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 .
DETAILED ACTION
Drawings
The drawings are objected to under 37 CFR 1.83(a). The drawings must show every feature of the invention specified in the claims. Therefore, the following feature(s) must be shown or the feature(s) canceled from the claim(s). No new matter should be entered.
“an open rotor gas turbine engine” in claim 19 having “an outlet disposed aft of a downstream end of a nacelle of the gas turbine engine” in claim 1 is not shown in Fig. 41 of the instant specification.
Corrected drawing sheets in compliance with 37 CFR 1.121(d) are required in reply to the Office action to avoid abandonment of the application. Any amended replacement drawing sheet should include all of the figures appearing on the immediate prior version of the sheet, even if only one figure is being amended. The figure or figure number of an amended drawing should not be labeled as “amended.” If a drawing figure is to be canceled, the appropriate figure must be removed from the replacement sheet, and where necessary, the remaining figures must be renumbered and appropriate changes made to the brief description of the several views of the drawings for consistency. Additional replacement sheets may be necessary to show the renumbering of the remaining figures. Each drawing sheet submitted after the filing date of an application must be labeled in the top margin as either “Replacement Sheet” or “New Sheet” pursuant to 37 CFR 1.121(d). If the changes are not accepted by the examiner, the applicant will be notified and informed of any required corrective action in the next Office action. The objection to the drawings will not be held in abeyance.
Claim Rejections - 35 USC § 112
The following is a quotation of the first paragraph of 35 U.S.C. 112(a):
(a) IN GENERAL.—The specification shall contain a written description of the invention, and of the manner and process of making and using it, in such full, clear, concise, and exact terms as to enable any person skilled in the art to which it pertains, or with which it is most nearly connected, to make and use the same, and shall set forth the best mode contemplated by the inventor or joint inventor of carrying out the invention.
The following is a quotation of the first paragraph of pre-AIA 35 U.S.C. 112:
The specification shall contain a written description of the invention, and of the manner and process of making and using it, in such full, clear, concise, and exact terms as to enable any person skilled in the art to which it pertains, or with which it is most nearly connected, to make and use the same, and shall set forth the best mode contemplated by the inventor of carrying out his invention.
Claim(s) 19 is rejected under 35 U.S.C. 112(a) or 35 U.S.C. 112 (pre-AIA ), first paragraph, as failing to comply with the written description requirement. The claim(s) contains subject matter which was not described in the specification in such a way as to reasonably convey to one skilled in the relevant art that the inventor or a joint inventor, or for applications subject to pre-AIA 35 U.S.C. 112, the inventor(s), at the time the application was filed, had possession of the claimed invention.
Claim 19 depends upon claim 1, thus also includes the amended limitation of “an outlet disposed aft of a downstream end of a nacelle of the gas turbine engine” of claim 1. However, claim 19 requires that the gas turbine engine is an open rotor gas turbine engine, where a nacelle and an outlet disposed aft of a downstream end of the nacelle are not shown in any of the drawings or described in the specification as part of an open rotor gas turbine engine. Therefore, claim 19 is interpreted to be new matter that was not described in the specification.
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.
Claim(s) 1-4, 6, 9-12, 15, and 20 is/are rejected under 35 U.S.C. 103 as being unpatentable over Glahn (US 20110247344 A1) in view of Stretton (US 20070245739 A1).
