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.
Specification
Amendment to the Specification filed 26 June 2026 is accepted and entered.
Claim Rejections - 35 USC § 112
The following is a quotation of 35 U.S.C. 112(b):
(b) CONCLUSION.—The specification shall conclude with one or more claims particularly pointing out and distinctly claiming the subject matter which the inventor or a joint inventor regards as the invention.
The following is a quotation of 35 U.S.C. 112 (pre-AIA ), second paragraph:
The specification shall conclude with one or more claims particularly pointing out and distinctly claiming the subject matter which the applicant regards as his invention.
Claims 14 and 21 are rejected under 35 U.S.C. 112(b) or 35 U.S.C. 112 (pre-AIA ), second paragraph, as being indefinite for failing to particularly point out and distinctly claim the subject matter which the inventor or a joint inventor, or for pre-AIA the applicant regards as the invention.
Regarding claims 14 and 21, both claims recite the same fan speed to fan-turbine radius ratio, twice, with two different ranges, wherein one range is narrower than the other. Thus, the scope of the claims is unclear as to which range defines the metes and bounds of the claim.
Dependent claims 15-19 are also rejected for depending upon at least one rejected claim above.
Claim Rejections - 35 USC § 102
The following is a quotation of the appropriate paragraphs of 35 U.S.C. 102 that form the basis for the rejections under this section made in this Office action:
A person shall be entitled to a patent unless –
(a)(1) the claimed invention was patented, described in a printed publication, or in public use, on sale, or otherwise available to the public before the effective filing date of the claimed invention.
(a)(2) the claimed invention was described in a patent issued under section 151, or in an application for patent published or deemed published under section 122(b), in which the patent or application, as the case may be, names another inventor and was effectively filed before the effective filing date of the claimed invention.
Claim(s) 14-19 is/are rejected under 35 U.S.C. 102(a)(1) as being anticipated by Gliebe1.
1 Philip R. Gliebe and Bangalore A. Janardan, Ultra-High Bypass Engine Aeroacoustic Study, October 2003, NASA/CR-2003-212525
Regarding Claim 14, Gliebe teaches a gas turbine engine for an aircraft (the S30 Engine #4 at bottom of Fig 4) comprising:
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an engine core having a core length (annotated Fig 4 below) and comprising, in continuous downstream axial flow series, a compressor, a combustor, and a turbine (annotated Fig 4 above);
the turbine comprising a lowest pressure rotor stage having a row of rotor blades (annotated Fig 4 above),
the turbine having a turbine diameter (2 × Rt in Fig 4 below) at the lowest pressure rotor stage, and each of the rotor blades extending radially and having a leading edge and a trailing edge (annotated Fig 4 above);
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438
1294
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a core shaft connecting the turbine to the compressor (required for each spool);
a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub (Fig 4 above), each of the plurality of fan blades having a leading edge (FLE, Fig 4 above) and a trailing edge (TFE in annotated Fig 4 above); and
a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft (annotated Fig 4 above),
wherein a fan radius (Rf, Fig 4 above) is defined as a point on a circle swept by an outermost tip of the leading edge of one of the plurality of fan blades, and a turbine radius (Rt, Fig 4 above) at the lowest pressure rotor stage is defined as a radial distance between a point on a circle swept by a radially outer tip of the trailing edge of each of the rotor blades of the lowest pressure rotor stage (interpreted to include Rf-Rt in Fig 4 above), and
wherein a fan-turbine radius difference is the difference between the fan and turbine radii, and a fan speed to fan-turbine radius ratio is defined as:
a
m
a
x
i
m
u
m
t
a
k
e
-
