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 .
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and the fee set forth in 37 CFR 1.17(e) has been timely paid, the finality of the previous Office action has been withdrawn pursuant to 37 CFR 1.114.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 7/9/2025 has been entered.
This is in response to the communication received on 1/16/2026.
Applicant elected without traverse Species B (Fig 6), Species 2 (Fig 10), Species X (Fig 4) and C1 (claim 1-16) in the reply filed on 6/12/2024.
Therefore Fig 5, 1, 11, 7, claims 17-20 are non-elected species.
Claim Rejections - 35 USC § 103
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 (i.e., changing from AIA to pre-AIA ) 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.
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 text of those sections of Title 35, U.S. Code not included in this action can be found in a prior Office action.
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.
This application currently names joint inventors. In considering patentability of the claims the examiner presumes that the subject matter of the various claims was commonly owned as of the effective filing date of the claimed invention(s) absent any evidence to the contrary. Applicant is advised of the obligation under 37 CFR 1.56 to point out the inventor and effective filing dates of each claim that was not commonly owned as of the effective filing date of the later invention in order for the examiner to consider the applicability of 35 U.S.C. 102(b)(2)(C) for any potential 35 U.S.C. 102(a)(2) prior art against the later invention.
Claim(s) 1, 3-4, 6, 9-14, 16, 21 is/are rejected under 35 U.S.C. 103 as being unpatentable over Khalid 20210108572, in view of Suciu 20150300293 Sibbach 20210107676, Thodiyil 20090108134 and Palmatier 2507044.
Regarding claim 1, Khalid teaches:
An assembly for an aircraft propulsion system (Fig. 1), comprising:
a propulsion module comprising an open propulsor rotor (Fig. 1);
a nose cone (image below) and a plurality of open guide vanes (18) arranged circumferentially about an axis (image below), the plurality of open guide vanes including a first open guide vane (20)
a gas generator (“FIG. 1 is a schematic, cross-sectional view of a gas turbine engine” [0011]) configured to drive rotation of the open propulsor rotor about an axis (“central longitudinal axis 14 of engine 10. The rotor assembly 12 is configured to rotate in the circumferential direction at an angular speed during operation, as is indicated by arrow 11.” [0026]), the gas generator including a flowpath (inter alia, 54), a compressor section (inter alia, 44, 34), a combustor section (40), a turbine section (inter alia, 36, 42) and an exhaust section (52), the flowpath extending through the compressor section, the combustor section, the turbine section and the exhaust section from an inlet into the flowpath to an exhaust from the flowpath (“a path for incoming atmospheric air to enter the turbomachinery flowpath 54 (and compressors 44, 34, combustion section 40, and turbines 36, 42) “ [0032], Fig. 1)
wherein the exterior surface of the propulsion module comprises at least one of the nose cone, the first open guide vane or the open propulsor rotor (image below).
Khalid is silent about the turbine section axially between the combustor section and the open propulsor rotor along the axis, as claimed.
However, Suciu a reverse core turbine engine for an aircraft (title), and:
a turbine section (118) axially between a combustor section (116) and a propulsor rotor (26) along the axis (Fig. 3)
It would have been obvious to a person having ordinary skills in the art before the effective filing date of the claimed invention to provide Khalid with Suciu's structure discussed above, so the turbine section axially between the combustor section and the open propulsor rotor along the axis, in order to provide a reverse core flow engine because ”The reverse core engine allows a more compact arrangement” as taught by Suciu [0004]
Khalid in view of Suciu is silent about
an anti-icing system configured to bleed gas from the flowpath in the exhaust section to provide heated gas, and the anti-icing system configured to direct the heated gas to the propulsion module to reduce or prevent ice accumulation on an exterior surface of the propulsion module
However, Sibbach an aircraft (abstract) with an anti-icing system (Fig. 7):
an anti-icing system (inter alia, 522 524, Fig 7) configured to bleed gas from the flowpath in the exhaust section to provide heated gas (the bypass portion 524 allows some of the auxiliary power unit exhaust flow to be routed from the auxiliary power unit exhaust duct 522, passing through 321, 328, Fig. 7), and the anti-icing system configured to direct the heated gas to the propulsion module to reduce or prevent ice accumulation on an exterior surface of the propulsion module (exhaust duct 522 to a portion of the nacelle 306 thereby providing de-ice capabilities using this configuration of the bypass portion 524 [0091], providing heat to elements along the path, inter alia, 328, 321, where the nacelle is part of the propulsion module in Sibbach).
