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 .
Drawings
The drawings were received on 9/23/2024. These drawings are not approved as they do not conform with the requirements of 37 CFR “1.84 Standards for drawings.”
“(l) Character of lines, numbers, and letters. All drawings must be made by a process which will give them satisfactory reproduction characteristics. Every line, number, and letter must be durable, clean, black (except for color drawings), sufficiently dense and dark, and uniformly thick and well-defined. The weight of all lines and letters must be heavy enough to permit adequate reproduction. This requirement applies to all lines however fine, to shading, and to lines representing cut surfaces in sectional views. Lines and strokes of different thicknesses may be used in the same drawing where different thicknesses have a different meaning.”
The drawings are objected to because the unlabeled rectangular box(es) shown in Figs. 10-12 should be provided with descriptive text labels, e.g. 801, 802, 803 in Fig. 10 etc.
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.
Specification
The disclosure is objected to because of the following informalities: pressure “in lb inches per second squared per inch squared” on page 105, lines 18-19 [excerpted below] does not make sense as a unit of pressure, which is understood to be force per area.
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.
Claims 1-15 are 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. Applicant specifies the unit core size on page 105 of the specification [see excerpt below]: but does not clearly indicate which each of: “s” “K” and “in” represent. For example, does K refer to Kelvin when T3 was already set as the temperature in Kelvin? Furthermore, it is not clear how applicant got from the core size equation on line 16 to the unit core size equation in line 22 as none of these variables are defined or otherwise spelled out. Accordingly, as the unit core size is referenced in claim 1, the claims fail to comply with the written description requirement. Moreover, pressure “in lb inches per second squared per inch squared” on lines 18-19 of the excerpt does not make sense as a unit of pressure, which is understood to be force per area.
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The following is a quotation of 35 U.S.C. 112(b):
(b) CONCLUSION.—The specification shall conclude with one or more claims particularly pointing out and distinctly claiming the subject matter which the inventor or a joint inventor regards as the invention.
Claims 1-15 are rejected under 35 U.S.C. 112(b) or 35 U.S.C. 112 (pre-AIA ), second paragraph, as being indefinite for failing to particularly point out and distinctly claim the subject matter which the inventor or a joint inventor (or for applications subject to pre-AIA 35 U.S.C. 112, the applicant), regards as the invention.
Claims 1, 15 claims the unit core size without defining it, and the specification does not clearly define it either, as set forth previously above. Claims 1, 15 [near the end]: “the fuel-flow nvPM emissions index ratio of the gas turbine engine is less than 0.08; and the gas turbine engine is configured to provide fuel comprising a sustainable aviation fuel (SAF) to the fuel spray nozzles.” However, the claim does not make it clear what relationship the SAF fuel has with the nvPM emissions, as it could be interpreted as the measured nvPM is for a standard fuel [e.g. JP-8] rather than the SAF fuel. The SAF could be utilized at any time as it has no defined relationship with the emissions.
Claim 11 “the, or each, ignitor” is unclear.
A broad range or limitation together with a narrow range or limitation that falls within the broad range or limitation (in the same claim) may be considered indefinite if the resulting claim does not clearly set forth the metes and bounds of the patent protection desired. See MPEP § 2173.05(c). In the present instance, claims 2-8, 12, 13 recites the broad recitation of a range, and the claim also recites a narrower “preferably …” statement of the range/limitation. The claim(s) are considered indefinite because there is a question or doubt as to whether the feature introduced by such narrower language is (a) merely exemplary of the remainder of the claim, and therefore not required, or (b) a required feature of the claims.
Claim Rejections - 35 USC § 102
The following is a quotation of the appropriate paragraphs of 35 U.S.C. 102 that form the basis for the rejections under this section made in this Office action:
A person shall be entitled to a patent unless –
(a)(1) the claimed invention was patented, described in a printed publication, or in public use, on sale, or otherwise available to the public before the effective filing date of the claimed invention.