o
f
f
r
o
t
a
t
i
o
n
a
l
s
p
e
e
d
o
f
t
h
e
f
a
n
i
n
r
p
m
f
a
n
-
t
u
r
b
i
n
e
r
a
d
i
u
s
d
i
f
f
e
r
e
n
c
e
(
i
n
m
m
)
in the range of 1.2 to 2.0 rpm/mm (MTO fan speed in rpm for engine S30 is 1691rpm per Table 11, while the fan radius is
130
2
=
65
i
n
. or 1651mm per Fig 4 and the turbine radius is ~25in or 635mm per Fig 4; which gives
1651
1691
-
635
≈
1.66
rpm/mm, which falls in the claimed range); and
PNG
media_image3.png
438
1294
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Greyscale
a nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside of the engine core (Fig 4 above), the bypass duct comprising a bypass exhaust nozzle having a bypass exhaust nozzle exit (Fig 4 above), the bypass exhaust nozzle having an outer radius measured as a radial distance between a centreline of the gas turbine engine and an inner surface of the nacelle at an axial position of a rearmost tip of the nacelle (Fig 4 above), wherein an outer bypass to fan ratio of
t
h
e
o
u
t
e
r
r
a
d
i
u
s
o
f
t
h
e
b
y
p
a
s
s
e
x
h
u
a
s
t
n
o
z
z
l
e
t
h
e
f
a
n
t
i
p
r
a
d
i
u
s
is in a range from 0.91-0.98 (
≈
60
65
≈
0.92); and
wherein a fan-turbine radius difference is defined as a radial distance between: a point on a circle swept by a radially outer tip of the trailing edge of one of the rotor blades of the lowest pressure rotor stage; and a point on a circle swept by the outermost tip of the leading edge of the one of the plurality of fan blades (interpreted to include Rf-Rt per annotated Fig 4 above), and
fan speed to fan-turbine radius ratio is defined as:
a
m
a
x
i
m
u
m
t
a
k
e
-
o
f
f
r
o
t
a
t
i
o
n
a
l
s
p
e
e
d
o
f
t
h
e
f
a
n
i
n
r
p
m
f
a
n
-
t
u
r
b
i
n
e
r
a
d
i
u
s
d
i
f
f
e
r
e
n
c
e
(
i
n
m
m
)
in the range of 1.50 to 1.7 rpm/mm (MTO fan speed in rpm for engine S30 is 1691rpm per Table 11, while the fan radius is
130
2
=
65
i
n
. or 1651mm per Fig 4 and the turbine radius is ~25in or 635mm per Fig 4; which gives
1691
1651
-
635
≈
1.66
rpm/mm, which falls in the claimed range).
Gliebe further teaches cruise conditions as 0.8M, 35000ft and standard day conditions (Tables 1-2, 4); component efficiency of the fan at cruise is 0.94 (94%; Table 2), fan pressure ratio (FPR) at cruise is 1.30 (Table 4), fan tip speed at cruise is 984fps or ~300m/s (Tables 4, 10).
Regarding claim 15, Gliebe teaches all the limitations of the claimed invention as discussed above. Gliebe further teaches the ratio of the fan radius to twice the turbine radius at the lowest pressure rotor stage is in the range of 0.9 to 1.7 (
65
2
×
25
≈
1.3
per Fig 4, which falls in the claimed range).
Regarding claim 16, Gliebe teaches all the limitations of the claimed invention as discussed above. Gliebe further teaches the ratio of the fan radius to the core length is in the range of 0.5 to 0.8 (
65
90
≈
0.72
from first compressor blade leading edge to last turbine blade trailing edge, per Fig 4, which falls in the claimed range).
Regarding claim 17, Gliebe teaches all the limitations of the claimed invention as discussed above. Gliebe further teaches the fan rotates at a rotational fan speed, and wherein the gas turbine engine is configured such that the maximum take-off rotational fan speed is in the range of 1450 rpm to 1910 rpm (1691rpm from Table 11 falls in the claimed range).
Regarding claims 18-19, Gliebe teaches all the limitations of the claimed invention as discussed above.
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Gliebe further teaches the gas turbine engine has an engine length and a centre of gravity (CGPos) position measured relative to the fan (Fig 4 above), and
a centre of gravity position ratio of
t
h
e
c
e
n
t
r
e
o
f
g
r
a
v
i
t
y
p
o
s
i
t
i
o
n
t
h
e
e
n
g
i
n
e
l
e
n
g
t
h
, and
wherein the fan has a maximum take-off rotational speed (1691rpm per Table 11), and
wherein a fan speed to centre of gravity ratio of
t
h
e
c
e
n
t
r
e
o
f
g
r
a
v
i
t
y
p
o
s
i
t
i
o
n
r
a
t
i
o
×
m
a
x
i
m
u
m
t
a
k
e
-
o
f
f
r
o
t
a
t
i
o
n
a
l
s
p
e
e
d
is in a range from 600 rpm to 910 rpm for claim 18, or in arrange of 650 rpm to 910 rpm for claim 19
47
120
×
1691
≈
662
r
p
m
.