Regarding claim 1
Glahn discloses a gas turbine engine (Fig 1), comprising:
a turbomachine comprising a compressor section (16, 18, Para 0014) having a low pressure compressor (16) and a high pressure compressor (18), a combustion section (20) defining a compressor discharge cavity (60, Para 0022 top), and a turbine section (22, 24),
the turbomachine defining a working gas flowpath (gas flowpath through compressor 16, 18, combustor 20, and turbine 22, 24) and further comprising a cooled cooling air (CCA) system, the CCA system comprising:
a cold side bleed assembly (compressor bleed manifold 56 and pipes 42, para 0020 top) defining an inlet (inlet of manifold 56) in fluid communication with the working gas flowpath (flowpath through compressor 16, 18, combustor 20, and turbine 22, 24) at a location between the low pressure compressor and the high pressure compressor (cold side bleed at a location in between LPC 16 and HPC 18),
a CCA heat exchanger (HX 50, Para 0022 top) in thermal communication with the cold side bleed assembly (56) downstream of the inlet of the cold side bleed assembly (HX 50 is downstream of inlet of cold side bleed 56); and
a hot side bleed assembly (high pressure compressor pipes 52, Figs 1-2) defining an inlet (inlet of pipes 52 at discharge cavity 60) in fluid communication with the working gas flowpath through the compressor section (HP compressor 18), at the compressor discharge cavity (60), or both,
the hot side bleed assembly (pipes 52, Fig 2) further in thermal communication with the CCA heat exchanger (HX 50) to cool an airflow through the hot side bleed assembly (HX 50 to cool the hot air flow in pipe 52),
the hot side bleed assembly (pipes 52, Fig 2) further in thermal communication with a hot component (rear hub 40 leading to HP turbine 22, Fig 2) of the turbomachine to cool the hot component of the turbomachine (bleed air from pipes 52 flows into the airstream 62 which is exhausted downstream of the rear hub 40 into the turbine cooling system and cools the hotter sections of the gas turbine engine 10, such as the HP turbine 22, Para 0024).
Glahn is silent on the cold bleed assembly having an outlet disposed aft of a downstream end of a nacelle of the gas turbine engine, wherein the gas turbine engine comprises a bifurcation defining a bifurcation cavity, wherein the turbomachine defines an under-cowl cavity, and wherein the CCA heat exchanger is located in the under-cowl cavity of the turbomachine, the bifurcation cavity, or both.
However, Stretton teaches a gas turbine engine having a cold bleed assembly (cold bleed air 23 in Fig 2a) defining an outlet (outlet 29 of duct 28 downstream of a heat exchanger 21, Para 0027 top) disposed aft of a downstream end of a nacelle (42, Fig 2a, Para 0025 middle, where outlet end 29 is located downstream of and outside an exit plane 112 of nacelle 42, Para 0027 middle) of the gas turbine engine,
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wherein the gas turbine engine comprises a bifurcation (a bifurcation 9, Fig 1B, Para 0002 middle) defining a bifurcation cavity (interior of bifurcation, annotated in Fig 2a), wherein the turbomachine defines an under-cowl cavity (ventilation zone 111 located in the under-cowl cavity, annotated in Fig 2a, Para 0027 top), and wherein the CCA heat exchanger (21, Para 0026) is located in the under-cowl cavity of the turbomachine, the bifurcation cavity, or both (heat exchanger 21 located in both the bifurcation cavity and the under-cowl cavity as seen in Fig 2a).
Therefore, it would have been obvious to one of ordinary skill in the art before the effective filing date to modify the cold side bleed assembly in Glahn to have an outlet disposed aft of a downstream end of a nacelle of the gas turbine engine, as suggested and taught by Stretton because having an outlet end at this location can maximize thrust recovery (Para 0027 bottom, Para 0030 bottom).
Regarding claim 2
Glahn in view of Stretton discloses the gas turbine engine of claim 1.
Glahn further discloses wherein the turbine section comprises a high pressure turbine (high pressure turbine 22, Fig 1, Para 0024), and wherein the hot component is the high pressure turbine (22).
Regarding claim 3
Glahn in view of Stretton discloses the gas turbine engine of claim 1.
Glahn further discloses wherein the hot component is a rotor bore of a turbine of the turbine section, an airfoil of the turbine (airstream 62 is exhausted downstream of the rear hub 40 into the turbine cooling system and cools the hotter sections, such as the high pressure turbine 22, Para 0024 where HP turbine 22 has a rotor bore where the turbine blades are rotating on).
Regarding claim 4
Glahn in view of Stretton discloses the gas turbine engine of claim 1.
Glahn in view of Stretton further discloses wherein the CCA heat exchanger is located in the under-cowl cavity of the turbomachine (Stretton teaches the heat exchanger 21 positioned in the under-cowl cavity 111, Fig 2a).
Regarding claim 6
Glahn in view of Stretton discloses the gas turbine engine of claim 1.