Claim(s) 1-7, 13-15 is/are rejected under 35 U.S.C. 102(a)(1) as being anticipated by Anderson et al, "Alternative Aviation Fuel Experiment" as evidenced by Stickles et al “TAPS II Technology Final Report – Technology Assessment Open Report” and Aubuchon et al. “CFM56-3 Turbofan Engine Description”. Anderson et al teach the CFM56-2C1 gas turbine engines were utilized for the data [see page 9]. Stickles et al [the manufacturer] teach that the CFM56 engine uses the single annular combustor design [SAC,] with rich quench lean (RQL) burn combustion [see bottom of page 8 and particularly pages 10-11]. Note Stickles et al teach the double annular combustor [DAC] design did not enter service until the much later generations of CFM56-5B and –7B engines.
“After many years of intermittent development, the DAC entered service in the CFM56-5B and –7B engines in the mid-1990s” [excerpted from page 11].
Hence, all previous generation engines [including the 2C1, 3 versions] utilized the single annular combustor design [SAC] design with the RQL. As for the number of fuel spray nozzles, these are well known for the SAC design to be 20 fuel spray nozzles [20 swirl fuel nozzles, e.g. from specifications cited by Aubuchon et al].
Anderson et al teach: A gas turbine engine for an aircraft, comprising: a rich burn, quick quench, lean burn (RQL) combustor [inherent, as evidenced by Stickles] having a number of fuel spray nozzles in the range 14-22 [20 as the published number for the SAC combustor of the CFM56 engine, e.g. from Aubuchon et al] or a number of fuel spray nozzles per unit engine core size in the range 2 to 6; and wherein: a fuel-flow nvPM emissions index ratio is defined as: EIidle×Wf,idle/EImaxTO×Wf,maxTO where: EIidle is the nvPM emissions index in mg/kg of the gas turbine engine if operating at around 7% available thrust for given operating conditions; and EImaxTO is the nvPM emissions index in mg/kg of the gas turbine engine if operating at around 100% available thrust for the given operating conditions [see page 51, Fig. 28, left Fig. which shows the nvPM at idle and at 100% thrust/power – see also page 172, Fig. 5 in Appendix G from which the data was sourced]; Wf,idle is the rate of fuel flow to the fuel spray nozzles in kg/s at around 7% available thrust for the given operating conditions; and Wf,maxTO is the rate of fuel flow to the fuel spray nozzles in kg/s1 at around 100% available thrust for the given operating conditions [see page 194, Fig. 7 which teaches that fuel flow rate is linear based on power at idle to at 100% thrust/power]; the fuel-flow nvPM emissions index ratio of the gas turbine engine is less than 0.08; and the gas turbine engine is configured to provide fuel comprising a sustainable aviation fuel (SAF) to the fuel spray nozzles [page 172, Fig. 5 in Appendix G, shows different SAF fuels, i.e. everything but the JP-8]. Note the ratio of EIidle×Wf,idle/EImaxTO×Wf,maxTO is the same as EIidle /EImaxTO× Wf,idle/ Wf,maxTO and examples of these values on Figs. 5 (pg. 172) and Fig. 7 (pg. 194) include about 2E12 / 2E14 x 1000/8000 = 1/100 x 1/8 = 0.0125. (2) wherein the fuel-flow nvPM emissions index ratio is less than 0.0798 and preferably less than 0.0731 and more preferably less than 0.0665. (3) wherein the fuel-flow nvPM emissions index ratio is less than or equal to 0.06 and preferably less than or equal to 0.04 and more preferably less than or equal to 0.02. (4) wherein the fuel-flow nvPM emissions index ratio is less than or equal to 0.0119 and preferably less than or equal to 0.0109 and more preferably less than or equal to 0.00986. (5) wherein the fuel-flow nvPM emissions index ratio is greater than or equal to 0.000993 and preferably greater than or equal to 0.00111 and more preferably greater than or equal to 0.00124. (6) wherein the fuel-flow nvPM emissions index ratio is in the range of 0.000993 to 0.0119 and preferably in the range of 0.00111 to 0.0109 and even more preferably in the range of 0.00124 to 0.00986. (7) wherein: a) Wf,idle is in the range of 0.0516 to 0.119 kg/s and preferably in the range of 0.0581 to 0.109 kg/s and more preferably in the range of 0.0645 kg/s to 0.0990 kg/s [Fig. 7 (pg. 194): 800 pounds / hour converts to 0.1 kg/s]; and/or b) Wf,maxTO is in the range 0.441 to 1.23 kg/s and preferably in the range of 0.496 to 1.13 kg/s and more preferably in the range of 0.551 kg/s to 1.03 kg/s [Fig. 7 (pg. 194): 8000 pounds/hour is 1kg/s]. (13) wherein the fuel provided to the combustor comprises a %SAF in the range of 