Claim Rejections - 35 USC § 103
The following is a quotation of 35 U.S.C. 103 which forms the basis for all obviousness rejections set forth in this Office action:
A patent for a claimed invention may not be obtained, notwithstanding that the claimed invention is not identically disclosed as set forth in section 102, if the differences between the claimed invention and the prior art are such that the claimed invention as a whole would have been obvious before the effective filing date of the claimed invention to a person having ordinary skill in the art to which the claimed invention pertains. Patentability shall not be negated by the manner in which the invention was made.
Claim(s) 11, 13, and 21 is/are rejected under 35 U.S.C. 103 as being unpatentable over Gliebe1 and Goulos2.
1 Philip R. Gliebe and Bangalore A. Janardan, Ultra-High Bypass Engine Aeroacoustic Study, October 2003, NASA/CR-2003-212525
2 Goulos et al., Aerodynamic Design of Separate-Jet Exhausts for Future Civil Aero-engines—
Part I: Parametric Geometry Definition and Computational Fluid Dynamics Approach, Journal of
Engineering for Gas Turbines and Power, Volume 138, August 2016, pages 081201-1 to 081201-14,
Downloaded From: http://gasturbinespower.asmedigitalcollection.asme.org/ on 03/16/2016
Regarding Claim 11, Gliebe teaches a gas turbine engine for an aircraft (the S30 Engine #4 at bottom of Fig 4) comprising:
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an engine core having a core length (annotated Fig 4 below) and comprising, in continuous downstream axial flow series, a compressor, a combustor, and a turbine (annotated Fig 4 above);
the turbine comprising a lowest pressure rotor stage having a row of rotor blades (annotated Fig 4 above),
the turbine having a turbine diameter (2 × Rt in Fig 4 below) at the lowest pressure rotor stage, and each of the rotor blades extending radially and having a leading edge and a trailing edge (annotated Fig 4 above);
PNG
media_image2.png
438
1294
media_image2.png
Greyscale
a core shaft connecting the turbine to the compressor (required for each spool);
a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub, each of the plurality of fan blades having a leading edge (FLE in Fig 4 above) and a trailing edge (FTE in Fig 4 above), the hub and the plurality of fan blades together defining a fan face having a fan face area (π × Rf2) and a fan tip radius (Rf in annotated Fig 4 above); and
a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft (annotated Fig 4 above),
wherein a fan-turbine radius difference is defined as a radial distance between: a point on a circle swept by a radially outer tip of the trailing edge of each of the rotor blades of the lowest pressure rotor stage; and a point on a circle swept by an outermost tip of the leading edge of one of the plurality of fan blades (interpreted to include Rf-Rt per annotated Fig 4 above), and
wherein a fan-turbine radius difference is the difference between the fan and turbine radii, and a fan speed to fan-turbine radius ratio is defined as:
a
m
a
x
i
m
u
m
t
a
k
e
-
o
f
f
r
o
t
a
t
i
o
n
a
l
s
p
e
e
d
o
f
t
h
e
f
a
n
i
n
r
p
m
f
a
n
-
t
u
r
b
i
n
e
r
a
d
i
u
s
d
i
f
f
e
r
e
n
c
e
(
i
n
m
m
)
in the range of 1.2 to 2.0 rpm/mm (MTO fan speed in rpm for engine S30 is 1691rpm per Table 11, while the fan radius is
130
2
=
65
i
n
. or 1651mm per Fig 4 and the turbine radius is ~25in or 635mm per Fig 4; which gives
1691
1651
-
635
≈
1.66
rpm/mm, which falls in the claimed range); and
PNG
media_image3.png
438
1294
media_image3.png
Greyscale
a nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside of the engine core (Fig 4 above), the bypass duct comprising a bypass exhaust nozzle having a bypass exhaust nozzle exit (Fig 4 above), the bypass exhaust nozzle having an outer radius measured as a radial distance between a centreline of the gas turbine engine and an inner surface of the nacelle at an axial position of a rearmost tip of the nacelle (Fig 4 above), wherein an outer bypass to fan ratio of
t
h
e
o
u
t
e
r
r
a
d
i
u
s
o
f
t
h
e
b
y
p
a
s
s
e
x
h
u
a
s
t
n
o
z
z
l
e
t
h
e
f
a
n
t
i
p
r
a
d
i
u
s
is in a range from 0.91-0.98 (
≈
60
65
≈
0.92);
the bypass exhaust nozzle having a bypass exhaust nozzle pressure ratio calculated using total pressure at the bypass exhaust nozzle exit under cruise conditions (the engine operates at cruise and the nozzle is characterized by a total pressure ratio in operation at cruise); and
a core exhaust nozzle pressure ratio (the core exhaust nozzle downstream of the low pressure turbine is characterized by a total pressure ratio in operation).