It would have been obvious to a person having ordinary skill the art before the effective filing date of the claimed invention to provide Khalid in view of Suciu with Sibbach's structure discussed above, providing an anti-icing system configured to bleed gas from the flowpath in the exhaust section to provide heated gas, and the anti-icing system configured to direct the heated gas to the propulsion module to reduce or prevent ice accumulation on an exterior surface of the propulsion module, “thereby providing de-ice capabilities using this configuration of the bypass portion 524” as taught by Sibbach [0091], this teaching being applicable to different portions of the propulsion module.
Additionally, Thodiyil teaches:
“The formation of ice on aircraft structures, for example engine inlets, wings, control surfaces, propellers, booster inlet vanes, inlet frames, etc., has been a formidable problem since the inception of heavier-than-air flight” [0002], and “Typically, icing protection is provided by heating the areas of the aircraft that are prone to icing. One of the most common anti-icing techniques is to disperse hot bleed air gases from the engine” [0003].
It would have been obvious to a person having ordinary skill the art before the effective filing date of the claimed invention to provide Khalid in view of Suciu with Sibbach and Thodiyil's teachings discussed above such that the heated gas discussed above is also provided to aircraft structures susceptible to icing as taught by Thodiyil [0002-0003].
Khalid teaches the exterior surface of the propulsion module comprises at least one of the nose cone, the first open guide vane or the open propulsor rotor, as already discussed above, and Khalid in view of Suciu, Sibbach and Thodiyil, teaches an anti-icing system configured to bleed gas from the flowpath in the exhaust section to provide heated gas, and the anti-icing system configured to direct the heated gas to the propulsion module to reduce or prevent ice accumulation on an exterior surface of the propulsion module, as also discussed above. Khalid in view of Suciu, Sibbach and Thodiyil does not teach preventing ice accumulation on at least one of the nose cone, the first open guide vane or the open propulsor rotor.
However, Palmatier teaches and aircraft de-icing system (Col 1, ll. 1-2), and:
the anti-icing system configured to direct the heated gas to the propulsion module to reduce or prevent ice accumulation on an exterior surface of the propulsion module (“stream of heated air may be utilized to heat any selected air craft surface to prevent the formation of ice thereon or to melt ice which has already been formed thereon” Col 5 ll. 10-16), the exterior surface of the propulsion module comprises at least one of the nose cone, the first open guide vane or the open propulsor rotor (“passes heated air into the spinner 42 and, particularly, into engagement with the interior surface of said spinner 42 from which such heated air may pass into heat-transferring relation with the blade cuffs 43 by way of the respective passages 47a and, further, heated air from the spinner 42 may pass into the hollow blades 414 by way of passages 41b” Col 5 ll. 35-43).
It would have been obvious to a person having ordinary skill the art before the effective filing date of the claimed invention to provide Khalid in view of Suciu, Sibbach and Thodiyil with Palmatier's structure discussed above so “stream of heated air may be utilized to heat any selected air craft surface to prevent the formation of ice thereon or to melt ice which has already been formed thereon” as taught by Palmatier Col 5 ll. 10-16.
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Regarding claim 6, Khalid in view of Suciu, Sibbach, Thodiyil and Palmatier teaches the invention as discussed for claim 1.
Khalid in view of Suciu, Sibbach, Thodiyil and Palmatier l teaches the open guide vanes as already discussed.
Khalid in view of Suciu, Sibbach, Thodiyil and Palmatier is silent about:
the turbine section is axially between the combustor section and the plurality of open guide vanes along the axis.
