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.
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and the fee set forth in 37 CFR 1.17(e) has been timely paid, the finality of the previous Office action has been withdrawn pursuant to 37 CFR 1.114.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 21 January 2026 has been entered.
Specification
Amendment to the Specification filed 21 January 2026 are not entered because it presents new matter. There does not appear to be any evidence in the original disclosure that the fan tip loading is a dimensionless value, particularly in light of its definition as dH/Utip2 which has units of
J
m
2
s
2
o
r
J
k
g
m
2
s
2
.
Claim Rejections - 35 USC § 112
The following is a quotation of the first paragraph of 35 U.S.C. 112(a):
(a) IN GENERAL.—The specification shall contain a written description of the invention, and of the manner and process of making and using it, in such full, clear, concise, and exact terms as to enable any person skilled in the art to which it pertains, or with which it is most nearly connected, to make and use the same, and shall set forth the best mode contemplated by the inventor or joint inventor of carrying out the invention.
The following is a quotation of the first paragraph of pre-AIA 35 U.S.C. 112:
The specification shall contain a written description of the invention, and of the manner and process of making and using it, in such full, clear, concise, and exact terms as to enable any person skilled in the art to which it pertains, or with which it is most nearly connected, to make and use the same, and shall set forth the best mode contemplated by the inventor of carrying out his invention.
Claims 13 and 15-20 are rejected under 35 U.S.C. 112(a) or 35 U.S.C. 112 (pre-AIA ), first paragraph, as failing to comply with the written description requirement. The claim(s) contains subject matter which was not described in the specification in such a way as to reasonably convey to one skilled in the relevant art that the inventor or a joint inventor, or for applications subject to pre-AIA 35 U.S.C. 112, the inventor(s), at the time the application was filed, had possession of the claimed invention.
Regarding claims 13 and 15-20, the combination of the fan speed to fan-turbine radius difference ratio of 1.2 - 2.0 rpm/mm with the following ranges:
Cl.13, 18: center of gravity position ratio multiplied by MTO fan speed: 600-1350rpm
Cl.15: fan tip radius / twice the turbine radius: 0.8-2.1
Cl.16: fan tip radius / core length: 0.3-1
Cl.17: MTO fan speed: 1450-3020rpm
Cl.19: center of gravity position ratio multiplied by MTO fan speed: 650-1276rpm
Cl.20: twice the fan tip radius / (twice the turbine radius multiplied by the center of gravity position ratio): 4.9-12.6
was not supported by the original disclosure because the originally filed Specification does not disclose the full scope of these ranges as being compatible with the limited scope of the ranges of claims 1 and 14. That is, pp.19 and 59 disclose the fan speed to fan-turbine radius ratio may be in range between 1.2 to 2.0 rpm/mm for an engine with fan tip radius in the range of 155-200cm. Whereas the corresponding fan radii and range values for the above parameters are as follows (per the originally filed Specification):
110-150cm
155-200cm
a
925-1325rpm
650-910rpm
b
0.8-2.1
0.9-1.7
c
0.4-0.9
0.5-0.8
d
1970-3020rpm
1450-1910rpm
e
925-1325rpm
650-910rpm
f
No disclosure
No disclosure
Thus, the combination of ratios claimed in claims 13 and 15-20 were not supported by the original disclosure under 112a.
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 13 and 15-20 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 13 and 15-20, the 112a new matter issue above renders the claim indefinite, because it is unclear whether the claims contain a typographical error (in which case it is further unclear which set of ranges, corresponding to which fan tip radius range, was intended by Applicant) or whether Applicant intended to claim the recited subject matter currently rejected under 112a.
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) 11 and 13-20 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 11, 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
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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 each of the rotor blades of the lowest pressure rotor stage (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).
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 14, 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 (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
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).
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 0.8 to 2.1 (
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 0.3 to 1 (
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 3020 rpm (1691rpm from Table 11 falls in the claimed range).
Regarding claims 13 and 18-19, Gliebe teaches all the limitations of the claimed invention as discussed above.
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media_image4.png
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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 1350 rpm for claims 13 and 18, or in arrange of 650rpm to 1276rpm for claim 19
47
120
×
1691
≈
662
r
p
m
.
Regarding claim 20, Gliebe teaches all the limitations of the claimed invention as discussed above.
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media_image4.png
438
1294
media_image4.png
Greyscale
Gliebe further teaches the gas turbine engine has an engine length and a centre of gravity position (CGPos) measured relative to the fan (Fig 4 above), and a ratio of:
2
×
t
h
e
f
a
n
t
i
p
r
a
d
i
u
s
2
×
t
h
e
t
u
r
b
i
n
e
r
a
d
i
u
s
a
t
t
h
e
l
o
w
e
s
t
p
r
e
s
s
u
r
e
r
o
t
o
r
s
t
a
g
e
×
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
is in the range of 4.9 to 12.6.
That is, per Fig 4:
2
×
65
2
×
25
×
47
120
≈
6.6
which falls in the claimed range.
Response to Arguments
Applicant’s arguments with respect to claim(s) 11 and 14 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.
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.
Examiner interviews are available via telephone, in-person, and video conferencing using a USPTO supplied web-based collaboration tool. To schedule an interview, applicant is encouraged to use the USPTO Automated Interview Request (AIR) at http://www.uspto.gov/interviewpractice.
If attempts to reach the examiner by telephone are unsuccessful, the examiner’s supervisor, Devon Kramer can be reached on (571-270-5426. The fax phone number for the organization where this application or proceeding is assigned is 571-273-8300.
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/STEPHANIE SEBASCO CHENG/Primary Examiner, Art Unit 3741