Gliebe further teaches cruise conditions as 0.8M, 35000ft and standard day conditions (Tables 1-2, 4); component efficiency of the fan at cruise is 0.94 (94%; Table 2), fan pressure ratio (FPR) at cruise is 1.30 (Table 4), fan tip speed at cruise is 984fps or ~300m/s (Tables 4, 10).
Gliebe does not teach the ratio of bypass exhaust nozzle pressure ratio to core exhaust nozzle pressure ratio is in the range from 1.3 to 1.6 under cruise conditions.
However, Goulos teaches a gas turbine engine (E1) for an aircraft (p.081201-8 section 3.3.1 Case Study Description) comprising:
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433
1105
media_image5.png
Greyscale
an engine core (Figs 1 and 12(a)) comprising a turbine, a compressor, a core shaft connecting the turbine to the compressor (conventional features of a turbofan core as depicted in Figs 1, 12(a)), and a core exhaust nozzle having a core exhaust nozzle exit (Figs 1, 12(a)),
PNG
media_image6.png
445
739
media_image6.png
Greyscale
the core exhaust nozzle having a core exhaust nozzle pressure ratio (core NPR) calculated using total pressure at the core nozzle exit (as is conventionally understood, see evidentiary reference: NASA, “Nozzle Performance”, by the Glen Research Center, defining NPR = pt8/p0, where position 8 is at the exit of the nozzle and position 0 is the free stream and defining pt8 = pt5, where position 5 is at the inlet of the nozzle due to no thermodynamic work being done in the nozzle; see p.081201-8, Col.2, Table 2, Col. E1 of Goulos);
a fan (Figs 1, 12(a)) located upstream of the engine core (Figs 1, 12(a)), the fan comprising a plurality of fan blades (required for turbofan operation),
wherein a fan tip radius of the fan (p.081201-13, Col.2 l.24) is measured between a centerline (about which Figs 1 and 12(a) are axisymmetric) of the gas turbine engine and an outermost tip of each fan blade (approximately within shaded region of Fig 12(a); clearance between fan tip and nacelle or fan engine case must be small in order to maintain fan efficiency); and
PNG
media_image7.png
357
1344
media_image7.png
Greyscale
a nacelle (Figs 1 and 12(a) above) surrounding the fan and the engine core and defining a bypass duct (Figs 1, 12(a)) located radially outside of the engine core (Figs 1, 12(a)),
the bypass duct comprising a bypass exhaust nozzle (Figs 1, 12(a)) having a bypass exhaust nozzle exit (Figs 1, 12(a)),
the bypass exhaust nozzle having an outer radius measured as a radial distance between the centerline of the gas turbine engine and an inner surface of the nacelle at an axial position of a rearmost tip of the nacelle (Fig 12(a) above),
wherein an outer bypass to fan ratio of:
t
h
e
o
u
t
e
r
r
a
d
i
u
s
o
f
t
h
e
b
y
p
a
s
s
e
x
h
a
u
s
t
n
o
z
z
l
e
t
h
e
f
a
n
t
i
p
r
a
d
i
u
s
is in a range from 0.65 to 1.00 (dashed line units in Annotated Fig 12(a) shows outer bypass to fan ratio being approximately
10.5
11
≈
0.95 to 1), and
the bypass exhaust nozzle having a bypass exhaust nozzle pressure ratio (bypass NPR) calculated using total pressure at the bypass nozzle exit (as is conventionally understood, see evidentiary reference: NASA, “Nozzle Performance”, by the Glen Research Center, defining NPR = pt8/p0, where position 8 is at the exit of the nozzle and position 0 is the free stream and defining pt8 = pt5, where position 5 is at the inlet of the nozzle due to no thermodynamic work being done in the nozzle; see p.081201-8, Col.2, Table 2, Col.E1 of Goulos); wherein a bypass to core ratio of:
b
y
p
a
s
s
e
x
h
a
u
s
t
n
o
z
z
l
e
p
r
e
s
s
u
r
e
r
a
t
i
o
[
c
o
r
e
e
x
h
a
u
s
t
n
o
z
z
l
e
p
r
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s
s
u
r
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r
a
t
i
o
]
is configured to be in the range from 1.3 to 1.6 under aircraft cruise conditions (p.081201-8, Col.2, Table 2, Col.E1, Ratio =
2.2
1.5
=
1.47
).
