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 .
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.
Claim Interpretation
Per Applicant’s Remarks filed 04 October 2024, claims 17-20 claim a method of operating a gas turbine engine with certain capabilities, wherein the method steps require operating the engine at the specific ISA temperatures and power levels claimed. For example, claim 17 requires operating the engine at ISA+40°C and 85%MTO, then operating the engine at ISA+40°C and 65%MTO. Then claim 20, further requires operating the engine at ISA+10°C and 85%MTO to achieve the claimed ratio.
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.
Claims 1-9, 12 and 17-22 is/are rejected under 35 U.S.C. 103 as being unpatentable over Schwarz 11053815 (hereinafter Schwarz815) in view of Niergarth 11187156 (hereinafter Niergarth156), and Schmidt 9982565.
Regarding claim 1, Schwarz815 teaches a gas turbine engine (20) for an aircraft (Fig 1 depicts an aircraft turbofan engine) comprising:
- an engine core (incl.29, 30, 31) comprising a compressor (29), a combustor (30), a turbine (31), and a core shaft (34, 36) connecting the turbine to the compressor (Fig 1);
- a fan (28) comprising a plurality of fan blades (connected to 44) and arranged upstream of the engine core (Fig 1);
- turbomachinery bearings (64);
- a power gearbox (32) adapted to drive the fan at a lower rotation speed than the turbine (col.3 ll.38-65); and
- a heat management system (Fig 4) configured to provide lubrication and cooling to the gearbox and turbomachinery bearings (lubricate and cool 96, 132; col.5 ll.32-33, col.6 ll.44-45),
and comprising a pipe assembly (incl. 97, 133) adapted to provide a lubricant flow to the gearbox and turbomachinery bearings (lubricant from 88, 124),
wherein the pipe assembly comprises a first lubricant circuit (incl. 97 or 133) and a second lubricant circuit (including the other of 97, 133), the first lubricant circuit and the second lubricant circuit being separate (Fig 4),
at least one air-lubricant heat exchanger (126, 90) arranged in the first lubricant circuit and configured to dissipate a first amount of heat to a first heat sink (air in 104, 140),
and at least one fuel-lubricant heat exchanger (92, 128) arranged in the second lubricant circuit and configured to dissipate a second amount of heat to a second heat sink (fuel from 76),
wherein the first heat sink is air and the second heat sink is fuel (Fig 4; col.4 ll.41-49, col.5 ll.11-16, col.6 ll.15-27).
Schwarz815 further teaches both circuits comprising respective air-lubricant heat exchangers (126, 90) and fuel-lubricant heat exchanger (128, 92).
Schwarz815 also teaches modulation of air flow via valves in order to adjust the amount of heat rejected to the first air heat sink (col.8 ll.16-19), and fuel pumps (80, 78) that adjust fuel flow through the fuel-lubricant heat exchangers to injectors (82). Thus, Schwarz815 is capable of modulating heat rejection and relative proportions of heat rejection (for example a proportion of heat rejected to air relative to the total amount of heat rejected to both air and fuel, a.k.a.
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to air and fuel from the lubrication system that is cooling and lubricating the gearbox and bearings. Additionally, because Schwarz815’s engine is an aircraft engine (Fig 1; col.4 l.45), it operates at environment temperatures relative to ISA, and defines a maximum take-off speed (MTO) relative to which it may operate at percentages of speed thereof (e.g. 65% of 85% of MTO). Thus, Schwarz815 being capable of a first proportion of heat rejection defined as
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a second proportion of heat rejection defined as
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Furthermore, Niergarth156 teaches a gas turbine engine (10) for an aircraft (Fig 1 depicts an aircraft turbofan engine) comprising:
- an engine core (incl. 22, 24, 26, 28, 30, 34, 36) comprising a compressor (22, 24), a combustor (26), a turbine (28, 30), and a core shaft (34, 36) connecting the turbine to the compressor (Fig 1);
- a fan (14) comprising a plurality of fan blades (40) and arranged upstream of the engine core (Fig 1);
- turbomachinery bearings (col.7 l.45);
- a power gearbox (46) adapted to drive the fan at a lower rotation speed than the turbine (col.6 ll.61-63); and
- a heat management system (100 in Fig 2) comprising a pipe assembly (incl. 102) adapted to provide a lubricant flow (oil; col.10 ll.17-31) through various heat sources and heat sinks (106A, 106B, 108A, 108B),
at least one air-lubricant heat exchanger (108B) to dissipate a first amount of heat to a first heat sink (air; col.19 ll.5-9),
and at least one fuel-lubricant heat exchanger (108A) to dissipate a second amount of heat to a second heat sink (fuel from 130),
wherein the first heat sink is air and the second heat sink is fuel (Fig 2; col.19 ll.5-9).
Niergarth156 further teaches modulation of lubrication flow and fuel flow around the various heat sink heat exchangers (108A, 108B) via bypass lines/valves (164, 166, 168, 170, 172, 174, 186, 188) per Fig 2, in order to precisely control the amount of heat rejection to the first air heat sink and second fuel heat sink (col.13 ll.52-64, col.14 ll.14-49; col.15 ll.1-12; col.16 ll.9-44; col.19 ll.5-14) including relative proportions thereof (for example a proportion of heat rejected to air relative to the total amount of heat rejected to both air and fuel, a.k.a.
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). Additionally, because Neirgarth156’s engine is an aircraft engine (Fig 1; col.7 ll.39-42; col.8 ll.4-7), it operates at environment temperatures relative to ISA, and defines a maximum take-off speed (MTO) relative to which it may operate at percentages of speed thereof (e.g. 65% of 85% of MTO).
Thus, Niergarth156 being capable of a first proportion of heat rejection defined as
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a second proportion of heat rejection defined as
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It would have been obvious to one of ordinary skill in the art before the effective filing date of the claimed invention to modify Schwarz815 with the bypass lines and valves of Niergarth156 in order to further precisely control the amount of heat rejection to the first air heat sink and second fuel heat sink (Niergarth156; col.13 ll.52-64, col.14 ll.14-49; col.15 ll.1-12; col.16 ll.9-44; col.19 ll.5-14).
Schwarz815 in view of Niergarth156 as discussed so far, does not specifically teach the gas turbine engine being configured to provide, at an environment temperature of ISA +40°C, the first proportion from 0.55 to 0.70 and the second proportion from 0.85 to 1; and the two lubricant circuits being commonly connected to the tank.
However, Schmidt 9982565 teaches two separate lubrication circuits (3, 5) to cool and lubricate respective heat loads (transmission 1, bearing 1’), the two lubrication circuits commonly connected to a tank (9) for thermally equilibrizing the two fluids, thereby increasing thermal efficiency and effectiveness of the system (col.4 l.32 – col.5 l.22). That is, whenever the heat rejection of the heat loads are different and using different lubrication circuits, the cooling efficiency of the hotter circuit for the hotter thermal load can be improved by transferring heat from the hotter circuit to the cooler circuit, and the cooler circuit can absorb the heat from the hotter circuit because its thermal load is cooler.
It would have been obvious to one of ordinary skill in the art before the effective filing date of the claimed invention to modify the separate lubrication circuits as taught by Schwarz815 and Niergarth156 to use a commonly connected tank as taught by Schmidt in order to thermally equilibrize the two lubrication circuits to improve system thermal efficiency (Schmidt, col.4 l.32 – col.5 l.22).
Schwarz815 in view of Niergarth156 and Schmidt as discussed so far, does not specifically teach the gas turbine engine being configured to provide, at an environment temperature of ISA +40°C, the first proportion from 0.55 to 0.70 and the second proportion from 0.85 to 1.
However, MPEP2112.01 provides that where the claimed and prior art products are identical or substantially identical in structure or composition (as mapped and discussed above, the engine of Schwarz815 in view of Niergarth156 and Schmidt is substantially identical to both the claimed engine and the disclosed, elected engine embodiment in Applicant’s Figs 1-3 and 8), claimed properties or functions (i.e. the achievable first and second proportions of heat rejection to air/fuel being 0.55-0.70 and 0.85-1) are presumed to be present (in this case, the prior art engine being substantially identical to Applicant’s claimed engine reasonably possesses the same characteristics and performance as the claimed engine when operated under the same conditions, e.g. at 85%MTO and 65%MTO in ISA + 40°C) and a prima facie case of either anticipation or obviousness has been established. That is, when the prior art structure and the claimed structure appear to be substantially identical, it is reasonable to conclude that they are capable of performing the same way under the same conditions. In this case, there is no structure, control, or mechanism claimed or disclosed by Applicant that is different from the prior art to enable Applicant’s engine to achieve the claimed performance in a way that the prior art could not.
Regarding claims 2-9, Schwarz815 in view of Niergarth156, and Schmidt teaches all the limitations of the claimed invention as discussed above (incl. first and second proportions at ISA + 40°C for 85%MTO and 65%MTO, respectively). Schwarz815 being a turbofan engine for an aircraft also teaches operation at cruise conditions to define a cruise proportion of heat (Fig 1).
Schwarz104 in view of Niergarth156, and Schmidt does not teach the following specific performances:
cl.2: the first proportion of heat generated by the gearbox and the turbomachinery and dissipated to air at 85% of a core shaft maximum take-off speed at an environment temperature of ISA +40°C is in from 0.60 to 0.70;
cl.3: the second proportion of heat generated by the gearbox and the turbomachinery and dissipated to air at 65% of the core shaft maximum take-off (MTO) speed at an environment temperature of ISA +40°C is from 0.90 to 1;
cl.4: the heat management system is configured to provide, at an environment temperature of ISA +10°C, first proportion from 0.40 to 0.60;
cl.5: the heat management system is configured to provide, at an environment temperature of ISA +10°C, the second proportion from 0.80 to 0.92;
cl.6: the heat management system is configured to provide, at an environment temperature of ISA -69°C, the first proportion from 0.20 to 0.40;
cl.7: the heat management system is configured to provide, at an environment temperature of ISA -69°C, the second proportion from 0.40 to 0.75;
cl.8: the heat management system is configured to provide, at an environment temperature of ISA +10°C, a ratio of the first proportion to the second proportion from 0.45 to 0.65.
cl.9: the heat management system is configured to provide the first amount of heat and the second amount of heat such that at the cruise conditions the cruise proportion of heat generated by the gearbox and the turbomachinery and dissipated to air is from 0.56 to 0.75.
Regarding the specific performance ranges, as discussed above, MPEP2112.01 provides that where the claimed and prior art products are identical or substantially identical in structure or composition (as mapped and discussed above, the engine of Schwarz815 in view of Niergarth156 and Schmidt is substantially identical to both the claimed engine and the disclosed, elected engine embodiment in Applicant’s Figs 1-3 and 8), claimed properties or functions (i.e. the achievable first, second, and cruise proportions of heat rejection to air/fuel being in the ranges: 0.60-0.70, 0.90-1, 0.40-0.60, 0.80-0.92, 0.20-0.40, 0.40-0.75, 0.45-0.65, and 0.56-0.75) are presumed to be present (in this case, the prior art engine being substantially identical to Applicant’s claimed engine reasonably possesses the same characteristics and performance as the claimed engine when operated under the same conditions e.g. at 85%MTO, 65%MTO, and cruise in ISA - 69°C, -10°C, and + 40°C) and a prima facie case of either anticipation or obviousness has been established. That is, when the prior art structure and the claimed structure appear to be substantially identical, it is reasonable to conclude that they are capable of performing the same way under the same conditions. In this case, there is no structure, control, or mechanism claimed or disclosed by Applicant that is different from the prior art to enable Applicant’s engine to achieve the claimed performance in a way that the prior art could not.
