DETAILED ACTION
The present application, filed on or after March 16, 2013, is being examined under the first inventor to file provisions of the AIA .
Continued Examination Under 37 CFR 1.114
A request for continued examination under 37 CFR 1.114, including the fee set forth in 37 CFR 1.17(e), was filed in this application after final rejection. Since this application is eligible for continued examination under 37 CFR 1.114, and the fee set forth in 37 CFR 1.17(e) has been timely paid, the finality of the previous Office action has been withdrawn pursuant to 37 CFR 1.114. Applicant's submission filed on 03/24/2025 has been entered.
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.
The factual inquiries for establishing a background for determining obviousness under 35 U.S.C. 103 are summarized as follows:
1. Determining the scope and contents of the prior art.
2. Ascertaining the differences between the prior art and the claims at issue.
3. Resolving the level of ordinary skill in the pertinent art.
4. Considering objective evidence present in the application indicating obviousness or nonobviousness.
Claim(s) 1, 3, 9-12 and 14 is/are rejected under 35 U.S.C. 103 as being unpatentable over Pub. No.: US 2020/0173369 A1 (Husband) in view of NPL “Civil Jet Aircraft Design” Online Database (Jenkinson).
Regarding claim 1, Husband discloses (see figs. 1 and 4) a gas turbine engine 10 for an aircraft (see par. 66), the gas turbine engine comprising, in axial flow sequence, a compressor module (14 and/or 15), a combustor module 16, and a turbine module (17 and/or 19), and further comprising a first electrical machine (54 for 56) rotationally connected to the turbine module (each electric machine is mounted to or otherwise coupled to respective shafts 26,27 wherein turbines 19,17 respectively drive such shafts; see pars. 111 and 113), the combustor module has a combustor volume V (m3) (see par. 178: “V is the volume of the combustion chamber 16”; this is consistent with applicant p. 58, ll. 9-11: “The term combustor volume 172 means the total volume of the combustor module 170 where fuel is burnt”; and wherein the combustion chamber is where fuel is burnt). Husband does not explicitly disclose wherein, in use, at a full power condition, the gas turbine engine has a maximum corrected core flow Q (kg/sec), and a ratio T of:
T
=
(
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(
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is in a range of between 500 and 1,500.
Jenkinson teaches a gas turbine (see title) and further teaches a Maximum Corrected Core Flow = Q. Jenkinson teaches the airflow and fuel flow through a variety of sample turbofan engines at takeoff conditions corrected for static conditions at International Standard Atmosphere (discussed below for Rolls Royce RB211-535E4/E4-B engine; see “TO (ISA SLS)” in upper left). It is known that BR=Fan flow rate/Core flow (i.e., Fan mass flow rate = BR X Core mass flow rate; this is consistent with applicant page 5, ll. 5-12) rate and that the Mass Flow through the engine is the sum of the Fan flow rate and the Core flow rate (i.e. Fan flow rate = Mass flow – Core flow rate). Thus Mass flow – Core flow rate = BR X Core flow rate therefore Mass flow = BR X Core flow rate + Core flow rate and thus Mass flow = Core flow rate X (BR +1). The Mass Flow for the instant RB211 that has a BR of 4.3 at takeoff is 1151 lb/s or in other words 523 Kg/s (this is interpreted as the air flow rate since the SFC is shown that is a measure of fuel flow). The Core flow rate therefore at takeoff maximum conditions is Core flow rate = Mass flow/(BR + 1) = 99 Kg/s. One of ordinary skill is knowledgeable that takeoff power is maximum or full power as discussed in the Pertinent Prior Art section infra. Jenkinson further teaches a SFC for the instant RB211 being 0.598 lb/hr/lb such that the fuel mass flow rate through the RB211 engine core = SFC X thrust = .598 lb/hr/lb X 43100 lbs = 25,773 lb/hr = 11,600 kg/hr = 3.2 kg/s. Therefore the total maximum corrected core flow is 99 Kg/s plus 3.2 Kg/s = 102.2 Kg/s in the scenario the claimed mass flow includes fuel (the claimed mass flow being discussed at applicant page 58, ll. 1-7). It is thought that applicant claimed Maximum Corrected Core Flow = Q represents an air mass flow in light of applicant page 5, ll. 5-12, page 38, ll. 10-15 and page 58, ll. 1-7 wherein there is not fuel discussed and additionally a similar disclosure to applicant disclosure discussed specific thrust as being with respect to air mass wherein specific thrust is discussed at applicant page 38, ll. 10-15. However, in case the fuel flow is included in the parameter Q, the prior art for this analysis includes fuel flow as a part of the mass flow Q.
