT second vane stage, wherein the cooling air passage includes a second cooling air passage segment; and the second cooling air passage segment is configured to direct said compressed bleed air to enter the HPT second vane stage from an outer radial position.
Knapp teaches bleed passage (5, fig 1) for cooling a gas turbine (10, fig 1), wherein a first portion (along line 5, fig 1) of said compressed bleed air is directed to the heat exchanger (7, fig 1) air passage, and a second portion (16, fig 1) of said compressed bleed air is directed to a control valve (blow off bleed 2, fig 1) and away from the turbine section (fig 1, the air is dumped into the exhaust diffuser).
It would have been obvious to one of ordinary skill in the art before the effective filing date of the claimed invention to have modified the bleed air passage disclosed by Mazeaud by having a second portion of bleed air directed to a control valve and away from the turbine section based on the teachings of Knapp. Doing so would allow for the gas turbine to operate at a more efficient temperature at part loads (par. 0029), as suggested by Knapp.
Huber teaches a bleed air system for a gas turbine, the system comprising an offtake (26, fig 1) for a cooling bleed passage (71, fig 1), wherein the bleed passage is aft of the compressor section (20, fig 1), but upstream of the combustor section (10, fig 1), the turbine including a vane stage that comprises a first HPT vane stage comprising a first HPT inlet guide vane (15, fig 1) and a first HPT rotor stage (41, fig 1), and a HPT second vane stage (16, fig 1) disposed aft of the HPT first rotor stage, and the cooling air passage is configured to direct the first portion compressed bleed air exiting the air outlet of a heat exchanger (40, fig 1) passage to the HPT second vane stage (branch off line 51 to 16, fig 1) from an outer radial position and into the first rotor stage from an inner radial position (102, fig 1) simultaneously.
It would have been obvious to one of ordinary skill in the art before the effective filing date of the claimed invention to have modified the bleed air offtake passage disclosed by Mazeaud by having the offtake passage be located aft of the compressor section while still being upstream of the combustor section based on the teachings of Huber. One of ordinary skill in the art would recognize that pulling bleed air from the diffusor section between the combustor and compressor sections would reduce pressure losses in the bleed pipe, due to the reduced swirl and flow velocity of the compressed air.
It would have been obvious to one of ordinary skill in the art before the effective filing date of the claimed invention to have modified the turbine disclosed by Mazeaud by having a second vane stage turbine stage based on the teachings of Huber. One of ordinary skill in the art would recognize that having multiple turbine blade stages would allow for more work to be extracted more efficiently from the fluid flow than a single stage.
It would have been obvious to one of ordinary skill in the art before the effective filing date of the claimed invention to have modified the bleed air system disclosed by Mazeaud by having a second cooling air passage segment which directs compressed air to a second vane stage from an outer radial direction based on the teachings of Huber. One of ordinary skill in the art would be motivated to cool these latter stages as they would recognize that even the later turbine stages would be susceptible to overheating.
Regarding claim 6, Mazeaud discloses wherein the system further includes a boost compressor (26, fig 1) configured to selectively increase the pressure of the first portion of said compressed bleed air exiting the air outlet of the heat exchanger air passage.
Mazeaud does not disclose wherein the boost compressor is in fluid communication with the first cooling air passage segment and the second cooling air passage segment.
Huber teaches having a boost compressor (50, fig 1) located before the air passage splits into a first and second cooling air passage segment.
It would have been obvious to one of ordinary skill in the art before the effective filing date of the claimed invention to have modified the location of the bleed air split disclosed by Mazeaud as previously modified by Huber by having the boost compressor be located upstream of both the first and second cooling passages based on the teachings of Huber. One of ordinary skill in the art would recognize the boost compressor would allow the cooled air to reach a high enough pressure to allow the cooling air to be expelled out of the turbine and prevent hot air from backflowing into the cooling circuit.
Regarding claim 7, Mazeaud discloses wherein the non-hydrocarbon fuel includes hydrogen (col 3, lines 5-10).
Regarding claim 8, Mazeaud discloses wherein the heat exchanger is configured to permit the passage of the non-hydrocarbon fuel through the fuel passage in liquid form (due to the language configured to permit, the fuel line doesn’t actually need to use the fuel in liquid form, it must merely be capable of doing so, and a fuel line would be able to use both gaseous and liquid hydrogen within).