Regarding claim 12, Schwarz815 in view of Niergarth156 and Schmidt teaches all the limitations of the claimed invention as discussed above. Schwarz815 further teaches the first heat sink is bypass air (Figs 1 and 4; col.5 ll.59-60), and the at least one air-lubricant heat exchanger is adapted to receive bypass air from a bypass duct (68; col.5 ll.59-60).
Regarding claim 21, Schwarz815 in view of Niergarth156 and Schmidt teaches all the limitations of the claimed invention as discussed above. Schwarz815 further teaches the second lubricant circuit provides lubrication and cooling to the gearbox (96, 32; Fig 4) and the first lubricant circuit provides lubrication and cooling to the turbomachinery bearings (132, 64; Fig 4).
Regarding claim 22, Schwarz815 in view of Niergarth156 and Schmidt teaches all the limitations of the claimed invention as discussed above. Schwarz815 further teaches at least two air-lubricant heat exchangers (90, 126) to dissipate the first amount of heat to the first heat sink (air, col.5 ll.59-60, col.6 l.67 – col.7 l.1; Fig 4), of which at least one is arranged in the first lubricant circuit and at least one is arranged in the second lubricant circuit (Fig 4).
Regarding claims 17-19, Schwarz815 teaches a method of operating a gas turbine engine (20) for an aircraft (Fig 1 depicts an aircraft turbofan engine), the method comprising providing the gas turbine engine comprising:
- an engine core (incl.29, 30, 31) comprising a compressor (29), a combustor (30), a turbine (31), and a core shaft (34, 36) connecting the turbine to the compressor (Fig 1);
- a fan (28) comprising a plurality of fan blades (connected to 44) and arranged upstream of the engine core (Fig 1);
- turbomachinery bearings (64);
- a power gearbox (32) adapted to drive the fan at a lower rotation speed than the turbine (col.3 ll.38-65); and
- a heat management system (Fig 4) configured to provide lubrication and cooling to the gearbox and turbomachinery bearings (lubricate and cool 64, 32),
and comprising a pipe assembly (incl. 97, 133) adapted to provide a lubricant flow to the gearbox and turbomachinery bearings (Fig 4),
wherein the pipe assembly comprises a first lubricant circuit (incl. one of 97, 133) and a second lubricant circuit (incl. other one of 97, 133), the first lubricant circuit and the second lubricant circuit being separate (Fig 4),
at least one air-lubricant heat exchanger (126, 90) arranged in the first lubricant circuit and configured to dissipate a first amount of heat to a first heat sink (air in 104, 140),
and at least one fuel-lubricant heat exchanger (192, 128) arranged in the second lubricant circuit and configured to dissipate a second amount of heat to a second heat sink (fuel from 76),
wherein the first heat sink is air and the second heat sink is fuel (Fig 4; col.4 ll.41-49, col.5 ll.11-16, col.6 ll.15-27), and
wherein the method further comprises operating the heat management system to provide the first amount of heat and the second amount of heat (via heat exchangers 90, 92, 126, 128). That is, Schwarz815 further teaches both circuits comprising respective air-lubricant heat exchangers (126, 90) and fuel-lubricant heat exchanger (128, 92).
Schwarz815 further teaches modulation of air flow via valves in order to adjust the amount of heat rejected to the first air heat sink (col.8 ll.16-19), and fuel pumps (78, 80) that adjust fuel flow through the fuel-lubricant heat exchangers to injectors (82). Thus, Schwarz815 is capable of modulating heat rejection and relative proportions of heat rejection (for example a proportion of heat rejected to air relative to the total amount of heat rejected to both air and fuel, a.k.a.
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Thus, Schwarz815 being capable of a first proportion of heat rejection defined as
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a second proportion of heat rejection defined as
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Furthermore, Niergarth156 teaches a method of operating a gas turbine engine (10) for an aircraft (Fig 1 depicts an aircraft turbofan engine), the method comprising providing the gas turbine engine comprising:
- an engine core (incl. 22, 24, 26, 28, 30, 34, 36) comprising a compressor (22, 24), a combustor (26), a turbine (28, 30), and a core shaft (34, 36) connecting the turbine to the compressor (Fig 1);
- a fan (14) comprising a plurality of fan blades (40) and arranged upstream of the engine core (Fig 1);
- turbomachinery bearings (col.7 l.45);
- a power gearbox (46) adapted to drive the fan at a lower rotation speed than the turbine (col.6 ll.61-63); and
- a heat management system (100 in Fig 2) comprising a pipe assembly (incl. 102) adapted to provide a lubricant flow (oil; col.10 ll.17-31) through various heat sources and heat sinks (106A, 106B, 108A, 108B),
at least one air-lubricant heat exchanger (108B) to dissipate a first amount of heat to a first heat sink (air; col.19 ll.5-9),
and at least one fuel-lubricant heat exchanger (108A) to dissipate a second amount of heat to a second heat sink (fuel from 130),
wherein the first heat sink is air and the second heat sink is fuel (Fig 2; col.19 ll.5-9);
and wherein the method further comprises operating the heat management system to provide the first amount of heat and the second amount of heat (via heat exchangers 108B, 108A).
Niergarth156 further teaches modulation of lubrication flow and fuel flow around the various heat sink heat exchangers (108A, 108B) via bypass lines/valves (164, 166, 168, 170, 172, 174, 186, 188) per Fig 2, in order to precisely control the amount of heat rejection to the first air heat sink and second fuel heat sink (col.13 ll.52-64, col.14 ll.14-49; col.15 ll.1-12; col.16 ll.9-44; col.19 ll.5-14) including relative proportions thereof (for example a proportion of heat rejected to air relative to the total amount of heat rejected to both air and fuel, a.k.a.
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Thus, Niergarth156 being capable of a first proportion of heat rejection defined as
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It would have been obvious to one of ordinary skill in the art before the effective filing date of the claimed invention to modify Schwarz815 with the bypass lines and valves of Niergarth156 in order to further precisely control the amount of heat rejection to the first air heat sink and second fuel heat sink (Niergarth156; col.13 ll.52-64, col.14 ll.14-49; col.15 ll.1-12; col.16 ll.9-44; col.19 ll.5-14).
Schwarz815 in view of Niergarth156 as discussed so far, does not specifically teach the method further comprises operating the engine at an environment temperature of ISA +40°C and at 85% of the core shaft maximum take-off speed, wherein the first proportion is from (claim 17) 0.55 to 0.70 or (claim 18) from 0.60 to 0.70, and operating the engine at the environment temperature of ISA +40°C and at 65% of the core shaft maximum take-off speed, wherein the second proportion is from (claim 17) 0.85 to 1 or (claim 19) from 0.90 to 1; and the two lubricant circuits being commonly connected to the tank.
However, Schmidt 9982565 teaches two separate lubrication circuits (3, 5) to cool and lubricate respective heat loads (transmission 1, bearing 1’), the two lubrication circuits commonly connected to a tank (9) for thermally equilibrizing the two fluids, thereby increasing thermal efficiency and effectiveness of the system (col.4 l.32 – col.5 l.22). That is, whenever the heat rejection of the heat loads are different and using different lubrication circuits, the cooling efficiency of the hotter circuit for the hotter thermal load can be improved by transferring heat from the hotter circuit to the cooler circuit, and the cooler circuit can absorb the heat from the hotter circuit because its thermal load is cooler.
It would have been obvious to one of ordinary skill in the art before the effective filing date of the claimed invention to modify the separate lubrication circuits as taught by Schwarz815 and Niergarth156 to use a commonly connected tank as taught by Schmidt in order to thermally equilibrize the two lubrication circuits to improve system thermal efficiency (Schmidt, col.4 l.32 – col.5 l.22).
Schwarz815 in view of Niergarth156 and Schmidt as discussed so far, does not specifically teach the method further comprises operating the engine at an environment temperature of ISA +40°C and at 85% of the core shaft maximum take-off speed, wherein the first proportion is from (claim 17) 0.55 to 0.70 or (claim 18) from 0.60 to 0.70, and operating the engine at the environment temperature of ISA +40°C and at 65% of the core shaft maximum take-off speed, wherein the second proportion is from (claim 17) 0.85 to 1 or (claim 19) from 0.90 to 1
However, MPEP2112.02(I) provides that if a prior art device (Schwarz815 in view of Niergarth156 and Schmidt), in its normal and usual operation (for flight, which requires takeoff, landing, cruise, etc. over a standard flight envelope under ambient conditions), would necessarily perform the method claimed (i.e. operating at 65%MTO and 85%MTO), then the method claimed will be considered to be anticipated by the prior art device. Note, that operation of an aircraft engine at ambient temperature up to ISA + 40°C was considered part of a normal/standard range of operating conditions as evidence by Applicant’s Admission that “engines are normally certified to operate in a range of environment temperatures between ISA (International Standard Atmosphere) -69°C and ISA +40°C” (on p.6, para.3 of Applicant’s Specification filed 15 February 2023) and typical Type-Certificate Data Sheets for engines registered with the EASA listing under “Climatic Operating Envelope” and engine usage “in ambient temperatures up to ISA +40°C” (e.g.: No. EASA E.047 for the RB211 Trent 800 series engines, dated 21 February 2015, Issue 04, pp.10-11).. MPEP2112.02(I) also provides that when the prior art device is the same as a device described in the specification for carrying out the claimed method, it can be assumed the device will inherently perform the claimed process. That is, when the prior art structure and the claimed structure appear to be substantially identical, it is reasonable to conclude that they perform the same way (i.e. the achievable first and second proportions of heat rejection to air/fuel being 0.55-0.70, 0.6-0.7, 0.85-1, and 0.90-1) under the same conditions (e.g. at 85%MTO and 65%MTO in ISA + 40°C). In this case, the engine of Schwarz815 in view of Niergarth156 and Schmidt is substantially identical to both the claimed engine and the disclosed, elected engine embodiment in Applicant’s Figs 1-3 and 8); and, based on the current record, there is no structure, control, or mechanism claimed or disclosed by Applicant that is different from the prior art to enable Applicant’s engine to achieve the claimed performance in a way that the prior art would not.
Regarding claim 20, Schwarz815 in view of Niergarth156 and Schmidt teaches all the limitations of the claimed invention as discussed above. Schwarz815 in view of Niergarth156 and Schmidt as discussed so far, does not specifically teach operating the heat management system at 85% of the core shaft maximum take-off speed at an environment temperature of ISA +10°C, wherein a ratio of the first proportion at 85% of the core shaft maximum take-off speed at the environment temperature of ISA + 40°C to the first proportion at 85% of the core shaft maximum take-off speed at the environment temperature of ISA +10°C is from 1.20 to 1.42.