Glahn further discloses wherein the gas turbine engine is a turbofan engine (Fig 1, engine 10 is a turbofan engine) comprising a fan (14) driven by the turbomachine (core engine 16, 18, 20, 22, 24) and an outer nacelle (fan case surrounding fan 14) surrounding the fan (14) and defining a bypass passage (bypass passage in between core engine casing 44 and fan 14 nacelle, Para 0020 middle, Fig 1) with the turbomachine.
Regarding claim 9
Glahn in view of Stretton discloses the gas turbine engine of claim 1.
Glahn further discloses wherein the inlet (inlet of manifold 56) is in direct fluid communication with the working gas flowpath (flowpath through compressor 16, 18, combustor 20, and turbine 22, 24).
Regarding claim 10
Glahn in view of Stretton discloses the gas turbine engine of claim 1.
Glahn further discloses wherein the CCA heat exchanger is a first CCA heat exchanger (four HX 50 shown in Fig 3a where the top HX 50 of the 1ST portion is interpreted to be the 1ST HX),
wherein the CCA system further comprises a second CCA heat exchanger (2ND HX 50 associated with 2ND portion annotated in Fig 3a),
wherein the cold side bleed assembly (56) comprises a first portion and a second portion arranged in parallel flow (four portions in four ducts 42),
wherein the first portion is in thermal communication with the first CCA heat exchanger and the second portion is in thermal communication with the second CCA heat exchanger (each portion 42 is associated with each HX 50).
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Regarding claim 11
Glahn in view of Stretton discloses the gas turbine engine of claim 10.
Glahn further discloses wherein the hot side bleed assembly (52, Fig 2) is a first hot side bleed assembly and is in thermal communication with the first CCA heat exchanger (struts 58 are integrated with the high pressure compressor pipes 52, Para 0026 middle, since there are four struts 58 shown in Fig 3a, this indicates that there are four hot side bleed assemblies in thermal communication with four CCA heat exchangers 50),
wherein the CCA system further comprises a second hot side bleed assembly in thermal communication with the second CCA heat exchanger (four HX 50 shown in Fig 3a, in thermal communication with four hot side bleed assemblies 58).
Regarding claim 12
Glahn in view of Stretton discloses the gas turbine engine of claim 1.
Glahn further discloses wherein the hot component (40, Fig 2) is a first hot component, wherein the CCA heat exchanger is a first CCA heat exchanger (Fig 3a, cooled air exits to HX 50 through a strut 58 and cools the rear rim 38 and the rear hub 40, where four HX 50 and struts 58 are shown, the top HX is interprete4d to be the first HX),
wherein the CCA system further comprises a third CCA heat exchanger (four HX 50 shown in Fig 3a),
wherein the cold side bleed assembly (56 Fig 3a) comprises a first portion and a third portion arranged in parallel flow (four portions in four ducts 42),
wherein the first portion is in thermal communication with the first CCA heat exchanger and the third portion is in thermal communication with the third CCA heat exchanger (each portion 42 is associated with each HX 50, annotated in Fig 3a), and
wherein the hot side bleed assembly is a first hot side bleed assembly and is in thermal communication with the first CCA heat exchanger (1ST hot side bleed associated with 1ST portion in Fig 3a),
wherein the CCA system further comprises a third hot side bleed assembly in thermal communication with the third CCA heat exchanger (four hot side bleed assemblies with four struts 58 annotated in Fig 3a) and a second hot component of the gas turbine engine separate from the first hot component (four struts 58 leading to four regions of the rear hub 40, interpreted to be the four separate hot components).
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Regarding claim 15
Glahn in view of Stretton discloses the gas turbine engine of claim 1.
Glahn further discloses wherein the inlet of the hot side bleed assembly (pipes 52, Figs 1-2) is in fluid communication with the working gas flowpath (flowpath from 16, 18, 20, 22, 24) at the compressor discharge cavity (60 Fig 2).
Regarding claim 20
Glahn in view of Stretton discloses the gas turbine engine of claim 1.
Glahn further discloses wherein the gas turbine engine is a turbofan engine (Fig 1, engine 10 is a turbofan engine).
Claim(s) 7-8 is/are rejected under 35 U.S.C. 103 as being unpatentable over Glahn in view of Stretton, as applied to claim 1 above, and further in view of Millhaem (US 20220235707 A1).