However, Suciu a reverse core turbine engine for an aircraft (title), and:
the turbine section (118) is axially between the combustor section (116) and the plurality of guide vanes (129) along the axis (Fig. 3).
Regarding claim 9, Khalid in view of Suciu, Sibbach, Thodiyil and Palmatier teaches the invention as discussed for claim 1. As already discussed for claim 1, Khalid in view of Suciu, Sibbach, Thodiyil and Palmatier teaches gas bled from the flowpath in the exhaust section.
Khalid in view of Suciu, Sibbach, Thodiyil and Palmatier, as discussed so far, is silent about:
wherein the anti-icing system includes a mixing valve configured to mix ambient air with the gas [bled from the flowpath in the exhaust section] to provide the heated gas
However, Palmatier teaches:
wherein the anti-icing system includes a mixing valve configured to mix ambient air with the gas to provide the heated gas (a heat transfer unit receiving air from said intake and delivering hot air, and means to dilute said hot air with ambient air to provide warm air, Claim 2).
Regarding claim 10, Khalid in view of Suciu, Sibbach, Thodiyil and Palmatier teaches the invention as discussed for claim 1. Khalid in view of Suciu, Sibbach, Thodiyil and Palmatier, as discussed so far, is silent about:
The assembly of claim 1, wherein the heated gas consists of the gas bled from the flowpath in the exhaust section.
However, Sibbach teaches:
the heated gas consists of the gas bled from the flowpath in the exhaust section (the bypass portion 524 allows some of the auxiliary power unit exhaust flow to be routed from the auxiliary power unit exhaust duct 522).
Regarding claim 11, Khalid in view of Suciu, Sibbach, Thodiyil and Palmatier teaches the invention as discussed for claim 1.
Khalid in view of Suciu, Sibbach, Thodiyil and Palmatier, as discussed for claim 1, teaches the open propulsor rotor.
However, Suciu a reverse core turbine engine for an aircraft (title), and:
wherein the turbine section comprises a free power turbine rotor (inter alia, 120, 128); and
the free power turbine rotor is coupled to and configured to drive the rotation (Fig 3) of the open propulsor rotor (Fig 3; examiner notes the open propulsor rotor has been discussed for claim 1).
Regarding claim 12, Khalid in view of Suciu, Sibbach, Thodiyil and Palmatier teaches the invention as discussed for claim 11.
Khalid in view of Suciu, Sibbach, Thodiyil and Palmatier, as discussed so far, is silent about the geartrain as claimed.
However, Suciu further teaches:
further comprising a geartrain (131) coupling the free power turbine rotor to the open propulsor rotor (Fig 3).
Regarding claim 13, Khalid in view of Suciu, Sibbach, Thodiyil and Palmatier teaches the invention as discussed for claim 11.
Khalid in view of Suciu, Sibbach, Thodiyil and Palmatier, as discussed so far, is silent about the free turbine is rotatable about the axis as claimed.
However, Suciu further teaches:
wherein the free power turbine rotor is rotatable about the axis (axis A, Fig 3).
Regarding claim 14, Khalid in view of Suciu, Sibbach, Thodiyil and Palmatier teaches the invention as discussed for claim 11. Khalid in view of Suciu, Sibbach, Thodiyil and Palmatier, as discussed so far, is silent about wherein the axis is a first axis and the free power turbine rotor is rotatable about a second axis that is offset from the first axis.
However, Suciu further teaches:
wherein the axis is a first axis (C, Fig 3) and the free power turbine rotor is rotatable about a second axis (A, Fig 3) that is offset from the first axis (Fig 3).
Regarding claim 16, Khalid in view of Suciu, Sibbach, Thodiyil and Palmatier teaches the invention as discussed for claim 1. Khalid in view of Suciu, Sibbach, Thodiyil and Palmatier, as discussed so far, is silent about the axis is a first axis, and the gas generator comprises a rotating assembly rotatable about a second axis offset from the first axis; and the rotating assembly includes a compressor rotor, a turbine rotor and a shaft coupling the turbine rotor to the compressor rotor, the compressor rotor is in the compressor section, and the turbine rotor is in the turbine section.