However, MPEP2112.02(I) provides that if a prior art device (Schwarz815 in view of Niergarth156 and Schmidt), in its normal and usual operation (for flight, which requires takeoff, landing, cruise, etc. over a standard flight envelope under ambient conditions), would necessarily perform the method claimed (i.e. operating at 85%MTO), then the method claimed will be considered to be anticipated by the prior art device. Note, that operation of an aircraft engine at ambient temperature up to ISA + 40°C (which also includes operation at ISA + 10°C) was considered part of a normal/standard range of operating conditions as evidence by Applicant’s Admission that “engines are normally certified to operate in a range of environment temperatures between ISA (International Standard Atmosphere) -69°C and ISA +40°C” (on p.6, para.3 of Applicant’s Specification filed 15 February 2023) and typical Type-Certificate Data Sheets for engines registered with the EASA listing under “Climatic Operating Envelope” and engine usage “in ambient temperatures up to ISA +40°C” including “maximum continuous” operation at “ISA + 10°C” (e.g.: No. EASA E.047 for the RB211 Trent 800 series engines, dated 21 February 2015, Issue 04, pp.10-11). MPEP2112.02(I) also provides that when the prior art device is the same as a device described in the specification for carrying out the claimed method, it can be assumed the device will inherently perform the claimed process. That is, when the prior art structure and the claimed structure appear to be substantially identical, it is reasonable to conclude that they perform the same way (i.e. the ratio of first proportions being in the range of 1.20-1.42) under the same conditions (e.g. at 85%MTO in ISA + 40°C and ISA + 10°C). In this case, the engine of Schwarz815 in view of Niergarth156 and Schmidt is substantially identical to both the claimed engine and the disclosed, elected engine embodiment in Applicant’s Figs 1-3 and 8); and, based on the current record, there is no structure, control, or mechanism claimed or disclosed by Applicant that is different from the prior art to enable Applicant’s engine to achieve the claimed performance in a way that the prior art would not.
Claims 10-11 is/are rejected under 35 U.S.C. 103 as being unpatentable over Schwarz815 in view of Niergarth156 and Schmidt, and further in view of Gaskell 20210190008.
Regarding claim 10, Schwarz815 in view of Niergarth156, and Schmidt teaches all the limitations of the claimed invention as discussed above. Schwarz815 in view of Niergarth156 and Schmidt as discussed so far, does not teach the core shaft maximum take-off speed is from 5500 rpm to 9500 rpm.
However, Gaskell teaches a gas turbine engine (Fig 1) for an aircraft (Abstract) comprising:
- an engine core (11) comprising a compressor (14, 15), a combustor (16), a turbine (17, 19), and a core shaft (26, 27) connecting the turbine to the compressor (Fig 1),
wherein the core shaft has a core shaft maximum take-off speed in the range of from 5500rpm to 9500rpm ([0023]; 5800-6200rpm falling in the claimed range);
- a fan (23) comprising a plurality of fan blades and arranged upstream of the engine core (Fig 1);
- turbomachinery bearings (Fig 1 below; and/or 26a, 26b, 26c in Fig 3); and
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- a power gearbox (30) adapted to drive the fan at a lower rotation speed than the turbine ([0010]).
Gaskell further teaches at least the bearings require lubrication ([0360, 367-368]); and the turbofan engine is operable in a speed range of 85% and 65% of the maximum take-off speed (engine operable between 1500-6200rpm, thus including 85% and 65% of the MTO speed of range 5800-6200rpm, the 85% speed being ≈ 4930-5270rpm and the 65% speed being ≈ 3770-4030rpm; [0021,23]). Note, Gaskell also teaches cruise speeds as 5400-5700rpm (or ≈ 87-92%MTO) and ground idle speed as 1500rpm (≈ 24-26%MTO).
Gaskell is silent as to the details of any lubrication and/or thermal management system for the turbofan engine.
It would have been obvious to one of ordinary skill in the art before the effective filing date of the claimed invention to modify the geared turbofan engine (comprising a power gearbox and turbomachinery bearings) of Gaskell to use the lubrication and thermal management system of Schwarz815 in view of Niergarth156 and Schmidt, in order to: provide improved lubrication and cooling to the multiple turbine engine components (power gearbox and turbomachinery bearings) of the turbofan (Schwarz815, col.1 ll.11-27); and precisely control the amount of heat rejection to the first air heat sink and second fuel heat sink (Niergarth156; col.13 ll.52-64, col.14 ll.14-49; col.15 ll.1-12; col.16 ll.9-44; col.19 ll.5-14).
Regarding claim 11, Schwarz815 in view of Niergarth156 and Schmidt teaches all the limitations of the claimed invention as discussed above. Schwarz815 in view of Niergarth156 and Schmidt as discussed so far, does not teach the fan has a fan rotational speed at MTO conditions from 1500 rpm to 2800 rpm.
However, Gaskell teaches a gas turbine engine (Fig 1) for an aircraft (Abstract) comprising:
- an engine core (11) comprising a compressor (14, 15), a combustor (16), a turbine (17, 19), and a core shaft (26, 27) connecting the turbine to the compressor (Fig 1),
wherein the core shaft has a core shaft maximum take-off speed in the range of from 5500rpm to 9500rpm ([0023]; 5800-6200rpm falling in the claimed range);
- a fan (23) comprising a plurality of fan blades and arranged upstream of the engine core (Fig 1),
wherein the fan has a fan rotational speed at MTO conditions in the range of from 1500 rpm to 2800 rpm (the gear ratio being greater than 3, to drive the fan, thus teaching the fan rpm at MTO of
5800
3
=
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=
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rpm; [0023] teaching core shaft MTO speed of 5800-6200rpm);
- turbomachinery bearings (Fig 1 below; and/or 26a, 26b, 26c in Fig 3); and
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- a power gearbox (30) adapted to drive the fan at a lower rotation speed than the turbine ([0010]).
Gaskell further teaches at least the bearings require lubrication ([0360, 367-368]); and the turbofan engine is operable in a speed range of 85% and 65% of the maximum take-off speed (engine operable between 1500-6200rpm, thus including 85% and 65% of the MTO speed of range 5800-6200rpm, the 85% speed being ≈ 4930-5270rpm and the 65% speed being ≈ 3770-4030rpm; [0021,23]). Note, Gaskell also teaches cruise speeds as 5400-5700rpm (or ≈ 87-92%MTO) and ground idle speed as 1500rpm (≈ 24-26%MTO).
Gaskell is silent as to the details of any lubrication and/or thermal management system for the turbofan engine.
It would have been obvious to one of ordinary skill in the art before the effective filing date of the claimed invention to modify the geared turbofan engine (comprising a power gearbox and turbomachinery bearings) of Gaskell to use the lubrication and thermal management system of Schwarz815 in view of Niergarth156 and Schmidt, in order to: provide improved lubrication and cooling to the multiple turbine engine components (power gearbox and turbomachinery bearings) of the turbofan (Schwarz815, col.1 ll.11-27); and precisely control the amount of heat rejection to the first air heat sink and second fuel heat sink (Niergarth156; col.13 ll.52-64, col.14 ll.14-49; col.15 ll.1-12; col.16 ll.9-44; col.19 ll.5-14).
Claim 13 is/are rejected under 35 U.S.C. 103 as being unpatentable over Schwarz815 in view of Niergarth156 and Schmidt, and further in view of Schwarz 20080028763 (Schwarz2008).
Regarding claim 13, Schwarz815 in view of Niergarth156 and Schmidt teaches all the limitations of the claimed invention as discussed above. Schwarz815 further teaches the heat management system includes a valve arranged in the air duct (104, 140) of the at least one air-lubricant heat exchanger adapted to vary a mass flow rate of the bypass air across the at least one air-lubricant heat exchanger, thereby varying the first amount of heat (col.5 ll.59-60, col.8 ll.16-19).
Schwarz815 in view of Niergarth156 and Schmidt does not teach the valve being downstream of the at least one air-lubricant heat exchanger.
However, Schwarz2008 teaches an air-oil heat exchanger (44; Figs 2-3) using bypass air (F) of a turbofan engine (Fig 1A) installed in a pylon of the turbofan (36), wherein flaps (52) that are used to modulate the air flow (and hence the heat transfer) through the heat exchanger (44) are placed downstream of the heat exchanger in order to simultaneously provide thrust recovery (Fig 3; [0008}).
It would have been obvious to one of ordinary skill in the art before the effective filing date of the claimed invention to modify the turbofan of Schwarz815 in view of Niergarth156 and Schmidt to use the pylon-integrated heat exchanger and valve of Schwarz2008, in order to simultaneously provide thrust recovery (Schwarz2008, [0008)).
Claim 16 is/are rejected under 35 U.S.C. 103 as being unpatentable over Schwarz815 in view of Niergarth156 and Schmidt, and further in view of Pilatis 20050097889 and Leese 9051056.
Regarding claim 16, Schwarz815 in view of Niergarth156 and Schmidt teaches all the limitations of the claimed invention as discussed above. Schwarz104 further teaches the combustor comprising a plurality of fuel spray nozzles (144).
Schwarz815 in view of Niergarth156 and Schmidt does not teach the combustor is a lean burn combustor, thus the plurality fuel spray nozzles being lean burn fuel spray nozzles; and each fuel spray nozzle comprising a pilot fuel injector and a main fuel injector.
However, Pilatis teaches a combustor (11) of a turbofan (Fig 1) that is a lean combustor (Fig 2; [0014, 22]) comprising a plurality of lean burn fuel spray nozzles (60 in Fig 2; [0021]), each fuel spray nozzle comprising a pilot fuel injector a main fuel injector ([0022-23]).
And Leese teaches adding a fuel-lubricant heat exchanger bypass line/valve (30, 29) and a fuel-air heat exchanger (25) to modify fuel/lubricant systems of turbofan engines (compare Figs 1 and 4) to accommodate the stricter fuel temperature requirements of lean burn combustors (col.2 ll.16-22; col.4 ll.42-64; col.5 l.65 – col.6 l.20).
It would have been obvious to one of ordinary skill in the art before the effective filing date of the claimed invention to modify the turbofan of Schwarz815 in view of Niergarth156 and Schmidt to accommodate a lean burn combustor as taught by Pilatis, in order to reduce NOx emissions (Pilatis, [0002]) and because Leese teaches methods of modifying the lubrication and fuel thermal systems to facilitate the lean burn combustion (Leese, col.2 ll.16-22; col.4 ll.42-64; col.5 l.65 – col.6 l.20).
Claims 17-20 is/are additionally rejected under 35 U.S.C. 103 as being unpatentable over Schwarz815 in view of Niergarth156 and Schmidt, and further in view of Gaskell, Appukuttan 20160281603 and Bosak 20210172375.
Regarding claims 17-20, Schwarz815 in view of Niergarth156 and Schmidt teaches all the limitations of the claimed invention as discussed above.
In addition, Gaskell teaches a method of operating a gas turbine engine (Fig 1) for an aircraft (Abstract),
the method comprising providing a gas turbine engine (Fig 1) comprising:
- an engine core (11) comprising a compressor (14, 15), a combustor (16), a turbine (17, 19), and a core shaft (26, 27) connecting the turbine to the compressor (Fig 1),
wherein the core shaft has a core shaft maximum take-off speed in the range of from 5500rpm to 9500rpm ([0023]; 5800-6200rpm falling in the claimed range);
- a fan (23) comprising a plurality of fan blades and arranged upstream of the engine core (Fig 1);
- turbomachinery bearings (Fig 1 below); and
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- a power gearbox (30) adapted to drive the fan at a lower rotation speed than the turbine ([0010]).