Goulos further teaches the BPR being “of the order of 16” (p.081201-8 Col.1 last paragraph).
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 bypass and core nozzles of Gliebe to achieve the bypass to core nozzle pressure ratio taught by Goulos because the nozzle pressure ratios of Houlos are specifically compatible with he high bypass ratio of ≈16 taught by Gliebe (Goulos, p.081201-1 col.2 para.2 to p.081201-2 col.1 para.1 and p,081201-8 col.1 last para.). Furthermore, it has been held that combining or simple substitution of prior art elements according to known methods to yield predictable results renders the limitation obvious (see MPEP 2141 (III)). In this case, combining the high bypass (BPR ≈ 16) gas turbine engine of Gliebe with the exhaust nozzle pressure ratio of Goulos’ engine E1 that has a bypass ratio of ≈ 16 would have resulted in an operable turbofan engine with BPR ≈ 16 and bypass to core exhaust nozzle pressure ratio of ≈ 1.47.
Regarding claim 13, Gliebe in view of Goulos teaches all the limitations of the claimed invention as discussed above.
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Gliebe further teaches the gas turbine engine has an engine length and a centre of gravity (CGPos) position measured relative to the fan (Fig 4 above), and
a centre of gravity position ratio of
t
h
e
c
e
n
t
r
e
o
f
g
r
a
v
i
t
y
p
o
s
i
t
i
o
n
t
h
e
e
n
g
i
n
e
l
e
n
g
t
h
, and
wherein the fan has a maximum take-off rotational speed (1691rpm per Table 11), and
wherein a fan speed to centre of gravity ratio of
t
h
e
c
e
n
t
r
e
o
f
g
r
a
v
i
t
y
p
o
s
i
t
i
o
n
r
a
t
i
o
×
m
a
x
i
m
u
m
t
a
k
e
-
o
f
f
r
o
t
a
t
i
o
n
a
l
s
p
e
e
d
is in a range from 600 rpm to 910 rpm
47
120
×
1691
≈
662
r
p
m
.
Regarding Claim 21, Gliebe teaches a gas turbine engine for an aircraft (the S30 Engine #4 at bottom of Fig 4) comprising:
PNG
media_image1.png
438
1294
media_image1.png
Greyscale
an engine core having a core length (annotated Fig 4 below) and comprising, in continuous downstream axial flow series, a compressor, a combustor, and a turbine (annotated Fig 4 above);
the turbine comprising a lowest pressure rotor stage having a row of rotor blades (annotated Fig 4 above),
the turbine having a turbine diameter (2 × Rt in Fig 4 below) at the lowest pressure rotor stage, and each of the rotor blades extending radially and having a leading edge and a trailing edge (annotated Fig 4 above);
PNG
media_image2.png
438
1294
media_image2.png
Greyscale
a core shaft connecting the turbine to the compressor (required for each spool);
a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub, each of the plurality of fan blades having a leading edge (FLE in Fig 4 above) and a trailing edge (FTE in Fig 4 above), the hub and the plurality of fan blades together defining a fan face having a fan face area (π × Rf2) and a fan tip radius (Rf in annotated Fig 4 above); and
a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft (annotated Fig 4 above),
wherein a fan-turbine radius difference is defined as a radial distance between: a point on a circle swept by a radially outer tip of the trailing edge of each of the rotor blades of the lowest pressure rotor stage; and a point on a circle swept by an outermost tip of the leading edge of one of the plurality of fan blades (interpreted to include Rf-Rt per annotated Fig 4 above), and