50% to 100%, preferably in the range 70% to 100%, and more preferably in the range 90% to 100% [up to 100% SAF tested, see page 10]. (14) A method of operating the gas turbine engine of claim 1, the method comprising providing fuel comprising a sustainable aviation fuel to the fuel spray nozzles [up to 100% SAF tested, see page 10]. (15) A method of operating a gas turbine engine, the gas turbine engine comprising: a rich burn, quick quench, lean burn (RQL) combustor having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit engine core size in the range 2 to 6; and wherein: a fuel-flow nvPM emissions index ratio is defined as: EIidle×Wf,idleEImaxTO×Wf,maxTOwhere: EIidle is the nvPM emissions index in mg/kg of the gas turbine engine if operating at around 7% available thrust for given operating conditions; and EImaxTO is the nvPM emissions index in mg/kg of the gas turbine engine if operating at around 100% available thrust for the given operating conditions; Wf,idle is the rate of fuel flow to the fuel spray nozzles in kg/s at around 7% available thrust for the given operating conditions; and Wf,maxTO is the rate of fuel flow to the fuel spray nozzles in kg/s at around 100% available thrust for the given operating conditions; the fuel-flow nvPM emissions index ratio of the gas turbine engine is less than 0.08; and the gas turbine engine is configured to provide fuel comprising a sustainable aviation fuel (SAF) to the fuel spray nozzles, wherein the method comprises providing fuel comprising a sustainable aviation fuel to the fuel spray nozzles. The method of claim 15 is treated analogously to claim 1.
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) 1-7, 12-15 is/are rejected under 35 U.S.C. 103 as being unpatentable over Anderson et al, "Alternative Aviation Fuel Experiment" (AAFEX), NASA, NASA/TM-2011-217059, (February 2011) in view of Stickles et al “TAPS II Technology Final Report and Aubuchon et al. “CFM56-3 Turbofan Engine Description”. Above, Anderson et al was cited as teaching the claimed ranges in a CFM56-2C1 gas turbine engines [see page 9]. Alternately, Stickles et al [the manufacturer of the engine] teach that the CFM56 engine uses the single annular combustor design [SAC,] with rich quench lean (RQL) burn combustion]. Note Stickles et al teach the double annular combustor design did not enter service until the much later generations of CFM56-5B and –7B engines.
“After many years of intermittent development, the DAC entered service in the CFM56-5B and –7B engines in the mid-1990s” [excerpted from page 11].
Hence, all previous generation engines [including the 2C1, 3 versions] utilized the single annular combustor design [SAC] design with the RQL. Furthermore, the number of fuel spray nozzles, these are well known for the SAC design to be 20 fuel spray nozzles [20 swirl fuel nozzles, e.g. from Aubuchon et al]. To the extent not already inherent, it would have been obvious to one of ordinary skill in the art to employ the single annular combustor design [SAC,] with rich quench lean (RQL) burn combustion, as taught by Stickles, as the conventional design used in the art by the manufacturer of the same gas turbine engine. It further would have been obvious to use a number of fuel spray nozzles in the range 14-22, as taught by Aubuchon et al, as the standard range utilized in the art for this engine/combustor design. It further would have been obvious to use the fuel flow rates of page 194 of Anderson to determine the claimed ratio of EIidle×Wf,idle/EImaxTO×Wf,maxTO, as an obvious matter of using the workable ranges in the art. Note the ratio of EIidle×Wf,idle/EImaxTO×Wf,maxTO is the same as EIidle /EImaxTO× Wf,idle/ Wf,maxTO and examples of these values on Figs. 5 (pg. 172) and Fig. 7 (pg. 194) include about 2E12 / 2E14 x 1000/8000 = 1/100 x 1/8 = 0.0125. The prior art do not specifically teach (12) wherein the number of fuel spray nozzles per unit engine core size is in the range 2.5 to 4.5 [nor the range 2 to 6 in claims 1, 16]; and more preferably in the range 3 to 4. As for the number of fuel spray nozzles per unit engine core size range, it does not appear that applicant has disclosed a parameter range [engine size] that is unknown in the art, but is rather claiming it in a specific way. Note that the unit engine core size, is an art effective variable. Applicant defines it in the following manner and each of these variables is known for each engine type, including those listed in the following excerpt from page 105 of the specification.