Gaskell further teaches at least the bearings require lubrication ([0360, 367-368]); and the turbofan engine is operable in a speed range of 85% and 65% of the maximum take-off speed (engine operable between 1500-6200rpm, thus including 85% and 65% of the MTO speed of range 5800-6200rpm, the 85% speed being ≈ 4930-5270rpm and the 65% speed being ≈ 3770-4030rpm; [0021,23]). Note, Gaskell also teaches cruise speeds as 5400-5700rpm (or ≈ 87-92%MTO) and ground idle speed as 1500rpm (≈ 24-26%MTO).
Gaskell is silent as to the details of any lubrication and/or thermal management system for the turbofan engine.
It would have been obvious to one of ordinary skill in the art before the effective filing date of the claimed invention to modify the geared turbofan engine (comprising a power gearbox and turbomachinery bearings) of Gaskell to use the lubrication and thermal management system of Schwarz815 in view of Niergarth156 and Schmidt, in order to: provide improved lubrication and cooling to the multiple turbine engine components (power gearbox and turbomachinery bearings) of the turbofan (Schwarz815, col.1 ll.11-27); and precisely control the amount of heat rejection to the first air heat sink and second fuel heat sink (Niergarth156; col.13 ll.52-64, col.14 ll.14-49; col.15 ll.1-12; col.16 ll.9-44; col.19 ll.5-14).
Furthermore, Appukuttan teaches a gas turbine engine (10) for an aircraft comprising:
- an engine core (incl. 14, 16, 18, 20, 22, 24) comprising a compressor (14, 16), a combustor (18), a turbine (22, 24), and a core shaft (28, 30) connecting the turbine to the compressor (Fig 1), wherein the core shaft has a core shaft maximum take-off speed (as is defined for every engine);
- a fan (12) comprising a plurality of fan blades (Fig 1) and arranged upstream of the engine core (Fig 1);
- turbomachinery bearings ([0027]);
- a heat management system (Figs 2-3) configured to provide lubrication and cooling to the turbomachinery bearings and other heat loads ([0027]), and
comprising a pipe assembly (incl. 1322, 116, 124) adapted to provide a lubricant flow (132) to the turbomachinery bearings ([0027]; Figs 2-3),
at least one air- lubricant heat exchanger (122) to dissipate a first amount of heat to a first heat sink (air; [0025]),
at least one fuel-lubricant heat exchanger (108) to dissipate a second amount of heat to a second heat sink (fuel, [0024]),
wherein the first heat sink is air and the second heat sink is fuel ([0024-25]),
wherein the heat management system is configured to provide the first amount of heat and the second amount of heat such that a first proportion of heat generated by the turbomachinery and other heat loads dissipated to air defined as
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at 85% of the core shaft maximum take-off speed (turbofan engines for aircraft operate between 0-100% of MTO over flight envelope) is in the range of from 0.55 to 0.70 or 0.60 to 0.70 (this is a desired-result/intended-use of the system; the turbofan engine having all the claimed structures is capable of achieving the claimed desired-result/intended-use of the system; e.g. by using valve 126; see also MPEP2112.01(I)); and
a second proportion of heat generated by the gearbox and the turbomachinery and dissipated to air is defined as
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Appukuttan further teaches that at lower operating power levels (i.e. cruise and below, including descent and ground idle; Fig 4), the fuel heat capacity is below the oil heat rejection requirement (Fig 4) and the FOHE is insufficient to provide the required cooling, such that the AOHE must provide the remaining proportion of cooling/heat-rejection to the lubricant ([0005-6]). According to Fig 4, the fraction of heat rejected to the AOHE may be at least
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-
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It would have been obvious to one of ordinary skill in the art before the effective filing date of the claimed invention to modify the method of Schwarz815 in view of Niergarth156, Schmidt, and Gaskell (in which Gaskell teaches 85%MTO is slightly less than cruise) to use a proportion of ≥ 0.25 for the air-oil heat exchanger heat rejection (first amount) relative to the fuel-oil heat exchanger rejection (second amount), as taught by Appukuttan, in order to maintain oil temperatures below the maximum even when the fuel heat capacity is low during cruise, descent, and ground idle (Appukuttan, [0004-6]).
Thus, Appukuttan also recognized the heat proportion to air versus fuel as a result-effective variable, i.e., a variable which achieves a recognized result, in the prior art, in this case the proportion of heat rejected from oil to air versus oil to fuel depending on the operating conditions and power level of the engine and resulting in different fuel and oil temperatures (Appukuttan, Fig 4 and [0004-6]), such that the determination of the optimum or workable ranges of said variable, in this case the first proportion at 85% being lower than the second proportion at 65% (0.55 or 0.60 to 0.70 versus 0.85 or 0.90 to 1), may have been characterized as routine experimentation. In re Antonie, 559 F.2d 618, 195 USPQ 6 (CCPA 1977); MPEP 2144.05(II)(B). Where the general conditions of a claim are disclosed in the prior art, in this case Schwarz815 in view of Niergarth156, Schmidt, Gaskell, and Appukuttan teaching all the claimed structures as discussed above, it has been held that the discovery of optimum or workable ranges by experimentation requiring only routine skill in the art, in this case the ability to adjust flow rates of fuel, oil, and air using pumps, valves, and bypass lines, would have been an obvious extension of prior art teachings in order to maintain oil temperatures below the maximum even when the fuel heat capacity is low during cruise, descent, and ground idle (Appukuttan, [0004-6]). In re Aller, 220 F.2d 454, 456, 105 USPQ 233, 235 (CCPA 1955); MPEP 2144.05(II)(A).
Furthermore, Bosak teaches a gas turbine (“gas turbine” [00011]) dissipating a first amount of heat (heat exchanged by 24; Fig. 1) to a first heat sink (air in Air-oil heat exchanger (AOHE) see Fig. 1 ) via an air-lubricant heat exchanger (AOHE, see Fig. 2), wherein the first heat sink is air (air in Air Oil Heat Exchanger (AOHE) see Fig. 1), and
dissipating a second amount of heat (heat exchanged by 22; Fig. 1) to a second heat sink (Fuel in Fuel Oil Heat Exchanger (FOHE) see Fig. 1) via a fuel-lubricant heat exchanger (FOHE, Fig. 1), wherein the second heat sink is fuel ((Fuel in Fuel Oil Heat Exchanger (FOHE) see Fig. 1); and
adjusting a proportion of heat dissipating by air (Figs. 2-5, where flow of oil in FOHE and AOHE is adjusted) at a certain speed of the engine (Figs. 2-5 different flight phase engine load) to adjust the temperature of oil and/or temperature of fuel and thus the fuel consumption (see Figs. 2-5).
Therefore, the proportion of heat dissipated to air at a certain speed of the engine and at certain ambient temperatures is recognized as a result-effective variable, i.e. a variable which achieves a recognized result. In re Antonie, 559 F.2d 618, 195 USPQ 6 (CCPA 1977); MPEP 2144.05(II)(B).
In this case, the recognized result is that the variation of this variable adjusts temperature of oil, fuel and fuel consumption.
Therefore, since the general conditions of the claim, i.e. that proportion of heat dissipated to air at a certain speed of the engine and certain ambient temperatures can be adjusted to adjust oil and fuel temperatures, as well as, fuel consumption, were disclosed in the prior art by Bosak, it is not inventive to discover the optimum workable range by routine experimentation, and it would have been obvious to one of ordinary skill in the art at the time of the invention to provide the claimed the proportion of heat dissipated to air, for both the first and second proportions, as taught by Bosak in order to achieve adjustment of oil temperature, fuel temperature, and fuel consumption. It has been held that “[W]here the general conditions of a claim are disclosed in the prior art, it is not inventive to discover the optimum or workable ranges by routine experimentation.” In re Aller, 220 F.2d 454, 456, 105 USPQ 233, 235 (CCPA 1955); MPEP 2144.05(II)(A).
Claims 1-9, 12 and 17-22 is/are rejected under 35 U.S.C. 103 as being unpatentable over Schwarz 10605104 (hereinafter Schwarz104) in view of Niergarth 11187156 (hereinafter Niergarth156), and Menczykalski 20190316488.
Regarding claim 1, Schwarz104 teaches a gas turbine engine (20) for an aircraft (Fig 1 depicts an aircraft turbofan engine) comprising:
- an engine core (incl.29, 30, 31) comprising a compressor (29), a combustor (30), a turbine (31), and a core shaft (34, 36) connecting the turbine to the compressor (Fig 1);
- a fan (28) comprising a plurality of fan blades (connected to 44) and arranged upstream of the engine core (Fig 1);
- turbomachinery bearings (64);
- a power gearbox (32) adapted to drive the fan at a lower rotation speed than the turbine (col.3 ll.38-65); and
- a heat management system (180, Fig 9; Fig 9 is a modification of Fig 4, and Fig 9 shares some features with Fig 4, thus descriptions and features of elements in Fig 4 also apply to the same elements present in Fig 9) configured to provide lubrication and cooling to the gearbox and turbomachinery bearings (Fig 9; lubricate and cool 64, 32),
and comprising a pipe assembly (incl. 76’’’, 78’’’) adapted to provide a lubricant flow to the gearbox and turbomachinery bearings (Fig 9),
wherein the pipe assembly comprises a first lubricant circuit (78’’’) and a second lubricant circuit (76’’’), each of the first lubricant circuit and the second lubricant circuit being separate except for a tank (72) and tank outlet (88) to which the first lubricant circuit and the second lubricant circuit are commonly connected (Fig 9),
at least one air-lubricant heat exchanger (106) arranged in the first lubricant circuit and configured to dissipate a first amount of heat to a first heat sink (air from 82 via 112),
and at least one fuel-lubricant heat exchanger (146 in place of, or in addition to 92; col.4 ll.62-end, col.5 ll.40-45) arranged in the second lubricant circuit and configured to dissipate a second amount of heat to a second heat sink (fuel from 138, which would be shared feature between Figs 7 and 9 per col.4 ll.62-end, col.5 ll.40-45),
wherein the first heat sink is air and the second heat sink is fuel (col.5 ll.1-13,46-65; col.7 ll.29-38).
Schwarz104 further teaches another air-lubricant heat exchanger (92) in the second lubrication circuit (Fig 9) and another fuel-lubricant heat exchanger (148) may be present in the first lubrication circuit (col.4 ll.62-end and col.5 ll.40-45).
Schwarz104 also teaches modulation of air flow via valves 80, 82 in order to adjust the amount of heat rejected to the first air heat sink (col.6 l.51 – col.7 l.28), and fuel pumps (140, 142) that adjust fuel flow through 146, 148 to injectors 144 (shared feature of Figs 7 and 9 as discussed above). Thus, Schwarz104 is capable of modulating heat rejection and relative proportions of heat rejection (for example a proportion of heat rejected to air relative to the total amount of heat rejected to both air and fuel, a.k.a.