wherein a fan-turbine radius difference is the difference between the fan and turbine radii, and a fan speed to fan-turbine radius ratio is defined as:
a
m
a
x
i
m
u
m
t
a
k
e
-
o
f
f
r
o
t
a
t
i
o
n
a
l
s
p
e
e
d
o
f
t
h
e
f
a
n
i
n
r
p
m
f
a
n
-
t
u
r
b
i
n
e
r
a
d
i
u
s
d
i
f
f
e
r
e
n
c
e
(
i
n
m
m
)
in the range of 1.2 to 2.0 rpm/mm (MTO fan speed in rpm for engine S30 is 1691rpm per Table 11, while the fan radius is
130
2
=
65
i
n
. or 1651mm per Fig 4 and the turbine radius is ~25in or 635mm per Fig 4; which gives
1691
1651
-
635
≈
1.66
rpm/mm, which falls in the claimed range); and
PNG
media_image3.png
438
1294
media_image3.png
Greyscale
a nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside of the engine core (Fig 4 above), the bypass duct comprising a bypass exhaust nozzle having a bypass exhaust nozzle exit (Fig 4 above), the bypass exhaust nozzle having an outer radius measured as a radial distance between a centreline of the gas turbine engine and an inner surface of the nacelle at an axial position of a rearmost tip of the nacelle (Fig 4 above), wherein an outer bypass to fan ratio of
t
h
e
o
u
t
e
r
r
a
d
i
u
s
o
f
t
h
e
b
y
p
a
s
s
e
x
h
u
a
s
t
n
o
z
z
l
e
t
h
e
f
a
n
t
i
p
r
a
d
i
u
s
is in a range from 0.91-0.98 (
≈
60
65
≈
0.92);
the bypass exhaust nozzle having a bypass exhaust nozzle pressure ratio calculated using total pressure at the bypass exhaust nozzle exit under cruise conditions (the engine operates at cruise and the nozzle is characterized by a total pressure ratio in operation at cruise); and
a core exhaust nozzle pressure ratio (the core exhaust nozzle downstream of the low pressure turbine is characterized by a total pressure ratio in operation).
wherein a fan-turbine radius difference is defined as a radial distance between: a point on a circle swept by a radially outer tip of the trailing edge of one of the rotor blades of the lowest pressure rotor stage; and a point on a circle swept by the outermost tip of the leading edge of the one of the plurality of fan blades (interpreted to include Rf-Rt per annotated Fig 4 above), and
fan speed to fan-turbine radius ratio is defined as:
a
m
a
x
i
m
u
m
t
a
k
e
-
o
f
f
r
o
t
a
t
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p
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m
f
a
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(
i
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m
m
)
in the range of 1.50 to 1.7 rpm/mm (MTO fan speed in rpm for engine S30 is 1691rpm per Table 11, while the fan radius is
130
2
=
65
i
n
. or 1651mm per Fig 4 and the turbine radius is ~25in or 635mm per Fig 4; which gives
1691
1651
-
635
≈
1.66
rpm/mm, which falls in the claimed range).
Gliebe further teaches cruise conditions as 0.8M, 35000ft and standard day conditions (Tables 1-2, 4); component efficiency of the fan at cruise is 0.94 (94%; Table 2), fan pressure ratio (FPR) at cruise is 1.30 (Table 4), fan tip speed at cruise is 984fps or ~300m/s (Tables 4, 10).
Gliebe does not teach the ratio of bypass exhaust nozzle pressure ratio to core exhaust nozzle pressure ratio is in the range from 1.3 to 1.6 under cruise conditions.