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As sizing the engine core, as well as the mass flow rates and pressures and temperatures from the core are typical variables used in the art, it would have been obvious to one of ordinary skill in the art to size the engine with the range of fuel nozzles per unit core size, as an obvious matter of using the workable ranges in the art.
Claim(s) 8-11 is/are rejected under 35 U.S.C. 103 as being unpatentable over Anderson et al "Alternative Aviation Fuel Experiment" as evidenced by {OR} in view of Stickles et al “TAPS II Technology Final Report and Aubuchon et al “CFM56-3 Turbofan Engine Description”, as applied above, and further in view of Hoke et al (2013/0125556). Anderson et al do not teach (8) wherein the fuel spray nozzles comprises one or more duplex nozzles and one or more single flow nozzles, and preferably the combustor comprises 10-14 duplex fuel spray nozzles and 4-8 single flow fuel spray nozzles; (9) wherein the duplex fuel spray nozzles are arranged in groups about the circumference of the combustor and optionally wherein the groups of duplex fuel spray nozzles comprise at least two groups arranged diametrically opposite each other; (10) wherein each group of duplex fuel spray nozzles comprises 2-8 nozzles; (11) wherein the combustor comprises one or more ignitors and the, or each, ignitor is arranged adjacent to one or more of the duplex fuel spray nozzles. Hoke et al teach a single annular combustor 66 which employs (8) wherein the fuel spray nozzles 86 comprises one or more duplex nozzles 86D [Fig. 4] and one or more single flow nozzles 86S, and preferably the combustor comprises 10-14 duplex fuel spray nozzles 86D and 4-8 single flow fuel spray 86S. (9) wherein the duplex fuel spray nozzles 86D are arranged in groups about the circumference of the combustor [Fig. 3, ¶ 0028] and optionally wherein the groups of duplex fuel spray nozzles 86D comprise at least two groups arranged diametrically opposite each other. (10) wherein each group of duplex fuel spray nozzles 86D comprises 2-8 nozzles [Fig. 3, ¶ 0028]. (11) wherein the combustor comprises one or more ignitors 124 and the, or each, ignitor 124 is arranged adjacent to one or more of the duplex fuel spray nozzles 86D.
Hoke et al teach using the duplex and single flow fuel nozzles allows for
“selectively forming a plurality of local circumferential zones with different fuel-air ratios within the combustor [¶ 0005]”
in a manner which controls combustor noise.
It would have been obvious to one of ordinary skill in the art to employ (8) wherein the fuel spray nozzles comprises one or more duplex nozzles and one or more single flow nozzles, and preferably the combustor comprises 10-14 duplex fuel spray nozzles and 4-8 single flow fuel spray nozzles; (9) wherein the duplex fuel spray nozzles are arranged in groups about the circumference of the combustor and optionally wherein the groups of duplex fuel spray nozzles comprise at least two groups arranged diametrically opposite each other; (10) wherein each group of duplex fuel spray nozzles comprises 2-8 nozzles; (11) wherein the combustor comprises one or more ignitors and the, or each, ignitor is arranged adjacent to one or more of the duplex fuel spray nozzles, as part of a single annular combustor design, as taught by Hoke et al, in order to “selectively form a plurality of local circumferential zones with different fuel-air ratios within the combustor” in a manner which controls combustor noise.