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a second proportion of heat rejection defined as
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at 65% of the core shaft maximum take-off (MTO) speed.
Furthermore, Niergarth156 teaches a gas turbine engine (10) for an aircraft (Fig 1 depicts an aircraft turbofan engine) comprising:
- an engine core (incl. 22, 24, 26, 28, 30, 34, 36) comprising a compressor (22, 24), a combustor (26), a turbine (28, 30), and a core shaft (34, 36) connecting the turbine to the compressor (Fig 1);
- a fan (14) comprising a plurality of fan blades (40) and arranged upstream of the engine core (Fig 1);
- turbomachinery bearings (col.7 l.45);
- a power gearbox (46) adapted to drive the fan at a lower rotation speed than the turbine (col.6 ll.61-63); and
- a heat management system (100 in Fig 2) comprising a pipe assembly (incl. 102) adapted to provide a lubricant flow (oil; col.10 ll.17-31) through various heat sources and heat sinks (106A, 106B, 108A, 108B),
at least one air-lubricant heat exchanger (108B) to dissipate a first amount of heat to a first heat sink (air; col.19 ll.5-9),
and at least one fuel-lubricant heat exchanger (108A) to dissipate a second amount of heat to a second heat sink (fuel from 130),
wherein the first heat sink is air and the second heat sink is fuel (Fig 2; col.19 ll.5-9).
Niergarth156 further teaches modulation of lubrication flow and fuel flow around the various heat sink heat exchangers (108A, 108B) via bypass lines/valves (164, 166, 168, 170, 172, 174, 186, 188) per Fig 2, in order to precisely control the amount of heat rejection to the first air heat sink and second fuel heat sink (col.13 ll.52-64, col.14 ll.14-49; col.15 ll.1-12; col.16 ll.9-44; col.19 ll.5-14) including relative proportions thereof (for example a proportion of heat rejected to air relative to the total amount of heat rejected to both air and fuel, a.k.a.
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). Additionally, because Neirgarth156’s engine is an aircraft engine (Fig 1; col.7 ll.39-42; col.8 ll.4-7), it operates at environment temperatures relative to ISA, and defines a maximum take-off speed (MTO) relative to which it may operate at percentages of speed thereof (e.g. 65% of 85% of MTO).
Thus, Niergarth156 being capable of a first proportion of heat rejection defined as
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at 85% of a core shaft maximum take-off speed; and
a second proportion of heat rejection defined as
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at 65% of the core shaft maximum take-off (MTO) speed.
It would have been obvious to one of ordinary skill in the art before the effective filing date of the claimed invention to modify Schwarz104 with the bypass lines and valves of Niergarth156 in order to further precisely control the amount of heat rejection to the first air heat sink and second fuel heat sink (Niergarth156; col.13 ll.52-64, col.14 ll.14-49; col.15 ll.1-12; col.16 ll.9-44; col.19 ll.5-14).
Schwarz104 in view of Niergarth156 as discussed so far, does not specifically teach the gas turbine engine being configured to provide, at an environment temperature of ISA +40°C, the first proportion from 0.55 to 0.70 and the second proportion from 0.85 to 1; and each of the first lubricant circuit and the second lubricant circuit being separate except for the tank (i.e. tank having two separate outlets).
However, MPEP2112.01 provides that where the claimed and prior art products are identical or substantially identical in structure or composition (as mapped and discussed above, the engine of Schwarz104 in view of Niergarth156 is substantially identical to both the claimed engine and the disclosed, elected engine embodiment in Applicant’s Figs 1-3 and 8), claimed properties or functions (i.e. the achievable first and second proportions of heat rejection to air/fuel being 0.55-0.70 and 0.85-1) are presumed to be present (in this case, the prior art engine being substantially identical to Applicant’s claimed engine reasonably possesses the same characteristics and performance as the claimed engine when operated under the same conditions, e.g. at 85%MTO and 65%MTO in ISA + 40°C) and a prima facie case of either anticipation or obviousness has been established. That is, when the prior art structure and the claimed structure appear to be substantially identical, it is reasonable to conclude that they are capable of performing the same way under the same conditions. In this case, there is no structure, control, or mechanism claimed or disclosed by Applicant that is different from the prior art to enable Applicant’s engine to achieve the claimed performance in a way that the prior art could not.
Schwarz104 in view of Niergarth156 still does not teach each of the first lubricant circuit and the second lubricant circuit being separate except for the tank (i.e. tank having two separate outlets).
However, Menczykalski teaches using two outlets (25, 29) to two lubrication circuits (43, 45) outside of a single tank (53) providing lubricant to turbomachinery bearings (68; [0072]) and a turbofan reduction gearbox (30), respectively. The two different outlets are offset in height within the tank (29 higher than 25), such that when oil levels are running low (e.g. due to a leak, [0004]) than the outlet (29) to the turbomachinery circuit can be used to determine the low oil level and provide a warning signal, thereby providing the pilot an opportunity to react to the low oil level and ensure lubricant flow to the gearbox to avoid damage thereof ([0109]).
It would have been obvious to one of ordinary skill in the art before the effective filing date of the claimed invention to modify Schwarz104 in view of Niergarth156 to include the two separate lubricant circuit outlets from the single tank as taught by Menczykalski in order to facilitate low oil level warning and protection of the gearbox ([0109]).
Regarding claims 2-9, Schwarz104 in view of Niergarth156, and Menczykalski teaches all the limitations of the claimed invention as discussed above (incl. first and second proportions at ISA + 40°C for 85%MTO and 65%MTO, respectively). Schwarz104 being a turbofan engine for an aircraft also teaches operation at cruise conditions to define a cruise proportion of heat (Fig 1; col.4 ll.14-28, col.7 l.33, col.8 ll.54-55, col.9 l.11).
Schwarz104 in view of Niergarth156, and Menczykalski does not teach the following specific performances:
cl.2: the first proportion of heat generated by the gearbox and the turbomachinery and dissipated to air at 85% of a core shaft maximum take-off speed at an environment temperature of ISA +40°C is in from 0.60 to 0.70;
cl.3: the second proportion of heat generated by the gearbox and the turbomachinery and dissipated to air at 65% of the core shaft maximum take-off (MTO) speed at an environment temperature of ISA +40°C is from 0.90 to 1;
cl.4: the heat management system is configured to provide, at an environment temperature of ISA +10°C, first proportion from 0.40 to 0.60;
cl.5: the heat management system is configured to provide, at an environment temperature of ISA +10°C, the second proportion from 0.80 to 0.92;
cl.6: the heat management system is configured to provide, at an environment temperature of ISA -69°C, the first proportion from 0.20 to 0.40;
cl.7: the heat management system is configured to provide, at an environment temperature of ISA -69°C, the second proportion from 0.40 to 0.75;
cl.8: the heat management system is configured to provide, at an environment temperature of ISA +10°C, a ratio of the first proportion to the second proportion from 0.45 to 0.65.
cl.9: the heat management system is configured to provide the first amount of heat and the second amount of heat such that at the cruise conditions the cruise proportion of heat generated by the gearbox and the turbomachinery and dissipated to air is from 0.56 to 0.75.
Regarding the specific performance ranges, as discussed above, MPEP2112.01 provides that where the claimed and prior art products are identical or substantially identical in structure or composition (as mapped and discussed above, the engine of Schwarz104 in view of Niergarth156, and Menczykalski is substantially identical to both the claimed engine and the disclosed, elected engine embodiment in Applicant’s Figs 1-3 and 8), claimed properties or functions (i.e. the achievable first, second, and cruise proportions of heat rejection to air/fuel being in the ranges: 0.60-0.70, 0.90-1, 0.40-0.60, 0.80-0.92, 0.20-0.40, 0.40-0.75, 0.45-0.65, and 0.56-0.75) are presumed to be present (in this case, the prior art engine being substantially identical to Applicant’s claimed engine reasonably possesses the same characteristics and performance as the claimed engine when operated under the same conditions e.g. at 85%MTO, 65%MTO, and cruise in ISA - 69°C, -10°C, and + 40°C) and a prima facie case of either anticipation or obviousness has been established. That is, when the prior art structure and the claimed structure appear to be substantially identical, it is reasonable to conclude that they are capable of performing the same way under the same conditions. In this case, there is no structure, control, or mechanism claimed or disclosed by Applicant that is different from the prior art to enable Applicant’s engine to achieve the claimed performance in a way that the prior art could not.
Regarding claim 12, Schwarz104 in view of Niergarth156, and Menczykalski teaches all the limitations of the claimed invention as discussed above. Schwarz104 further teaches the first heat sink is bypass air (Figs 1 and 4-6; col.6 ll.28-62), and the at least one air-lubricant heat exchanger is adapted to receive bypass air (via 80, 82) from a bypass duct (68).
Regarding claim 21, Schwarz104 in view of Niergarth156, and Menczykalski teaches all the limitations of the claimed invention as discussed above. Schwarz104 further teaches the second lubricant circuit provides lubrication and cooling to the gearbox (32; Fig 9) and the first lubricant circuit provides lubrication and cooling to the turbomachinery bearings (64; Fig 9).
Regarding claim 22, Schwarz104 in view of Niergarth156, and Menczykalski teaches all the limitations of the claimed invention as discussed above. Schwarz104 further teaches at least two air-lubricant heat exchangers (92, 106) to dissipate the first amount of heat to the first heat sink (shared feature of Fig 4 and 9; col.4 ll.62-end, col.5 ll.40-45, col.6 ll.37-44), of which at least one is arranged in the first lubricant circuit and at least one is arranged in the second lubricant circuit (shared feature of Figs 4 and 9; col.4 ll.62-end, col.5 ll.40-45; a fuel-lubricant heat exchanger in each circuit and an air-lubricant heat exchanger in each circuit).
Regarding claims 17-19, Schwarz104 teaches a method of operating a gas turbine engine (20) for an aircraft (Fig 1 depicts an aircraft turbofan engine), the method comprising providing the gas turbine engine comprising:
- an engine core (incl.29, 30, 31) comprising a compressor (29), a combustor (30), a turbine (31), and a core shaft (34, 36) connecting the turbine to the compressor (Fig 1);
- a fan (28) comprising a plurality of fan blades (connected to 44) and arranged upstream of the engine core (Fig 1);
- turbomachinery bearings (64);
- a power gearbox (32) adapted to drive the fan at a lower rotation speed than the turbine (col.3 ll.38-65); and
- a heat management system (180, Fig 9; Fig 9 is a modification of Fig 4, and Fig 9 shares some features with Fig 4, thus descriptions and features of elements in Fig 4 also apply to the same elements present in Fig 9) configured to provide lubrication and cooling to the gearbox and turbomachinery bearings (Fig 9; lubricate and cool 64, 32),
and comprising a pipe assembly (incl. 76’’’, 78’’’) adapted to provide a lubricant flow to the gearbox and turbomachinery bearings (Fig 9),
wherein the pipe assembly comprises a first lubricant circuit (78’’’) and a second lubricant circuit (76’’’), each of the first lubricant circuit and the second lubricant circuit being separate except for a tank (72) and tank outlet (88) to which the first lubricant circuit and the second lubricant circuit are commonly connected (Fig 9),
at least one air-lubricant heat exchanger (106) arranged in the first lubricant circuit and configured to dissipate a first amount of heat to a first heat sink (air from 82 via 112),
and at least one fuel-lubricant heat exchanger (146 in place of, or in addition to 92; col.4 ll.62-end, col.5 ll.40-45) arranged in the second lubricant circuit and configured to dissipate a second amount of heat to a second heat sink (fuel from 138, which would be shared feature between Figs 7 and 9 per col.4 ll.62-end, col.5 ll.40-45),
wherein the first heat sink is air and the second heat sink is fuel (col.5 ll.1-13,46-65; col.7 ll.29-38), and
wherein the method further comprises operating the heat management system to provide the first amount of heat and the second amount of heat (via heat exchangers 92, 106, 146, 148). That is, Schwarz104 further teaches another air-lubricant heat exchanger (92) in the second lubrication circuit (Fig 9) and another fuel-lubricant heat exchanger (148) may be present in the first lubrication circuit (col.4 ll.62-end and col.5 ll.40-45).