Regarding claim 10, Mazeaud discloses a method of cooling an aircraft turbine engine, the turbine engine including a compressor section (6,7, and 9, fig 1), a combustor section (4, fig 1), a turbine section (5, fig 1), and a fuel source (col 3, lines 12-20), the method comprising: providing a heat exchanger (21, fig 1) having an air passage that includes an air inlet and an air outlet (passage that attaches 20a and 20b via heat exchanger 21, fig 1), the air passage configured to permit a passage of air therethrough, and a fuel passage configured to permit a passage of a fuel therethrough (col 3, lines 5-10, the heat exchanger brings the fuel and air into heat exchange without mixing, meaning that they would each have separate passages within the heat exchanger), wherein the air passage and the fuel passage are fluidically separate; bleeding compressed air off of a core flow path (taking air from path 10 into passage 20a, fig 1) at a position aft of the compressor and upstream of the combustor section (the compressor rotor ends at 7, and the air is taken out of 10 which is defined as a duct which contains stationary vanes 11 to align the flow, it is known to have a diffusion duct after the compressor to align the flow after exiting the compressor and before entering the combustor, so the bleed is after the compressor and upstream of the combustor) and directing a first portion of the compressed bleed air through the heat exchanger air passage; directing the fuel through the heat exchanger fuel passage; and directing the first portion of compressed bleed air exiting the air outlet of the heat exchanger to both a vane stage of the turbine section and to a rotor stage of the turbine section (col 3, lines 25-40), the vane stage including a high pressure turbine (HPT) inlet guide vane stage (17 is an inlet guide vane stage), the HPT inlet guide vane stage disposed aft of the combustor section, the rotor stage including an HPT first rotor stage (18 is the first rotor), and the HPT first rotor stage disposed aft of the HPT inlet guide vane stage, and a first cooling air passage (20b, fig 1) configured to receive a first portion of said compressed air exiting the air outlet of the heat exchanger air passage and configured to direct said first portion of compressed bleed air exiting the air outlet of the heat exchanger air passage to the HPT inlet guide vane stage and to the HPT first rotor stage (col 3, lines 25-40) from a radially inward position.
Mazeaud does not disclose wherein the bleed passage is aft of the compressor section, a first portion of said compressed bleed air is directed to the heat exchanger air passage, and a second portion of said compressed bleed air is directed to a control valve and away from the turbine section, wherein the vane stage further includes an HPT second vane stage, wherein the HPT second vane stage is disposed aft of the HPT first rotor stage; and wherein the step of directing the compressed bleed air includes directing the compressed bleed air to the HPT second vane stage wherein the compressed bleed air is directed to the HPT second vane stage from an outer radial position.
Knapp teaches bleed passage (5, fig 1) for cooling a gas turbine (10, fig 1), wherein a first portion (along line 5, fig 1) of said compressed bleed air is directed to the heat exchanger (7, fig 1) air passage, and a second portion (16, fig 1) of said compressed bleed air is directed to a control valve (blow off bleed 2, fig 1) and away from the turbine section (fig 1, the air is dumped into the exhaust diffuser).
It would have been obvious to one of ordinary skill in the art before the effective filing date of the claimed invention to have modified the bleed air passage disclosed by Mazeaud by having a second portion of bleed air directed to a control valve and away from the turbine section based on the teachings of Knapp. Doing so would allow for the gas turbine to operate at a more efficient temperature at part loads (par. 0029), as suggested by Knapp.
Huber teaches a bleed air system for a gas turbine, the system comprising an offtake (26, fig 1) for a cooling bleed passage (71, fig 1), wherein the bleed passage is aft of the compressor section (20, fig 1), but upstream of the combustor section (10, fig 1), the turbine including a vane stage that comprises a first HPT vane stage comprising a first HPT inlet guide vane (15, fig 1) and a first HPT rotor stage (41, fig 1), and a HPT second vane stage (16, fig 1) disposed aft of the HPT first rotor stage, and the cooling air passage is configured to direct the first portion compressed bleed air exiting the air outlet of a heat exchanger (40, fig 1) passage to the HPT second vane stage (branch off line 51 to 16, fig 1) from an outer radial position and into the first rotor stage from an inner radial position (102, fig 1) simultaneously.
It would have been obvious to one of ordinary skill in the art before the effective filing date of the claimed invention to have modified the bleed air offtake passage disclosed by Mazeaud by having the offtake passage be located aft of the compressor section while still being upstream of the combustor section based on the teachings of Huber. One of ordinary skill in the art would recognize that pulling bleed air from the diffusor section between the combustor and compressor sections would reduce pressure losses in the bleed pipe, due to the reduced swirl and flow velocity of the compressed air.
It would have been obvious to one of ordinary skill in the art before the effective filing date of the claimed invention to have modified the turbine disclosed by Mazeaud by having a second vane stage turbine stage based on the teachings of Huber. One of ordinary skill in the art would recognize that having multiple turbine blade stages would allow for more work to be extracted more efficiently from the fluid flow than a single stage.