Claim(s) 1-15 is/are rejected under 35 U.S.C. 103 as being unpatentable over Burd et al (2007/0125093) in view of Anderson et al "Alternative Aviation Fuel Experiment" as evidenced by {OR} in view of Stickles et al “TAPS II Technology Final Report and Aubuchon et al “CFM56-3 Turbofan Engine Description”, as applied above, and for claims 8-11 further in view of Hoke et al (2013/0125556) .
Burd et al teach (1) A gas turbine engine for an aircraft, comprising: a rich burn, quick quench, lean burn (RQL) combustor 26 having a number of fuel spray nozzles [in 41] in the range 14-22 [see Fig. 3 for the number or ¶ 0031] or a number of fuel spray nozzles per unit engine core size in the range 2 to 6. Burd et al do not teach:
wherein: a fuel-flow nvPM emissions index ratio is defined as: EIidle×Wf,idle/EImaxTO×Wf,maxTO where: EIidle is the nvPM emissions index in mg/kg of the gas turbine engine if operating at around 7% available thrust for given operating conditions; and EImaxTO is the nvPM emissions index in mg/kg of the gas turbine engine if operating at around 100% available thrust for the given operating conditions; Wf,idle is the rate of fuel flow to the fuel spray nozzles in kg/s at around 7% available thrust for the given operating conditions; and Wf,maxTO is the rate of fuel flow to the fuel spray nozzles in kg/s at around 100% available thrust for the given operating conditions; the fuel-flow nvPM emissions index ratio of the gas turbine engine is less than 0.08; and the gas turbine engine is configured to provide fuel comprising a sustainable aviation fuel (SAF) to the fuel spray nozzles. (2) wherein the fuel-flow nvPM emissions index ratio is less than 0.0798 and preferably less than 0.0731 and more preferably less than 0.0665. (3) wherein the fuel-flow nvPM emissions index ratio is less than or equal to 0.06 and preferably less than or equal to 0.04 and more preferably less than or equal to 0.02. (4) wherein the fuel-flow nvPM emissions index ratio is less than or equal to 0.0119 and preferably less than or equal to 0.0109 and more preferably less than or equal to 0.00986. (5) wherein the fuel-flow nvPM emissions index ratio is greater than or equal to 0.000993 and preferably greater than or equal to 0.00111 and more preferably greater than or equal to 0.00124. (6) wherein the fuel-flow nvPM emissions index ratio is in the range of 0.000993 to 0.0119 and preferably in the range of 0.00111 to 0.0109 and even more preferably in the range of 0.00124 to 0.00986. (7) wherein: a) Wf,idle is in the range of 0.0516 to 0.119 kg/s and preferably in the range of 0.0581 to 0.109 kg/s and more preferably in the range of 0.0645 kg/s to 0.0990 kg/s; and/or b) Wf,maxTO is in the range 0.441 to 1.23 kg/s and preferably in the range of 0.496 to 1.13 kg/s and more preferably in the range of 0.551 kg/s to 1.03 kg/s. (13) wherein the fuel provided to the combustor comprises a %SAF in the range of 50% to 100%, preferably in the range 70% to 100%, and more preferably in the range 90% to 100%. (14) A method of operating the gas turbine engine of claim 1, the method comprising providing fuel comprising a sustainable aviation fuel to the fuel spray nozzles. (15) A method of operating a gas turbine engine, the gas turbine engine comprising: a rich burn, quick quench, lean burn (RQL) combustor having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit engine core size in the range 2 to 6; and wherein: a fuel-flow nvPM emissions index ratio is defined as: EIidle×Wf,idleEImaxTO×Wf,maxTOwhere: EIidle is the nvPM emissions index in mg/kg of the gas turbine engine if operating at around 7% available thrust for given operating conditions; and EImaxTO is the nvPM emissions index in mg/kg of the gas turbine engine if operating at around 100% available thrust for the given operating conditions; Wf,idle is the rate of fuel flow to the fuel spray nozzles in kg/s at around 7% available thrust for the given operating conditions; and Wf,maxTO is the rate of fuel flow to the fuel spray nozzles in kg/s at around 100% available thrust for the given operating conditions; the fuel-flow nvPM emissions index ratio of the gas turbine engine is less than 0.08; and the gas turbine engine is configured to provide fuel comprising a sustainable aviation fuel (SAF) to the fuel spray nozzles, wherein the method comprises providing fuel comprising a sustainable aviation fuel to the fuel spray nozzles.