Regarding claim 7
Glahn in view of Stretton discloses the gas turbine engine of claim 1.
Glahn in view of Stretton is silent on wherein the turbomachine comprises an operability bleed assembly defining an operability bleed duct extending between an inlet in fluid communication with the working gas flowpath at a location between the low pressure compressor and the high pressure compressor and an outlet, wherein the inlet of the cold side bleed assembly is in fluid communication with the operability bleed duct.
However, Millhaem teaches a gas turbine engine having a cold side bleed assembly (cooling duct 62 having inlet 64 connected between LPC 14 and HPC 16, Figs 1-2, Para 0049), and
an operability bleed assembly (duct 46, Para 0036 bottom, 0047) defining an operability bleed duct (46) extending between an inlet (inlet of operability bleed duct 46 annotated in Fig 2) in fluid communication with the working gas flowpath (gas flowpath from LPC 14 to HPC 16 to combustor 18) at a location between the low pressure compressor (14) and the high pressure compressor (16) and an outlet (48),
wherein the inlet (64) of the cold side bleed assembly (62) is in fluid communication with the operability bleed duct (46).
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Therefore, it would have been obvious to one of ordinary skill in the art before the effective filing date to add an operability bleed assembly, taught by Millhaem, to the cold side bleed assembly in Glahn in view of Stretton, to define an operability bleed duct extending between an inlet in fluid communication with the working gas flowpath at a location between the low pressure compressor and the high pressure compressor and an outlet, wherein the inlet of the cold side bleed assembly is in fluid communication with the operability bleed duct, as suggested and taught by Millhaem, because the operability bleed assembly functions as a bleed vent for controlling low pressure compressor stall margin (Para 0036).
Regarding claim 8
Glahn in view of Stretton and Millhaem discloses the gas turbine engine of claim 7.
Glahn in view of Stretton and Millhaem further discloses wherein the operability bleed assembly (Millhaem teaches an operability bleed assembly 46, Fig 2, Para 0036 bottom) further comprises a variable bleed valve (42, Para 0036) in communication with the operability bleed duct (46) at a location upstream of the inlet (64) of the cold side bleed assembly (62).
Claim(s) 13-14 is/are rejected under 35 U.S.C. 103 as being unpatentable over Glahn in view of Stretton, as applied to claim 1 above, and further in view of Briley (US 20220381181 A1).
Regarding claim 13
Glahn in view of Stretton discloses the gas turbine engine of claim 1.
Glahn in view of Stretton is silent on wherein the CCA system further comprises a flow control valve in operable communication with the cold side bleed assembly.
However, Briley teaches a gas turbine engine having a CCA system (bleed system 100 with pre-cooler 126, Fig 2, Para 0032 top) including a flow control valve (valve 156 connected to a cold side bleed fan air 128, Para 0041 middle) in operable communication with the cold side bleed assembly (fan air 128).
Therefore, it would have been obvious to one of ordinary skill in the art before the effective filing date to add a flow control valve, taught by Briley, to the cold side bleed assembly in Glahn in view of Stretton, to automatically and precisely control the amount of air bleeding through the cold side.
Regarding claim 14
Glahn in view of Stretton discloses the gas turbine engine of claim 1.
Glahn in view of Stretton is silent on wherein the CCA system further comprises a flow control valve in operable communication with the hot side bleed assembly.
However, Briley teaches a gas turbine engine having a CCA system (bleed system 100 with pre-cooler 126, Fig 2, Para 0032 top) including a flow control valve (148, Para 0041 middle) in operable communication with the hot side bleed assembly (high pressure valve 148).
Therefore, it would have been obvious to one of ordinary skill in the art before the effective filing date to add a flow control valve, taught by Briley, to the hot side bleed assembly in Glahn in view of Stretton, to automatically and precisely control the amount of air bleeding through the hot side.
Claim(s) 16 is/are rejected under 35 U.S.C. 103 as being unpatentable over Glahn in view of Stretton, as applied to claim 1 above, and further in view of Kubota (US 20100154434 A1).
Regarding claim 16
Glahn in view of Stretton discloses the gas turbine engine of claim 1.