However, Suciu a reverse core turbine engine for an aircraft (title), and:
wherein the axis is a first axis (C, Fig 3), and the gas generator comprises a rotating assembly rotatable about a second axis (A, Fig 3) offset from the first axis (Fig 3); and
the rotating assembly includes a compressor rotor (114),
a turbine rotor (118) and a shaft coupling the turbine rotor to the compressor rotor (The core engine 130 is mounted on a centerline C [0037]), the compressor rotor is in the compressor section (Fig 3), and the turbine rotor is in the turbine section (Fig 3).
Regarding claim 21, Khalid in view of Suciu, Sibbach, Thodiyil and Palmatier teaches the invention as discussed for claim 3.
Khalid in view of Suciu, Sibbach, Thodiyil and Palmatier, as already discussed, teaches the passage.
Khalid further teaches:
the first open guide vane (20, as discussed) comprises an external skin forming an outer surface of the first open guide vane (exterior surface of 20, Fig. 1).
Regarding forming a peripheral boundary of the passage, as already discussed for Claim 3, Khalid in view of Suciu, Sibbach, Thodiyil and Palmatier teaches the passage extends into the first open inlet guide vane. Since the skin forms an exterior surface of the first open guide vane as discussed above, one of ordinary skill would understand the skin would form a peripheral boundary of the passage. This also taught by Sibbach in the image below:
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Claim(s) 22 is/are rejected under 35 U.S.C. 103 as being unpatentable over Khalid 20210108572, in view of Suciu 20150300293 Sibbach 20210107676, Charier 20230271695,
Regarding claim 22, Khalid teaches:
An assembly for an aircraft propulsion system (Fig. 1), comprising:
a propulsion module comprising an open propulsor rotor (Fig. 1) and a plurality of open guide vanes (18) arranged circumferentially about an axis (image below), the plurality of open guide vanes including a first open guide vane (20);
a gas generator (“FIG. 1 is a schematic, cross-sectional view of a gas turbine engine” [0011]) configured to drive rotation of the open propulsor rotor about the axis (“central longitudinal axis 14 of engine 10. The rotor assembly 12 is configured to rotate in the circumferential direction at an angular speed during operation, as is indicated by arrow 11.” [0026]), the gas generator including a flowpath (inter alia, 54), a compressor section (inter alia, 44, 34), a combustor section (40), a turbine section (inter alia, 36, 42) and an exhaust section (52), the flowpath extending through the compressor section, the combustor section, the turbine section and the exhaust section from an inlet into the flowpath to an exhaust from the flowpath (“a path for incoming atmospheric air to enter the turbomachinery flowpath 54 (and compressors 44, 34, combustion section 40, and turbines 36, 42) “ [0032], Fig. 1),
an anti-icing system configured to bleed gas from the flowpath in the exhaust section to provide heated gas, the anti-icing system configured to direct the heated gas to reduce or prevent ice accumulation on an exterior surface of the propulsion module, and the exterior surface of the propulsion module comprising the first open guide vane (image below).
Khalid is silent about the turbine section axially between the combustor section and the open propulsor rotor along the axis, as claimed.