It would have been obvious to one of ordinary skill in the art before the effective filing date of the claimed invention to have modified the bleed air system disclosed by Mazeaud by having a second cooling air passage segment which directs compressed air to a second vane stage from an outer radial direction based on the teachings of Huber. One of ordinary skill in the art would be motivated to cool these latter stages as they would recognize that even the later turbine stages would be susceptible to overheating.
Regarding claim 11, Mazeaud discloses wherein the fuel includes a non-hydrocarbon fuel (col 3, lines 12-20).
Regarding claim 13, Mazeaud discloses wherein the first portion of the compressed bleed air is directed to the HPT inlet guide vane stage and to the HPT first rotor stage from an inner radial position (27, 29, fig 1).
Regarding claim 16, Mazeaud discloses using a boost compressor to selectively increase the pressure of the first portion of the compressed bleed air exiting the air outlet of the heat exchanger air passage (26, fig 1).
Mazeaud does not disclose wherein the boost compressor is in fluid communication with the first cooling air passage segment and the second cooling air passage segment.
Huber teaches having a boost compressor (50, fig 1) located before the air passage splits into a first and second cooling air passage segment.
It would have been obvious to one of ordinary skill in the art before the effective filing date of the claimed invention to have modified the location of the bleed air split disclosed by Mazeaud as previously modified by Huber by having the boost compressor be located upstream of both the first and second cooling passages based on the teachings of Huber. One of ordinary skill in the art would recognize the boost compressor would allow the cooled air to reach a high enough pressure to allow the cooling air to be expelled out of the turbine and prevent hot air from backflowing into the cooling circuit.
Regarding claim 17, Mazeaud discloses wherein the non-hydrocarbon fuel includes hydrogen (col 3, lines 12-20).
Regarding claim 20, Mazeaud discloses wherein the cooling air passage comprises channels disposed in a secondary air system of the aircraft propulsion system (an engine secondary air system is the part of the engine that takes air for cooling and other non-combustion purposes, since the cooling air passage is part of the secondary air system, and the air passage has a channel which air passes through, this means that the cooling air passage meets the claimed limitation).
Claim 18 is rejected under 35 U.S.C. 103 as being unpatentable over Mazeaud in view of Knapp and Huber as applied to claim 17 above, and further in view of Rambo (11692479)
Regarding claim 18, Mazeaud as modified by Takata does not disclose wherein the non-hydrocarbon fuel is in liquid form.
Rambo teaches a turbine engine using liquid hydrogen (col 9, lines 45-55) as a hydrogen fuel source.
It would have been obvious to one of ordinary skill in the art before the effective filing date of the claimed invention to have modified the fuel source disclosed by Mazeaud as modified by Takata by using the fuel source in its liquid form based on the teachings of Rambo. One of ordinary skill in the art would recognize that using the liquid hydrogen would reduce storage space requirements due to the liquids increased density.
Response to Arguments
Applicant’s arguments, see remarks, filed 11/04/2025, with respect to the rejection(s) of claim(s) 1 and 10 under Mazeaud in view of Knapp, Turco and Takata have been fully considered and are persuasive. Therefore, the rejection has been withdrawn. However, upon further consideration, a new ground(s) of rejection is made in view of Mazeaud as modified by Knapp and Huber.
Applicant’s arguments, see remarks, filed 11/04/2025, with respect to the 112b rejection have been fully considered and are persuasive. The 112b of claim 10 has been withdrawn.
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 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 SEAN V MEILLER whose telephone number is (571)272-9229. The examiner can normally be reached on 7:30am-5pm.
Examiner interviews are available via telephone, in-person, and video conferencing using a USPTO supplied web-based collaboration tool. To schedule an interview, applicant is encouraged to use the USPTO Automated Interview Request (AIR) at http://www.uspto.gov/interviewpractice.
If attempts to reach the examiner by telephone are unsuccessful, the examiner’s supervisor, Todd E. Manahan can be reached on 571-272-4713. The fax phone number for the organization where this application or proceeding is assigned is 571-273-8300.
Information regarding the status of an application may be obtained from the Patent Application Information Retrieval (PAIR) system. Status information for published applications may be obtained from either Private PAIR or Public PAIR. Status information for unpublished applications is available through Private PAIR only. For more information about the PAIR system, see https://ppair-my.uspto.gov/pair/PrivatePair. Should you have questions on access to the Private PAIR system, contact the Electronic Business Center (EBC) at 866-217-9197 (toll-free). If you would like assistance from a USPTO Customer Service Representative or access to the automated information system, call 800-786-9199 (IN USA OR CANADA) or 571-272-1000.
/SEAN V MEILLER/Examiner, Art Unit 3741
/DEVON C KRAMER/Supervisory Patent Examiner, Art Unit 3741