Glahn further discloses wherein the hot side bleed assembly (52 Fig 1) comprises a first portion (portion 52) in thermal communication with the CCA heat exchanger (50).
Glahn in view of Stretton is silent on wherein the hot side bleed assembly comprises a bypass portion bypassing the CCA heat exchanger.
However, Kubota teaches a gas turbine engine (Fig 1) wherein the hot side bleed assembly (compressed air that has been extracted from midstream of the compressor section 2 to a branch point 11, Para 0121 middle, 123 middle) comprises a first portion (1ST portion passes through HX 7, Para 0122) in thermal communication with the CCA heat exchanger (7) and a bypass portion (bypass portion 8, Para 0122 top) bypassing the CCA heat exchanger (7).
Therefore, it would have been obvious to one of ordinary skill in the art before the effective filing date to add to the hot side bleed assembly in Glahn in view of Stretton, a bypass portion bypassing the CCA heat exchanger, as suggested and taught by Kubota, because having a bypass allows for flexibility in controlling the temperature of the compressed air from the hot side bleed assembly, in the case where no cooling is required.
Claim(s) 17-18 is/are rejected under 35 U.S.C. 103 as being unpatentable over Glahn in view of Stretton, as applied to claim 1 above, and further in view of Armstrong (US 20230242264 A1).
Regarding claim 17
Glahn in view of Stretton discloses the gas turbine engine of claim 1.
Glahn further discloses that the gas turbine engine is a turbofan engine (Fig 1, engine 10 is a turbofan engine).
Glahn in view of Stretton is silent on wherein the gas turbine engine defines an overall pressure ratio greater than or equal to 50:1 and less than or equal to 70:1 when operated at a rated speed during standard day operating conditions.
However, Armstrong teaches a gas turbine engine being a turbofan (Fig 1) having an overall pressure ratio greater than or equal to 50:1 and less than or equal to 70:1 (overall pressure ratio may be in an inclusive range bounded by any two of the values, for example in the range of 50 to 65, Para 0076 middle) when operated at a rated speed during standard day operating conditions (overall pressure at cruise, Para 0076 middle, cruise is interpreted to be a rated speed during standard operating condition of the engine).
Therefore, it would have been obvious to one of ordinary skill in the art before the effective filing date to operate the engine in Glahn in view of Stretton, to have an overall pressure ratio greater than or equal to 50:1 and less than or equal to 70:1 when operated at a rated speed during standard day operating conditions, as suggested and taught by Armstrong, because this results in a balance between operability and engine efficiency of the aircraft (Abstract).
Regarding claim 18
Glahn in view of Stretton discloses the gas turbine engine of claim 1.
Glahn further discloses that the gas turbine engine is a turbofan engine (Fig 1, engine 10 is a turbofan engine) including a fan section (14) having a fan driven by the turbomachine (turbine 24).
Glahn in view of Stretton is silent on wherein the fan defines a fan pressure ratio less than or equal to 1.6 when the gas turbine engine is operated at a cruise condition.
However, Armstrong teaches a gas turbine engine being a turbofan (Fig 1) having a fan defines a fan pressure ratio less than or equal to 1.6 when the gas turbine engine is operated at a cruise condition (fan pressure ratio may be no greater than 1.5 at cruise conditions, Para 0077).
Therefore, it would have been obvious to one of ordinary skill in the art before the effective filing date to operate the engine in Glahn in view of Stretton, to have a fan pressure ratio less than or equal to 1.6 when the gas turbine engine is operated at a cruise condition, as suggested and taught by Armstrong, because this results in a balance between operability and engine efficiency of the aircraft (Abstract).
Response to Arguments
Applicant’s arguments with respect to claim(s) 1 have been considered but are moot because the new ground of rejection does not rely on any reference applied in the prior rejection of record for any teaching or matter specifically challenged in the argument.
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 Thuyhang Nguyen whose telephone number is (571)272-5317. The examiner can normally be reached Monday-Friday 8am-5pm EST.
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If attempts to reach the examiner by telephone are unsuccessful, the examiner’s supervisor, Edward F. Landrum can be reached at (571) 272-5567. The fax phone number for the organization where this application or proceeding is assigned is 571-273-8300.
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/Thuyhang N Nguyen/Examiner, Art Unit 3761
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