However, Suciu a reverse core turbine engine for an aircraft (title), and:
a turbine section (118) axially between a combustor section (116) and a propulsor rotor (26) along the axis (Fig. 3)
It would have been obvious to a person having ordinary skills in the art before the effective filing date of the claimed invention to provide Khalid with Suciu's structure discussed above, so the turbine section axially between the combustor section and the open propulsor rotor along the axis, in order to provide a reverse core flow engine because ”The reverse core engine allows a more compact arrangement” as taught by Suciu [0004]
Khalid in view of Suciu is silent about:
an anti-icing system configured to bleed gas from the flowpath in the exhaust section to provide heated gas, and the anti-icing system configured to direct the heated gas to the propulsion module to reduce or prevent ice accumulation on an exterior surface of the propulsion module
However, Sibbach an aircraft (abstract) with an anti-icing system (Fig. 7):
an anti-icing system (inter alia, 522 524, Fig 7) configured to bleed gas from the flowpath in the exhaust section to provide heated gas (the bypass portion 524 allows some of the auxiliary power unit exhaust flow to be routed from the auxiliary power unit exhaust duct 522, passing through 321, 328, Fig. 7), and the anti-icing system configured to direct the heated gas to the propulsion module to reduce or prevent ice accumulation on an exterior surface of the propulsion module (exhaust duct 522 to a portion of the nacelle 306 thereby providing de-ice capabilities using this configuration of the bypass portion 524 [0091], providing heat to elements along the path, inter alia, 328, 321, where the nacelle is part of the propulsion module in Sibbach).
It would have been obvious to a person having ordinary skill the art before the effective filing date of the claimed invention to provide Khalid in view of Suciu with Sibbach's structure discussed above, providing an anti-icing system configured to bleed gas from the flowpath in the exhaust section to provide heated gas, and the anti-icing system configured to direct the heated gas to the propulsion module to reduce or prevent ice accumulation on an exterior surface of the propulsion module, “thereby providing de-ice capabilities using this configuration of the bypass portion 524” as taught by Sibbach [0091], this teaching being applicable to different portions of the propulsion module.
Khalid in view of Suciu and Sibbach teaches the open guide vane and preventing ice accumulation on exterior surface as already discussed, but is silent about
direct the heated gas to first open guide vane [to reduce or prevent ice accumulation on an exterior surface of the propulsion module, and the exterior surface of the propulsion module comprising the first open guide vane].
However, Charier teaches a “Turbine engines with at least one unducted propeller are known as “open rotor” or “unducted fan”” [0002], and
direct the heated gas (“hot air from the primary vein 20” ) to first open guide vane (23, fig. 2) [to reduce or prevent ice accumulation on an exterior surface of the propulsion module, and the exterior surface of the propulsion module comprising the first open guide vane] (“The de-icing device 60 is designed to take hot air from the primary vein 20 and distribute it to the retaining means (retaining member 37 and arm 50) of the stator vanes 23” [0057]).
It would have been obvious to a person having ordinary skills in the art before the effective filing date of the claimed invention to provide Khalid in view of Suciu and Sibbach with Charier's teachings discussed above in order “to prevent ice from forming” as taught by Charier [0057]
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Claim(s) 1 is/are rejected under 35 U.S.C. 103 as being unpatentable over Suciu 20150300293 in view of Khalid 20210108572, Sibbach 20210107676 and Thodiyil 20090108134 (SECOND REJECTION)
Regarding claim 1, Suciu teaches:
An assembly for an aircraft propulsion system, comprising:
a propulsion module (comprising elements in Fig 3, inter alia, 24, fan 26, compressor 114, combustor 116, 144, core engine 130, 124, vanes 129, gear reduction 131, shaft 128, vanes 129, duct 111, free turbine 120, inlet duct 28, etc) comprising a propulsor rotor (inter alia, fan 26), a nose cone (Fig. 3, nose cone on the left side of the image along centerline A)
a gas generator (inter alia, 114, 116, 118) configured to drive rotation of the propulsor rotor (free turbine drives the fan rotor [0019]) about an axis (A), the gas generator including a flowpath (starting at the arrow on the bottom right end of Fig 3, exiting at 124), a compressor section (114), a combustor section (116), a turbine section (inter alia, 118, 120) and an exhaust section (exhaust nozzle [0006]), the flowpath extending through the compressor section (Fig 3), the combustor section (Fig 3), the turbine section (Fig 3) and the exhaust section from an inlet into the flowpath to an exhaust from the flowpath (Fig 3), and the turbine section axially between the combustor section and the open propulsor rotor along the axis (Fig 3);
wherein the exterior surface of the propulsion module comprises at least one of the nose cone (Fig. 3, nose cone on the left side of the image along centerline A)
Suciu is silent about:
the propulsion module being an open propulsor rotor;
and a plurality of open guide vanes arranged circumferentially about an axis, the plurality of open guide vanes including a first open guide vane;
However, Khalid teaches:
an open propulsor rotor (Fig 1)
a plurality of open guide vanes arranged circumferentially about an axis, the plurality of open guide vanes including a first open guide vane (20)
It would have been obvious to a person having ordinary skill the art before the effective filing date of the claimed invention to provide Suciu with Khalid's structure discussed above in order to provide “an open rotor engine” [0004] because it “permits the use of larger rotor blades able to act upon a larger volume of air than for a traditional turbofan engine, potentially improving propulsive efficiency over conventional turbofan engine designs” [0004].