It is further noted that “when a patent claims a structure already known in the prior art that is altered by the mere substitution of one element for another known in the field, the combination must do more than yield a predictable result.” KSR International Co. v. Teleflex Inc., 82 USPQ2d 1385 at 1395 (U.S. 2007) (MPEP 2143 I.B.).
It would have been obvious to one of ordinary skill in the art before the effective filing date of the current invention to substitute the value of Maximum Corrected Core Flow = Q = 102.2 Kg/s (or alternatively Q = 99 Kg/s if fuel is not included in the mass flow) of Jenkinson for the value of Maximum Corrected Core Flow = Q of Husband for the purpose of substituting one known element for another in order to provide the expected result of providing a mass flow through a turbofan engine to produce thrust of the combination. For example, Husband states in par. 60 that the engine of Husband may have a variety of mass flows through the engine corresponding to the range of thrust levels of the engine of Husband. In par. 61, states 190 kN for maximum thrust at takeoff. This is consistent with Jenkinson’s RB211 thrust of 189 kN (i.e., 43,100 lb thrust).
The teachings of Husband in view of Jenkinson thus far include for example
T
=
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=
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=
102.2
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)
(
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=
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and the claim requires T is in the range between 500 and 1,500. Another way of communicating this scenario is the boundaries for combustor volume to meet the claim limitations is
(
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=
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=
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or in other words
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500
t
o
1,500
=
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V
or in other words the boundaries of V are
(
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=
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=
102.2
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T
=
500
and
(
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=
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=
102.2
K
g
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T
=
1,500
. Thus in order to meet the claim limitations V is between 102.2/500 m3 and 102.2/1,500 m3 or in other words V is between .204 m3 and .068 m3. Thus Husband in view of Jenkinson teach the current invention as claimed and discussed above except wherein
C
o
m
b
u
s
t
o
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v
o
l
u
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e
V
is between and .068 m3 and .204 m3.
Husband teaches in par. 180 that the volume of a gas turbine combustor is a design point of the combustor. For example, the combustor volume must be sufficient enough to permit an in-flight restart. If the volume is too small a ground start may be able to be accomplished however a small volume combustor can preclude a mid-air restart. If the volume is too large than the combustor would take up extra space within the engine and add additional weight resulting in decreased fuel efficiency. Thus, the combustor volume must be optimized for safety and one of ordinary skill would understand that the combustor volume V is a result effective variable. For example, not optimizing combustor volume would not be acceptable because a mid-air restart may not be able to be accomplished (such mid-air restart being accommodated by optimizing combustor volume).
Therefore, an ordinary skilled worker would recognize that the Combustor volume V is a result-effective variable that controls starting features of an engine. Thus, the claimed wherein
C
o
m
b
u
s
t
o
r
v
o
l
u
m
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V
being between .068 m3 and .204 m3 is found to be an obvious optimization of the prior art obtainable by an ordinary skilled worker through routine experimentation.
Therefore, since the general conditions of the claim, i.e. a combustor volume of a turbofan engine were disclosed in the prior art by Husband in view of Jenkinson, 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 before the effective filing date of the claimed invention to modify Husband’s invention to include wherein
C
o
m
b
u
s
t
o
r
v
o
l
u
m
e
V
being between .068 m3 and .204 m3 such that
T
=
(
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m
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d
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=
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(
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is in a range of between 500 and 1,500 in order to provide an engine being able to accomplish an in-flight restart as suggested and taught by Husband and Jenkinson discussed above. It has been held “where 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).