Schwarz104 further teaches modulation of air flow via valves 80, 82 in order to adjust the amount of heat rejected to the first air heat sink (col.6 l.51 – col.7 l.28), and fuel pumps (140, 142) that adjust fuel flow through 146, 148 to injectors 144 (Fig 7). Thus, Schwarz104 is capable of modulating heat rejection and relative proportions of heat rejection (for example a proportion of heat rejected to air relative to the total amount of heat rejected to both air and fuel, a.k.a.
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) to air and fuel from the lubrication system that is cooling and lubricating the gearbox and bearings. Additionally, because Schwarz104’s engine is an aircraft engine (Fig 1; col.4 ll.14-28, col.7 l.33, col.8 ll.54-55, col.9 l.11), it operates at environment temperatures relative to ISA, and defines a maximum take-off speed (MTO) relative to which it may operate at percentages of speed thereof (e.g. 65% of 85% of MTO).
Thus, Schwarz104 being capable of a first proportion of heat rejection defined as
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at 85% of a core shaft maximum take-off speed; and
a second proportion of heat rejection defined as
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Furthermore, Niergarth156 teaches a method of operating a gas turbine engine (10) for an aircraft (Fig 1 depicts an aircraft turbofan engine), the method comprising providing the gas turbine engine comprising:
- an engine core (incl. 22, 24, 26, 28, 30, 34, 36) comprising a compressor (22, 24), a combustor (26), a turbine (28, 30), and a core shaft (34, 36) connecting the turbine to the compressor (Fig 1);
- a fan (14) comprising a plurality of fan blades (40) and arranged upstream of the engine core (Fig 1);
- turbomachinery bearings (col.7 l.45);
- a power gearbox (46) adapted to drive the fan at a lower rotation speed than the turbine (col.6 ll.61-63); and
- a heat management system (100 in Fig 2) comprising a pipe assembly (incl. 102) adapted to provide a lubricant flow (oil; col.10 ll.17-31) through various heat sources and heat sinks (106A, 106B, 108A, 108B),
at least one air-lubricant heat exchanger (108B) to dissipate a first amount of heat to a first heat sink (air; col.19 ll.5-9),
and at least one fuel-lubricant heat exchanger (108A) to dissipate a second amount of heat to a second heat sink (fuel from 130),
wherein the first heat sink is air and the second heat sink is fuel (Fig 2; col.19 ll.5-9);
and wherein the method further comprises operating the heat management system to provide the first amount of heat and the second amount of heat (via heat exchangers 108B, 108A).
Niergarth156 further teaches modulation of lubrication flow and fuel flow around the various heat sink heat exchangers (108A, 108B) via bypass lines/valves (164, 166, 168, 170, 172, 174, 186, 188) per Fig 2, in order to precisely control the amount of heat rejection to the first air heat sink and second fuel heat sink (col.13 ll.52-64, col.14 ll.14-49; col.15 ll.1-12; col.16 ll.9-44; col.19 ll.5-14) including relative proportions thereof (for example a proportion of heat rejected to air relative to the total amount of heat rejected to both air and fuel, a.k.a.
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). Additionally, because Neirgarth156’s engine is an aircraft engine (Fig 1; col.7 ll.39-42; col.8 ll.4-7), it operates at environment temperatures relative to ISA, and defines a maximum take-off speed (MTO) relative to which it may operate at percentages of speed thereof (e.g. 65% of 85% of MTO).
Thus, Niergarth156 being capable of a first proportion of heat rejection defined as
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at 85% of a core shaft maximum take-off speed; and
a second proportion of heat rejection defined as
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It would have been obvious to one of ordinary skill in the art before the effective filing date of the claimed invention to modify Schwarz104 with the bypass lines and valves of Niergarth156 in order to further precisely control the amount of heat rejection to the first air heat sink and second fuel heat sink (Niergarth156; col.13 ll.52-64, col.14 ll.14-49; col.15 ll.1-12; col.16 ll.9-44; col.19 ll.5-14).
Schwarz104 in view of Niergarth156 as discussed so far, does not specifically teach the method further comprises operating the engine at an environment temperature of ISA +40°C and at 85% of the core shaft maximum take-off speed, wherein the first proportion is from (claim 17) 0.55 to 0.70 or (claim 18) from 0.60 to 0.70, and operating the engine at the environment temperature of ISA +40°C and at 65% of the core shaft maximum take-off speed, wherein the second proportion is from (claim 17) 0.85 to 1 or (claim 19) from 0.90 to 1; each of the first lubricant circuit and the second lubricant circuit being separate except for the tank (i.e. tank having two separate outlets).
However, MPEP2112.02(I) provides that if a prior art device (Schwarz104 in view of Niergarth156), in its normal and usual operation (for flight, which requires takeoff, landing, cruise, etc. over a standard flight envelope under ambient conditions), would necessarily perform the method claimed (i.e. operating at 65%MTO and 85%MTO), then the method claimed will be considered to be anticipated by the prior art device. Note, that operation of an aircraft engine at ambient temperature up to ISA + 40°C was considered part of a normal/standard range of operating conditions as evidence by Applicant’s Admission that “engines are normally certified to operate in a range of environment temperatures between ISA (International Standard Atmosphere) -69°C and ISA +40°C” (on p.6, para.3 of Applicant’s Specification filed 15 February 2023) and typical Type-Certificate Data Sheets for engines registered with the EASA listing under “Climatic Operating Envelope” and engine usage “in ambient temperatures up to ISA +40°C” (e.g.: No. EASA E.047 for the RB211 Trent 800 series engines, dated 21 February 2015, Issue 04, pp.10-11).. MPEP2112.02(I) also provides that when the prior art device is the same as a device described in the specification for carrying out the claimed method, it can be assumed the device will inherently perform the claimed process. That is, when the prior art structure and the claimed structure appear to be substantially identical, it is reasonable to conclude that they perform the same way (i.e. the achievable first and second proportions of heat rejection to air/fuel being 0.55-0.70, 0.6-0.7, 0.85-1, and 0.90-1) under the same conditions (e.g. at 85%MTO and 65%MTO in ISA + 40°C). In this case, the engine of Schwarz104 in view of Niergarth156 is substantially identical to both the claimed engine and the disclosed, elected engine embodiment in Applicant’s Figs 1-3 and 8); and, based on the current record, there is no structure, control, or mechanism claimed or disclosed by Applicant that is different from the prior art to enable Applicant’s engine to achieve the claimed performance in a way that the prior art would not.
Schwarz104 in view of Niergarth156 still does not teach each of the first lubricant circuit and the second lubricant circuit being separate except for the tank (i.e. tank having two separate outlets).
However, Menczykalski teaches using two outlets (25, 29) to two lubrication circuits (43, 45) outside of a single tank (53) providing lubricant to turbomachinery bearings (68; [0072]) and a turbofan reduction gearbox (30), respectively. The two different outlets are offset in height within the tank (29 higher than 25), such that when oil levels are running low (e.g. due to a leak, [0004]) than the outlet (29) to the turbomachinery circuit can be used to determine the low oil level and provide a warning signal, thereby providing the pilot an opportunity to react to the low oil level and ensure lubricant flow to the gearbox to avoid damage thereof ([0109]).
It would have been obvious to one of ordinary skill in the art before the effective filing date of the claimed invention to modify Schwarz104 in view of Niergarth156 to include the two separate lubricant circuit outlets from the single tank as taught by Menczykalski in order to facilitate low oil level warning and protection of the gearbox ([0109]).
Regarding claim 20, Schwarz104 in view of Niergarth156, and Menczykalski teaches all the limitations of the claimed invention as discussed above. Schwarz104 in view of Niergarth156, and Menczykalski as discussed so far, does not specifically teach operating the heat management system at 85% of the core shaft maximum take-off speed at an environment temperature of ISA +10°C, wherein a ratio of the first proportion at 85% of the core shaft maximum take-off speed at the environment temperature of ISA + 40°C to the first proportion at 85% of the core shaft maximum take-off speed at the environment temperature of ISA +10°C is from 1.20 to 1.42.
However, MPEP2112.02(I) provides that if a prior art device (Schwarz104 in view of Niergarth156, and Menczykalski), in its normal and usual operation (for flight, which requires takeoff, landing, cruise, etc. over a standard flight envelope under ambient conditions), would necessarily perform the method claimed (i.e. operating at 85%MTO), then the method claimed will be considered to be anticipated by the prior art device. Note, that operation of an aircraft engine at ambient temperature up to ISA + 40°C (which also includes operation at ISA + 10°C) was considered part of a normal/standard range of operating conditions as evidence by Applicant’s Admission that “engines are normally certified to operate in a range of environment temperatures between ISA (International Standard Atmosphere) -69°C and ISA +40°C” (on p.6, para.3 of Applicant’s Specification filed 15 February 2023) and typical Type-Certificate Data Sheets for engines registered with the EASA listing under “Climatic Operating Envelope” and engine usage “in ambient temperatures up to ISA +40°C” including “maximum continuous” operation at “ISA + 10°C” (e.g.: No. EASA E.047 for the RB211 Trent 800 series engines, dated 21 February 2015, Issue 04, pp.10-11). MPEP2112.02(I) also provides that when the prior art device is the same as a device described in the specification for carrying out the claimed method, it can be assumed the device will inherently perform the claimed process. That is, when the prior art structure and the claimed structure appear to be substantially identical, it is reasonable to conclude that they perform the same way (i.e. the ratio of first proportions being in the range of 1.20-1.42) under the same conditions (e.g. at 85%MTO in ISA + 40°C and ISA + 10°C). In this case, the engine of Schwarz104 in view of Niergarth156, and Menczykalski is substantially identical to both the claimed engine and the disclosed, elected engine embodiment in Applicant’s Figs 1-3 and 8); and, based on the current record, there is no structure, control, or mechanism claimed or disclosed by Applicant that is different from the prior art to enable Applicant’s engine to achieve the claimed performance in a way that the prior art would not.
Claims 10-11 is/are rejected under 35 U.S.C. 103 as being unpatentable over Schwarz104 in view of Niergarth156 and Menczykalski, and further in view of Gaskell 20210190008.