Anderson et al teach using the sustainable aviation fuel (SAF) to the fuel spray nozzles and that doing so results in the claimed ranges for fuel-flow nvPM emissions index ratio in a comparable annular combustor. The full teachings of Anderson et al are set forth above and not repeated for conciseness. Anderson et al teach using the sustainable aviation fuel (SAF), including the nvPM emissions index ratio for idle and full power conditions, teaching the SAF fuel greatly reduces the nvPM emissions for idle conditions compared to full 100% available thrust. He also teaches the nvPM emissions are overall significantly reduced compared to using conventional jet engine fuels (JP8). Note also that nvPM emissions are well known to have health risks for those exposed on the ground. It would have been obvious to one of ordinary skill in the art to one of ordinary skill in the art to achieve the claimed ranges of the nvPM emissions by using SAF fuel, as taught by Anderson et al, in order to
1) lower the emissions, especially of nvPM during operation of the gas turbine engine at idle power compared to 100% available thrust, 2) to reduce the emissions compared to conventional fuels and 3) to reduce health risks of personnel exposed to nvPM.
Anderson et al further teach the claimed ranges of fuel flow Wf,idle and Wf,maxTO are in the typical range of operation. It would have been obvious to one of ordinary skill in the art to obvious to one of ordinary skill in the art to employ the claimed ranges of fuel flow, as an obvious matter of using the working ranges in the art – noting that the combustor types of Burd and Anderson et al are established as analogous Rich-Quench-Lean combustor and single annular in nature [see e.g. Stickles et al]. Burd et al do not teach (8) wherein the fuel spray nozzles comprises one or more duplex nozzles and one or more single flow nozzles, and preferably the combustor comprises 10-14 duplex fuel spray nozzles and 4-8 single flow fuel spray nozzles; (9) wherein the duplex fuel spray nozzles are arranged in groups about the circumference of the combustor and optionally wherein the groups of duplex fuel spray nozzles comprise at least two groups arranged diametrically opposite each other; (10) wherein each group of duplex fuel spray nozzles comprises 2-8 nozzles; (11) wherein the combustor comprises one or more ignitors and the, or each, ignitor is arranged adjacent to one or more of the duplex fuel spray nozzles. Hoke et al teach an analogous [to Burd et al] single annular combustor 66 which employs (8) wherein the fuel spray nozzles 86 comprises one or more duplex nozzles 86D [Fig. 4] and one or more single flow nozzles 86S, and preferably the combustor comprises 10-14 duplex fuel spray nozzles 86D and 4-8 single flow fuel spray 86S. (9) wherein the duplex fuel spray nozzles 86D are arranged in groups about the circumference of the combustor [Fig. 3, ¶ 0028] and optionally wherein the groups of duplex fuel spray nozzles 86D comprise at least two groups arranged diametrically opposite each other. (10) wherein each group of duplex fuel spray nozzles 86D comprises 2-8 nozzles [Fig. 3, ¶ 0028]. (11) wherein the combustor comprises one or more ignitors 124 and the, or each, ignitor 124 is arranged adjacent to one or more of the duplex fuel spray nozzles 86D.
Hoke et al teach using the duplex and single flow fuel nozzles allows for
“selectively forming a plurality of local circumferential zones with different fuel-air ratios within the combustor [¶ 0005]”
in a manner which controls the noise from the combustor.
