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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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.
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 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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