Regarding claim 10, Schwarz104 in view of Niergarth156, and Menczykalski teaches all the limitations of the claimed invention as discussed above. Schwarz104 in view of Niergarth156, and Menczykalski as discussed so far, does not teach the core shaft maximum take-off speed is from 5500 rpm to 9500 rpm.
However, Gaskell teaches a gas turbine engine (Fig 1) for an aircraft (Abstract) comprising:
- an engine core (11) comprising a compressor (14, 15), a combustor (16), a turbine (17, 19), and a core shaft (26, 27) connecting the turbine to the compressor (Fig 1),
wherein the core shaft has a core shaft maximum take-off speed in the range of from 5500rpm to 9500rpm ([0023]; 5800-6200rpm falling in the claimed range);
- a fan (23) comprising a plurality of fan blades and arranged upstream of the engine core (Fig 1);
- turbomachinery bearings (Fig 1 below; and/or 26a, 26b, 26c in Fig 3); and
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- a power gearbox (30) adapted to drive the fan at a lower rotation speed than the turbine ([0010]).
Gaskell further teaches at least the bearings require lubrication ([0360, 367-368]); and the turbofan engine is operable in a speed range of 85% and 65% of the maximum take-off speed (engine operable between 1500-6200rpm, thus including 85% and 65% of the MTO speed of range 5800-6200rpm, the 85% speed being ≈ 4930-5270rpm and the 65% speed being ≈ 3770-4030rpm; [0021,23]). Note, Gaskell also teaches cruise speeds as 5400-5700rpm (or ≈ 87-92%MTO) and ground idle speed as 1500rpm (≈ 24-26%MTO).
Gaskell is silent as to the details of any lubrication and/or thermal management system for the turbofan engine.
It would have been obvious to one of ordinary skill in the art before the effective filing date of the claimed invention to modify the geared turbofan engine (comprising a power gearbox and turbomachinery bearings) of Gaskell to use the lubrication and thermal management system of Schwarz104 in view of Niergarth156, and Menczykalski, in order to: provide improved lubrication and cooling to the multiple turbine engine components (power gearbox and turbomachinery bearings) of the turbofan (Schwarz104, col.1 ll.13-30), which cooling is adjustable based on the changing cooling needs of the components (Schwarz104, col.7 ll.2-7); and precisely control the amount of heat rejection to the first air heat sink and second fuel heat sink (Niergarth156; col.13 ll.52-64, col.14 ll.14-49; col.15 ll.1-12; col.16 ll.9-44; col.19 ll.5-14).
Regarding claim 11, Schwarz104 in view of Niergarth156, and Menczykalski teaches all the limitations of the claimed invention as discussed above. Schwarz104 in view of Niergarth156, and Menczykalski as discussed so far, does not teach the fan has a fan rotational speed at MTO conditions from 1500 rpm to 2800 rpm.
However, Gaskell teaches a gas turbine engine (Fig 1) for an aircraft (Abstract) comprising:
- an engine core (11) comprising a compressor (14, 15), a combustor (16), a turbine (17, 19), and a core shaft (26, 27) connecting the turbine to the compressor (Fig 1),
wherein the core shaft has a core shaft maximum take-off speed in the range of from 5500rpm to 9500rpm ([0023]; 5800-6200rpm falling in the claimed range);
- a fan (23) comprising a plurality of fan blades and arranged upstream of the engine core (Fig 1),
wherein the fan has a fan rotational speed at MTO conditions in the range of from 1500 rpm to 2800 rpm (the gear ratio being greater than 3, to drive the fan, thus teaching the fan rpm at MTO of
5800
3
=
1933
to
6200
3
=
2067
rpm; [0023] teaching core shaft MTO speed of 5800-6200rpm);
- turbomachinery bearings (Fig 1 below; and/or 26a, 26b, 26c in Fig 3); and
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- a power gearbox (30) adapted to drive the fan at a lower rotation speed than the turbine ([0010]).
Gaskell further teaches at least the bearings require lubrication ([0360, 367-368]); and the turbofan engine is operable in a speed range of 85% and 65% of the maximum take-off speed (engine operable between 1500-6200rpm, thus including 85% and 65% of the MTO speed of range 5800-6200rpm, the 85% speed being ≈ 4930-5270rpm and the 65% speed being ≈ 3770-4030rpm; [0021,23]). Note, Gaskell also teaches cruise speeds as 5400-5700rpm (or ≈ 87-92%MTO) and ground idle speed as 1500rpm (≈ 24-26%MTO).
Gaskell is silent as to the details of any lubrication and/or thermal management system for the turbofan engine.
It would have been obvious to one of ordinary skill in the art before the effective filing date of the claimed invention to modify the geared turbofan engine (comprising a power gearbox and turbomachinery bearings) of Gaskell to use the lubrication and thermal management system of Schwarz104 in view of Niergarth156, and Menczykalski, in order to: provide improved lubrication and cooling to the multiple turbine engine components (power gearbox and turbomachinery bearings) of the turbofan (Schwarz104, col.1 ll.13-30), which cooling is adjustable based on the changing cooling needs of the components (Schwarz104, col.7 ll.2-7); and precisely control the amount of heat rejection to the first air heat sink and second fuel heat sink (Niergarth156; col.13 ll.52-64, col.14 ll.14-49; col.15 ll.1-12; col.16 ll.9-44; col.19 ll.5-14).
Claim 13 is/are rejected under 35 U.S.C. 103 as being unpatentable over Schwarz104 in view of Niergarth156 and Menczykalski, and further in view of Schwarz 20080028763 (Schwarz2008).
Regarding claim 13, Schwarz104 in view of Niergarth156, and Menczykalski teaches all the limitations of the claimed invention as discussed above. Schwarz104 further teaches the heat management system includes a flow restriction valve (80, 82) arranged upstream of the at least one air-lubricant heat exchanger (Figs 5-6) adapted to vary a mass flow rate of the bypass air across the at least one air-lubricant heat exchanger (Figs 5-6; col.6 l.28 – col.7 l.28), thereby varying the first amount of heat (col.7 ll.2-7, 24-28).
Schwarz104 in view of Niergarth156, and Menczykalski does not teach the valve being downstream of the at least one air-lubricant heat exchanger.
However, Schwarz2008 teaches an air-oil heat exchanger (44; Figs 2-3) using bypass air (F) of a turbofan engine (Fig 1A) installed in a pylon of the turbofan (36), wherein flaps (52) that are used to modulate the air flow (and hence the heat transfer) through the heat exchanger (44) are placed downstream of the heat exchanger in order to simultaneously provide thrust recovery (Fig 3; [0008}).
It would have been obvious to one of ordinary skill in the art before the effective filing date of the claimed invention to modify the turbofan of Schwarz104 in view of Niergarth156, and Menczykalski to use the pylon-integrated heat exchanger and valve of Schwarz2008, in order to simultaneously provide thrust recovery (Schwarz2008, [0008)).
Claim 16 is/are rejected under 35 U.S.C. 103 as being unpatentable over Schwarz104 in view of Niergarth156 and Menczykalski, and further in view of Pilatis 20050097889 and Leese 9051056.
Regarding claim 16, Schwarz104 in view of Niergarth156, and Menczykalski teaches all the limitations of the claimed invention as discussed above. Schwarz104 further teaches the combustor comprising a plurality of fuel spray nozzles (144).
Schwarz104 in view of Niergarth156, and Menczykalski does not the combustor is a lean burn combustor, thus the plurality fuel spray nozzles being lean burn fuel spray nozzles; and each fuel spray nozzle comprising a pilot fuel injector and a main fuel injector.
However, Pilatis teaches a combustor (11) of a turbofan (Fig 1) that is a lean combustor (Fig 2; [0014, 22]) comprising a plurality of lean burn fuel spray nozzles (60 in Fig 2; [0021]), each fuel spray nozzle comprising a pilot fuel injector a main fuel injector ([0022-23]).
And Leese teaches adding a fuel-lubricant heat exchanger bypass line/valve (30, 29) and a fuel-air heat exchanger (25) to modify fuel/lubricant systems of turbofan engines (compare Figs 1 and 4) to accommodate the stricter fuel temperature requirements of lean burn combustors (col.2 ll.16-22; col.4 ll.42-64; col.5 l.65 – col.6 l.20).
It would have been obvious to one of ordinary skill in the art before the effective filing date of the claimed invention to modify the turbofan of Schwarz104 in view of Niergarth156, and Menczykalski to accommodate a lean burn combustor as taught by Pilatis, in order to reduce NOx emissions (Pilatis, [0002]) and because Leese teaches methods of modifying the lubrication and fuel thermal systems to facilitate the lean burn combustion (Leese, col.2 ll.16-22; col.4 ll.42-64; col.5 l.65 – col.6 l.20).
Claims 17-20 is/are additionally rejected under 35 U.S.C. 103 as being unpatentable over Schwarz104 in view of Niergarth156 and Menczykalski, and further in view of Gaskell, Appukuttan 20160281603 and Bosak 20210172375.
Regarding claims 17-20, Schwarz104 in view of Niergarth156, and Menczykalski teaches all the limitations of the claimed invention as discussed above.
In addition, Gaskell teaches a method of operating a gas turbine engine (Fig 1) for an aircraft (Abstract),
the method comprising providing a gas turbine engine (Fig 1) comprising:
- an engine core (11) comprising a compressor (14, 15), a combustor (16), a turbine (17, 19), and a core shaft (26, 27) connecting the turbine to the compressor (Fig 1),
wherein the core shaft has a core shaft maximum take-off speed in the range of from 5500rpm to 9500rpm ([0023]; 5800-6200rpm falling in the claimed range);
- a fan (23) comprising a plurality of fan blades and arranged upstream of the engine core (Fig 1);
- turbomachinery bearings (Fig 1 below); and
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- a power gearbox (30) adapted to drive the fan at a lower rotation speed than the turbine ([0010]).
Gaskell further teaches at least the bearings require lubrication ([0360, 367-368]); and the turbofan engine is operable in a speed range of 85% and 65% of the maximum take-off speed (engine operable between 1500-6200rpm, thus including 85% and 65% of the MTO speed of range 5800-6200rpm, the 85% speed being ≈ 4930-5270rpm and the 65% speed being ≈ 3770-4030rpm; [0021,23]). Note, Gaskell also teaches cruise speeds as 5400-5700rpm (or ≈ 87-92%MTO) and ground idle speed as 1500rpm (≈ 24-26%MTO).
Gaskell is silent as to the details of any lubrication and/or thermal management system for the turbofan engine.
It would have been obvious to one of ordinary skill in the art before the effective filing date of the claimed invention to modify the geared turbofan engine (comprising a power gearbox and turbomachinery bearings) of Gaskell to use the lubrication and thermal management system of Schwarz104 in view of Niergarth156, and Menczykalski, in order to: provide improved lubrication and cooling to the multiple turbine engine components (power gearbox and turbomachinery bearings) of the turbofan (Schwarz104, col.1 ll.13-30), which cooling is adjustable based on the changing cooling needs of the components (Schwarz104, col.7 ll.2-7); and precisely control the amount of heat rejection to the first air heat sink and second fuel heat sink (Niergarth156; col.13 ll.52-64, col.14 ll.14-49; col.15 ll.1-12; col.16 ll.9-44; col.19 ll.5-14).