It would have been obvious to one of ordinary skill in the art to employ (8) wherein the fuel spray nozzles comprises one or more duplex nozzles and one or more single flow nozzles, and preferably the combustor comprises 10-14 duplex fuel spray nozzles and 4-8 single flow fuel spray nozzles; (9) wherein the duplex fuel spray nozzles are arranged in groups about the circumference of the combustor and optionally wherein the groups of duplex fuel spray nozzles comprise at least two groups arranged diametrically opposite each other; (10) wherein each group of duplex fuel spray nozzles comprises 2-8 nozzles; (11) wherein the combustor comprises one or more ignitors and the, or each, ignitor is arranged adjacent to one or more of the duplex fuel spray nozzles, as part of a single annular combustor design, as taught by Hoke et al, in order to “selectively form a plurality of local circumferential zones with different fuel-air ratios within the combustor” in a manner which controls the noise from the combustor. The prior art do not specifically teach (12) wherein the number of fuel spray nozzles per unit engine core size is in the range 2.5 to 4.5 [nor the range 2 to 6 in claims 1, 16]; and more preferably in the range 3 to 4. As for the number of fuel spray nozzles per unit engine core size range, it does not appear that applicant has disclosed a parameter range [engine size] that is unknown in the art, but is rather claiming it in a specific way. Note that the unit engine core size is an art effective variable. Each of these variables is known for each engine type, including those listed in the excerpt above from page 105 of the specification. As sizing the engine core, as well as the mass flow rates and pressures and temperatures from the core are typical variables used in the art, it would have been obvious to one of ordinary skill in the art to size the engine with the range of fuel nozzles per unit core size, as an obvious matter of using the workable ranges in the art.
Prior Art
The following references: 20240209800, 20240210039, and 20240209801 have at least one inventor in common with the current application and could have been applied against the claims. These listed prior art may be applied in future actions should applicant not disqualify the art by the standards set below. Based upon the earlier effectively filed date of these references, these references constitute prior art under 35 U.S.C. 102(a)(2). This prior art might be overcome by: (1) a showing under 37 CFR 1.130(a) that the subject matter disclosed in the reference was obtained directly or indirectly from the inventor or a joint inventor of this application and is thus not prior art in accordance with 35 U.S.C. 102(b)(2)(A); (2) a showing under 37 CFR 1.130(b) of a prior public disclosure under 35 U.S.C. 102(b)(2)(B) if the same invention is not being claimed; or (3) a statement pursuant to 35 U.S.C. 102(b)(2)(C) establishing that, not later than the effective filing date of the claimed invention, the subject matter disclosed in the reference and the claimed invention were either owned by the same person or subject to an obligation of assignment to the same person or subject to a joint research agreement.
This is a continuation of applicant's earlier Application No. 18/892,678. All claims are identical to, patentably indistinct from, or have unity of invention with the invention claimed in the earlier application (that is, restriction (including lack of unity) would not be proper) and could have been finally rejected on the grounds and art of record in the next Office action if they had been entered in the earlier application. Accordingly, THIS ACTION IS MADE FINAL even though it is a first action in this case. See MPEP § 706.07(b). 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.
Contact Information
Any inquiry concerning this communication or earlier communications from the Examiner should be directed to TED KIM whose telephone number is 571-272-4829. The Examiner can be reached on regular business hours before 5:00 pm, Monday to Thursday and every other Friday.
The fax number for the organization where this application is assigned is 571-273-8300.
If attempts to reach the examiner by telephone are unsuccessful, the examiner’s supervisor, Devon Kramer, can be reached at 571-272-7118. Alternate inquiries to Technology Center 3700 can be made via 571-272-3700.
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/Ted Kim/
Telephone
571-272-4829
Primary Examiner
Fax
571-273-8300
September 9, 2026
1 Note that while the unit is kg/s, since the ratio divides fuel flows at idle by fuel flow at 100% thrust, the units do not matter since the units cancel. In other words using the conversion factor to convert fuel flows from e.g. lbs/hour to kg/s affects both the numerator and denominator equally and the conversion factor cancels from both the numerator and denominator. Accordingly, since only the ratio of these quantities is being used, it is irrelevant what units are used as long as they are the same since any conversion factors.
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