However, Goulos teaches a gas turbine engine (E1) for an aircraft (p.081201-8 section 3.3.1 Case Study Description) comprising:
PNG
media_image5.png
433
1105
media_image5.png
Greyscale
an engine core (Figs 1 and 12(a)) comprising a turbine, a compressor, a core shaft connecting the turbine to the compressor (conventional features of a turbofan core as depicted in Figs 1, 12(a)), and a core exhaust nozzle having a core exhaust nozzle exit (Figs 1, 12(a)),
PNG
media_image6.png
445
739
media_image6.png
Greyscale
the core exhaust nozzle having a core exhaust nozzle pressure ratio (core NPR) calculated using total pressure at the core nozzle exit (as is conventionally understood, see evidentiary reference: NASA, “Nozzle Performance”, by the Glen Research Center, defining NPR = pt8/p0, where position 8 is at the exit of the nozzle and position 0 is the free stream and defining pt8 = pt5, where position 5 is at the inlet of the nozzle due to no thermodynamic work being done in the nozzle; see p.081201-8, Col.2, Table 2, Col. E1 of Goulos);
a fan (Figs 1, 12(a)) located upstream of the engine core (Figs 1, 12(a)), the fan comprising a plurality of fan blades (required for turbofan operation),
wherein a fan tip radius of the fan (p.081201-13, Col.2 l.24) is measured between a centerline (about which Figs 1 and 12(a) are axisymmetric) of the gas turbine engine and an outermost tip of each fan blade (approximately within shaded region of Fig 12(a); clearance between fan tip and nacelle or fan engine case must be small in order to maintain fan efficiency); and
PNG
media_image7.png
357
1344
media_image7.png
Greyscale
a nacelle (Figs 1 and 12(a) above) surrounding the fan and the engine core and defining a bypass duct (Figs 1, 12(a)) located radially outside of the engine core (Figs 1, 12(a)),
the bypass duct comprising a bypass exhaust nozzle (Figs 1, 12(a)) having a bypass exhaust nozzle exit (Figs 1, 12(a)),
the bypass exhaust nozzle having an outer radius measured as a radial distance between the centerline of the gas turbine engine and an inner surface of the nacelle at an axial position of a rearmost tip of the nacelle (Fig 12(a) above),
wherein an outer bypass to fan ratio of:
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a
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o
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t
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e
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z
l
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a
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t
i
p
r
a
d
i
u
s
is in a range from 0.65 to 1.00 (dashed line units in Annotated Fig 12(a) shows outer bypass to fan ratio being approximately
10.5
11
≈
0.95 to 1), and
the bypass exhaust nozzle having a bypass exhaust nozzle pressure ratio (bypass NPR) calculated using total pressure at the bypass nozzle exit (as is conventionally understood, see evidentiary reference: NASA, “Nozzle Performance”, by the Glen Research Center, defining NPR = pt8/p0, where position 8 is at the exit of the nozzle and position 0 is the free stream and defining pt8 = pt5, where position 5 is at the inlet of the nozzle due to no thermodynamic work being done in the nozzle; see p.081201-8, Col.2, Table 2, Col.E1 of Goulos); wherein a bypass to core ratio of:
b
y
p
a
s
s
e
x
h
a
u
s
t
n
o
z
z
l
e
p
r
e
s
s
u
r
e
r
a
t
i
o
[
c
o
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e
e
x
h
a
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o
z
z
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p
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s
s
u
r
e
r
a
t
i
o
]
is configured to be in the range from 1.3 to 1.6 under aircraft cruise conditions (p.081201-8, Col.2, Table 2, Col.E1, Ratio =
2.2
1.5
=
1.47
).
Goulos further teaches the BPR being “of the order of 16” (p.081201-8 Col.1 last paragraph).
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 bypass and core nozzles of Gliebe to achieve the bypass to core nozzle pressure ratio taught by Goulos because the nozzle pressure ratios of Houlos are specifically compatible with he high bypass ratio of ≈16 taught by Gliebe (Goulos, p.081201-1 col.2 para.2 to p.081201-2 col.1 para.1 and p,081201-8 col.1 last para.). Furthermore, it has been held that combining or simple substitution of prior art elements according to known methods to yield predictable results renders the limitation obvious (see MPEP 2141 (III)). In this case, combining the high bypass (BPR ≈ 16) gas turbine engine of Gliebe with the exhaust nozzle pressure ratio of Goulos’ engine E1 that has a bypass ratio of ≈ 16 would have resulted in an operable turbofan engine with BPR ≈ 16 and bypass to core exhaust nozzle pressure ratio of ≈ 1.47.
Response to Arguments
Applicant’s arguments with respect to pressure ratio limitations of claim(s) 11 and 21 have been considered but are moot because the new ground of rejection does not rely on any reference applied in the prior rejection of record for any teaching or matter specifically challenged in the argument.
Applicant’s arguments with respect to the new range of fan speed to fan turbine radius ratio in claims 14 and 21 is merely that it is a limitation that has not been previously considered. However, the range is merely a narrower range of the previous limitation, and the teachings of Gliebe fall in the narrower range. So it is still anticipated by Gliebe.
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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