Regarding claim 3, Husband in view of Jenkinson teach the current invention as claimed and discussed above. Husband further discloses (see fig. 1) the gas turbine engine 10 is a turbofan engine 10 comprising, in axial flow sequence, a fan assembly 23, a compressor module (14 and/or 15), a combustor module 16, and a turbine module (17 and/or 19).
Regarding claims 9 and 10, Husband in view of Jenkinson teach the current invention as claimed and discussed above. Husband further discloses (see fig. 1) in use, the maximum dry thrust produced by the engine 10 is in the range of between 30kN and 170kN (see 160 kN in par. 61); and in use, the maximum dry thrust produced by the engine is in the range of between 170kN and 500kN (see e.g., 300 kN in par. 61). The term dry thrust in the gas turbine arts can be referred to as thrust without use of an afterburner. No such afterburner is disclosed by Husband and thus the instant thrust is dry thrust.
Regarding claim 11, Husband in view of Jenkinson teach the current invention as claimed and discussed above. Husband further discloses (see fig. 1) the aircraft (see par. 66) comprising a gas turbine engine 23 as claimed in Claim 1.
Regarding claim 12, Husband discloses (see fig. 2) a method of operating a gas turbine engine 10 for an aircraft (see par. 66), the gas turbine engine comprising, in axial flow sequence, a compressor module (14 and/or 15), a combustor module 16, and a turbine module (17 and/or 19); wherein the method comprises the steps of: (i) providing, in axial flow sequence, the compressor module, the combustor module, and the turbine module (see fig. 2); (ii) sizing the combustor module to have a combustor volume V (m3) (see par. 180). Husband does not explicitly disclose
(iii) operating the gas turbine engine at a full power condition in which the gas turbine engine has a maximum corrected core flow Q (kg/sec), and a ratio T of:
T
=
(
M
a
x
i
m
u
m
C
o
r
r
e
c
t
e
d
C
o
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e
F
l
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w
=
Q
)
(
C
o
m
b
u
s
t
o
r
v
o
l
u
m
e
=
V
)
is in a range of between 500 and 1,500.
Jenkinson teaches a gas turbine (see title) and further teaches a Maximum Corrected Core Flow = Q. Jenkinson teaches the airflow and fuel flow through a variety of sample turbofan engines at takeoff conditions corrected for static conditions at International Standard Atmosphere (discussed below for Rolls Royce RB211-535E4/E4-B engine; see “TO (ISA SLS)” in upper left). It is known that BR=Fan flow rate/Core flow (i.e., Fan mass flow rate = BR X Core mass flow rate; this is consistent with applicant page 5, ll. 5-12) rate and that the Mass Flow through the engine is the sum of the Fan flow rate and the Core flow rate (i.e. Fan flow rate = Mass flow – Core flow rate). Thus Mass flow – Core flow rate = BR X Core flow rate therefore Mass flow = BR X Core flow rate + Core flow rate and thus Mass flow = Core flow rate X (BR +1). The Mass Flow for the instant RB211 that has a BR of 4.3 at takeoff is 1151 lb/s or in other words 523 Kg/s (this is interpreted as the air flow rate since there the SFC is shown that is a measure of fuel flow). The Core flow rate therefore at takeoff maximum conditions is Core flow rate = Mass flow/(BR + 1) = 99 Kg/s. One of ordinary skill is knowledgeable that takeoff power is maximum or full power as discussed in the Pertinent Prior Art section infra. Jenkinson further teaches a SFC for the instant RB211 being 0.598 lb/hr/lb such that the fuel mass flow rate through the RB211 engine core = SFC X thrust = .598 lb/hr/lb X 43100 lbs = 25,773 lb/hr = 11,600 kg/hr = 3.2 kg/s. Therefore the total maximum corrected core flow is 99 Kg/s plus 3.2 Kg/s = 102.2 Kg/s in the scenario the claimed mass flow includes fuel (the claimed mass flow being discussed at applicant page 58, ll. 1-7). It is thought that applicant claimed Maximum Corrected Core Flow = Q represents an air mass flow in light of applicant page 5, ll. 5-12, page 38, ll. 10-15 and page 58, ll. 1-7 wherein there is not fuel discussed and additionally a similar disclosure to applicant disclosure discussed specific thrust as being with respect to air mass wherein specific thrust is discussed at applicant page 38, ll. 10-15. However, in case the fuel flow is included in the parameter Q, the prior art for this analysis includes fuel flow as a part of the mass flow Q.