Suciu in view of Khalid is silent about:
an anti-icing system configured to bleed gas from the flowpath in the exhaust section to provide heated gas, and the anti-icing system configured to direct the heated gas to the propulsion module to reduce or prevent ice accumulation on an exterior surface of the propulsion module.
However, Sibbach teaches:
an anti-icing system (inter alia, 522 524, Fig 7) configured to bleed gas from the flowpath in the exhaust section to provide heated gas (the bypass portion 524 allows some of the auxiliary power unit exhaust flow to be routed from the auxiliary power unit exhaust duct 522, passing through 321, 328, Fig. 7), and the anti-icing system configured to direct the heated gas to the propulsion module to reduce or prevent ice accumulation on an exterior surface of the propulsion module (exhaust duct 522 to a portion of the nacelle 306 thereby providing de-ice capabilities using this configuration of the bypass portion 524 [0091], providing heat to elements along the path, inter alia, 328, 321, where the nacelle is part of the propulsion module in Sibbach).
It would have been obvious to a person having ordinary skill the art before the effective filing date of the claimed invention to provide Suciu in view of Khalid with Sibbach's structure discussed above, providing an anti-icing system configured to bleed gas from the flowpath in the exhaust section to provide heated gas, and the anti-icing system configured to direct the heated gas to the propulsion module to reduce or prevent ice accumulation on an exterior surface of the propulsion module, “thereby providing de-ice capabilities using this configuration of the bypass portion 524” as taught by Sibbach [0091], this teaching being applicable to different portions of the propulsion module.
Additionally, Thodiyil teaches:
“The formation of ice on aircraft structures, for example engine inlets, wings, control surfaces, propellers, booster inlet vanes, inlet frames, etc., has been a formidable problem since the inception of heavier-than-air flight” [0002], and “Typically, icing protection is provided by heating the areas of the aircraft that are prone to icing. One of the most common anti-icing techniques is to disperse hot bleed air gases from the engine” [0003].
It would have been obvious to a person having ordinary skill the art before the effective filing date of the claimed invention to provide Suciu in view of Khalid and Sibbach with Thodiyil's structure discussed above such that the heated gas discussed above is also provided to aircraft structures susceptible to icing as taught by Thodiyil [0002-0003].
Claim(s) 9 is/are rejected under 35 U.S.C. 103 as being unpatentable over Suciu 20150300293 in view of Khalid 20210108572, Sibbach 20210107676, Thodiyil 20090108134 in view of Palmatier 2507044 (SECOND REJECTION).
Regarding claim 9, Suciu in view of Khalid, Sibbach and Thodiyil teaches the invention as discussed for claim 1. As already discussed for claim 1, Suciu in view of Khalid, Sibbach and Thodiyil teaches gas bled from the flowpath in the exhaust section.
Suciu in view of Khalid, Sibbach and Thodiyil, as discussed so far, is silent about:
wherein the anti-icing system includes a mixing valve configured to mix ambient air with the gas [bled from the flowpath in the exhaust section] to provide the heated gas
However, Palmatier teaches:
wherein the anti-icing system includes a mixing valve configured to mix ambient air with the gas to provide the heated gas (a heat transfer unit receiving air from said intake and delivering hot air, and means to dilute said hot air with ambient air to provide warm air, Claim 2).