Furthermore, Appukuttan teaches a gas turbine engine (10) for an aircraft comprising:
- an engine core (incl. 14, 16, 18, 20, 22, 24) comprising a compressor (14, 16), a combustor (18), a turbine (22, 24), and a core shaft (28, 30) connecting the turbine to the compressor (Fig 1), wherein the core shaft has a core shaft maximum take-off speed (as is defined for every engine);
- a fan (12) comprising a plurality of fan blades (Fig 1) and arranged upstream of the engine core (Fig 1);
- turbomachinery bearings ([0027]);
- a heat management system (Figs 2-3) configured to provide lubrication and cooling to the turbomachinery bearings and other heat loads ([0027]), and
comprising a pipe assembly (incl. 1322, 116, 124) adapted to provide a lubricant flow (132) to the turbomachinery bearings ([0027]; Figs 2-3),
at least one air- lubricant heat exchanger (122) to dissipate a first amount of heat to a first heat sink (air; [0025]),
at least one fuel-lubricant heat exchanger (108) to dissipate a second amount of heat to a second heat sink (fuel, [0024]),
wherein the first heat sink is air and the second heat sink is fuel ([0024-25]),
wherein the heat management system is configured to provide the first amount of heat and the second amount of heat such that a first proportion of heat generated by the turbomachinery and other heat loads dissipated to air defined as
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M
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at 85% of the core shaft maximum take-off speed (turbofan engines for aircraft operate between 0-100% of MTO over flight envelope) is in the range of from 0.55 to 0.70 or 0.60 to 0.70 (this is a desired-result/intended-use of the system; the turbofan engine having all the claimed structures is capable of achieving the claimed desired-result/intended-use of the system; e.g. by using valve 126; see also MPEP2112.01(I)); and
a second proportion of heat generated by the gearbox and the turbomachinery and dissipated to air is defined as
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at the 65% of the core shaft maximum take-off (MTO) speed (turbofan engines for aircraft operate between 0-100% of MTO over flight envelope) is in the range of from 0.85 to 1 or 0.90 to 1 (this is a desired-result/intended-use of the system; the turbofan engine having all the claimed structures is capable of achieving the claimed desired-result/intended-use of the system; e.g. by using valve 126; see also MPEP2112.01(I)).
Appukuttan further teaches that at lower operating power levels (i.e. cruise and below, including descent and ground idle; Fig 4), the fuel heat capacity is below the oil heat rejection requirement (Fig 4) and the FOHE is insufficient to provide the required cooling, such that the AOHE must provide the remaining proportion of cooling/heat-rejection to the lubricant ([0005-6]). According to Fig 4, the fraction of heat rejected to the AOHE may be at least
200
-
150
200
k
W
k
W
≈ 0.25 at cruise and at least
100
-
40
100
k
W
k
W
≈ 0.6 at descent.
It would have been obvious to one of ordinary skill in the art before the effective filing date of the claimed invention to modify the method of Schwarz104 in view of Niergarth156, Menczykalski, and Gaskell (in which Gaskell teaches 85%MTO is slightly less than cruise) to use a proportion of ≥ 0.25 for the air-oil heat exchanger heat rejection (first amount) relative to the fuel-oil heat exchanger rejection (second amount), as taught by Appukuttan, in order to maintain oil temperatures below the maximum even when the fuel heat capacity is low during cruise, descent, and ground idle (Appukuttan, [0004-6]).
Thus, Appukuttan also recognized the heat proportion to air versus fuel as a result-effective variable, i.e., a variable which achieves a recognized result, in the prior art, in this case the proportion of heat rejected from oil to air versus oil to fuel depending on the operating conditions and power level of the engine and resulting in different fuel and oil temperatures (Appukuttan, Fig 4 and [0004-6]), such that the determination of the optimum or workable ranges of said variable, in this case the first proportion at 85% being lower than the second proportion at 65% (0.55 or 0.60 to 0.70 versus 0.85 or 0.90 to 1), may have been characterized as routine experimentation. In re Antonie, 559 F.2d 618, 195 USPQ 6 (CCPA 1977); MPEP 2144.05(II)(B). Where the general conditions of a claim are disclosed in the prior art, in this case Schwarz104 in view of Niergarth156, Menczykalski, Gaskell, and Appukuttan teaching all the claimed structures as discussed above, it has been held that the discovery of optimum or workable ranges by experimentation requiring only routine skill in the art, in this case the ability to adjust flow rates of fuel, oil, and air using pumps, valves, and bypass lines, would have been an obvious extension of prior art teachings in order to maintain oil temperatures below the maximum even when the fuel heat capacity is low during cruise, descent, and ground idle (Appukuttan, [0004-6]). In re Aller, 220 F.2d 454, 456, 105 USPQ 233, 235 (CCPA 1955); MPEP 2144.05(II)(A).
Furthermore, Bosak teaches a gas turbine (“gas turbine” [00011]) dissipating a first amount of heat (heat exchanged by 24; Fig. 1) to a first heat sink (air in Air-oil heat exchanger (AOHE) see Fig. 1 ) via an air-lubricant heat exchanger (AOHE, see Fig. 2), wherein the first heat sink is air (air in Air Oil Heat Exchanger (AOHE) see Fig. 1), and
dissipating a second amount of heat (heat exchanged by 22; Fig. 1) to a second heat sink (Fuel in Fuel Oil Heat Exchanger (FOHE) see Fig. 1) via a fuel-lubricant heat exchanger (FOHE, Fig. 1), wherein the second heat sink is fuel ((Fuel in Fuel Oil Heat Exchanger (FOHE) see Fig. 1); and
adjusting a proportion of heat dissipating by air (Figs. 2-5, where flow of oil in FOHE and AOHE is adjusted) at a certain speed of the engine (Figs. 2-5 different flight phase engine load) to adjust the temperature of oil and/or temperature of fuel and thus the fuel consumption (see Figs. 2-5).
Therefore, the proportion of heat dissipated to air at a certain speed of the engine and at certain ambient temperatures is recognized as a result-effective variable, i.e. a variable which achieves a recognized result. In re Antonie, 559 F.2d 618, 195 USPQ 6 (CCPA 1977); MPEP 2144.05(II)(B).
In this case, the recognized result is that the variation of this variable adjusts temperature of oil, fuel and fuel consumption.
Therefore, since the general conditions of the claim, i.e. that proportion of heat dissipated to air at a certain speed of the engine and certain ambient temperatures can be adjusted to adjust oil and fuel temperatures, as well as, fuel consumption, were disclosed in the prior art by Bosak, it is not inventive to discover the optimum workable range by routine experimentation, and it would have been obvious to one of ordinary skill in the art at the time of the invention to provide the claimed the proportion of heat dissipated to air, for both the first and second proportions, as taught by Bosak in order to achieve adjustment of oil temperature, fuel temperature, and fuel consumption. It has been held that “[W]here the general conditions of a claim are disclosed in the prior art, it is not inventive to discover the optimum or workable ranges by routine experimentation.” In re Aller, 220 F.2d 454, 456, 105 USPQ 233, 235 (CCPA 1955); MPEP 2144.05(II)(A).
Response to Arguments
Applicant's arguments filed 01 August 2025 have been carefully considered, but they are not persuasive. To the extent possible, Applicant’s arguments were addressed in the rejections above at the relevant locations.
In summary:
Applicant asserts that “[t]he Office Action alleges that Menczkalski discloses “piping of each of the first lubricant circuit and the second lubricant circuit being separate outside of a common tank,” as previously recited, based upon lubrication circuits 43 and 45 and tank 53. Office Action p. 7”.
However, it is clarified, that the previous office action (mailed 14 May 2025) does not allege that Menczkalski, alone, teaches the above cited limitation. Rather, p.7 of the previous office action states:
“Menczykalski teaches using two outlets (25, 29) to two lubrication circuits (43, 45) outside of a shared common tank (53) providing lubricant to turbomachinery bearings (68; [0072]) and a turbofan reduction gearbox (30), respectively. The two different outlets are offset in height within the tank (29 higher than 25), such that when oil levels are running low (e.g. due to a leak, [0004]) than the outlet (29) to the turbomachinery circuit can be used to determine the low oil level and provide a warning signal, thereby providing the pilot an opportunity to react to the low oil level and ensure lubricant flow to the gearbox to avoid damage thereof ([0109]).”
This paragraph is meant to explain that Menczkalski specifically teaches two outlets from a lubrication tank supplying two respective circuits (not necessarily mutually exclusive circuits). The rest of the claimed lubrication circuit details are taught by Schwarz104 in view of Niergarth156 as discussed in the previous section(s).
Applicant is correct that Menczkalski teaches two lubricant circuits that supply different components, but also supply at least one common component (gearbox). However, this fact does not change the applicability of the teachings of separate outlets taught by Menczkalski as further discussed below.
Applicant further asserts that “the low oil level warning and protection of the gearbox would only function if both circuits commonly lubricate the gearbox” based on a citation to paragraph [0109] of Menczkalski relating protection of the engine against “single point failure” to the offset oil tank outlets.
However, Menczkalski, as a whole, teaches cross feeding and scavenging of lubrication to protect the journal bearings of the gearbox from single-point failure ([0014, 72-73, 77-78]). The use of offset outlets to provide low oil level warning merely provides extra time to the pilot to determine a course of action ([0071, 0109] “In the event of low oil levels in the oil tank 53 the turbomachinery circuit 45 will generate a low oil pressure warning or another warning signal first without the first circuit 43 having yet been affected. The level of offset can be designed such that, under all reasonable rates of oil loss, the pilot has reacted to the warning before the first oil circuit 43 feed to the journal bearings is lost or reduced”). Although this feature does further facilitate avoiding single point failure (using cross feeding and scavenging of lubrication), it also facilitates avoiding failure in general (regardless of whether cross feeding and scavenging are implemented). That is, regardless of whether the second oil circuit also feeds the journal bearings (fed by the first circuit), offsetting the second oil circuit tank outlet higher than the first oil circuit tank outlet provides the low oil pressure/level warning before the first oil circuit is affected, as desired by Menczkalski. Note, MPEP2123(I) provides that the use of patents as references is not limited to what the patentees describe as their own inventions or to the problems with which they are concerned; they are part of the literature of the art, relevant for all they contain. In this case, the use of Menczkalski as a secondary reference in a 103 non-obviousness rejection is not limited to what Menczkalski describes as their own invention (i.e. for the purpose of avoiding single-point failure). The technical solution of offset lubrication tank outlets for providing the pilot more time to react to a low oil level/pressure warning as taught by Menczkalski is useful and applicable to other prior art systems regardless of whether the other prior art systems are subject to single-point failure.
Nevertheless, additional prior art rejections are also presented above (in light of the claim amendments pertaining specifically to the first and second lubricant circuits being separate except for a tank to which the first and second lubricant circuits are commonly connected) to address Applicant’s concerns.
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.
Correspondence
Any inquiry concerning this communication or earlier communications from the examiner should be directed to STEPHANIE SEBASCO CHENG whose telephone number is (469)295-9153. The examiner can normally be reached on 1000-1600 ET.
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/STEPHANIE SEBASCO CHENG/Primary Examiner, Art Unit 3741
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