It is further noted that “when a patent claims a structure already known in the prior art that is altered by the mere substitution of one element for another known in the field, the combination must do more than yield a predictable result.” KSR International Co. v. Teleflex Inc., 82 USPQ2d 1385 at 1395 (U.S. 2007) (MPEP 2143 I.B.).
It would have been obvious to one of ordinary skill in the art before the effective filing date of the current invention to substitute the value of Maximum Corrected Core Flow = Q = 102.2 Kg/s (or alternatively Q = 99 Kg/s if fuel is not included in the mass flow) of Jenkinson for the value of Maximum Corrected Core Flow = Q of Husband for the purpose of substituting one known element for another in order to provide the expected result of providing a mass flow through a turbofan engine to produce thrust of the combination. For example, Husband states in par. 60 that the engine of Husband may have a variety of mass flows through the engine corresponding to the range of thrust levels of the engine of Husband. In par. 61, states 190 kN for maximum thrust at takeoff. This is consistent with Jenkinson’s RB211 thrust of 189 kN (i.e., 43,100 lb thrust).
The teachings of Husband in view of Jenkinson thus far include for example
T
=
(
M
a
x
i
m
u
m
C
o
r
r
e
c
t
e
d
C
o
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F
l
o
w
=
Q
=
102.2
K
g
/
s
)
(
C
o
m
b
u
s
t
o
r
v
o
l
u
m
e
=
V
)
and the claim requires T is in the range between 500 and 1,500. Another way of communicating this scenario is the boundaries for combustor volume to meet the claim limitations is
(
M
a
x
i
m
u
m
C
o
r
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e
c
t
e
d
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o
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=
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=
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(
C
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=
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)
=
T
or in other words
(
M
a
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o
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c
t
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=
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a
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f
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m
500
t
o
1,500
=
C
o
m
b
u
s
t
o
r
v
o
l
u
m
V
or in other words the boundaries of V are
(
M
a
x
i
m
u
m
C
o
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c
t
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d
C
o
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F
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w
=
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=
102.2
K
g
/
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)
T
=
500
and
(
M
a
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m
u
m
C
o
r
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e
c
t
e
d
C
o
r
e
F
l
o
w
=
Q
=
102.2
K
g
/
s
)
T
=
1,500
. Thus in order to meet the claim limitations V is between 102.2/500 m3 and 102.2/1,500 m3 or in other words V is between .204 m3 and .068 m3. Thus Husband in view of Jenkinson teach the current invention as claimed and discussed above except wherein
C
o
m
b
u
s
t
o
r
v
o
l
u
m
e
V
is between and .068 m3 and .204 m3.
Husband teaches in par. 180 that the volume of a gas turbine combustor is a design point of the combustor. For example, the combustor volume must be sufficient enough to permit an in-flight restart. If the volume is too small a ground start may be able to be accomplished however a small volume combustor can preclude a mid-air restart. If the volume is too large than the combustor would take up extra space within the engine and add additional weight resulting in decreased fuel efficiency. Thus, the combustor volume must be optimized for safety and one of ordinary skill would understand that the combustor volume V is a result effective variable. For example, not optimizing combustor volume would not be acceptable because a mid-air restart may not be able to be accomplished (such mid-air restart being accommodated by optimizing combustor volume).
Therefore, an ordinary skilled worker would recognize that the Combustor volume V is a result-effective variable that controls starting features of an engine. Thus, the claimed wherein
C
o
m
b
u
s
t
o
r
v
o
l
u
m
e