It would have been obvious to a person having ordinary skill the art before the effective filing date of the claimed invention to provide Suciu in view of Khalid, Sibbach and Thodiyil with Palmatier's structure discussed above such that the anti-icing system includes a mixing valve configured to mix ambient air with the gas bled from the flowpath in the exhaust section to provide the heated gas in order to “ to dilute said hot air with ambient air to provide warm air to said supply pipe at a controlled temperature” as taught by Palmatier, Claim 2.
Claim(s) 10-14, 16 is/are rejected under 35 U.S.C. 103 as being unpatentable over Suciu 20150300293 in view of Khalid 20210108572, Sibbach 20210107676 and Thodiyil 20090108134 (SECOND REJECTION).
Regarding claim 10, Suciu in view of Khalid, Sibbach and Thodiyil teaches the invention as discussed for claim 1. Suciu in view of Khalid, Sibbach and Thodiyil, as discussed so far, is silent about:
The assembly of claim 1, wherein the heated gas consists of the gas bled from the flowpath in the exhaust section.
However, Sibbach teaches:
the heated gas consists of the gas bled from the flowpath in the exhaust section (the bypass portion 524 allows some of the auxiliary power unit exhaust flow to be routed from the auxiliary power unit exhaust duct 522).
Regarding claim 11, Suciu in view of Khalid, Sibbach and Thodiyil teaches the invention as discussed for claim 1.
Suciu in view of Khalid, Sibbach and Thodiyil, as discussed for claim 1, teaches the open propulsor rotor.
Suciu further teaches:
wherein the turbine section comprises a free power turbine rotor (inter alia, 120, 128); and
the free power turbine rotor is coupled to and configured to drive the rotation (Fig 3) of the open propulsor rotor (Fig 3; examiner notes the open propulsor rotor has been discussed for claim 1).
Regarding claim 12, Suciu in view of Khalid, Sibbach and Thodiyil teaches the invention as discussed for claim 11. Suciu further teaches:
further comprising a geartrain (131) coupling the free power turbine rotor to the open propulsor rotor (Fig 3).
Regarding claim 13, Suciu in view of Khalid, Sibbach and Thodiyil teaches the invention as discussed for claim 11. Suciu further teaches:
wherein the free power turbine rotor is rotatable about the axis (axis A, Fig 3).
Regarding claim 14, Suciu in view of Khalid, Sibbach and Thodiyil teaches the invention as discussed for claim 11. Suciu further teaches:
wherein the axis is a first axis (C, Fig 3) and the free power turbine rotor is rotatable about a second axis (A, Fig 3) that is offset from the first axis (Fig 3).
Regarding claim 16, Suciu in view of Khalid, Sibbach and Thodiyil teaches the invention as discussed for claim 1. Suciu further teaches:
wherein the axis is a first axis (C, Fig 3), and the gas generator comprises a rotating assembly rotatable about a second axis (A, Fig 3) offset from the first axis (Fig 3); and
the rotating assembly includes a compressor rotor (114),
a turbine rotor (118) and a shaft coupling the turbine rotor to the compressor rotor (The core engine 130 is mounted on a centerline C [0037]), the compressor rotor is in the compressor section (Fig 3), and the turbine rotor is in the turbine section (Fig 3).
Response to Arguments/Remarks
Applicant’s arguments have been considered, but they are not persuasive because they do not apply to the new combination of references, i.e., adding a new reference to the old combination of references, that was necessitated by applicant’s amendment. However, to the extent possible, applicant’s arguments have been addressed in the body of the rejections above, at the appropriate location.
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 Roberto T. Igue whose telephone number is (303)297-4389. The examiner can normally be reached Monday-Friday 7:30-4:30 PT.
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, Phutthiwat Wongwian 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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/ROBERTO TOSHIHARU IGUE/ Examiner, Art Unit 3741
/PHUTTHIWAT WONGWIAN/Supervisory Patent Examiner, Art Unit 3741